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EP-4283096-B1 - CMC VANE WITH FLANGE HAVING SLOPED RADIAL FACE

EP4283096B1EP 4283096 B1EP4283096 B1EP 4283096B1EP-4283096-B1

Inventors

  • Kim, Russell
  • LILES, HOWARD J.

Dates

Publication Date
20260513
Application Date
20230530

Claims (7)

  1. A vane arc segment (60) comprising: a ceramic matrix composite (CMC) fairing (62) having first and second platforms (64, 66) and an airfoil section (68) extending in a radial direction (RD) there between, each of the first and second platforms (64, 66) including axially-facing leading and trailing sides (70c, 70d), a core gaspath side (70a), and a non-core gaspath side (70b), the non-core gaspath side (70b) of the first platform (64) having a flange (72) projecting radially there from and extending adjacent the trailing side (70d), the flange (72) having a radial face (72a) that is sloped with respect to the radial direction (RD); characterised in that : the CMC fairing (62) is comprised of fiber plies (74) disposed in a ceramic matrix, and two fiber plies (74) are turned-up back-to-back to form the flange (72), and each of the turned-up fiber plies (74) has a terminal end face (74a) that forms a portion of the radial face (72a).
  2. The vane arc segment (60) as recited in claim 1, wherein the radial face (72a) is a frustoconic arc segment.
  3. The vane arc segment (60) as recited in claim 1 or 2, wherein the radial face (72a) defines a reference surface (RS) that, when infinitely extended, is non-intersecting with the airfoil section (68).
  4. The vane arc segment (60) as recited in claim 1, 2 or 3, wherein the flange (72) is elongated in a circumferential direction (CD).
  5. The vane arc segment (60) as recited in any preceding claim, wherein the flange (72) is a lone flange on the non-core gaspath side (70b) of the first platform (64).
  6. The vane arc segment (60) as recited in any preceding claim, wherein the radial face (72a) is sloped at an angle of 30 degrees to 60 degrees relative to the radial direction (RD).
  7. A gas turbine engine (20) comprising: the vane arc segment (60) of any preceding claim; and first and second supports (61a, 61b) radially between which the vane arc segment (60) is held, the first support (61a) supporting the vane arc segment (60) at the radial face (72a), and the second support (61b) supporting the vane arc segment (60) via the second platform (66).

Description

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. EP 3683405 A1 discloses a prior art vane arc segment according to the preamble of claim 1. SUMMARY A vane arc segment according to an aspect of the present invention is provided according to claim 1. In a further embodiment of the foregoing embodiment, 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. The present disclosure may include any one or more of the individual features disclosed above and/or below alone or in any combination thereof. BRIEF DESCRIPTION OF THE DRAWINGS The various features and advantages of the present disclosure will become apparent to those skilled in the art from the following detailed description. The drawings that accompany the detailed description can be briefly described as follows. Figure 1 illustrates a gas turbine engine.Figure 2 illustrates a portion of the turbine section of the engine.Figure 3 illustrates a CMC fairing of the turbine section.Figure 4 illustrates a flange of the CMC fairing.Figure 5A illustrates an example fiber ply configuration of the flange.Figure 5B illustrates another example fiber ply configuration of the flange.Figure 5C illustrates another example fiber ply configuration of the flange.Figure 5D illustrates another example fiber ply configuration of the flange. 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 Figure 1 schematically illustrates a gas turbine engine 20. The gas turbine engine 20 is disclosed herein as a two-spool turbofan that generally incorporates a fan section 22, a compressor section 24, a combustor section 26 and a turbine section 28. The fan section 22 drives air along a bypass flow path B in a bypass duct defined within a housing 15 such as a fan case or nacelle, and also drives air along a core flow path C for compression and communication into the combustor section 26 then expansion through the turbine section 28. 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 three-spool architectures. The exemplary engine 20 generally includes a low speed spool 30 and a high speed spool 32 mounted for rotation about an engine central longitudinal axis A relative to an engine static structure 36 via several bearing systems 38. It should be understood that various bearing systems 38 at various locations may alternatively or additionally be provided, and the location of bearing systems 38 may be varied as appropriate to the application. The low speed spool 30 generally includes an inner shaft 40 that interconnects, a first (or low) pressure compressor 44 and a first (or low) pressure turbine 46. The inner shaft 40 is connected to the fan 42 through a speed change mechanism, which in exemplary gas turbine engine 20 is illustrated as a geared architecture 48 to drive a fan 42 at a lower speed than the low speed spool 30. The high speed spool 32 includes an outer shaft 50 that interconnects a second (or high) pressure compressor 52 and a second (or high) pressure turbine 54. A com