Stator vane for a turbine of a turbomachine

Abstract

A stator vane (3) for a turbine (50c) of a turbomachine (50), the stator vane having a stator vane airfoil (3c), an inner shroud (3a) and an outer shroud (3b), the inner shroud (3a) and the outer shroud (3b) bounding an annular space (2), in which working gas (51) is conveyed during operation, radially with respect to a longitudinal axis (52) of the turbomachine (50), and the stator vane airfoil (3c) having a stator vane airfoil channel (3d) extending through its interior between a radially inner inlet (6) and a radially outer outlet (7). A characteristic features is that the inlet (6) is disposed in such a manner that a gas (8) flowing through the stator vane airfoil channel (3d) during operation is at least partially formed of the working gas (51) conveyed in the annular space (2), and thus the working gas is redistributed from radially inward to radially outward.

Claims

1. A stator vane for a turbine of a turbomachine, the stator vane comprising: a stator vane airfoil; an inner shroud; and an outer shroud, the inner shroud and the outer shroud bounding an annular space radially with respect to a longitudinal axis of the turbomachine, working gas conveyed in the annular space during operation; the stator vane airfoil having a stator vane airfoil channel extending through an interior between a radially inner inlet and a radially outer outlet, the inlet being disposed in such a manner that a gas flowing through the stator vane airfoil channel during operation is at least partially formed of the working gas conveyed in the annular space so that the gas including the working gas is redistributed from radially inward to radially outward; wherein the inlet of the stator vane airfoil channel is disposed at a leading edge of the inner shroud of the stator vane, the inlet on the leading edge facing in an upstream direction relative to the flow of the working gas through the annular space.

2. The stator vane as recited in claim 1 wherein the outlet of the stator vane airfoil channel is disposed radially outwardly of the outer shroud of the stator vane.

3. The stator vane as recited in claim 2 wherein the outlet of the stator vane airfoil channel is offset from a trailing edge of the stator vane airfoil in a downstream direction relative to the flow of the working gas through the annular space.

4. A turbine module comprising the stator vane as recited in claim 1.

5. The turbine module as recited in claim 4 further comprising a rotor blade disposed upstream of the stator vane relative to the flow of the working gas through the annular space, the rotor blade having a rotor blade inner shroud and a rotor blade airfoil, a downstream-pointing trailing edge of the rotor blade inner shroud having an axial overlap with the leading edge of the inner shroud of the stator vane in order to form a labyrinth seal.

6. The turbine module as recited in claim 5 further comprising a sealing fin disposed radially inwardly of the inner shroud of the stator vane, the sealing fin being provided, as part of the labyrinth seal, radially inwardly of the inner shroud of the stator vane and having an axial overlap therewith.

7. The turbine module as recited in claim 5 wherein the turbine module is designed so that sealing fluid flows through the labyrinth seal from radially inward to radially outward during operation, the sealing fluid at least partially being suctioned off through the inlet of the stator vane airfoil channel and flowing through the stator vane airfoil channel as part of the gas.

8. The turbine module as recited in claim 4 further comprising a rotor blade disposed downstream of the stator vane relative to the flow of the working gas through the annular space, the rotor blade having a rotor blade airfoil as well as a rotor blade inner shroud and a rotor blade outer shroud, the outlet of the stator vane airfoil channel is disposed in such a manner that the gas flowing through the stator vane airfoil channel is at least partially by-passed radially outwardly of the rotor blade outer shroud.

9. The turbine module as recited in claim 8 wherein an amount of the gas that is by-passed radially outwardly of the rotor blade outer shroud is selected such that the amount of gas blocks the working gas from flowing directly out of the annular space and over the outer shroud of the rotor blade.

10. The turbine module as recited in claim 8 wherein the outlet of the stator vane airfoil channel is provided in such a manner that the gas flowing through the stator vane airfoil channel exits divergently from a direction of rotation of the rotor blade.

11. The turbine module as recited in claim 8 wherein the outlet of the stator vane airfoil channel is provided in such a manner that the gas flowing through the stator vane airfoil channel exits at a different velocity or in a different direction than a working gas velocity or direction that the working gas conveyed in the annular space at the outlet.

12. The turbine module as recited in claim 4 further comprising a rotor blade downstream of the stator vane and having a rotor blade airfoil made of a forged material.

13. The turbine module as recited in claim 4 further comprising a rotor blade downstream of the stator vane and part of a disk with integral rotor blades, the disk being made of a forged material.

14. A method for operating the turbine module as recited in claim 4 comprising conveying the working gas in the annular space, and flowing the gas from radially inward to radially outward through the stator vane airfoil channel, the gas being at least partially formed of the working gas conveyed in the annular space so that the gas including the working gas is redistributed from radially inward to radially outward.

Description

BRIEF DESCRIPTION OF THE DRAWINGS

(1) The present invention will now be explained in more detail with reference to an exemplary embodiment. The individual features may also be essential to the invention in other combinations within the scope of the other independent claims, and, as above, no distinction is specifically made between different claim categories.

(2) In the drawing,

(3) FIG. 2 shows an axial cross-sectional view of a turbine module having a stator vane provided with a stator vane airfoil channel, according to the present invention;

(4) FIG. 1 shows, in comparison to FIG. 2, a variant without a stator vane airfoil channel to illustrate the advantages achieved by the present invention;

(5) FIG. 3 shows a diagram illustrating the radial temperature profile;

(6) FIG. 4 shows a diagram illustrating the radial efficiency profile;

(7) FIG. 5 shows an axial cross-sectional view of a turbomachine having a turbine module as shown in FIG. 2.

DETAILED DESCRIPTION OF PREFERRED EMBODIMENTS

(8) FIG. 2 shows, in axial cross-sectional view, a portion of a turbine module 1. During operation, working gas traveling from the combustor (located to the left of turbine module 1) to the nozzle (located to the right thereof) flows through an annular space 2 formed by turbine module 1 (see also FIG. 5 for illustration). Disposed in this annular space 2 is a stator vane 3 having an inner shroud 3a, an outer shroud 3b, and a stator vane airfoil 3c therebetween. A rotor blade 4 is disposed upstream of stator vane 3; a rotor blade 5 is disposed downstream thereof. Stator vane 3 is shown in cross-section. A stator vane airfoil channel 3d extends from radially inward to radially outward through stator vane airfoil 3c. The inlet 6 into stator vane airfoil channel 3d is located at inner shroud 3a of stator vane 3, and specifically at the upstream leading edge thereof. The outlet 7 of stator vane airfoil channel 3d is disposed radially outwardly of the outer shroud 3b and is axially offset in the downstream direction from trailing edge 3ca of stator vane airfoil 3c.

(9) Due to the pressure difference across stator vane 3, suctioning occurs at a radially inner position, at inlet 6, and blowing out occurs at a radially outer position, at outlet 7. Inlet 6 is disposed such that the gas 8 flowing through stator vane airfoil channel 3d is at least partially formed of the working gas conveyed in annular space 2. Specifically, an endwall boundary layer 10 of the mainstream flow is suctioned off. This is advantageous from an aerodynamic standpoint alone, and, in addition, the temperatures in the annular space are lower radially inwardly than radially outwardly, and thus an excessive temperature gradient can be prevented by the redistribution.

(10) Furthermore, a sealing fluid 11, which is introduced in the radially inner region to shield the hub region and flows through a labyrinth seal 12, is also partially suctioned in through inlet 6. The labyrinth seal is formed by an axial overlap of a sealing fin 13, inner shroud 4a of rotor blade 4, and specifically the trailing edge thereof, and inner shroud 3a of stator vane 3, and specifically the leading edge thereof. This sealing fluid 11 is significantly cooler compressor air, whose radially outward redistribution through stator vane airfoil channel 3d is advantageous with regard to preventing excessive temperature gradients.

(11) In comparison, FIG. 1 shows a turbine module 1 having an analogously configured labyrinth seal 12. However, unlike FIG. 2, stator vane airfoil 3c is not provided with a stator vane airfoil channel 3d. Accordingly, sealing fluid 11 flows into annular space 2, disturbing the mainstream flow therein. In addition, endwall boundary layers 10 generally suffer from aerodynamic issues anyway; i.e., overall, flow losses and efficiency losses are likely to occur (compared to the variant shown in FIG. 2). FIG. 1 further illustrates that there is also a leakage flow 20 in the radially outward region, the leakage flow flowing over outer shrouds 4b, 5b of rotor blades 4, 5. This, too, results in a disturbance of the mainstream flow.

(12) In the inventive design, this is avoided by positioning outlet 7 of stator vane airfoil channel 3d in such a way that the gas 8 conveyed radially outward flows over outer shroud 5b of rotor blade 5. The amount is selected such that no working gas from annular space 2 flows over outer shroud 5b. As can be seen FIG. 2, this applies analogously to the upstream turbine stage. However, for the sake of clarity, the description refers to the interaction of stator vane 3 with rotor blade 5.

(13) FIG. 3 illustrates a radial temperature profile as arises in a turbine module 1 according to FIG. 1; i.e., without redistribution through stator vane airfoil channel 3d. Temperature T is plotted on the x-axis; the radius taken in a direction away from the inner shroud is plotted on the y-axis. The solid line represents the temperature of the working gas, which is primarily determined by the temperature profile at the combustor exit. The temperature increases radially outwardly (see also the introductory part of the description).

(14) FIG. 4 illustrates the efficiency q (x-axis) in relation to radius R (y-axis). A drop in efficiency in the radially inner region and in the radially outer region, inter alia, occurs because of boundary layer flow 10 and leakage flow 20. In addition to this, a disturbance is caused by the sealing fluid 11 flowing into the annular space in the radially inner region. A can be seen from FIG. 3, sealing fluid 11 has a significantly lower temperature than the working gas there (see point T.sub.11 on the x-axis). Thus, when sealing fluid 11 flows into annular space 2, a mixture temperature T.sub.Mix arises there, so that temperature gradient (ΔT.sub.(a-Mix)) is even greater than when considering the working gas alone (ΔT.sub.(a-i)).

(15) As explained above, with the approach of the present invention, the cooler sealing fluid 11 and, in addition, cooler working gas are redistributed from radially inward to radially outward, so that the temperature gradients can be reduced. As a result of the reduced disturbance of the mainstream flow in the radially inner and radially outer regions, an improved efficiency profile can be achieved as well.

(16) FIG. 5 shows, in axial cross-sectional view, a turbomachine 50, specifically a jet engine. Turbomachine 50 is functionally divided into a compressor 50a, a combustor 50b and a turbine 50c. Both compressor 50a and turbine 50c are made up of a plurality of components or stages, each stage being composed of a stator vane ring and a rotor blade ring. The rotor blade rings are driven by working gas 51 and rotate about longitudinal axis 52 of turbomachine 50. The aforedescribed turbine module 1 is part of turbine 50c, and specifically forms the low-pressure turbine.

LIST OF REFERENCE NUMERALS

(17) turbine module 1 annular space 2 stator vane 3 inner shroud 3a outer shroud 3b stator vane airfoil 3c trailing edge 3ca stator vane airfoil channel 3d rotor blade (upstream) 4 inner shroud 4a outer shroud 4b rotor blade airfoil 4c rotor blade (downstream) 5 inner shroud 5a outer shroud 5b rotor blade airfoil 5c inlet 6 outlet 7 gas 8 endwall boundary layer/boundary layer flow   sealing fluid 11 labyrinth seal 12 sealing fin 13 leakage flow 20 turbomachine 50 compressor 50a combustor 50b turbine 50c working gas 51 longitudinal axis 52 temperature T radius R efficiency η