Ceramic matrix composite component manufacturing
A gas turbine engine is provided. The gas turbine engine includes a stage of airfoil segments having a first ceramic matrix composite (CMC) airfoil segment, the first CMC airfoil segment includes: an end band having a radial surface defining a cold side surface area, A CS ; and an airfoil coupled to or formed with the end band, the airfoil and the end band defining a joint therebetween, the joint including one or more radial interfaces defining a total radial interface surface area, A RI , and one or more axial interfaces defining a total axial interface surface area, A AI . The first CMC airfoil segment defines a first ratio of the cold side surface area, A CS , to the total axial interface surface area, A AI , and a second ratio of the total axial interface surface area, A AI to the total radial interface surface area, A RI .
1. A gas turbine engine, the gas turbine engine comprising:
a turbomachine comprising a stage of rotating rotor blades and a stage of airfoil segments adjacent the stage of the rotating rotor blades, the stage of airfoil segments comprising a first ceramic matrix composite (CMC) airfoil segment, the first CMC airfoil segment comprising:
an end band having a radial surface defining a cold side surface area, A CS ; and
an airfoil coupled to or formed with the end band, the airfoil and the end band defining a joint therebetween, the joint comprising one or more radial interfaces defining a total radial interface surface area, A RI , and one or more axial interfaces defining a total axial interface surface area, A AI , the first CMC airfoil segment defining a first ratio of the cold side surface area, A CS , to the total axial interface surface area, A AI , greater than or equal to 2.3 and less than or equal to 10 and a second ratio of the total axial interface surface area, A AI to the total radial interface surface area, A RI , greater than or equal to 1.8 and less than or equal to 10.
2. The gas turbine engine of claim 1 , wherein the turbomachine comprises a high pressure turbine, and wherein the stage of rotating rotor blades is a first stage of rotor blades of the high pressure turbine.
3. The gas turbine engine of claim 1 , wherein the end band is an inner band along a radial direction.
4. The gas turbine engine of claim 1 , wherein the end band is an outer band along a radial direction.
5. The gas turbine engine of claim 1 , wherein the total axial interface surface area, A AI , is a sum of a plurality of axial interfaces of the joint, and wherein the total radial interface surface area, A RI , is a sum of a plurality of radial interfaces of the joint.
6. The gas turbine engine of claim 1 , wherein the airfoil comprises a body and a base, wherein the base includes an extension having a first surface extending perpendicularly from the base, a second surface opposite the first surface also extending perpendicularly from the base, and a third surface, wherein the first surface and the second surface contribute to the total axial interface surface area, A AI , and wherein the third surface contributes to the total radial interface surface area, A RI .
7. The gas turbine engine of claim 6 , wherein the extension defines a non-rectangular shape.
8. The gas turbine engine of claim 6 , wherein the extension defines a dovetail shape or a dovetail-shaped slot.
9. The gas turbine engine of claim 6 , wherein the airfoil defines a suction side, and wherein the extension is positioned on the suction side of the airfoil.
10. The gas turbine engine of claim 1 , wherein the cold side surface area, A CS , of the end band is a total footprint surface area of a cold side of the end band.
11. A ceramic matrix composite (CMC) component comprising:
an end band having a radial surface defining a cold side surface area, A CS ; and
an airfoil coupled to or formed with the end band, the airfoil and the end band defining a joint therebetween, the joint comprising one or more radial interfaces defining a total radial interface surface area, A RI , and one or more axial interfaces defining a total axial interface surface area, A AI , the CMC component defining a first ratio of the cold side surface area, A CS , to the total axial interface surface area, A AI , greater than or equal to 2.3 and less than or equal to 10 and a second ratio of the total axial interface surface area, A AI to the total radial interface surface area, A RI , greater than or equal to 1.8 and less than or equal to 10.
12. The CMC component of claim 11 , wherein the end band is an inner band along a radial direction.
13. The CMC component of claim 11 , wherein the end band is an outer band along a radial direction.
14. The CMC component of claim 11 , wherein the total axial interface surface area, A AI , is a sum of a plurality of axial interfaces of the joint.
15. The CMC component of claim 11 , wherein the total radial interface surface area, A RI , is a sum of a plurality of radial interfaces of the joint.
16. The CMC component of claim 11 , wherein the airfoil comprises a body and a base, wherein the base includes an extension having a first surface extending perpendicularly from the base, a second surface opposite the first surface also extending perpendicularly from the base, and a third surface, wherein the first surface and the second surface contribute to the total axial interface surface area, A AI , and wherein the third surface contributes to the total radial interface surface area, A RI .
17. The CMC component of claim 16 , wherein the extension defines a non-rectangular shape.
18. The CMC component of claim 16 , wherein the extension defines a dovetail shape or a dovetail-shaped slot.
19. The CMC component of claim 16 , wherein the airfoil defines a suction side, and wherein the extension is positioned on the suction side of the airfoil.
20. The CMC component of claim 11 , wherein the cold side surface area, A CS , of the end band is a total footprint surface area of a cold side of the end band.