Thermally coupled CMC combustor liner

A combustor for a gas turbine engine is provided, which includes: a CMC liner having a forward end and an aft end; an annular dome comprising a metal and defining an annular slot within its end defined between an outer arm and an inner arm; a feather seal extending from an annularly exterior surface of the annular dome to an annularly exterior surface of the liner; and a plurality of pin members. The forward end of the liner defines a plurality of fingers and a plurality of axial slots, and is fitted between the outer arm and the inner arm within the annular slot. Each pin member extending through an aperture in the feather seal, through an aperture in the outer arm of the annular dome, through an opening defined by the liner, and through an aperture in the inner arm of the annular dome.

FIELD OF THE INVENTION

The present invention relates generally to the use of Ceramic Matrix Composite (CMC) liners in a gas turbine engine combustor and, in particular, to the mounting of such CMC liners to the dome and cowl of the combustor so as to accommodate differences in thermal growth therebetween.

BACKGROUND OF THE INVENTION

It will be appreciated that the use of non-traditional high temperature materials, such as Ceramic Matrix Composites (CMC), are being studied and utilized as structural components in gas turbine engines. There is particular interest, for example, in making combustor components which are exposed to extreme temperatures from such material in order to improve the operational capability and durability of the engine. However, substitution of materials having higher temperature capabilities than metals has been difficult in light of the widely disparate coefficients of thermal expansion when different materials are used in adjacent components of the combustor. This mismatch can result in binding with adjacent components and subsequent failure unless sufficient clearance is available.

Accordingly, various schemes have been employed to address problems that are associated with mating parts having differing thermal expansion properties. As seen in U.S. Pat. No. 5,291,732 to Halila, U.S. Pat. No. 5,291,733 to Halila, and U.S. Pat. No. 5,285,632 to Halila, an arrangement is disclosed which permits a metal heat shield to be mounted to a liner made of CMC so that radial expansion therebetween is accommodated. This involves positioning a plurality of circumferentially spaced mount pins through openings in the heat shield and liner so that the liner is able to move relative to the heat shield.

U.S. Pat. No. 6,397,603 to Edmondson et al. also discloses a combustor having a liner made of Ceramic Matrix Composite materials, where the liner is mated with an intermediate liner dome support member in order to accommodate differential thermal expansion without undue stress on the liner. The Edmondson et al. patent further includes the ability to regulate part of the cooling air flow through the interface joint.

While each of the aforementioned patents reveals mounting arrangements for a CMC liner which are useful for their particular combustor designs, none involve a liner made of CMC materials being connected directly to the dome and cowl portions of the combustor in a single mounting arrangement. Thus, it would be desirable for a simple mounting assembly to be developed for a liner having a different coefficient of thermal expansion than the components to which it is mated. It would also be desirable for such mounting assembly to be efficiently sized such that clearances with adjacent hardware are not required.

BRIEF DESCRIPTION OF THE INVENTION

A combustor for a gas turbine engine is generally provided. In one embodiment, the combustor comprises: a liner comprising a ceramic matrix composite material and having a forward end and an aft end; an annular dome comprising a metal and defining an annular slot within its end defined between an outer arm and an inner arm; a feather seal extending from an annularly exterior surface of the annular dome to an annularly exterior surface of the liner; and a plurality of pin members. The forward end of the liner defines a plurality of fingers and a plurality of axial slots, and is fitted between the outer arm and the inner arm within the annular slot. Each pin member extending through an aperture in the feather seal, through an aperture in the outer arm of the annular dome, through an opening defined by the liner, and through an aperture in the inner arm of the annular dome.

A gas turbine engine is also generally provided, which comprises a compressor; a combustor; and a turbine. The combustor generally comprises: a liner comprising a ceramic matrix composite material and having a forward end and an aft end; an annular dome comprising a metal and defining an annular slot within its end defined between an outer arm and an inner arm; a feather seal extending from an annularly exterior surface of the annular dome to an annularly exterior surface of the liner; and a plurality of pin members. The forward end of the liner defines a plurality of fingers and a plurality of axial slots, and is fitted between an outer arm and an inner arm within the annular slot. Each pin member extending through an aperture in the feather seal, through an aperture in the outer arm of the annular dome, through an opening defined by the liner, and through an aperture in the inner arm of the annular dome.

A liner of a combustor is also generally provided. In one embodiment, the liner comprises a ceramic matrix composite material, with the liner having a forward end that defines a plurality of fingers and a plurality of axial slots.

DETAILED DESCRIPTION OF THE INVENTION

Referring now to the drawings,FIG. 1illustrates a cross-sectional view of one embodiment of a gas turbine engine10that may be utilized within an aircraft in accordance with aspects of the present subject matter, with the engine10being shown having a longitudinal or axial centerline axis12extending therethrough for reference purposes. In general, the engine10may include a core gas turbine engine (indicated generally by reference character14) and a fan section16positioned upstream thereof. The core engine14may generally include a substantially tubular outer casing18that defines an annular inlet20. In addition, the outer casing18may further enclose and support a booster compressor22for increasing the pressure of the air that enters the core engine14to a first pressure level. A high pressure, multi-stage, axial-flow compressor24may then receive the pressurized air from the booster compressor22and further increase the pressure of such air. The pressurized air exiting the high-pressure compressor24may then flow to a combustor26within which fuel is injected into the flow of pressurized air, with the resulting mixture being combusted within the combustor26. The high energy combustion products are directed from the combustor26along the hot gas path of the engine10to a first (high pressure) turbine28for driving the high pressure compressor24via a first (high pressure) drive shaft30, and then to a second (low pressure) turbine32for driving the booster compressor22and fan section16via a second (low pressure) drive shaft34that is generally coaxial with first drive shaft30. After driving each of turbines28and32, the combustion products may be expelled from the core engine14via an exhaust nozzle36to provide propulsive jet thrust.

It should be appreciated that each turbine28,30may generally include one or more turbine stages, with each stage including a turbine nozzle (not shown inFIG. 1) and a downstream turbine rotor (not shown inFIG. 1). As will be described below, the turbine nozzle may include a plurality of vanes disposed in an annular array about the centerline axis12of the engine10for turning or otherwise directing the flow of combustion products through the turbine stage towards a corresponding annular array of rotor blades forming part of the turbine rotor. As is generally understood, the rotor blades may be coupled to a rotor disk of the turbine rotor, which is, in turn, rotationally coupled to the turbine's drive shaft (e.g., drive shaft30or34).

Additionally, as shown inFIG. 1, the fan section16of the engine10may generally include a rotatable, axial-flow fan rotor38that configured to be surrounded by an annular fan casing40. In particular embodiments, the (LP) drive shaft34may be connected directly to the fan rotor38such as in a direct-drive configuration. In alternative configurations, the (LP) drive shaft34may be connected to the fan rotor38via a speed reduction device37such as a reduction gear gearbox in an indirect-drive or geared-drive configuration. Such speed reduction devices may be included between any suitable shafts/spools within engine10as desired or required.

It should be appreciated by those of ordinary skill in the art that the fan casing40may be configured to be supported relative to the core engine14by a plurality of substantially radially-extending, circumferentially-spaced outlet guide vanes42. As such, the fan casing40may enclose the fan rotor38and its corresponding fan rotor blades44. Moreover, a downstream section46of the fan casing40may extend over an outer portion of the core engine14so as to define a secondary, or by-pass, airflow conduit48that provides additional propulsive jet thrust.

During operation of the engine10, it should be appreciated that an initial air flow (indicated by arrow50) may enter the engine10through an associated inlet52of the fan casing40. The air flow50then passes through the fan blades44and splits into a first compressed air flow (indicated by arrow54) that moves through conduit48and a second compressed air flow (indicated by arrow56) which enters the booster compressor22. The pressure of the second compressed air flow56is then increased and enters the high pressure compressor24(as indicated by arrow58). After mixing with fuel and being combusted within the combustor26, the combustion products60exit the combustor26and flow through the first turbine28. Thereafter, the combustion products60flow through the second turbine32and exit the exhaust nozzle36to provide thrust for the engine10.

Referring now toFIG. 2, a cross-sectional view is provided of the combustion section26of the exemplary turbofan engine10ofFIG. 1. More particularly,FIG. 2provides a perspective, cross-sectional view of a combustor assembly100, which may be positioned in the combustion section26of the exemplary turbofan engine10ofFIG. 1, in accordance with an exemplary embodiment of the present disclosure. Notably,FIG. 2provides a perspective, cross-sectional view of the combustor assembly100having an outer combustor casing removed for clarity.

As shown, the combustor assembly100generally includes an inner liner102extending between and aft end104and a forward end106generally along the axial direction, as well as an outer liner108also extending between and aft end110and a forward end112generally along the axial direction. The inner and outer liners102,108together at least partially define a combustion chamber114therebetween. The inner and outer liners102,108are each attached to an annular dome111. More particularly, the combustor assembly100includes an inner portion116of the annular dome111attached to the forward end106of the inner liner102and an outer portion118of the annular dome111attached to the forward end112of the outer liner108. As will be discussed in greater detail below, the inner and outer portions116,118of the annular dome111each include an enclosed surface120defining an annular slot122for receipt of the forward ends106,112of the respective inner and outer liners102,108.FIG. 3shows this orientation in greater detail, using the outer liner108and outer portion118of the annular dome111as representative, though the present disclosure is not limited to the outer liner108and may be applied similarly to the inner liner102.

The combustor assembly100further includes a plurality of fuel and air mixers124spaced along a circumferential direction within the outer portion118of the annular dome111. More particularly, the plurality of fuel air mixers124are disposed between the outer portion118of the annular dome111and the inner portion116of the annular dome111along the radial direction. Compressed air from the compressor section of the turbofan engine10flows into or through the fuel air mixers124, where the compressed air is mixed with fuel and ignited to create the combustion gases within the combustion chamber114. The inner and outer domes116,118are configured to assist in providing such a flow of compressed air from the compressor section into or through the fuel air mixers124. For example, the outer portion118of the annular dome111includes an outer cowl126at a forward end128and the inner portion116of the annular dome111similarly includes an inner cowl130at a forward end132. The outer cowl126and inner cowl130may assist in directing the flow of compressed air from the compressor section26into or through one or more of the fuel air mixers124.

Moreover, the inner and outer domes116,118can each include attachment portions configured to assist in mounting the combustor assembly100within the turbofan engine10. For example, the outer portion118of the annular dome111can include an attachment extension configured to be mounted to an outer combustor casing and the inner portion116of the annular dome111can include a similar attachment extension configured to attach to an annular support member within the turbofan engine10. In certain exemplary embodiments, the inner portion116of the annular dome111may be formed integrally as a single annular component, and similarly, the outer portion118of the annular dome111may also be formed integrally as a single annular component. It should be appreciated, however, that in other exemplary embodiments, the inner portion116of the annular dome111and/or the outer portion118of the annular dome111may be formed by one or more components joined in any suitable manner. For example, with reference to the outer portion118of the annular dome111, in certain exemplary embodiments, the outer cowl126may be formed separately from the outer portion118of the annular dome111and attached to outer portion118of the annular dome111using, e.g., a welding process. Similarly, any attachment extension may also be formed separately from the outer dam118and attached to the outer portion118of the annular dome111using, e.g., a welding process. Additionally, or alternatively, the inner portion116of the annular dome111may have a similar configuration.

Referring still toFIG. 2, the exemplary combustor assembly100further includes a heat shield142positioned around each fuel air mixer124, arrange circumferentially. The heat shields142, for the embodiment depicted, are attached to and extend between the outer portion118of the annular dome111and the inner portion116of the annular dome111. The heat shields142are configured to protect certain components of the turbofan engine10from the relatively extreme temperatures of the combustion chamber114.

For the embodiment depicted, the inner liner102and outer liner108are each comprised of a ceramic matrix composite (CMC) material, which is a non-metallic material having high temperature capability. Exemplary CMC materials utilized for such liners102,108may include silicon carbide, silicon, silica or alumina matrix materials and combinations thereof. Ceramic fibers may be embedded within the matrix, such as oxidation stable reinforcing fibers including monofilaments like sapphire and silicon carbide (e.g., Textron's SCS-6), as well as rovings and yarn including silicon carbide (e.g., Nippon Carbon's NICALON®, Ube Industries' TYRANNO®, and Dow Corning's SYLRAMIC®), alumina silicates (e.g., Nextel's 440 and 480), and chopped whiskers and fibers (e.g., Nextel's 440 and SAFFIL®), 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). CMC materials may have coefficients of thermal expansion in the range of about 1.3×10−6in/in/° F. to about 3.5×106in/in/° F. in a temperature of approximately 1000-1200° F.

By contrast, the inner portion116of the annular dome111and outer portion118of the annular dome111, including the inner cowl130and outer cowl126, respectively, may be formed of a metal, such as a nickel-based superalloy (having a coefficient of thermal expansion of about 8.3-8.5×10−6in/in/° F. in a temperature of approximately 1000-1200° F.) or cobalt-based superalloy (having a coefficient of thermal expansion of about 7.8-8.1×10−6in/in/° F. in a temperature of approximately 1000-1200° F.). Thus, the inner and outer liners102,108may be better able to handle the extreme temperature environment presented in the combustion chamber114. However, attaching the inner and outer liners102,108to the respective inner and outer domes116,118presents a problem due to the differing mechanical characteristics of the components. Accordingly, as will be discussed below, a specially designed mounting assembly144is utilized to attach the forward end106of the inner liner102to the inner portion116of the annular dome111, as well as to attach the forward end112of the outer liner108to the outer portion118of the annular dome111. The mounting assemblies144are configured to accommodate the relative thermal expansion between the inner and outer domes116,118and the inner and outer liners102,108, respectively, along the radial direction.

Referring now particularly toFIG. 3, a close up, cross-sectional view of an attachment point where the forward end112of the outer liner108is attached to the outer annular dome118is depicted. As stated, to allow for a relative thermal expansion of the outer liner108and outer portion118of the annular dome111, the mounting assemblies144are provided extending through the annular slots122defined by the inner surface120between an outer arm200and an inner arm202. More particularly, referring specifically to the outer portion118of the annular dome111and forward end112of the outer liner108depicted inFIG. 3, the outer portion118of the annular dome111includes an outer arm200and an inner arm202that extend substantially parallel to one another, which for the embodiment depicted is a direction substantially parallel to the axial direction of the turbofan engine10.

For the embodiment depicted, the mounting assembly144includes a pin member166and an optional bushing168that extend through apertures201,203defined in the outer arm200and the inner arm202, respectively. The pin member166includes a head170and a nut174is attached to a distal end of the pin member166. In certain exemplary embodiments, the pin member166may be configured as a bolt and the nut174may be rotatably engaged with the pin member166for tightening the mounting assembly144. Alternatively, however, in other exemplary embodiments, the pen member166and nut174may have any other suitable configuration. For example, in other exemplary embodiments, the pin166may include a body172defining a substantially smooth cylindrical shape in the nut174may be configured as a clip. Additionally, the bushing168is generally cylindrical in shape and positioned around the pin member166.

Referring toFIG. 4, the forward end112of the outer liner108includes a plurality of fingers113. The fingers113are spaced apart from each other to define a slot109between adjacent fingers113. Thus, a plurality of slots109are defined annularly on the outer liner108. As show, each finger113defines a pair of longitudinal edges115. In the array of fingers113, at least a portion of oppositely facing longitudinal edges of adjacent fingers113have an indentation117therein to as to define an opening119for receipt of a pin member or bushing therethrough. That is, the indentations117on adjacent fingers113substantially align to receive the pin member168therethrough. The indentation117and the pin member166(or bushing166) can be sized so as to fit together such that the outer liner108is secured in place while allowing for some movement in the axial direction to account for differences in the thermal expansion discussed above.FIG. 5shows a similar embodiment where at least one finger113defines an opening119between the pair of longitudinal edges115(i.e., within the body of the finger113) for receipt of a pin member or bushing therethrough. Of course, features from bothFIGS. 4 and 5may be combined, if desired.

Referring again toFIG. 3, a terminal end112of each finger113extends into the annular slot122and can form a gap123between an inner surface120of the annular slot122of the annular dome118and the terminal end112of each finger113. In particular embodiments, the outer arm200of the annular slot122of the annular dome118defines a slot length (L), and wherein the gap123defined from the inner surface120of the annular slot122of the annular dome118to the terminal end112of each finger112has a length of about 1% to about 25% of the slot length (L) at room temperature (i.e., about 25° C.), such as about 1% to about 10%. In other embodiments, the terminal end112of each finger113can contact the inner surface120of the annular slot122of the annular dome118.

FIG. 3also shows a feather seal210extending from an annularly exterior surface209of the annular dome118to an annularly exterior surface219of the outer liner108. The feather seal210is, in the embodiment shown, in a spring loaded contact with the annularly exterior surface209of the outer liner108. In one embodiment, the feather seal210comprises a metal with a wear coating thereon such that the wear coating contacts the annularly exterior surface209of the outer liner108. The feather seal210generally forms a fluid-tight barrier between the internal combustion chamber114and the space external of the inner liner102and outer liner108, and inhibits the flow of gas therethrough.

In particular embodiments, the outer liner108defines a tapered portion211. That is, the outer liner108has a thickness in its body portion213that is greater than the thickness of the fingers113and/or at its forward end112. In the embodiment shown inFIG. 3, the annularly exterior surface209defines a taper211. However, in other embodiments, the tapered surface can be on the annularly inner surface opposite of the annularly exterior surface209.

Each pin member166extends through an aperture in the feather seal211, through an aperture in the outer arm200of the annular dome118, through an axial slot109in the outer liner108, and through an aperture in the inner arm202of the annular dome118to secure the components together. The number of pin members166annularly securing the outer annular dome118may be the same as the number of slots109(i.e., one pin member166extending through each slot109); may be less than the number of slots109; or more than the number of slots109. That is, the plurality of axial slots109can be greater in number than the plurality of pin members116, to allow for radial expansion and contraction of the outer liner108in certain embodiments. However, in other embodiments, the plurality of axial slots109can be lesser in number than the plurality of pin members116(e.g., when using wider and/or longer fingers, more than 1 pin member166may be utilized per finger).

A combustor in accordance with an exemplary embodiment of the present disclosure assembly having a cap positioned over an inner liner or an outer liner may be capable of controlling an airflow from a relatively high pressure plenum or a relatively high pressure inner passage into a combustion chamber through an attachment point between the inner or outer liners and an inner or outer dome. Moreover, such a combustor assembly may be capable of controlling an airflow from a relatively high pressure plenum or a relatively high pressure inner passage into a combustion chamber through an attachment point between the inner or outer liners and an inner or outer dome while still accommodating a relative thermal expansion between the inner or outer liners and inner or outer domes.