CMC vane with flange having sloped radial face

A vane arc segment includes a ceramic matrix composite fairing that has first and second platforms and an airfoil section that extends in a radial direction there between. Each of the first and second platforms includes axially-facing leading and trailing sides, a core gaspath side, and a non-core gaspath side. The non-core gaspath side of the first platform has a flange that projects radially there from and extends adjacent the trailing side. The flange has a radial face that is sloped with respect to the radial direction.

BACKGROUND

A gas turbine engine typically includes 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-pressure and temperature exhaust gas flow. The high-pressure and temperature exhaust gas flow expands through the turbine section to drive the compressor and the fan section. The compressor section may include low and high pressure compressors, and the turbine section may also include low and high pressure turbines.

Airfoils in the turbine section are typically formed of a superalloy and may include thermal barrier coatings to extend temperature capability and lifetime. Ceramic matrix composite (“CMC”) materials are also being considered for airfoils. Among other attractive properties, CMCs have high temperature resistance. Despite this attribute, however, there are unique challenges to implementing CMCs in airfoils.

SUMMARY

A vane arc segment according to an example of the present disclosure includes a ceramic matrix composite (CMC) fairing that has first and second platforms and an airfoil section that extends in a radial direction there between. Each of the first and second platforms includes axially-facing leading and trailing sides, a core gaspath side, and a non-core gaspath side. The non-core gaspath side of the first platform has a flange that projects radially there from and extends adjacent the trailing side. The flange has a radial face that is sloped with respect to the radial direction.

In a further embodiment of any of the foregoing embodiments, the radial face is a frustoconic arc segment.

In a further embodiment of any of the foregoing embodiments, the radial face defines a reference surface that, when infinitely extended, is non-intersecting with the airfoil section

In a further embodiment of any of the foregoing embodiments, the flange is elongated in a circumferential direction.

In a further embodiment of any of the foregoing embodiments, the flange is a lone flange on the non-core gaspath side of the first platform.

In a further embodiment of any of the foregoing embodiments, the radial face is sloped at an angle of 30 degrees to 60 degrees relative to the radial direction.

In a further embodiment of any of the foregoing embodiments, the CMC fairing is comprised of fiber plies disposed in a ceramic matrix, and one of the fiber plies in the first platform is turned up to form at least a portion of the flange.

In a further embodiment of any of the foregoing embodiments, the flange includes a radial stack of the fiber plies adjacent to the one of the fiber plies that is turned up.

In a further embodiment of any of the foregoing embodiments, the one of the fiber plies that is turned up has a terminal end face that forms a portion of the radial face.

In a further embodiment of any of the foregoing embodiments, the CMC fairing is comprised of fiber plies disposed in a ceramic matrix, and multiple ones of the fiber plies in the first platform are turned up to form the flange.

In a further embodiment of any of the foregoing embodiments, each of the fiber plies that are turned up has a terminal end face that forms a portion of the radial face.

A gas turbine engine according to an example of the present disclosure includes a vane arc segment including a ceramic matrix composite (CMC) fairing that has first and second platforms and an airfoil section that extends in a radial direction there between. Each of the first and second platforms includes axially-facing leading and trailing sides, a core gaspath side, and a non-core gaspath side. The non-core gaspath side of the first platform has a flange that projects radially there from and extends adjacent the trailing side. The flange has a radial face that is sloped with respect to the radial direction. First and second supports radially between which the vane arc segment is held, the first support supporting the vane arc segment at the radial face, and the second support supporting the vane arc segment via the second platform.

In a further embodiment of any of the foregoing embodiments, the radial face is a frustoconic arc segment.

In a further embodiment of any of the foregoing embodiments, the radial face defines a reference surface that, when infinitely extended, is non-intersecting with the airfoil section

In a further embodiment of any of the foregoing embodiments, the flange is elongated in a circumferential direction, and the flange is a lone flange on the non-core gaspath side of the first platform.

In a further embodiment of any of the foregoing embodiments, the CMC fairing is comprised of fiber plies disposed in a ceramic matrix, and one of the fiber plies in the first platform is turned up to form at least a portion of the flange.

In a further embodiment of any of the foregoing embodiments, the flange includes a radial stack of the fiber plies adjacent to the one of the fiber plies that is turned up.

In a further embodiment of any of the foregoing embodiments, the one of the fiber plies that is turned up has a terminal end face that forms a portion of the radial face.

In a further embodiment of any of the foregoing embodiments, the CMC fairing is comprised of fiber plies disposed in a ceramic matrix, and multiple ones of the fiber plies in the first platform are turned up to form the flange.

In a further embodiment of any of the foregoing embodiments, each of the fiber plies that are turned up has a terminal end face that forms a portion of the radial face.

In this disclosure, like reference numerals designate like elements where appropriate and reference numerals with the addition of one-hundred or multiples thereof designate modified elements that are understood to incorporate the same features and benefits of the corresponding elements.

DETAILED DESCRIPTION

The engine20in one example is a high-bypass geared aircraft engine. In a further example, the engine20bypass ratio is greater than about six (6), with an example embodiment being greater than about ten (10), and can be less than or equal to about 18.0, or more narrowly can be less than or equal to 16.0. The geared architecture48is an epicyclic gear train, such as a planetary gear system or other gear system, with a gear reduction ratio of greater than about 2.3. The gear reduction ratio may be less than or equal to 4.0. The low pressure turbine46has a pressure ratio that is greater than about five. The low pressure turbine pressure ratio can be less than or equal to 13.0, or more narrowly less than or equal to 12.0. In one disclosed embodiment, the engine20bypass ratio is greater than about ten (10:1), the fan diameter is significantly larger than that of the low pressure compressor44, and the low pressure turbine46has a pressure ratio that is greater than about five 5:1. Low pressure turbine46pressure ratio is pressure measured prior to an inlet of low pressure turbine46as related to the pressure at the outlet of the low pressure turbine46prior to an exhaust nozzle. The geared architecture48may be an epicycle gear train, such as a planetary gear system or other gear system, with a gear reduction ratio of greater than about 2.3:1 and less than about 5:1. It should be understood, however, that the above parameters are only exemplary of one embodiment of a geared architecture engine and that the present invention is applicable to other gas turbine engines including direct drive turbofans.

A significant amount of thrust is provided by the bypass flow B due to the high bypass ratio. The fan section22of the engine20is designed for a particular flight condition—typically cruise at about 0.8 Mach and about 35,000 feet (10,668 meters). The flight condition of 0.8 Mach and 35,000 ft (10,668 meters), with the engine at its best fuel consumption—also known as “bucket cruise Thrust Specific Fuel Consumption (‘TSFC’)”—is the industry standard parameter of lbm of fuel being burned divided by lbf of thrust the engine produces at that minimum point. The engine parameters described above and those in this paragraph are measured at this condition unless otherwise specified. “Low fan pressure ratio” is the pressure ratio across the fan blade alone, without a Fan Exit Guide Vane (“FEGV”) system. The low fan pressure ratio as disclosed herein according to one non-limiting embodiment is less than about 1.45, or more narrowly greater than or equal to 1.25. “Low corrected fan tip speed” is the actual fan tip speed in ft/sec divided by an industry standard temperature correction of [(Tram ° R)/(518.7° R)]0.5. The “Low corrected fan tip speed” as disclosed herein according to one non-limiting embodiment is less than about 1150.0 ft/second (350.5 meters/second), and can be greater than or equal to 1000.0 ft/second (304.8 meters/second).

FIG.2illustrates an axial view of an example of a portion of the turbine section28of the engine20. The turbine section28includes vane arc segments60radially disposed in an annulus defined between first (outer) and second (inner) supports61a/61bthat support the vane arc segments60. The supports61a/61bare shown schematically, but each one may be a continuous full hoop ring, an engine case, one or more intermediate structures that attach to an engine case, a series of spars, or other static structure or structures in the engine20.

FIG.3illustrates a representative sectioned view through one of the vane arc segments60(only a portion of the first support61ais shown). Each vane arc segment60is comprised of a ceramic matrix composite (CMC) fairing62. The CMC fairing62has several sections, including first (outer) and second (inner) platforms64/66and an airfoil section68that extends between the platforms64/66. The airfoil section68has that has an internal through-cavity68afor conveying cooling air, such as bleed air from the compressor section24. The platforms64/66provide radially outer and inner bounds of the core gas path C. Each of the platforms64/66defines a first radial side70a(core gas path side), an opposed second radial side70b(non-core gas path side), an axially forward-facing leading side70c, and an axially aft-facing trailing side70d, as well as circumferential sides (not shown).

The terms such as “inner” and “outer” refer to location with respect to the central engine axis A, i.e., radially inner or radially outer. Moreover, the terminology “first” and “second” as used herein is to differentiate that there are two architecturally distinct components or features. It is to be further understood that the terms “first” and “second” are interchangeable in the embodiments herein in that a first component or feature could alternatively be termed as the second component or feature, and vice versa.

The CMC material from which the fairings62are made is comprised of a ceramic reinforcement, which is usually ceramic fibers, in a ceramic matrix. Example ceramic matrices of the CMC are silicon-containing ceramic, such as but not limited to, a silicon carbide (SiC) matrix or a silicon nitride (Si3N4) matrix. Example ceramic reinforcement of the CMC are silicon-containing ceramic fibers, such as but not limited to, silicon carbide (SiC) fiber or silicon nitride (Si3N4) fibers. The ceramic matrix composite may be, but is not limited to, a SiC/SiC ceramic matrix composite in which SiC fiber plies are disposed within a SiC matrix. The fiber plies have a fiber architecture, which refers to an ordered arrangement of the fiber tows relative to one another, such as a 2D woven ply or a 3D structure. The CMC fairings62may be one-piece structures in which at least a portion of the fiber plies are continuous from the first platform64, through the airfoil section68, and into the second platform66.

Structural vane arc segments require axial, radial, and circumferential constraints to inhibit motion when loaded by gas path and/or secondary flow forces. Attachment of CMC fairings in an engine, however, is challenging. Attachment features, such as hooks, that are typically used for metal alloy vanes can result in inefficient loading if employed in CMCs, which may be sensitive to stress directionality and distress conditions that differ from those of metal alloy components. Additionally, features such as feather seal slots, variable thickness walls, buttresses, gussets, weldments, complex-geometry investment casting cores, bare machined surfaces, etc. that may be used in metal alloy components are generally not acceptable or attainable with CMC materials.

To facilitate addressing the considerations above the first platform64of the includes a flange72that projects radially from the non-core gaspath side70b. In this example, the flange72is a lone, exclusive flange of the first platform64in that it is the only radial projection of the platform64on non-core gaspath side70b. As shown, the flange72is a radially upstanding wall that is located adjacent the axially-facing trailing side70d. As shown inFIG.4, the flange72is generally elongated in the circumferential direction CD. For instance, the flange72fully spans or substantially fully spans the circumferential sides of the platform64.

The flange72has a radial face72athat is sloped with respect to the radial direction RD. The radial face72aarcs about the axis A, but in cross-section taken along a radial plane can be flat, concave toward the axis A, or convex away from the axis A. For example, the radial face72ais sloped at an angle AG of 30 degrees to 60 degrees relative to the radial direction. If the radial face72ais concave or convex, the angle AG is taken relative to a reference line that intersects the forward-most and aft-most points of the radial face72a. In a further example, the angle AG is 40 degrees to 50 degrees.

The CMC fairing62is supported via the flange72and, in particular, via interface with the radial face72a. The slope of the radial face72afocuses load transmission along a line of actionLOAthat extends across the airfoil section68to the forward portion of the platform66. In terms of the orientation of the radial face72a, this means that the radial face72afaces in aft radially outwardly direction such that it defines a reference surface RS that, when infinitely extended, is non-intersecting with the airfoil section68.

The locations at which the loads are primarily borne are indicated inFIG.3with block arrows. The slope of the radial face72aproduces the line of actionLOA, which in turn causes the airfoil fairing62to be loaded in compression. As CMC material is relatively strong in compression, as compared to loading in tension, this loading scheme provided by the radial face72aproduces a favorable stress state for the CMC fairing62.

In a further example, to facilitate load distribution, the radial face72ahas the geometry of a frustoconic arc segment. A conic frustum is a truncated cone and the outer surface of the frustum is frustoconical. If the conic frustum is sectioned off along a plane that is non-intersecting with the central axis of the frustum, the shape of the frustoconical surface sectioned off is a frustoconic arc segment, which could also be referred to as a frustic surface.

As mentioned above, the CMC fairing62is made of fiber plies disposed in a ceramic matrix.FIGS.5A,5B,5C, and5Dshow example ply configurations of the fiber plies74in the platform64and flange72, although the CMC fairing62is not limited to these. In each of the examples, the radial face72amay be formed after formation of the flange72by machining the flange72. Alternatively, the radial face72amay be formed in full or in part during a ply lay-up process by cutting prepreg fiber plies.

FIG.5Adepicts a double-J configuration in which two fiber plies74are turned-up back-to-back to form the flange72. Each of the turned-up fiber plies has a terminal end face74athat forms a portion of the radial face72a. In this configuration, the loads in the flange72are primarily borne along the in-plane direction of the turned-up fiber plies74, i.e., the fiber plies are compressed along their length.

FIG.5Bdepicts a stacked configuration in which a stack74bof fiber plies74forms the flange72. One or more of the fiber plies74in the stack74bhas a terminal end face74cthat forms a portion of the radial face72a. In this configuration, the loads in the flange72are primarily borne along the out-of-plane direction of the stacked fiber plies74, i.e., the stack74bis compressed, although there may also be some shear along the fiber ply interfaces.

FIG.5Cdepicts a single-J stacked configuration in which one of the fiber plies74is upturned and backs against a stack74bof the fiber plies74. The terminal end face74aof the upturned fiber ply74forms a portion of the radial face72a, and the terminal end face74cof one of the stacked fiber plies74forms another portion of the radial face72a. In this configuration, the loads in the flange72are primarily borne along the in-plane direction in the upturned fiber ply74and the out-of-plane direction of the stacked fiber plies74, i.e., the upturned fiber ply74is compressed along its length and the stack74bis compressed, although there may also be some shear along the fiber ply interfaces.

FIG.5Dis a reverse single-J stacked configuration. It is the same as the configuration inFIG.5Cexcept that the position of the stack74band the upturned fiber ply74are swapped. In this case, it is only the terminal end faces74cof the fiber plies74of the stack74bthat form the radial face72a. However, depending on the angle of the radial face72aand thickness of the fiber plies74it is also possible that the radial face72aintersects the upturned fiber ply74such that its terminal end face74aforms a portion of the radial face72a.

As discussed above, the radial face72afacilitates a favorable loading scheme for the CMC fairing62. Furthermore, the flange72is readily manufacturable via ply lay-up and machining, if used, is minimal. Additionally, the use of the flange72permits the walls of the platform64to be entirely substantially uniform in thickness. For instance, if a sloped face were to be formed directly into the platform, the platform would need to be made thicker to provide volume to machine away to form the sloped face. As a result, the platform would be substantially thicker in some areas than others, which could exacerbate thermal gradients and may be difficult to manufacture with a uniform matrix density.