Improved gas turbine fixed ring cooling system

Abstract

Improved fixed ring cooling system for gas turbine. Fixed ring (4) extending around a central axis (X), suitable for mounting on a turbomachine casing (2) to surround a gas passage (14) of the turbomachine (1), the casing (2) delimiting an annular cooling chamber (30) disposed radially outside the ring (4), said ring (4) comprising a first internal cooling cavity (10) suitable for communicating fluidly with the annular cooling chamber (30) via a first set of orifices (40), and a second internal cooling cavity (20) communicating fluidly with the first internal cooling cavity (10) via a second set of orifices (50), the second internal cooling cavity (20) being disposed closer to a downstream end (162) than to an upstream end (161) of the ring sector (16). Figure for the summary: Fig. 2.

Description

Title of the invention: Improved gas turbine fixed ring cooling system Technical field

[0001] The present 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 ring sectors of the turbine 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 provide 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 efficient and uniform cooling of the ring sectors to be obtained, in particular 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, all the more marked when the cavity to be cooled is large and the number of orifices is large, reduce the efficiency of the cooling.

[0007] The lifetime of the ring sectors is therefore affected. Furthermore, and for 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 making it possible to limit at least in part the aforementioned drawbacks and making it possible to improve the efficiency of cooling 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 communicating fluidly with the annular cooling chamber via a first set of orifices, and a second internal cooling cavity communicating fluidly 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 the present 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 according to a normal direction of flow of the gases 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 orifices, and then flows from the first internal cooling cavity to the second internal cooling cavity via the second set of orifices.

[0013] The second internal cooling cavity being arranged closer to a downstream end than to an upstream end of the ring sector, that is to say close to a downstream axial end of the ring sector, it is possible to cool initially an upstream portion of the ring sector by means of the first internal cavity, then a downstream portion of the ring sector by means of of 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 the cooling efficiency. In other words, this arrangement makes it possible to reduce the cooling air flow rate required to obtain the same thermal performance on the ring. This also makes it possible to increase the temperature of the hot air flow stream and therefore the performance of the gas turbine and the turbomachine.

[0016] In some embodiments, the first internal cavity comprises an upstream portion and a downstream portion, the upstream portion being offset radially inwardly relative to 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, relative to the downstream portion of the first internal cavity. This arrangement makes it possible to improve the efficiency of the cooling 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 orifices, and the downstream portion of the first internal cavity communicates with the second internal cavity via the second set of orifices.

[0021] Considering 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 impacts the upstream portion of the wall of the ring sector 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 wall of the sector 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 makes it possible to further improve 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 orifices are arranged at the same radial position.

[0026] It is understood that the orifices of the first and second sets 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 orifices of the second set of orifices are those which have the limiting section for the passage of the cooling air flow and ensures the correct supply and flow of air in the two cavities. In particular, at constant flow rate, the smallest drilling section corresponds to the fastest fluid speed, and therefore the lowest static pressure. Placing the smallest section between the first and second internal cavities makes it possible to ensure a lower static pressure in the second cavity 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 thus providing it with enough energy to impact the internal wall in the second internal cavity. Thus, this improves the cascade flow effect of the cooling air, and therefore improves 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 discharge orifices is larger than a diameter of the orifices of the second set of orifices. This makes it possible to limit the effects of air blocking in the second internal cavity, which would increase the pressure in the latter. Larger orifices also make it easier to manufacture these discharge orifices, since cooling downstream of the latter, and therefore acceleration of the fluid, is not necessary.

[0031] In some embodiments, a section of the discharge 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 discharge orifices have a diameter between 1 and 1.5 mm.

[0034] These values ​​make it possible 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] 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.

[0036] The present disclosure also relates to a high pressure turbine comprising an assembly defined above. Brief description of the drawings

[0037] 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:

[0038] [Fig-1] [Fig.l] schematically represents a part of a gas turbine illustrating the location of a fixed ring relative to that of the moving blades,

[0039] [Fig.2] [Fig.2] is a longitudinal sectional view of a ring sector according to one embodiment of the invention,

[0040] [Fig.3] [Fig.3] is a perspective view of the ring sector alone of [Fig.2],

[0041] [Fig.4] [Fig.4] is a perspective and sectional view of the ring sector of the [Fig.3] according to a section plane A,

[0042] [Fig.5] Figures 5 is a perspective and sectional view of the ring sector of [Fig.3] along a sectional plane B. Description of the embodiments

[0043] An embodiment of the invention will be described with reference to Figures 1 to 5.

[0044] The terms “upstream” and “downstream” are subsequently defined in relation to the meaning 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.

[0045] [Fig.l] schematically represents a part of a high-pressure turbine 1 of a turbomachine.

[0046] 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.

[0047] Generally, the turbine ring 4 is composed 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 and the same continuous part. The present invention applies equally to a single turbine ring and to a turbine ring sector.

[0048] Referring to [Fig. 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.

[0049] The internal wall 17 is opposite the passage 14 for the flow of hot gases F. Each ring sector 16 further 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.

[0050] 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.

[0051] 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.

[0052] The first internal cavity 10 communicates fluidly 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.

[0053] 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.

[0054] [Fig.3], [Fig.4] (representing the ring sector 16 of [Fig.3] according to a section plane A) and [Fig.5] (representing the ring sector 16 of [Fig.3] according to 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.

[0055] 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 inward 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.

[0056] 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.

[0057] 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.

[0058] 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.

[0059] 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.

[0060] 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.

[0061] Given this arrangement, the cooling air follows a cooling circuit illustrated by the arrows in [Fig. 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 comes impact 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.

[0062] 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.

[0063] 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.

[0064] Given this arrangement, the cooling air jets injected into the first internal cavity 10 by the first set 40 of orifices are not disturbed by the cooling air jets injected into the second internal cavity 20 by the second set 50 of orifices, and vice versa.

[0065] 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 arranging 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 various 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 first internal cooling cavity (10) suitable for fluidly communicating with the annular cooling chamber (30) via a first set of orifices (40), and 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 the ring sector (16).

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, comprising a plurality of ring sectors (16) arranged circumferentially end-to-end around the central axis (X), each ring sector (16) comprising a radially inner wall (17) and a radially outer wall (18) delimiting at least in part the first internal cavity (10), 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 inner wall (17), the second internal cavity (20).

6. Ring (4) according to claim 5, wherein the first set of orifices (40) is formed in the radially outer wall (18), and the second set of orifices (50) is formed in the intermediate wall (19).

7. Ring (4) according to any one of claims 1 to 6, in which the first and second set of orifices (40, 50) are arranged at the same radial position.

8. Ring (4) according to any one of claims 1 to 7, 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).

9. Ring (4) according to any one of claims 1 to 8, 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).

10. The ring (4) of claim 9, wherein a diameter of the discharge ports (62) is larger than a diameter of the ports (52) of the second set of ports (50).

11. Ring (4) according to any one of claims 1 to 10, 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 of less than 1 mm.

12. Assembly comprising a turbomachine casing (2) and a ring (4) according to any one of claims 1 to 11 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).

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

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