Patent Description:
The compressor section can include rotors that carry airfoils to compress the air entering the compressor section. A shaft may be coupled to the rotors to rotate the airfoils.

<CIT> describes a vane arc segment that includes radially inner and outer platforms and an airfoil mechanically clamped between the platforms. The airfoil has an airfoil section that extends radially between radially inner and outer fairing platforms. At least one of the fairing platforms includes forward and aft sides, circumferential sides, and a gas path side and an opposed radial side. The radial side includes a plurality of protrusions that have faces that are oriented substantially normal to, respectively, radial, tangential, and axial load transmission directions of the airfoil such that the faces, respectively, primarily bear radial, tangential, and axial load transmissions of the airfoil.

According to a first aspect of the present invention, there is provided a vane arc segment for a gas turbine engine comprising a continuous airfoil piece that defines first and second platforms and an airfoil section that extends between the first and second platforms. The airfoil section has a pressure side and a suction side. The first platform defines axially-sloped suction and pressure side circumferential mate faces, first and second axial sides, a gaspath side, a non-gaspath side, and a flange that projects from the non-gaspath side. The flange extends along at least a portion the axially-sloped suction side circumferential mate face. The axially-sloped suction and pressure side circumferential mate faces form an angle of at least <NUM>° relative to an axial direction of the vane arc segment. The axial direction is a central longitudinal axis of the gas turbine engine.

In an embodiment of any of the above embodiments, the airfoil piece is formed of a ceramic matrix composite.

In an embodiment of any of the above embodiments, the ceramic matrix composite has a plurality of fiber plies that are continuous through the flange and first platform, the airfoil section, and the second platform.

In an embodiment of any of the above embodiments, the flange is coextensive with the axially-sloped suction side circumferential mate face.

In an embodiment of any of the above embodiments, the flange is flush with the axially-sloped suction side circumferential mate face.

In an embodiment of any of the above embodiments, the flange includes a radial face that has a curved profile.

In an embodiment of any of the above embodiments, the curved profile is a cylindrical surface segment.

In an embodiment of any of the above embodiments, the flange is elongated along a flange axis, and the flange axis is substantially perpendicular to a total aerodynamic load vector of the airfoil piece.

In an embodiment of any of the above embodiments, one of the faces of the flange is a radial face, the radial face has a curved profile, and the curved profile is a cylindrical surface segment.

Although depicted as a two-spool turbofan gas turbine engine in the disclosed non-limiting embodiment, it should be understood that the concepts described herein are not limited to use with two-spool turbofans as the teachings may be applied to other types of turbine engines including but not limited to three-spool architectures.

Terms such as "axial," "radial," "circumferential," and variations of these terms are made with reference to the engine central axis A.

In a further example, the engine <NUM> bypass ratio is greater than about <NUM>:<NUM>, with an example embodiment being greater than about <NUM>:<NUM>, the geared architecture <NUM> is an epicyclic gear train, such as a planetary gear system or other gear system, with a gear reduction ratio of greater than about <NUM>:<NUM> and the low pressure turbine <NUM> has a pressure ratio that is greater than about <NUM>:<NUM>. In one disclosed embodiment, the engine <NUM> bypass ratio is greater than about <NUM>:<NUM>, the fan diameter is significantly larger than that of the low pressure compressor <NUM>, and the low pressure turbine <NUM> has a pressure ratio that is greater than about <NUM>:<NUM>. The low pressure turbine <NUM> pressure ratio is pressure measured prior to the inlet of low pressure turbine <NUM> as related to the pressure at the outlet of the low pressure turbine <NUM> prior to an exhaust nozzle. 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 but not limited to direct drive turbofans.

The low fan pressure ratio as disclosed herein according to one non-limiting embodiment is less than about <NUM>:<NUM>.

<FIG> illustrates a representative portion of a vane ring assembly from the turbine section <NUM> of the engine <NUM>. The vane ring assembly is made up of a plurality of vane arc segments <NUM> that are situated in a circumferential row about the engine central axis A. <FIG> illustrate isolated views from different angles of a representative one of the vane arc segments <NUM>. Although the vane arc segments <NUM> are shown and described with reference to application in the turbine section <NUM>, it is to be understood that the examples herein are also applicable to structural vanes in other sections of the engine <NUM>.

The vane arc segment <NUM> is comprised of an airfoil piece <NUM>. The airfoil piece <NUM> includes several sections, including first and second platforms <NUM>/<NUM> and an airfoil section <NUM> that extends between the first and second platforms <NUM>/<NUM>. The airfoil section <NUM> defines a leading end 66a, a trailing end 66b, and pressure and suction sides 66c/66d.

In this example, the first platform <NUM> is a radially outer platform and the second platform <NUM> is a radially inner platform. The first platform <NUM> defines axially-sloped suction and pressure side circumferential mate faces 63a/63b, first and second axial sides 63c/63d, a gaspath side 63e, and a non-gaspath side 63f. Likewise, the second platform <NUM> defines axially-sloped suction and pressure side circumferential mate faces 64a/64b, first and second axial sides 64c/64d, a gaspath side 64e, and a non-gaspath side 64f. The first axial sides 63c/64c are axially forward sides, and the second axial sides 63d/64d are axially trailing sides.

The mate faces 63a/63b/64a/64b are sloped with respect to the central engine axis A. As an example, the mate faces 63a/63b/64a/64b form angles, represented at <NUM>, with the axis A, and the angles are at least <NUM>°. In some examples, where the mate faces 63a/63b/64a/64b are straight, it is the planes of the mate faces 63a/63b/64a/64b that form the angle <NUM>. In other examples, where the mate faces 63a/63b/64a/64b are curvilinear, the angle <NUM> may be represented using the line between the forward and trailing corners of the mate face 63a/63b/64a/64b.

The first platform <NUM> further includes a flange <NUM> that projects (radially) from the non-gaspath side 63f. The flange <NUM> is generally elongated and runs along at least a portion of the extent of the circumferential mate face 63a of the first platform <NUM>. In the illustrated example, the flange <NUM> is co-extensive with the mate face 63a, although in modified examples the flange <NUM> may be somewhat shorter than the mate face 63a. Most typically, the flange <NUM> runs along at least <NUM>%, at least <NUM>%, or at least <NUM>% of the extent of the mate face 63a.

The flange <NUM> defines a radial face 68a, an outer side face 68b, an inner side face 68c, a forward face 68d, and a trailing face 68e. Although the position of the flange <NUM> could vary somewhat, in the illustrated example the outer side face 68b is flush with the mate face 63a, and the trailing face 68e is flush with the axial side 63d.

The radial face 68a may have planar profile or a curved profile. In the illustrated example, the radial face 68a has a curved profile, represented at 68f. As will be described further below, the profile 68f may be used to facilitate load transmission. In the illustrated example, the curved profile 68f is that of a cylindrical surface segment. A cylindrical surface segment is a surface, here the radial face 68a, that has the shape of a section of a surface of a cylinder. For example, as shown in <FIG>, which is an axial view looking forward, the cylindrical surface segment of the profile 68f is that of a reference cylinder (RS) that has its main axis parallel to the central engine axis A. For instance, the main axis of the reference cylinder is co-linear with the central engine axis A.

The airfoil piece <NUM> is continuous in that the platforms <NUM>/<NUM> and airfoil section <NUM> constitute a single, uninterrupted body. As an example, the airfoil piece may be formed of a ceramic material, such as a ceramic matrix composite. In the illustrated example, referring to cutaway section B in <FIG>, the airfoil piece <NUM> is formed of a ceramic matrix composite in which ceramic fibers 70a are disposed in a ceramic matrix 70b. The ceramic matrix composite may be, but is not limited to, a SiC/SiC ceramic matrix composite in which SiC fibers are disposed within a SiC matrix. The ceramic fibers 70a may be provided in fiber plies, represented at 70c. The plies 70c may be woven or unidirectional and may collectively include plies of different fiber weave configurations. The fiber plies 70c may be continuous through the first platform <NUM>, including the flange <NUM>, the airfoil section <NUM>, and the second platform <NUM>. In one example, the fiber plies 70c are laid-up in a laminate configuration. The use of the flange <NUM> is primarily directed to ceramic airfoil pieces <NUM>, however, it is to be appreciated that this disclosure may also be applicable to certain stress-limited metallic alloys, such as refractory metallic alloys (e.g., molybdenum-based alloys) or other alloys that have high temperature resistance but low strength.

In general, support schemes for mounting structural vane segments formed of ceramic matrix composites or stress-limited metallic alloys are challenging due to lower material stress limits in comparison to high strength superalloys used for some traditional vane segments. For instance, traditional support schemes that utilize hooks or a series of rails can concentrate stresses, create aerodynamic loads, and/or create thermal stresses which may exceed material limits of ceramic matrix composites or stress-limited metallic alloys. Moreover, traditional hooks and rails often have complex geometries that are challenging to manufacture from ceramic matrix composites. Therefore, even though ceramic matrix composites or stress-limited metallic alloys may have many potential benefits, such benefits cannot be realized without a suitable support scheme. In this regard, as will be discussed, the vane arc segment <NUM> is designed to facilitate a low-stress mounting scheme.

The vane arc segment <NUM> may be mounted in the engine <NUM> between inner and outer support structures. The support structures are not particularly limited and may be cases, intermediate carriers, or the like, and are typically formed of metallic alloys that can bear the loads received. During operation of the engine <NUM> combustion gases flow across the airfoil section <NUM> and gaspath sides 63e/64e of the platforms <NUM>/<NUM>. The flow causes aerodynamic loads on the vane arc segment <NUM>. The aerodynamic loads are transmitted through the vane arc segment <NUM> to the support structure. In this regard, the flange <NUM> serves as the primary load-bearing feature to transmit such loads.

The aerodynamic loads may be summed as a bulk, or total, aerodynamic load vector. For instance, the aerodynamic load vector may be optimized or design for a maximum pressure condition and may be understood as the force applied at the center of pressure on the vane arc segment <NUM>. As shown in <FIG>, the total aerodynamic load vector is represented at <NUM>. In this case, the total aerodynamic load vector <NUM> intersects the flange <NUM> and, in this example, is orthogonal within +/- <NUM>° to the length direction of the flange <NUM>. The orthogonal relationship between the total aerodynamic load vector <NUM> and the direction of the flange <NUM> facilitates establishing an efficient load path for transmission of the loads from the vane arc segment <NUM>, which minimizes pressure-driven stresses on the arc vane segment <NUM> without imparting a twist between the airfoil section <NUM> and the first platform <NUM>. By comparison, a less efficient load path would have twist and, therefore, a longer, less direct load path.

The orthogonal orientation of the flange <NUM> is governed by the desired orientation relative to the total aerodynamic load vector <NUM>. The slope angles of the circumferential mate faces 63a/63b/64a/64b of the platforms <NUM>/<NUM> are then designed to the orientation of the flange <NUM>. Here, with the orthogonal orientation of the flange <NUM>, the slope of the circumferential mate face 63a is highly angled at slope angle of <NUM>° or more.

The flange <NUM> transmits tangential, axial, and radial loads. For example, at least the faces 68a, 68b, and 68e of the flange <NUM> are in contact with a corresponding portion of the support structure to transmit loads to. Tangential and axial loads are transmitted through the outer side face 68b and axial face 68e, as well as the axial side 63d of the platform <NUM>. The radial loads are transmitted through the radial face 68a. For example, all radial loads from the total aerodynamic load vector <NUM> that are transmitted through the radially outer diameter of the vane arc segment <NUM> are transmitted through the radial face 68a (radial loads may also be transmitted through the radially inner diameter at the second platform <NUM>). In this regard, the afore-mentioned curved profile 68f of the radial face 68a may be used to facilitate a more uniform load transfer. For instance, the radial face 68a may mate with a complimentary face on the support structure. The total aerodynamic load vector <NUM> may cause the vane arc segment <NUM> to tilt slightly in the direction of the total aerodynamic load vector <NUM>. The profile 68f enables the radial face 68a to be at a single radius and thereby serve as a true radial face to uniformly transmit radial loads, to the complimentary face on the support structure.

The flange <NUM> also has a geometry that can be readily manufactured from CMC or from stress-limited alloys. For instance, the flange <NUM> is generally straight and does not contain hooks or high-radius turns. For CMCs, the flange <NUM> is also of relatively low-profile and may therefore also facilitate lowering thermal gradients. For instance, the lack of hooks and high-radius enables the flange <NUM> to be formed from a layered CMC composite by simply bending the layers of the CMC of the platform radially outwards, which also enables the platform to be thinner than it otherwise would need to be if more complex, larger reaction features were used. The thinness may facilitate reducing thermal gradients because there is less insulating CMC material. As a result, without the disclosed geometry of the flange <NUM>, thicker features would be required to react out the aerodynamic loads and there would be higher thermal gradients due to the thickness. Additionally, without hooks or the like, the mounting of the vane arc segment <NUM> is relatively simple.

Claim 1:
A vane arc segment (<NUM>) for a gas turbine engine comprising:
a continuous airfoil piece (<NUM>) defining first and second platforms (<NUM>, <NUM>) and an airfoil section (<NUM>) extending between the first and second platforms (<NUM>, <NUM>),
the airfoil section (<NUM>) having a pressure side and a suction side (66c, 66d),
the first platform (<NUM>) defining axially-sloped suction and pressure side circumferential mate faces (63a/63b), first and second axial sides (63c, 63d), a gaspath side (63e), a non-gaspath side (63f), and a flange (<NUM>) projecting from the non-gaspath side (63f),
the flange (<NUM>) extending along at least a portion the axially-sloped suction side circumferential mate face (63a);
characterized in that
the axially-sloped suction and pressure side circumferential mate faces (63a/63b) form an angle (<NUM>) of at least <NUM>° relative to an axial direction (A), wherein the axial direction (A) is a central longitudinal axis of the gas turbine engine .