Controlled convergence compressor flowpath for a gas turbine engine

Abstract

A controlled convergence compressor flowpath (10) configured to better distribute the limited flowpath (10) convergence within compressors (12) in turbine engines (14) is disclosed. The compressor (12) may have a flowpath (10) defined by circumferentially extending inner and outer boundaries (16, 18) that having portions in which the rate of convergence changes to better distribute fluid flow therethrough. The rate of convergence may increase at surfaces (20, 22) adjacent to roots (24) of airfoils (26) and decrease near airfoil tips (68) and in the axial gaps (28) between airfoil rows (30). In at least one embodiment, the compressor flowpath (10) between leading and trailing edges (44, 46) of a first compressor blade (42) may increase convergence moving downstream to a trailing edge (46) of the first compressor blade (42) due to increased convergence of the inner compressor surface (22). The compressor flowpath (10) between leading and trailing edges (32, 34) of a first compressor vane (36) immediately downstream from the first compressor blade (42) may increase convergence moving downstream due to increased convergence of the outer compressor surface (20).

Claims

1. A gas turbine engine comprising: a compressor formed from a rotor assembly and a stator assembly; wherein the compressor comprises an inner compressor surface and an outer compressor surface; wherein the rotor assembly is formed from a plurality of radially outward extending compressor blades from the inner compressor surface aligned into a plurality of circumferentially extending rows and wherein the rotor assembly is rotatable; wherein the stator assembly is formed from a plurality of radially inward extending compressor vanes from the outer compressor surface aligned into a plurality of circumferentially extending rows, wherein the stator assembly is fixed relative to the rotatable rotor assembly and wherein the rows of compressor vanes alternate with the rows of compressor blades moving in a downstream direction; wherein the inner and outer compressor surfaces form a compressor flowpath; wherein the compressor flowpath converges moving downstream; wherein a rate of the convergence of the compressor flowpath increases at the inner and outer surfaces adjacent to roots of the blades and vanes; and wherein the rate of the convergence of the compressor flowpath reduces at the inner and outer surfaces adjacent to tips of the blades and vanes.

2. The gas turbine engine of claim 1, wherein the rate of the convergence of the compressor flowpath at the inner compressor surface between a leading edge and a trailing edge of a first compressor blade increases aft of a point of maximum thickness of a root of the first compressor blade.

3. The gas turbine engine of claim 1, wherein the rate of the convergence of the compressor flowpath at the inner compressor surface radially aligned with and between a leading edge and a trailing edge of a first compressor blade is nonlinear.

4. The gas turbine engine of claim 1, wherein the rate of the convergence of the compressor flowpath at the inner compressor surface radially aligned with and between a leading edge and a trailing edge of a first compressor blade curves radially inward moving downstream.

5. The gas turbine engine of claim 1, wherein the rate of the convergence of the compressor flowpath at the inner compressor surface between a trailing edge of a first compressor blade and a leading edge of a first compressor vane immediately downstream from the first compressor blade is linear.

6. The gas turbine engine of claim 1, wherein the rate of the convergence of the compressor flowpath at the outer compressor surface between a trailing edge of a first compressor blade and a leading edge of a first compressor vane immediately downstream from the first compressor blade is linear.

7. The gas turbine engine of claim 1, wherein the rate of the convergence of the compressor flowpath at the outer compressor surface between a leading edge and a trailing edge of a first compressor vane immediately downstream from a first compressor blade increases moving downstream.

8. The gas turbine engine of claim 1, wherein the rate of the convergence of the compressor flowpath at the outer compressor surface between a leading edge and a trailing edge of a first compressor vane increases moving downstream due to increased convergence of the outer compressor surface.

9. The gas turbine engine of claim 1, wherein the rate of the convergence of the compressor flowpath at the outer compressor surface between a leading edge and a trailing edge of a first compressor vane increases aft of a point of maximum thickness of a root of the first compressor vane.

10. The gas turbine engine of claim 1, wherein the rate of the convergence of the compressor flowpath at the inner compressor surface between a leading edge and a trailing edge of a first compressor vane reduces radially inwardly.

11. The gas turbine engine of claim 1, wherein the rate of the convergence of the compressor flowpath at the outer compressor surface radially aligned with and between a leading edge and a trailing edge of a first compressor vane is nonlinear.

12. The gas turbine engine of claim 1, wherein the rate of the convergence of the compressor flowpath at the outer compressor surface radially aligned with and between a leading edge and a trailing edge of a first compressor vane curves radially inward moving downstream.

13. The gas turbine engine of claim 1, wherein the rate of the convergence of the compressor flowpath at the outer compressor surface between a trailing edge of a first compressor vane and a leading edge of a compressor blade immediately downstream from the first compressor vane reduces from the rate of the convergence of the compressor flowpath at the outer compressor surface between a leading edge and the trailing edge of the first compressor vane.

14. The gas turbine engine of claim 1, wherein the rate of the convergence of the compressor flowpath at the inner compressor surface between a trailing edge of a first compressor blade and a leading edge of a first compressor vane immediately downstream from the first compressor blade reduces from the rate of the convergence of the compressor flowpath at the inner compressor surface between a leading edge and the trailing edge of the first compressor blade.

15. The gas turbine engine of claim 1, wherein the rate of the convergence of the compressor flowpath at the inner and outer surfaces transitions from linear over the tips of the blades and vanes to nonlinear over the roots of the blades and vanes.

Description

BRIEF DESCRIPTION OF THE DRAWINGS

(1) The accompanying drawings, which are incorporated in and form a part of the specification, illustrate embodiments of the presently disclosed invention and, together with the description, disclose the principles of the invention.

(2) FIG. 1 is a perspective view of a gas turbine engine with a partial cross-sectional view with a compressor.

(3) FIG. 2 is a cross-sectional side view of a portion of the compressor

DETAILED DESCRIPTION OF THE INVENTION

(4) As shown in FIGS. 1-2, a controlled convergence compressor flowpath 10 configured to better distribute the limited flowpath convergence within compressors 12 in turbine engines 14 is disclosed. The compressor 12 may have a flowpath 10 defined by circumferentially extending inner and outer boundaries 16, 18 that have portions in which the rate of convergence changes to better distribute fluid flow therethrough. The rate of convergence may increase at surfaces 20, 22 adjacent to roots 24 of airfoils 26 and decrease near airfoil tips 68 and in the axial gaps 28 between airfoil rows 30. In at least one embodiment, the rate of convergence may increase at surfaces 20, 22 adjacent to roots 24 of airfoils 26 and aft of a location of maximum thickness of the roots 24 and may reduce convergence near airfoil tips 68 and in the axial gaps 28 between airfoil rows 30. In at least one embodiment, the compressor flowpath 10 between leading and trailing edges 44, 46 of a first compressor blade 42 may increase convergence moving downstream to the trailing edge 46 of the first compressor blade 42 due to increased convergence of an inner compressor surface 22 aft of a point 60 of maximum thickness of a root 24 of the first compressor blade 42. The compressor flowpath 10 within the vaneless axial gap 28 between rows 30 of compressor blades 42 and rows 30 of compressor vanes 36 may have reduced convergence compared to the row 30 of compressor blades 42 immediately upstream. The compressor flowpath between leading and trailing edges 32, 34 of a first compressor vane 36 immediately downstream from the first compressor blade 42 may increase convergence moving downstream relative to the axial gap 28 upstream of the first compressor vane 36 due to increased convergence of the outer compressor surface 20 aft of a point 62 of maximum thickness of a root 24 of the first compressor vane 36.

(5) In at least one embodiment, the gas turbine engine 14 may include one or more compressors 12 formed from a rotor assembly 48 and a stator assembly 50. The rotor assembly 48 may be formed from a plurality of radially outward extending compressor blades 42 aligned into a plurality of circumferentially extending rows 30. The rotor assembly 48 may be rotatable about an axis of the turbine engine 14. The stator assembly 50 may be formed from a plurality of radially inward extending compressor vanes 36 aligned into a plurality of circumferentially extending rows 30. The stator assembly 50 may be fixed relative to the rotatable rotor assembly 48. The rows 30 of compressor vanes 36 may alternate with the rows 30 of compressor blades 42 moving in a downstream direction.

(6) The inner compressor surface 22 may define a circumferential inner boundary surface 54 of the compressor 12, and the outer compressor surface 20 may define a circumferential outer boundary surface 56 of the compressor 12 whereby the inner and outer compressor surfaces 22, 20 form the compressor flowpath 10. The compressor flowpath 10 may converge moving downstream from an inlet 58 of the compressor 12 to an outlet 59.

(7) In at least one embodiment, the compressor flowpath 10 radially outward of, such as at the OD, and between the leading edge 44 and the trailing edge 46 of one or more first compressor blades 42 forming a row 30 of compressor blades 42, otherwise known as a stage when positioned adjacent a row of turbine vanes, may increase convergence moving downstream to the trailing edge 46 of the first compressor blade 42 relative to a rate of convergence immediately upstream from the first compressor blade 42. In at least one embodiment, the compressor flowpath 10 radially outward of and between the leading edge 44 and the trailing edge 46 of the first compressor blade 42 may increase convergence moving downstream to the trailing edge 44 of the first compressor blade 42 due to increased convergence of the inner compressor surface 22 aft of a point 60 of maximum thickness of a root 24 of the first compressor blade 42. The slope of convergence of the controlled convergence compressor flowpath 10 proximate to a blade tip 68 at the OD 64 may be reduced and the slope of convergence may be increased proximate to the airfoil root at the ID 66 so that, at the location of largest thickness of the blade 42 near the root, the convergence of the flowpath increases to prevent flow separation from occurring aft of the airfoil maximum thickness point. Blade tips 68 are typically thinner than blade roots, thus area convergence within the blade row 30 is less effective proximate to the blade tip 68. The inner compressor surface 22 radially aligned with and between the leading edge 44 and the trailing edge 46 of the first compressor blade 42 may be nonlinear. In at least one embodiment, the inner compressor surface 22 radially aligned with and between the leading edge 44 and the trailing edge 46 of the first compressor blade 42 curves radially inward moving downstream.

(8) The compressor flowpath 10 in the axial gap 28 radially outward of and between the trailing edge 46 of the first compressor blade 42 and the leading edge 32 of a first compressor vane 36 immediately downstream from the first compressor blade 42 reduces convergence from a rate of convergence between the leading and trailing edges 44, 46 of the first compressor blade 42. In at least one embodiment, the rate of convergence in the vaneless axial gaps 28 between the compressor blades 42 and compressor vanes 36 at the inner compressor surface 22 and at the outer compressor surface 20 may be equal. In at least one embodiment, the inner compressor surface 22 between the trailing edge 46 of the first compressor blade 42 and the leading edge 32 of a first compressor vane 36 immediately downstream from the first compressor blade 42 may be linear. The outer compressor surface 20 between the trailing edge 46 of the first compressor blade 42 and the leading edge 32 of a first compressor vane 36 immediately downstream from the first compressor blade 42 may be linear.

(9) The compressor flowpath 10 between the leading edge 32 and the trailing edge 34 of the first compressor vane 36 immediately downstream from the first compressor blade 42 may increase convergence moving downstream. In at least one embodiment, the compressor flowpath 10 between the leading edge 32 and the trailing edge 34 of the first compressor vane 36 may increase convergence moving downstream due to increased convergence of the outer compressor surface 20 aft of a point 62 of maximum thickness of a root 24 of the first compressor vane 36. The outer compressor surface 20 radially aligned with and between the leading edge 32 and the trailing edge 34 of the first compressor vane 36 may be nonlinear. In at least one embodiment, the outer compressor surface 20 radially aligned with and between the leading edge 32 and the trailing edge 34 of the first compressor vane 36 may curve radially inward moving downstream, thereby increasing convergence. The compressor flowpath 10 between the trailing edge 34 of the first compressor vane 36 and a leading edge 44 of a compressor blade immediately downstream from the first compressor vane 36 reduces convergence from a rate of convergence between the leading and trailing edges 32, 34 of the first compressor vane 36.

(10) The foregoing is provided for purposes of illustrating, explaining, and describing embodiments of this invention. Modifications and adaptations to these embodiments will be apparent to those skilled in the art and may be made without departing from the scope or spirit of this invention.