Improved cooling system for a fixed ring of a gas turbine

Abstract

The invention relates to a fixed ring (4) extending around a central axis (X), suitable for being mounted on a turbine engine casing (2) in order to surround a gas passage (14) of the turbine engine (1), the casing (2) defining an annular cooling chamber (30) arranged radially outside the ring (4), the ring (4) comprising a first internal cooling cavity (10) capable of fluidically communicating with the annular cooling chamber (30) via a first set of openings (40), and a second internal cooling cavity (20) fluidly communicating with the first internal cooling cavity (10) via a second set of openings (50), the second internal cooling cavity (20) being arranged closer to a downstream end (162) than to an upstream end (161) of the ring sector (16).

Description

Description Title of invention: Improved gas turbine fixed ring cooling system Technical Field

[0001] This disclosure relates to fixed rings surrounding gas passages of gas turbines, and more particularly to the cooling of fixed gas turbine rings. Prior art

[0002] A gas turbine, in particular a high-pressure turbine of a turbomachine, typically comprises a plurality of fixed blades arranged axially alternately with a plurality of moving blades in the passage of hot gases from the combustion chamber of the turbomachine. The fixed and moving blades of the turbine extend circumferentially around a central axis of rotation of the gas turbine and are surrounded over the entire circumference by a fixed ring which is generally formed of a plurality of ring sectors. These ring sectors partly define the passage for the flow of hot gases through the blades of the turbine.

[0003] The turbine ring sectors are thus subjected to the high temperatures of the hot gases coming from the combustion chamber of the turbomachine. For the mechanical and thermal resistance of the turbine ring, it is therefore necessary to equip the ring sectors with cooling devices.

[0004] In addition to applying a thermal protection layer to the face of the ring in contact with the hot gases, one of the known cooling methods consists of supplying cooling air to an impact plate mounted on the body of the ring sectors. The plate is provided with a plurality of orifices for the passage of air which, under the pressure difference on either side of the plate, cools the ring sector by impact.

[0005] The cooling air is then evacuated into the hot gas passage through holes made through the ring sector.

[0006] Such a method does not allow for efficient and uniform cooling of the ring sectors, particularly at the upstream end of the ring sector, which is an area particularly exposed to hot gases. Indeed, there may be disturbances between the jets coming from adjacent orifices, and between the jets from the orifices located upstream and downstream of the device. These disturbances, which are all the more marked when the cavity to be cooled is large and the number of orifices is high, reduce the cooling efficiency.

[0007] The service life of the ring sectors is therefore affected. Furthermore, and to limit the disadvantages mentioned above linked to disturbances between jets, this technology requires a significant intake of cooling air, which reduces the performance of the turbine.

[0008] There is therefore a need for a solution to at least partially limit the aforementioned drawbacks and to improve the cooling efficiency of fixed turbine rings. Statement of the invention

[0009] The present disclosure relates to a fixed ring extending around a central axis, capable of being mounted on a turbomachine casing to surround a gas passage of the turbomachine, the casing delimiting an annular cooling chamber arranged radially outside the ring, said ring comprising a first internal cooling cavity capable of fluidly communicating with the annular cooling chamber via a first set of orifices, and a second internal cooling cavity fluidly communicating with the first internal cooling cavity via a second set of orifices, the second internal cooling cavity being arranged closer to a downstream end than to an upstream end of the ring sector.

[0010] In this disclosure, an axial direction is considered along the central axis of the gas turbine, which is the axis of rotation of the rotating parts of the turbine. In addition, the terms "upstream" and "downstream" are considered along a normal direction of gas flow in the turbine along the central axis, in particular from the combustion chamber to the high-pressure turbine.

[0011] Furthermore, the terms "internal" and "external" are considered in a radial direction perpendicular to the central axis. Thus, the external wall of the combustion chamber, for example, is further from the central axis than its internal wall in the radial direction.

[0012] It is understood that the first and second internal cavities are arranged in a cascade arrangement, in which cooling air from the annular cooling chamber first flows into the first internal cooling cavity via the first set of ports, and then flows from the first internal cooling cavity to the second internal cooling cavity via the second set of ports.

[0013] The second internal cooling cavity being arranged closer to a downstream end than to an upstream end of the ring sector, i.e. close to a downstream axial end of the ring sector, it is possible to cool initially an upstream portion of the ring sector via the first internal cavity, then a downstream portion of the ring sector via the second internal cavity.

[0014] The first and second internal cavities being distinct from each other, but communicating with each other, it is thus possible to cool an upstream portion of the ring sector, without the jets coming from the first set of orifices being disturbed by the jets coming from the second set of orifices, making it possible to cool a downstream portion of the ring sector. In addition, this makes it possible to homogenize the cooling air in the first internal cavity, before its reinjection into the second internal cavity.

[0015] Thus, for the same cooling air flow rate, it is possible to improve cooling efficiency. In other words, this arrangement allows the cooling air flow rate required to achieve the same thermal performance on the ring to be reduced. This also increases the temperature of the hot air flow path and therefore the performance of the gas turbine and turbomachine.

[0016] In some embodiments, the first internal cavity includes an upstream portion and a downstream portion, the upstream portion being radially inwardly offset from the downstream portion.

[0017] It is understood that the upstream portion and the downstream portion communicate with each other, for example by means of a shoulder offsetting these portions relative to each other. This makes it possible to bring the upstream portion closer to the hot air flow stream, and therefore to improve the cooling of the wall in contact with said hot gases, on the upstream portion of the ring sector.

[0018] In some embodiments, the second internal cavity is disposed radially below the downstream portion of the first internal cavity.

[0019] In other words, according to this arrangement, the upstream portion of the first internal cavity, and the second internal cavity, are each arranged closer to the hot air flow stream, compared to the downstream portion of the first internal cavity. This arrangement makes it possible to improve the cooling efficiency of the ring sector, by cooling on the one hand an upstream portion of the wall in contact with the hot gases via the upstream portion of the first internal cavity, and on the other hand a downstream portion of the wall in contact with the hot gases via the second internal cavity.

[0020] In some embodiments, the upstream portion of the first internal cavity is capable of fluidly communicating with the annular cooling chamber via the first set of ports, and the downstream portion of the first internal cavity communicates with the second internal cavity via the second set of ports.

[0021] Given this arrangement in which the second internal cavity is arranged radially below the downstream portion of the first internal cavity, the cooling air from the annular cooling chamber comes impact the upstream portion of the ring sector wall in contact with the hot gases via the orifices of the first set of orifices, then the cooling air from the first internal cavity impacts the downstream portion of the sector wall in contact with the hot gases via the orifices of the second set of orifices. This further improves the cooling efficiency.

[0022] In some embodiments, the ring comprises a plurality of ring sectors arranged circumferentially end-to-end around the central axis, each ring sector comprising a radially inner wall and a radially outer wall at least partially delimiting the first internal cavity, and an intermediate wall disposed between the radially inner wall and the radially outer wall and delimiting, with a portion of the radially inner wall, the second internal cavity.

[0023] It is understood that an inner face of the radially inner wall delimits the flow path of the hot gases, the radially inner wall and the radially outer wall being spaced from each other. The radially inner wall delimits both a lower portion of the first inner cavity and the second inner cavity. The cooling air injected through the first and second sets of orifices respectively impacts the radially inner wall, which further improves the cooling efficiency.

[0024] In some embodiments, the first set of ports is formed in the radially outer wall, and the second set of ports is formed in the intermediate wall.

[0025] In some embodiments, the first and second sets of ports are disposed at a common radial position.

[0026] It is understood that the orifices of the first and second set of orifices are all arranged in the same curved radial plane whose center of curvature is the central axis, the second set of orifices being downstream of the first set of orifices. This arrangement makes it possible to homogenize the cooling of the upstream and downstream portions of the radially internal wall of the ring sector.

[0027] In some embodiments, the orifices of the first set of orifices have a larger diameter than the orifices of the second set of orifices.

[0028] This ensures that the holes in the second set of holes have the limiting cross-section for the passage of the cooling air flow and ensures the correct supply and flow of air in both cavities. In particular, at constant flow rate, the smallest bore cross-section corresponds to the fastest fluid velocity, and therefore the lowest static pressure. Placing the smallest cross-section between the first and second internal cavities ensures that the static pressure in the second cavity is lower than in the first, thus facilitating the air flow. Furthermore, accelerating the fluid after it has lost energy following its impact with the internal wall in the first cavity, and therefore providing it with enough energy to impact the internal wall in the second internal cavity.This helps to improve the cascading flow effect of the cooling air, and thus improve the cooling efficiency.

[0029] In certain embodiments, the downstream end of the ring comprises discharge orifices capable of fluidly connecting the second internal cavity with the gas passage of the turbomachine.

[0030] In some embodiments, a diameter of the exhaust orifices is larger than a diameter of the orifices of the second set of orifices. This limits the effects of air blocking in the second internal cavity, which would increase the pressure in the latter. Larger orifices also facilitate the manufacture of these exhaust orifices, since cooling downstream of the latter, and therefore acceleration of the fluid, is not necessary.

[0031] In some embodiments, a section of the exhaust ports is larger than a section of the ports of the first set.

[0032] In some embodiments, the orifices of the first set of orifices and the orifices of the second set of orifices have a diameter of less than 1 mm.

[0033] In some embodiments, the vent holes have a diameter of between 1 and 1.5 mm.

[0034] These values ​​allow to optimize the cascade flow effect of the cooling air, from the annular cooling chamber to the hot gas flow vein, passing through the first internal cavity then the second internal cavity.

[0035] In some embodiments, the radially inner wall comprises a plurality of disruptors projecting from an upstream portion of said radially inner wall toward the radially outer wall and / or from a downstream portion of said radially inner wall toward the intermediate wall.

[0036] In some embodiments, each of the disruptors is arranged so as not to be radially opposite one of the orifices of the first set of orifices and / or one of the orifices of the second set of orifices.

[0037] The present disclosure also relates to an assembly comprising a turbomachine casing and a ring according to any one of the preceding embodiments, mounted on the casing, the casing surrounding the ring and delimiting an annular cooling chamber arranged radially outside the ring.

[0038] This disclosure also relates to a high pressure turbine comprising an assembly defined above. Brief description of the drawings

[0039] The invention and its advantages will be better understood upon reading the detailed description given below of different embodiments of the invention given as non-limiting examples. This description refers to the appended pages of figures, in which:

[0040] [Fig. 1] Figure 1 schematically represents a part of a gas turbine illustrating the location of a fixed ring relative to that of the moving blades,

[0041] [Fig. 2] Figure 2 is a longitudinal sectional view of a ring sector according to one embodiment of the invention,

[0042] [Fig. 3] Figure 3 is a perspective view of the ring sector alone of Figure 2,

[0043] [Fig. 4] Figure 4 is a perspective and sectional view of the ring sector of Figure 3 along a section plane A,

[0044] [Fig. 5] Figures 5 is a perspective and sectional view of the ring sector of Figure 3 along a section plane B.

[0045] [Fig. 6] Figure 6 is a detailed view of an upstream portion of the ring sector of Figure 2, according to a modified example of the embodiment of the invention,

[0046] [Fig. 7] Figure 7 schematically represents, seen from above, an example of the arrangement of the disruptors of the ring sector of Figure 6, and their positions relative to the orifices of the first set of orifices. Description of the embodiments

[0047] An embodiment of the invention will be described with reference to Figures 1 to 5.

[0048] The terms “upstream” and “downstream” are subsequently defined in relation to the direction of flow of hot gases through a turbomachine, indicated by the arrow F in Figures 1 and 2. Furthermore, the terms “internal” and “external” are considered in a radial direction perpendicular to the central axis X. Thus, the radially external wall 18 of the ring sector 16, for example, is further from the central axis X than its radially internal wall 17 in the radial direction.

[0049] Figure 1 schematically represents a part of a high-pressure turbine 1 of a turbomachine.

[0050] The high-pressure turbine 1 comprises in particular a fixed annular housing 2 forming a casing of the turbomachine. A fixed turbine ring 4 is fixed to this housing 2 and surrounds a plurality of moving blades 6 of the turbine. These moving blades 6 are arranged downstream of fixed blades 8 relative to the flow direction F of the hot gases coming from a combustion chamber 12 of the turbomachine and passing through the turbine. Thus, the turbine ring 4 surrounds a passage 14 (or vein) for the flow of hot gases.

[0051] Generally, the turbine ring 4 consists of a plurality of ring sectors arranged circumferentially around the central axis X of the turbine so as to form a circular and continuous surface. However, it is also possible to imagine that the turbine ring is composed of only one single continuous piece. The present invention applies equally to a single turbine ring and to a turbine ring sector.

[0052] Referring to Figure 2, illustrating in section one of the ring sectors 16 forming the fixed ring 4, it can be seen that each ring sector 16 has a radially internal wall 17 (hereinafter “internal wall 17”) which is annular and a radially external wall 18 (hereinafter “external wall 18”) which is offset radially outwards relative to the internal wall 17.

[0053] The internal wall 17 faces the passage 14 for the flow of hot gases F. Each ring sector 16 also has, at its upstream axial end 161, an upstream hook 22 and, at its downstream axial end 162, a downstream hook 24. The upstream 22 and downstream 24 hooks allow the ring sector 16 to be fixed to the fixed annular housing 2 (or casing) of the turbine.

[0054] The fixed annular housing 2 and the turbine ring 4 formed by the ring sectors 16 define between them an annular cooling chamber 30 which is supplied with cooling air via at least one cooling air supply orifice 32, passing through the fixed annular housing 2. The cooling air supplying this annular cooling chamber 30 typically comes from a portion of the outside air which passes through a fan and bypasses the combustion chamber 12 of the turbomachine.

[0055] According to the invention, each ring sector 16 is provided with a first internal cavity 10 and a second internal cavity 20. These internal cavities 10, 20 make it possible to ensure cooling of the ring sectors 16 by jet impact and thermal convection.

[0056] The first internal cavity 10 fluidly communicates with the annular cooling chamber 30 via a first set 40 of orifices, comprising a plurality of orifices 42 passing through the external wall 18. The second internal cavity 20 communicates fluidly with the first internal cavity 10 via a second set 50 of orifices, comprising a plurality of orifices 52 passing through an intermediate wall 19, disposed between the internal wall 17 and the external wall 18. The second internal cavity 20 also communicates fluidly with the passage 14 via discharge orifices 62.

[0057] According to this embodiment, the first set 40 of orifices 42 and the second set 50 of orifices 52 each comprise 36 (thirty-six) orifices 42, 52 respectively, distributed at regular intervals and arranged according to a 6*6 matrix, making it possible to standardize the cooling of the ring sector 16. In addition, twelve discharge orifices 62 are formed on the downstream end 162 of the ring sector, and are distributed at regular intervals and circumferentially around the central axis X.

[0058] Figure 3, Figure 4 (representing the ring sector 16 of Figure 3 along a section plane A) and Figure 5 (representing the ring sector 16 of Figure 3 along a section plane B), illustrate the arrangement of the orifices 42, 52. However, this example is not limiting, the orifices 42, 52, 62 being able to be arranged differently, and present in different numbers, without departing from the scope of the invention.

[0059] The first internal cavity 10 comprises an upstream portion 11, located closer to the upstream end 161 of the ring sector 16, and a downstream portion 13, located closer to the downstream end 162 of the ring sector 16. The upstream portion 11 and the downstream portion 13 are connected to each other by a shoulder 15. Thus, the upstream portion 11 is offset radially inwards relative to the downstream portion 13, and is therefore closer to the passage 14 than the downstream portion 13. In particular, the upstream portion 11 of the first internal cavity 10 is separated from the passage 14 only by the thickness of the internal wall 17.

[0060] This shape of the first internal cavity 10 is obtained thanks to the structure of the external wall 18 and the intermediate wall 19. More precisely, the external wall 18 comprises an upstream external wall portion 181 and a downstream external wall portion 182, connected to each other by an external wall shoulder 183, such that the upstream external wall portion 181 is offset radially inwards relative to the downstream external wall portion 182.

[0061] Furthermore, the intermediate wall 19 comprises an annular portion 191 arranged radially between the internal wall 17 and the downstream portion of the external wall 182, and a radial portion 192 connecting the annular portion 191 and the internal wall 17.

[0062] Thus, the upstream portion 11 of the first internal cavity 10 is formed between the internal wall 17 and the upstream portion of the external wall 181, the downstream portion 13 of the first internal cavity 10 is formed between the annular portion 191 of the intermediate wall 19 and the downstream portion of the external wall 182, and the shoulder 15 of the first internal cavity 10 is formed between the radial portion 192 of the intermediate wall 19 and the shoulder of the external wall 183.

[0063] The second internal cavity 20 is thus delimited by the internal wall 17, the intermediate wall 19, and the downstream end 162 of the ring sector 16, and is arranged radially under the downstream portion 13 of the first internal cavity 10. In particular, the second internal cavity 20 is separated from the passage 14 only by the thickness of the internal wall 17.

[0064] Furthermore, the first set 40 of orifices is formed through the upstream portion of the outer wall 181, and the second set 50 of orifices is formed through the annular portion 191 of the intermediate wall 19, such that the first set 40 of orifices and the second set 50 of orifices are arranged substantially at the same radial position. In other words, the upstream portion of the outer wall 181 and the annular portion 191 of the intermediate wall 19 are arranged at the same distance from the central axis X, the annular portion 191 being arranged downstream of the upstream portion of the outer wall 181.

[0065] Given this arrangement, the cooling air follows a cooling circuit illustrated by the arrows in Figure 2. This cooling circuit comprises the entry of the cooling air into the annular cooling chamber 30 via the orifices 32. The cooling air present in the chamber 30 is then injected into the first internal cavity 10, more precisely into its upstream portion 11 via the orifices 42 of the first set 40 of orifices, and impacts the internal wall 17 directly in contact with the hot gases F, in particular an upstream portion of said internal wall 17. The cooling air then flows into the shoulder 15 then into the downstream portion 13 of the first internal cavity 10.

[0066] The cooling air thus standardized in the first internal cavity 10 after a first injection into said internal cavity 10, is then injected again into the second internal cavity 20 via the orifices 52 of the second set 50 of orifices, and thus again impacts the internal wall 17 directly in contact with the hot gases F, in particular a downstream portion of said internal wall 17. The cooling air present in the second internal cavity 20 is then evacuated to the outside of the ring sector 16, in particular towards the passage 14, via the evacuation orifices 62.

[0067] To allow this flow, the orifices 52 of the second set 50 of orifices have a smaller cross-section than that of the orifices 42 of the first set 40 of orifices, and than that of the discharge orifices 62. Typically, the orifices 42 of the first set 40 of orifices have a diameter of between 0.5 and 1 mm, the orifices 52 of the second set 50 of orifices have a diameter of between 0.3 and 0.5 mm, and the discharge orifices 62 have a diameter of between 1 and 1.5 mm.

[0068] Given this arrangement, the cooling air jets injected into the first internal cavity 10 through the first set 40 of orifices are not disturbed by the cooling air jets injected into the second internal cavity 20 through the second set 50 of orifices, and vice versa.

[0069] A modified example of the embodiment described above will then be described with reference to Figures 6 and 7. Figure 6 represents a portion of the ring sector 16 of Figure 2, in particular the upstream portion thereof, such that only the upstream portion 11 of the first internal cavity 10 is visible.

[0070] This modified example differs from the embodiment described above in that the ring sector 16 comprises disruptors 60. In this non-limiting example, the disruptors 60 are portions of material having the shape of a peak projecting into the first internal cavity 10 from the internal wall 17 towards the external wall 18, more precisely the upstream portion of the internal wall 17 towards the upstream portion of the external wall 181.

[0071] The presence of these disruptors makes it possible to increase the heat exchanges by disturbing the flow after the jet impact on the internal wall 17 and during the transfer phases by thermal convection. In particular, the disruptors 60 make it possible to increase and homogenize the heat exchanges on either side of the jet impacts.

[0072] It will be noted that the disruptors 60 could alternatively, or additionally, be arranged in the second internal cavity 20, projecting from the downstream portion of the internal wall 17 towards the intermediate wall 19. Furthermore, the disruptors 60 could have a different shape, for example a cylindrical or parallelepiped shape.

[0073] Preferably, the disruptors 60 are arranged so as not to be opposite the orifices 42, 52. More precisely, the disruptors 60 are arranged so that each disruptor 60 is not radially aligned with one of the orifices 42, 52. In other words, each orifice 42, 52 comprises a central axis, and each disruptor 60 is offset from these central axes such that none of these central axes passes through a disruptor 60.

[0074] Figure 7 schematically represents an example of arrangement of the disruptors 60 in a top view, that is to say parallel to the radial direction, in the case where the disruptors 60 are arranged on the portion upstream of the inner wall 17, opposite the first set of orifices 40. Although the outer wall 18 is not shown, the orifices 42 of the first set of orifices 40 are superimposed on the inner wall 17 to visualize their positioning. As can be seen in Figure 7, each orifice 42 is surrounded by disruptors 60 in this top view, but none of the orifices 42 is arranged above a disruptor 60.

[0075] This arrangement makes it possible not to hinder, or even cancel out, the beneficial effect of jet impact cooling. In other words, this arrangement makes it possible to benefit from jet impact cooling as described above, while benefiting from an improvement in convection cooling during the passage of the cooling air into the first internal cavity 10, then into the second internal cavity 20.

[0076] Although the present invention has been described with reference to specific exemplary embodiments, it is evident that modifications and changes may be made to these examples without departing from the general scope of the invention as defined by the claims. In particular, instead of extending over the entire circumference of the ring around the central axis X, the internal cavities 10, 20 could be segmented, for example by having radial walls distributed circumferentially around the central axis X. This would limit parasitic effects and thus further improve thermal performance. Furthermore, individual features of the different illustrated / mentioned embodiments may be combined in additional embodiments. Therefore, the description and drawings should be considered in an illustrative rather than restrictive sense.

Claims

Claims

1. Fixed ring (4) extending around a central axis (X), suitable for being mounted on a turbomachine casing (2) to surround a gas passage (14) of the turbomachine (1), the casing (2) delimiting an annular cooling chamber (30) arranged radially outside the ring (4), said ring (4) comprising a plurality of ring sectors (16) arranged circumferentially end-to-end around the central axis (X), ring sector (16) comprising a radially inner wall (17) and a radially outer wall (18) delimiting at least in part a first internal cooling cavity (10) suitable for fluidly communicating with the annular cooling chamber (30) via a first set of orifices (40), and an intermediate wall (19) arranged between the radially inner wall (17) and the radially outer wall (18) and delimiting, with a portion of the radially internal wall (17),a second internal cooling cavity (20) fluidly communicating with the first internal cooling cavity (10) via a second set of orifices (50), the second internal cooling cavity (20) being arranged closer to a downstream end (162) than to an upstream end (161) of each ring sector (16), the first set of orifices (40) being formed in the radially outer wall (18), and the second set of orifices (50) being formed in the intermediate wall (19), the first and second sets of orifices (40, 50) being arranged at the same radial position.,

2. A ring (4) according to claim 1, wherein the first internal cavity (10) comprises an upstream portion (11) and a downstream portion (13), the upstream portion (11) being offset radially inwardly relative to the downstream portion (13).

3. Ring (4) according to claim 2, in which the second internal cavity (20) is arranged radially under the downstream portion (13) of the first internal cavity (10).

4. Ring (4) according to any one of claims 1 to 3, in which the upstream portion (11) of the first internal cavity (10) is capable of fluidly communicating with the annular cooling chamber (30) via the first set of orifices (40), and the downstream portion (13) of the first internal cavity (10) fluidly communicates with the second internal cavity (20) via the second set of orifices (50).

5. Ring (4) according to any one of claims 1 to 4, wherein the orifices (42) of the first set of orifices (40) have a larger diameter than the orifices (52) of the second set of orifices (50).

6. Ring (4) according to any one of claims 1 to 5, in which the downstream end (162) of the ring (4) comprises discharge orifices (62) capable of fluidly connecting the second internal cavity (20) with the gas passage (14) of the turbomachine (1).

7. The ring (4) of claim 6, wherein a diameter of the discharge ports (62) is larger than a diameter of the ports (52) of the second set of ports (50).

8. Ring (4) according to any one of claims 1 to 7, in which the orifices (42) of the first set of orifices (40) and the orifices (52) of the second set of orifices (50) have a diameter less than

9. Ring (4) according to any one of claims 1 to 8, wherein the radially inner wall (17) comprises a plurality of disruptors (60) projecting from an upstream portion of said radially inner wall (17) towards the radially outer wall (18) and / or from a downstream portion of said radially inner wall (17) towards the intermediate wall (19).

10. Ring (4) according to claim 9, wherein each of the disruptors (60) is arranged so as not to be radially opposite one of the orifices (42) of the first set of orifices (40) and / or one of the orifices (52) of the second set of orifices (50).

11. Assembly comprising a turbomachine casing (2) and a ring (4) according to any one of claims 1 to 10 mounted on the casing (2), the casing (2) surrounding the ring (4) and delimiting an annular cooling chamber (30) arranged radially outside the ring (4).

12. A high pressure turbine (1) comprising an assembly according to claim 11.

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