Turbine nozzle guide vane including a ring seal element

Abstract

The invention relates to a nozzle guide vane (3) for a turbine (2), the nozzle guide vane having an axis of rotation (X) and comprising: - a flange (6) extending radially with respect to the axis (X); - an annular seal (12) mounted on the flange and comprising an inner face (15a) configured to carry a wear element (13); - a ventilation cavity (17) delimited radially on the inside by an outer face (15b) of the annular seal (12) and on the downstream by an upstream face (6a) of the flange (6), - a baffle configured to direct a flow of air towards the ventilation cavity (17); and - at least one through opening (20) formed in the annular seal (12) and configured to put the ventilation cavity (17) in fluid communication with the inner face (15a) of the annular seal (12).

Description

Technical Field

[0001] The present invention relates to a turbojet engine, and more generally to a turbojet engine including a turbine nozzle having a radially inner surface supporting an annular sealing element of a labyrinth seal, the annular sealing element being made of a wear-resistant material and designed to engage with a blade edge carried by a rotor facing the annular sealing element. Background Technology

[0002] A gas turbine engine, in the direction of gas flow through the turbine engine from upstream to downstream, typically includes: a fan, one or more compressor stages (e.g., a low-pressure compressor and a high-pressure compressor), a combustion chamber, one or more turbine stages (e.g., a high-pressure turbine and a low-pressure turbine), and an exhaust nozzle 8.

[0003] In a manner known per se, turbines are manufactured in the form of a series of stages, each stage comprising a moving impeller (rotor) rotating in front of a fixed bladed wheel (stator or nozzle guide vanes). This moving impeller is configured to convert combustion energy into driving energy, specifically driving the compressor stage and the fan. The turbine's rotor disk is centered on the longitudinal axis X of the turbine engine and is typically assembled together via shell rings at upstream and downstream disks. Each assembly of the nozzle guide vanes comprises multiple fixed blades circumferentially distributed around the axis.

[0004] The nozzle guide vane is a fixed vane, which is configured to accelerate and divert the combustion gas flow in the direction of the moving impeller directly downstream.

[0005] It is known to place dynamic seals radially inside the turbine nozzle guide vanes to reduce gas flow outside the main annular duct in which the blades and vanes extend. The seals typically comprise an annular seal and one or more blade edges, the annular seal supporting a wear-resistant element fixed to the nozzle guide vane, and the blade edges fixed to the rotor.

[0006] When a turbojet engine is running, the nozzle guide vanes are exposed to a higher temperature than the moving wheel disc, and the thermal inertia of the nozzle guide vanes is typically lower than that of the disc. This causes variations in the clearance between the sealing ring and the blade edge. When this clearance increases, it leads to increased leakage flow, and when the clearance decreases, it leads to accelerated wear of the wear-resistant components.

[0007] Document FR 3 027 343 discloses a component scheme that allows radial movement between an annular seal and a nozzle guide vane. However, the seal of this component is highly sensitive to temperature changes because the expansion of the annular seal causes the gap between the wear-resistant element and the facing blade edge to open, thereby degrading engine performance. The cooling flow injected between the wear-resistant element and the blade edge does not mix with the hot air obtained from the airflow leaking between the nozzle guide vane and the directly adjacent rotating disk. This means that the temperature change of the annular seal depends on the free expansion of the hot air flow leaking from the airflow. During the transition phase, this means that the annular seal expands under the influence of temperature rise, which translates into an opening of the sealing gap and increased permeability of the seal, thus reducing the seal's performance.

[0008] Finally, the large temperature difference between the upstream and downstream (both sides of the seal) of the ventilation cavity creates a temperature difference between the annular fixed flange and causes differential thermal expansion, which may prevent the annular seal from moving radially relative to the turbine nozzle guide vane. Summary of the Invention

[0009] One object of the present invention is to overcome the above-mentioned disadvantages by proposing a nozzle guide vane that includes a seal having properties that are less sensitive to temperature changes, particularly when the seal comprises an annular seal mounted on the nozzle guide vane with a radial clearance and having an improved lifespan.

[0010] For this purpose, in a first aspect of the invention, a turbine nozzle guide vane is provided, the turbine nozzle guide vane having a rotation axis and comprising:

[0011] - A flange that extends radially relative to the axis;

[0012] - An annular seal, which is mounted on a flange and includes an inner surface configured to carry a wear-resistant element;

[0013] - A ventilation cavity, which is radially defined on its inner side by the outer surface of an annular seal and downstream by the upstream surface of a flange;

[0014] - A deflector configured to guide airflow toward the ventilation cavity; and

[0015] - At least one through opening, said at least one through opening being formed in the annular seal and configured to allow the ventilation cavity to be in fluid communication with the inner surface of the annular seal.

[0016] Some preferred, but non-limiting, features of the turbine nozzle guide vanes according to the first aspect are as follows, and these features are employed individually or in combination:

[0017] - The deflector is mounted on one of the flange and the annular seal, such that the deflector extends substantially radially within the ventilation cavity;

[0018] - The deflector is fixed to the curved upstream edge of the annular seal;

[0019] - The upstream edge extends radially inward;

[0020] - The turbine nozzle guide vane includes a curved metal strip fixed to the upstream edge of an annular seal, with at least one opening passing through the deflector;

[0021] - The deflector consists of a generally flat or truncated conical sheet of metal;

[0022] - The deflector extends radially relative to the axis;

[0023] - The deflector is brazed to one of the flange and the annular seal, or the deflector is integral with the annular seal; and / or

[0024] - The annular seal is installed so that it can move radially relative to the flange.

[0025] In a second aspect, the present invention provides a turbine comprising turbine nozzle guide vanes according to the first aspect.

[0026] In a third aspect, the present invention provides a turbine engine comprising a turbine according to the second aspect, and further comprising an auxiliary turbine having a rotational speed greater than that of the turbine.

[0027] In a fourth aspect, the present invention proposes an aircraft comprising a turbine engine according to the third aspect. Attached Figure Description

[0028] Other features, objects, and advantages of the invention will become apparent from the following non-limiting description, which is merely illustrative and will be read in conjunction with the accompanying drawings, in which:

[0029] Figure 1 This is a cross-sectional view of an example turbine including nozzle guide vanes according to an embodiment of the present invention, the view schematically showing different fluids (gases) flowing below the nozzle guide vanes;

[0030] Figure 2 This is a partial perspective view of an example of an annular seal and a deflector according to an embodiment of the present invention;

[0031] Figure 3 This is a cross-sectional view of an example of a conventional turbine including nozzle guide vanes, schematically showing the different fluids flowing below the nozzle guide vanes; and

[0032] Figure 4 An aircraft comprising two turbine engines is shown according to an embodiment of the present invention.

[0033] In all the accompanying drawings, similar elements have the same reference numerals. Detailed Implementation

[0034] In this application, upstream and downstream are defined relative to the normal direction of gas flow in the turbine engine 1 (specifically, the aircraft 100), and therefore along the nozzle guide vanes 3 of the low-pressure turbine 2. The axis X of the nozzle guide vanes 3 is the axis around which the nozzle guide vanes 3 extend, and corresponds to the axis of rotation X of the rotor 7 of the turbine 2. The axial direction corresponds to the direction of the axis X, and the radial direction is the direction perpendicular to and through the axis X. Furthermore, the circumferential direction corresponds to the direction perpendicular to the axis X but not through the axis X. Unless otherwise specified, inner (or internal) and outer (or external) are used with reference to the radial direction, such that the inner portion or inner surface of an element is closer to the axis X than the outer portion or outer surface of the same element.

[0035] In a manner known per se, particularly the nozzle guide vanes 3 of the low-pressure turbine 2, include a hub platform 4, which is provided with a plurality of fixed blades 5 circumferentially distributed around the hub platform 4. The hub platform 4 specifically includes a flange 6 that extends radially inward relative to the blades 5.

[0036] The nozzle guide vane 3 can be segmented and includes multiple guide vane segments 3 fixed end-to-end around axis X. As a variation, the nozzle guide vane 3 can be integrally formed as a single piece, in which case the segments of the nozzle guide vane 3 are integral.

[0037] The turbine 2 also includes a rotor 7, which includes a plurality of disks 8, each disk including a periphery 9 and a blade 10. The blade 10 is positioned and fixed in a groove 11 formed in the periphery 9 of the respective disk 8 and extends radially from the periphery 9.

[0038] The nozzle guide vane 3 extends between two adjacent disks 8 of the rotor 7 of the turbine 2. Preferably, the turbine 2 comprises a series of stages, each stage including a disk 8 of the rotor 7 rotating in front of the associated nozzle guide vane 3.

[0039] The turbine 2 also includes a seal comprising an annular seal 12 and a second sealing element 14. The annular seal is slidably connected to the flange 6 and configured to carry a first sealing element 13. The second seal 14 is configured to mate with the first sealing element 13 and may be carried, for example, by a support fixed between two adjacent disks 8 of the rotor 7 of the turbine 2. The first sealing element 13 may include a wear-resistant element 13 (e.g., a honeycomb structure) fixed to the inner surface 15a of the annular seal 12. The second sealing element 14 may then include a blade edge 14 extending radially from the support. Examples of seals including wear-resistant elements 13 and blade edges 14 are described in the remainder of this document. However, this is not limiting, and the invention can be applied to seals including other sealing elements.

[0040] More specifically, the flange 6 extends generally radially from the hub platform 4 in the direction of the second sealing element (i.e., opposite to the blade) and has generally radial upstream surface 6a and downstream surface 6b.

[0041] An annular seal 12 can be added and connected to the flange 6 via a sliding connection. The annular seal includes a metal sheet comprising an axial portion 15 and a radial portion 16. The axial portion 15 has an inner surface 15a configured to receive a wear-resistant element 13 and an outer surface 15b opposite the inner surface 15, configured to extend towards the hub platform 4. The wear-resistant element 13 is directly fixed to the inner surface 15a of the axial portion 15. The axial portion 15 also has an upstream edge 15c and a downstream edge 15d.

[0042] The radial portion 16 extends from the outer surface 15b of the axial portion 15 and is configured to be mounted on the flange 6. The radial portion 16 may be centered relative to the axial portion 15.

[0043] In one embodiment, the radial portion 16 is mounted on the flange 6 to provide a radial clearance between the annular seal 12 and the flange 6. For this purpose, the radial portion 16 may include an upstream end plate 16a and a downstream end plate 16b, configured to contact the upstream surface 6a and the downstream surface 6b of the flange 6, respectively. Further details regarding the radial clearance mounting of the annular seal 12 on the flange 6 can be found in FR 3 027 343.

[0044] Ring 12 and flange 6 together define an upstream ventilation cavity 17 and a downstream ventilation cavity 18. The upstream cavity 17 is defined upstream by the rotating disk 8 of the rotor 7 directly upstream, downstream by the upstream surface 6a of flange 6, radially inward by the outer surface 15b of the axial portion of ring 12, and radially outward by the inner surface of hub platform 4. The downstream cavity 18 is defined upstream by the downstream surface 6b of flange, downstream by the rotating disk 8 of the rotor 7 directly downstream, radially inward by the outer surface 15b of the axial portion of ring 12, and radially outward by the inner surface of hub platform 4.

[0045] The ring 12 is mounted on the flange 6 with a gap, allowing air to pass between the end plates 16a, 16b of the ring 12 and the flange 6 (region 16c), thereby enabling fluid communication between the upstream cavity 17 and the downstream cavity 18 for ventilation.

[0046] In a manner known per se, a cooling gas flow F1 (particularly from the upstream portion of the turbine engine 1, such as from the compressor section) passes through the bottom of the groove 11 and is injected between the wear-resistant element 13 and the blade edge 14 of the seal to cool the seal and ensure its tightness.

[0047] Furthermore, it is possible that gas F2 can leak from the outer surface of the periphery 9 of the directly upstream disk 8 and be injected into the upstream cavity 17. This gas F2 is relatively hot, thus having the effect of heating the upstream cavity 17.

[0048] To improve the tightness and lifespan of the seal, the nozzle guide vane 3 also includes a deflector 19 configured to guide airflow toward the upstream cavity 17. Specifically, the deflector 19 may be mounted on one of the flange 6 and the ring 12, such that the deflector extends substantially radially within the upstream cavity 17. Furthermore, at least one through opening 20 is formed in the annular seal 12 to allow the portion of the upstream cavity 17 located between the outer surface 15b of the deflector 19 and the ring 12 to fluidly communicate with the inner surface 15a of the annular seal 12.

[0049] Therefore, the deflector 19 and the through opening 20 enable the cooling fluid F1 (gas) from the groove 11 of the rotor 7 disk 8 directly upstream of the nozzle guide vane 3 to mix with the hot gases F3 and F2 from the airflow and from the periphery 9 of the disk 8, thereby reducing the temperature gradient between the upstream cavity 17 and the downstream cavity 18. The rotation of the rotor 7 of the turbine 2 radially outward drives the cooling fluid F1, causing the fluid F1 to move upward along the deflector 19 until it reaches the upstream cavity 17, where it encounters the hot gases F2 and F3. The mixing of the cooling fluid F1 with the hot gases F2 and F3 thus allows for the cooling of the upstream cavity 17 (and the upstream end plate 16a). If appropriate, a portion of the resulting mixture F1a can then pass through the channel 16a provided for this purpose between the end plates 16a, 16b and the flange 6, thereby cooling the downstream cavity 18, while another portion F1b moves downward toward the blade edge 14 via the through opening to ensure a tight seal.

[0050] The deflector 19 is substantially radial relative to the axis X. "Substantially radial" here means that the deflector 19 extends in a plane capable of forming an angle between -20° and +20° with a plane perpendicular to the axis X. In one embodiment, the deflector is located in a plane perpendicular to the axis X to provide effective guidance of cooling fluid toward the upstream cavity 17.

[0051] Preferably, a plurality of through openings 20 are formed in the ring 12 to ensure sufficient passage for gas flow from the upstream cavity 17. As a variation, an oblong orifice 20 may be formed in the deflector 19 and / or the ring 12. Since the seal is being calibrated, the cross-sectional area of ​​the passage of the orifice 20 need not be involved in the calibration of the gas flow in the seal.

[0052] In the first embodiment, the deflector 19 is fixed to the ring 12, for example, by brazing or mechanical connection (bolting, etc.). The deflector 19 thus includes an inner radial end disposed near the annular seal 20 and an outer radial end extending into the upstream cavity 17.

[0053] For example, the deflector 19 can be secured to the upstream edge 15c of the ring 12 and can extend radially outward in the upstream cavity 17. To facilitate securing the deflector 19 to the ring 12, at least one of the ring 12 and the deflector 19 includes a curved edge, and the other of the deflector 19 and the ring 12 is attached to this curved edge. For example, Figure 1 and Figure 2The illustration shows a ring 12 including a curved upstream edge 15c, with a deflector 19 then fixed to the curved portion of the upstream edge 15c. Preferably, the curved portion of the upstream edge 15c of the ring 12 extends radially inward (i.e., toward the axis X), preferably extending to the outlet of the groove 11 (i.e., the end of the upstream edge 15c faces the outlet of the groove 11), to form a barrier that blocks the cold cooling fluid F1 from the groove 11 of the directly upstream disk 8. With this configuration, the cooling fluid F1 can be further forced upward toward the upstream cavity 17, rather than flowing directly toward the seal.

[0054] The deflector 19 comprises a substantially planar metal sheet including a radially inner edge 19a and a radially outer edge 19b. The radially inner edge 19a and the radially outer edge 19b are preferably annular. The deflector 19 may be a single piece, or it may comprise multiple segments attached end-to-end (potentially overlapping).

[0055] In this embodiment, the through opening 20 is formed in the axial portion 15a of the ring 12, near the bend. Preferably, the opening 20 is offset relative to the wear-resistant element 13: therefore, the opening is formed in the portion of the ring 12 that does not have the wear-resistant element, i.e., near the bend of the upstream edge 15c. For example, the opening 20 may be formed in the bend of the upstream edge 15c and extend radially within the bend to open along the deflector 19 (see...). Figure 2 Therefore, the cooling fluid F1b flows along the deflector 19 through the opening 20 before reaching the seal.

[0056] If the deflector 19 has a curved portion, it further includes an annular piece integral with the radial inner edge of the metal sheet. The deflector 19 can then be added to and secured to the outer surface 15b of the ring 12 near the upstream edge 15c of the ring. An opening 20 is then formed such that it mates with the axial portion 15a of the deflector 19 and the ring 12 at the point of overlap.

[0057] As a variation, the deflector 19 is integral with the ring 12 and can be obtained by additive manufacturing. An opening 20 is then formed at the junction near the axial portion 15a of the ring 12. The deflector 19 can simply be bent to the axial portion 15a of the ring 12, or, as a variation, the deflector can include a radially outer portion and a radially inner portion, the radially outer portion extending from the upstream edge 15c of the ring 12 along the direction of the upstream cavity 17, and the radially inner portion extending from the upstream edge 15c of the ring 12 along the direction of the axis X to form a barrier preventing the passage of cold fluid F1 from the groove 11.

[0058] In the second embodiment (not shown), the deflector 19 is fixed to the radial portion 16 of the ring 12 or to the flange 6. The deflector 19 then includes a perforation to allow fluid F1 to flow toward the opening 20 of the ring 12, and optionally (depending on its position relative to the end plates 16a, 16b) toward the gas passage 16c at the end plates 16a, 16b. Furthermore, the deflector 19 includes an inner radial end extending within the upstream cavity 17 and an outer radial end disposed near the radial portion 16 of the annular seal 20.

[0059] The deflector 19 may include a metal sheet in the form of a channel, which includes a first planar sheet configured to be fixed to a radial portion 16 of the flange 6 or ring 12, a second planar sheet extending substantially radially in the upstream cavity 17, and a central sheet connecting the first and second metal sheets together.

[0060] As a variation, the radial portion 16 of the flange 6 or ring 12 may include a bent metal sheet that can be brazed to the radial portion or flange 6. The deflector 19 may then include a bent metal sheet that is fixed (e.g., by brazing) to a bent metal sheet of the flange 6 or the radial portion.

[0061] In another variation, the deflector 19 may include a radially inwardly bent metal sheet integral with the radial portion 16 of the flange 6 or ring 12.

[0062] Figure 4 The diagram illustrates the flow of different airflows F1, F2, and F3 in a conventional nozzle guide vane. Hot gas F3, originating from the airflow, leaks into the area between the hub platform 4 of the nozzle guide vane 3 and the disk 8 of the rotor 7 directly upstream. Simultaneously, cooling airflow F1, arriving from the slot 11 of the disk 8, is injected between the wear-resistant element 13 and the blade edge 14 to cool the seal. Finally, the hot leakage flow F2 near the periphery 9 of the disk 8 can be injected into the upstream cavity 17. As can be seen from the diagram, the cooling fluid F1 does not move upwards to the upstream cavity 17, thus the upstream cavity 17 is very hot and only cools the seal. It can also be seen that there are significant temperature differences between the upstream cavity 17 and the downstream cavity 18, and between the upstream end plate 16a and the downstream end plate 16b of the ring 12.

[0063] For comparison, the airflow of F1, F2, and F3 is as follows: Figure 1 As shown, Figure 1 A nozzle guide vane 3 according to an embodiment of the present invention is shown. Hot gas F3 from the airflow mixes in the upstream cavity 17 with cooling fluid F1 moving upward along the deflector 19 and with heat leakage flow F2 from the periphery 9 of the disk 8. Therefore, the total temperature of the upstream cavity 17 is much lower than... Figure 4The temperature difference between the upstream cavity 17 and the downstream cavity 18, as well as between the upstream end plate 16a and the downstream end plate 16b of the ring 12, is also low. The similarly low-temperature gas F1a flowing between the end plates 16a, 16b and the flange 6 has already mixed with the fluid F1 arriving from the groove 11. Finally, a portion of the cold fluid F1b returns to the seal via the opening 20 to ensure the seal is cooled.

[0064] Because the upstream cavity 17 is continuously supplied with a strong cooling flow, the ring 12 experiences relatively small temperature changes during the service life of the turbine engine 1. This is further attributed to the mixing of gases F1, F2, and F3 in the upstream cavity 17. Therefore, the ring 12 is less subject to thermal expansion.

[0065] Because the temperature of the upstream cavity 17 and the downstream cavity 18 is more uniform, the differential thermal expansion between the two end plates 16a and 16b is reduced (or even eliminated). This avoids the risk of the ring 12 being blocked (especially the risk of radial movement of the ring relative to the flange 6).

[0066] The deflector 19 is preferably metallic and can be made of the same material as the flange 6 or ring 12. For example, the deflector 19 may include an aluminum-steel alloy, etc.

[0067] This invention can be applied to any axial turbine 2 consisting of a series of moving stages (rotor 7) and static stages (nozzle guide vanes 3), particularly to turbines 1 (and their derivatives) and gas turbines. The ring 12 of the turbine 2 can be free, i.e., having a radial clearance with the flange 6, or, as a variation, the ring 12 can be fixed relative to the flange 6. This invention is particularly suitable for turbines 2 with significant leakage flow from the airflow.

Claims

1. A nozzle guide vane (3) of a turbine (2), the nozzle guide vane having a rotation axis (X) and comprising: -Flange (6), the flange extending radially relative to the axis of rotation (X); - An annular seal (12), which is mounted on the flange (6) and has an inner surface (15a) configured to support the wear-resistant element (13). - Ventilation cavity (17), which is radially defined on the inner side by the outer surface (15b) of the annular seal (12) and downstream by the upper surface (6a) of the flange (6); The nozzle guide vane (3) is characterized in that it further includes: - A deflector configured to guide airflow toward the ventilation cavity (17), the deflector (19) being fixed to the curved upstream edge (15c) of the annular seal (12), the curved upstream edge extending radially inward; and - At least one through opening (20) is formed in the annular seal (12) and is configured to allow the ventilation cavity (17) to be in fluid communication with the inner surface (15a) of the annular seal (12).

2. The nozzle guide vane (3) according to claim 1, wherein, The deflector (19) is mounted on one of the flange (6) and the annular seal (12) so as to extend substantially radially within the ventilation cavity (17).

3. The nozzle guide vane (3) according to claim 1 or 2, wherein, The upstream edge (15c) extends upward into a groove (11) formed in the periphery (9) of the disk (8) of the rotor of the turbine (2) adjacent to the nozzle guide vane.

4. The nozzle guide vane (3) according to claim 1 or 2, wherein, The at least one through opening (20) is configured to allow airflow from the ventilation cavity (17) toward the inner surface (15a) of the annular seal (12).

5. The nozzle guide vane (3) according to claim 1 or 2, wherein, The nozzle guide vane (3) includes a bent metal sheet fixed to the upstream edge (15c) of the annular seal (12), and the at least one through opening (20) passes through the deflector (19).

6. The nozzle guide vane (3) according to claim 1 or 2, wherein, The deflector (19) comprises a substantially planar or truncated conical metal sheet.

7. The nozzle guide vane (3) according to claim 1 or 2, wherein, The deflector (19) extends radially relative to the axis of rotation (X).

8. The nozzle guide vane (3) according to claim 1 or 2, wherein, The deflector (19) is brazed to one of the flange (6) and the annular seal (12), or the deflector is integral with the annular seal (12).

9. The nozzle guide vane (3) according to claim 1 or 2, wherein, The annular seal (12) is mounted to be radially movable relative to the flange (6).

10. A turbine (2), the turbine comprising a nozzle guide vane (3) according to any one of claims 1 to 9.

11. A turbine engine (1), comprising a turbine (2) according to claim 10, and comprising an additional turbine having a rotational speed greater than that of the turbine (2).

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