Contoured endwall for a gas turbine engine

Patent metadata

Field	Value
Publication number	US-10920599-B2
Application number	US-201916263063-A
Country	US
Kind code	B2
Filing date	Jan 31, 2019
Priority date	Jan 31, 2019
Publication date	Feb 16, 2021
Grant date	Feb 16, 2021

Abstract

A vane for a gas turbine engine according to an example of the present disclosure includes, among other things, first and second endwalls each having a radially facing surface that bounds a gas path, an airfoil section extending in a radial direction between the first and second endwalls, extending in an axial direction between an airfoil leading edge and an airfoil trailing edge, and extending in a circumferential direction between pressure and suction sides. The radially facing surface of each of the first and second endwalls is axially sloped such that the gas path converges in the axial direction between the airfoil leading and trailing edges. The first endwall has an axisymmetric contour at least partially swept in the circumferential direction from each of the pressure and suction sides.

First claim

What is claimed is:

1. A vane for a gas turbine engine comprising:
first and second endwalls each including a radially facing surface that bounds a gas path;
an airfoil section extending in a radial direction between the first and second endwalls, extending in an axial direction between an airfoil leading edge and an airfoil trailing edge, and extending in a circumferential direction between pressure and suction sides;
wherein the radially facing surface of each of the first and second endwalls is axially sloped such that the gas path converges in the axial direction between the airfoil leading and trailing edges;
wherein the first endwall includes an axisymmetric contour at least partially swept in the circumferential direction from each of the pressure and suction sides; and
wherein the first endwall extends in the circumferential direction between opposed mate faces of the outer endwall, and the axisymmetric contour is swept in the circumferential direction from the pressure and suction sides to respective ones of the opposed mate faces.

2. The vane as recited in claim 1 , wherein the axisymmetric contour is a depression in the radially facing surface.

3. The vane as recited in claim 2 , wherein the axisymmetric contour has an arcuate cross sectional geometry.

4. The vane as recited in claim 3 , wherein the arcuate cross sectional geometry includes an apex that is skewed in the axial direction toward one of the airfoil leading and trailing edges.

5. The vane as recited in claim 1 , wherein the axisymmetric contour is a protrusion that extends outwardly from the radially facing surface and into the gas path.

6. The vane as recited in claim 5 , wherein the axisymmetric contour has an arcuate cross sectional geometry.

7. The vane as recited in claim 1 , wherein the axisymmetric contour has a sinusoidal cross sectional geometry.

8. The vane as recited in claim 7 , wherein the sinusoidal cross sectional geometry includes a concave portion and a convex portion, the concave portion extends inwardly from the radially facing surface with respect to the radial direction, the convex portion extends outwardly from the radially facing surface with respect to the radially direction, and the concave portion is defined between the airfoil leading edge and the convex portion with respect to the axial direction.

9. The vane as recited in claim 1 , wherein the vane is a fan stator.

10. A section for a gas turbine engine comprising:
a rotor carrying an array of blades that extend into a gas path, the rotor rotatable about a longitudinal axis; and
an array of vanes distributed about the longitudinal axis, wherein each of the vanes comprises:
an airfoil section extending in a radial direction between inner and outer endwalls, extending in an axial direction between an airfoil leading edge and an airfoil trailing edge, and extending in a circumferential direction between pressure and suction sides;
wherein the inner and outer endwalls each includes a radially facing surface dimensioned such that the gas path converges in the axial direction at the airfoil trailing edge relative to the airfoil leading edge;
wherein the outer endwall includes an axisymmetric contour at least partially swept in the circumferential direction from the pressure and suction sides; and
wherein the radially facing surface extends in the circumferential direction between opposed mate faces of the outer endwall, and the axisymmetric contour is swept in the circumferential direction from the pressure and suction sides to respective ones of the opposed mate faces.

11. The section as recited in claim 10 , wherein the array of vanes are axially forward of the array of blades relative to the longitudinal axis such that the array of vanes and the array of blades comprise adjacent stages of the section.

12. The section as recited in claim 11 , wherein the axisymmetric contour has an arcuate cross sectional geometry.

13. The section as recited in claim 11 , wherein the axisymmetric contour has a sinusoidal cross sectional geometry.

14. The section as recited in claim 13 , wherein the axisymmetric contour includes an arcuate depression in the radially facing surface of the outer endwall and that extends inwardly from a reference plane, the reference plane extends between junctions of the leading and trailing edges and the outer endwall, and the axisymmetric contour includes an arcuate protrusion that establishes the apex, the sinusoidal cross sectional geometry established by the depression and the protrusion, and the protrusion extending outwardly from the radially facing surface and into the gas path such that an apex of the depression and the apex of the protrusion are on opposed sides of the reference plane.

15. A gas turbine engine comprising:
a fan section;
a combustor in fluid communication with the fan section;
a turbine section rotationally coupled to the fan section; and
wherein the fan section includes a row of blades rotatable about an engine longitudinal axis, a stator assembly including a row of vanes adjacent the row of blades, and wherein each of the vanes comprises:
an airfoil section extending in a radial direction between inner and outer endwalls that bound a gas path, extending in an axial direction between an airfoil leading edge and an airfoil trailing edge, and extending in the circumferential direction between pressure and suction sides; and
wherein the inner and outer endwalls converge in the axial direction to define a converging portion of the gas path;
wherein the stator assembly includes an axisymmetric contour swept in the circumferential direction along the outer endwall between each of the vanes to bound the converging portion of the gas path; and
wherein the axisymmetric contour is swept in the circumferential direction from the pressure side of each of the vanes to the suction side of adjacent ones of the vanes.

16. The gas turbine engine as recited in claim 15 , wherein the row of blades and the row of vanes comprise an axially forwardmost stage of the gas turbine engine relative to the engine longitudinal axis.

17. The gas turbine engine as recited in claim 15 , wherein radially facing surfaces of the inner and outer endwalls are axially sloped in the axial direction between the airfoil leading and trailing edges to define the converging portion of the gas path.

18. The gas turbine engine as recited in claim 17 , wherein the axisymmetric contour has an arcuate cross sectional geometry including an apex that is skewed in the axial direction toward one of the airfoil leading and trailing edges.

19. The gas turbine engine as recited in claim 18 , wherein the axisymmetric contour includes an arcuate depression in the radially facing surface of the outer endwall and that extends inwardly from a reference plane, the reference plane extends between junctions of the leading and trailing edges and the outer endwall, and the axisymmetric contour includes an arcuate protrusion that establishes the apex, the axisymmetric contour having a sinusoidal cross sectional geometry established by the depression and the protrusion, and the protrusion extending outwardly from the radially facing surface and into the gas path such that an apex of the depression and the apex of the protrusion are on opposed sides of the reference plane.

20. The gas turbine engine as recited in claim 19 , wherein the radially facing surface of the outer endwall slopes radially inward in the axial direction from the airfoil leading edge towards the airfoil trailing e

Assignees

• United Technologies Corp
• Raytheon Tech Corp

Inventors

• Adhate Abhir A
• Prior Frank M
• Wagner Kalman V

Classifications

• F05D2250/70

  Shape · CPC title

• F04D29/542

  Bladed diffusers (fixing blades to stators F01D9/042) · CPC title

• Y02T50/60

  Efficient propulsion technologies, e.g. for aircraft · CPC title

• F05D2250/72

  symmetric · CPC title

• F05D2250/713

  inflexed · CPC title