COMBUSTOR ASSEMBLY FOR A TURBINE ENGINE

1 . A combustor assembly for a gas turbine engine defining an axial direction and a radial direction, the combustor assembly comprising:

a dome comprising a forward surface and an inner surface, the forward surface and the inner surface at least partially defining a slot; and

a liner at least partially defining a combustion chamber and extending between an aft end and a forward end, the forward end of the liner received within the slot of the dome, the forward end of the liner comprising an axial interface surface and a radial interface surface, the axial interface surface defining a radial gap with the inner surface of the dome and the radial interface surface defining an axial gap with the forward surface of the dome, at least one of the radial gap or axial gap being less than about 0.150 inches during operating conditions of the combustor assembly.

2 . The combustor assembly of claim 1 , wherein at least one of the radial gap or axial gap is less than about 0.020 inches during operating conditions of the combustor assembly.

3 . The combustor assembly of claim 2 , wherein the radial gap is less than about 0.020 inches during operating conditions of the combustor assembly.

4 . The combustor assembly of claim 3 , wherein the radial gap is between 0.005 and 0.015 inches during operating conditions of the combustor assembly.

5 . The combustor assembly of claim 3 , wherein the axial gap is less than about 0.150 inches during operating conditions of the combustor assembly.

6 . The combustor assembly of claim 5 , wherein the axial gap is between 0.050 and 0.100 inches during operating conditions of the combustor assembly.

7 . The combustor assembly of claim 1 , wherein the radial interface surface is a machined surface.

8 . The combustor assembly of claim 1 , wherein the axial interface surface is a machined surface.

9 . The combustor assembly of claim 1 , wherein the forward surface of the dome extends linearly along the radial direction.

10 . The combustor assembly of claim 1 , wherein the liner is an outer liner and wherein the annular dome is an outer dome section.

11 . The combustor assembly of claim 1 , wherein the liner is an inner liner and wherein the annular dome is an inner dome section.

12 . The combustor assembly of claim 1 , wherein the liner is comprised of a ceramic matrix composite material, and wherein the annular dome is comprised of a metal material.

13 . A method for manufacturing a combustor assembly of a gas turbine engine, the combustor assembly comprising a liner and a dome, the dome comprising a forward surface and an inner surface, the method comprising:

forming a liner of a ceramic matrix composite material to include a baseline geometry;

removing material from the liner to change the baseline geometry to include an interface surface; and

mounting the liner to the dome.

14 . The method of claim 13 , wherein forming a liner to include a baseline geometry comprises forming a liner to include a forward end comprising the baseline geometry, wherein removing material from the liner to change the baseline geometry to include an interface surface comprises removing material from the forward end of the liner to change the baseline geometry to include an axial interface surface and a radial interface surface.

15 . The method of claim 14 , wherein mounting the liner to the dome comprises mounting the forward end of the liner to the dome such that the axial interface surface defines a radial gap with the inner surface of the dome and the radial interface surface defines an axial gap with the forward surface of the dome, at least one of the axial gap or the radial gap being less than about 0.150 inches during operating conditions of the combustor assembly

16 . The method of claim 14 , wherein removing material from the forward end of the liner comprises machining the forward end of the liner to define the axial interface surface.

17 . The method of claim 14 , wherein removing material from the forward end of the liner comprises machining the forward end of the liner to define the radial interface surface.

18 . The method of claim 13 , wherein removing material from the liner to change the baseline geometry to include an interface surface comprises removing material from at least one of a midspan region of the liner or an aft end of the liner.

19 . A method for cooling a combustor assembly of a gas turbine engine, the combustor assembly comprising a liner and a dome, the dome comprising a forward surface and an inner surface at least partially defining a slot, the liner comprising a forward end received within the slot, the method comprising:

providing a cooling airflow to the slot defined by the forward surface and the inner surface of the dome;

providing the cooling airflow through an axial gap defined between the forward end of the liner and the forward surface of the dome, the axial gap being less than about 0.150 inches;

providing the cooling airflow through a radial gap defined between the forward end of the liner and the inner surface of the dome, the radial gap being less than about 0.020 inches; and

providing the cooling airflow to a combustion chamber defined at least in part by the liner and the dome.

20 . The method of claim 18 , wherein the axial gap is between 0.050 and 0.100 inches during operating conditions of the combustor assembly, and wherein the radial gap is between 0.005 and 0.015 inches during operating conditions of the combustor assembly.