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EP-4549701-B1 - COMBINED GAS TURBINE ENGINE VANE AND BLADE OUTER AIR SEAL ASSEMBLY AND GAS TURBINE ENGINE

EP4549701B1EP 4549701 B1EP4549701 B1EP 4549701B1EP-4549701-B1

Inventors

  • ROACH, JAMES T.
  • BANHOS, Jonas
  • Kim, Russell

Dates

Publication Date
20260513
Application Date
20241028

Claims (8)

  1. A combined gas turbine engine vane and blade outer air seal assembly comprising: a vane (122, 222, 322, 422) having an airfoil (127) extending from a leading edge (124) to a trailing edge (108), and having an outer platform (105), the outer platform (105) having a cooling channel (112) that extends into the airfoil (127) to receive cooling air; the outer platform (105) extending to an integral blade outer air seal (130) to be positioned radially outwardly of a turbine blade (102) in a gas turbine engine (20); and at least a portion of said vane (122, 222, 322, 422) and said blade outer air seal (130) formed of ceramic matrix composite materials, characterised in that : the outer platform (105) and the blade outer air seal (130) are formed of a plurality of laminae of ceramic matrix composite materials; and there are cooling air passages (232) within said outer platform (105) and said blade outer air seal (130) and connected into the cooling channel (112) such that air can be communicated from the cooling channel (112) into said blade outer air seal (130).
  2. The assembly as set forth in claim 1, wherein at least some of the cooling air passages (232) extend from the cooling channel (112) to at least one outlet (334) at an aft end spaced from said trailing edge (108) of the airfoil (127).
  3. The assembly as set forth in claim 1 or 2, wherein the cooling air passages (232) extend from the cooling channel (112) to at least one outlet (234) extending radially inwardly of the blade outer air seal (130).
  4. The assembly as set forth in any preceding claim, wherein the blade outer air seal (130) has a radially outwardly extending portion (433) with the cooling air passages (232) extending through the radially outwardly extending portion (433) to at least one outlet (334) at a radially outward end of the blade outer air seal (130).
  5. The assembly as set forth in any preceding claim, wherein the cooling air passages (232) extend to an or the radially outwardly extending portion (433) and then into outlets (434) extending in an axially aft direction.
  6. The assembly as set forth in claim 5, wherein there is an enclosed portion of the blade outer air seal (130) radially outward of the outlets (434) extending in an axially aft direction.
  7. The assembly as set forth in any preceding claim, wherein an attachment structure (150, 152, 154) is attached to the blade outer air seal (130) portion on a radially outward side.
  8. A gas turbine engine comprising a compressor section (24), a combustor section (26), and a turbine section (28), the turbine section (28) including the combined vane and blade outer air seal assembly of any preceding claim, wherein the integral blade outer air seal (130) is positioned radially outwardly of a turbine blade (102) in the turbine section (28).

Description

BACKGROUND OF THE INVENTION This invention relates to the technical field of gas turbine engine vanes that have a radially outer platform which includes an integrated blade outer air seal. More precisely, the invention relates to a combined gas turbine engine vane and blade outer air seal assembly, and to a gas turbine engine. Gas turbine engines are known, and typically include a compressor section delivering compressed air into a combustor where it is mixed with fuel and ignited. Products of the combustion pass downstream over turbine rotors driving them to rotate. The turbine rotors in turn drive compressor rotors, and in some cases a propulsor rotor such as a fan or propeller. It is known that the products of combustion are quite hot, and thus it is desirable to have cooling air provided to components in the turbine section. Moreover, it is desirable to control the flow of the products of combustion through the turbine section to provide a maximum efficiency to the turbine section. Thus, it is known to place static vanes having airfoils which direct flow of the products of combustion toward the plurality of turbine rotors. Moreover, it is known to have blade outer air seals positioned radially outwardly of turbine blades to ensure that the products of combustion do not pass over the turbine section without driving the turbine rotors. Recently it has been proposed to utilize ceramic matrix composites ("CMCs") for several components in a gas turbine engine. The CMCs can withstand higher temperatures than many other materials. US 11 073 039 B1 discloses a prior art combined gas turbine engine vane and blade outer air seal assembly as set forth in the preamble of claim 1. US 2021/231022 A1 discloses a prior art turbine engine with reused secondary cooling flow. US 2021/189901 A1 discloses a prior art ceramic matrix composite component including counterflow channels and a method of producing such components. SUMMARY OF THE INVENTION From one aspect, there is provided a combined gas turbine engine vane and blade outer air seal assembly as recited in claim 1. There is also provided a gas turbine engine as recited in claim 8. Features of embodiments are set forth in the dependent claims. These and other features of the present invention can be best understood from the following specification and drawings, the following of which is a brief description. BRIEF DESCRIPTION OF THE DRAWINGS Figure 1 schematically shows a gas turbine engine.Figure 2 schematically shows a turbine section including a combined turbine vane and blade outer air seal according to one way of carrying out the invention claimed.Figure 3 shows a first arrangement for illustration purposes.Figure 4 shows a subsequent arrangement for illustration purposes.Figure 5A shows a first alternative embodiment for the Figure 4 embodiment.Figure 5B shows an alternative embodiment.Figure 5C shows an alternative embodiment.Figure 5D shows an alternative embodiment.Figure 6A shows a first support embodiment for the blade outer air seal.Figure 6B shows an alternative support.Figure 6C shows another alternative support.Figure 7 schematically shows a layup of a plurality of plies to form the integrated vane and blade outer air seal of one example of this invention.Figure 8A shows a first embodiment of an air supply.Figure 8B shows an alternative embodiment. 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 may include a single-stage fan 42 having a plurality of fan blades 43. The fan blades 43 may have a fixed stagger angle or may have a variable pitch to direct incoming airflow from an engine inlet. The fan 42 drives air along a bypass flow path B in a bypass duct 13 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. A splitter 29 aft of the fan 42 divides the air between the bypass flow path B and the core flow path C. The housing 15 may surround the fan 42 to establish an outer diameter of the bypass duct 13. The splitter 29 may establish an inner diameter of the bypass duct 13. Although depicted as a two-spool turbofan gas turbine engine in the disclosed nonlimiting 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 engine 20 may incorporate a variable area nozzle for varying an exit area of the bypass flow path B and/or a thrust reverser for generating reverse thrust. The exemplary engine 20 generally includes a low speed spool 30 and a high speed spool 32 mounted for rotation about an eng