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) arranged 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) through a first set of orifices (40), and a second internal cooling cavity (20) communicating fluidly with the first internal cooling cavity (10) through 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). Figure for the summary: Fig. 2.

Description

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

[0001] The present exposition relates to fixed rings surrounding gas passages of gas turbines, and more particularly to the cooling of fixed gas turbine rings. Previous technique

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

[0003] The turbine ring sectors are thus subjected to the high temperatures of the hot gases from the turbomachine's combustion chamber. For the mechanical and thermal integrity 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 known cooling method 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 homogeneous cooling of the ring sectors, particularly at the upstream end of the ring sector, which is an area especially exposed to hot gases. Indeed, disturbances can occur between jets from adjacent orifices, and between jets from orifices located upstream and downstream of the device. These disturbances, which are all the more pronounced when the cavity to be cooled is large and the number of orifices is significant, reduce the cooling efficiency.

[0007] The lifespan of the ring sectors is therefore affected. Furthermore, and to limit the aforementioned drawbacks related to jet interference, this technology requires a significant intake of cooling air, which reduces turbine performance.

[0008] There is therefore a need for a solution which makes it possible to limit at least in part the aforementioned disadvantages and to improve the efficiency of the cooling of the fixed turbine rings. Description of the invention

[0009] The present description relates to a fixed ring extending around a central axis, suitable for mounting 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 suitable for communicating fluidly with the annular cooling chamber through a first set of orifices, and a second internal cooling cavity communicating fluidly with the first internal cooling cavity through 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 description, 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. Furthermore, the terms "upstream" and "downstream" are considered in the normal direction of gas flow in the turbine along the central axis, specifically 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 the cooling air from the annular cooling chamber flows first into the first internal cooling cavity through the first set of orifices, and then flows from the first internal cooling cavity to the second internal cooling cavity through the second set of orifices.

[0013] Since the second internal cooling cavity is located closer to a downstream end than to an upstream end of the ring sector, i.e., near a downstream axial end of the ring sector, it is possible to initially cool 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] Since the first and second internal cavities are distinct from each other but communicate with one another, it is thus possible to cool an upstream portion of the ring sector without the jets from the first set of orifices being disturbed by the jets from the second set of orifices, thereby cooling a downstream portion of the ring sector. Furthermore, this allows the cooling air to be homogenized in the first internal cavity before being reinjected 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 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 radially offset inwards relative to the downstream portion.

[0017] It is understood that the upstream and downstream portions communicate with each other, for example by means of a shoulder offsetting these portions relative to each other. This allows the upstream portion to be brought closer to the hot air flow, and thus improves 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 arranged radially under 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 positioned closer to the hot air flow duct, relative to the downstream portion of the first internal cavity. This arrangement improves 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 certain embodiments, the upstream portion of the first internal cavity is able to communicate fluidly 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] Given this arrangement, in which the second internal cavity is positioned radially below the downstream portion of the first internal cavity, the cooling air from the annular cooling chamber impacts the upstream portion of the ring sector wall in contact with the hot gases via the orifices of the first set of orifices, and 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 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 internal wall and a radially external wall delimiting at least in part the first internal cavity, and an intermediate wall disposed between the radially internal wall and the radially external wall and delimiting, with a portion of the radially internal wall, the second internal cavity.

[0023] It is understood that an inner face of the radially internal wall delimits the flow path of the hot gases, the radially internal wall and the radially external wall being spaced apart. The radially internal wall delimits both a lower portion of the first internal cavity and the second internal cavity. The cooling air injected through the first and second sets of orifices respectively impacts the radially internal wall, thereby further improving the cooling efficiency.

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

[0025] In some embodiments, the first and second sets of orifices are arranged in 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 have the limiting cross-section for the passage of the cooling airflow and guarantees proper air supply and flow in both cavities. In particular, at a constant flow rate, the smallest drilling cross-section This 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 a lower static pressure in the second cavity than in the first, thus facilitating airflow. Furthermore, accelerating the fluid after it has lost energy upon impact with the internal wall in the first cavity provides it with sufficient energy to impact the internal wall in the second internal cavity. This enhances the cascading effect of the cooling air, thereby improving cooling efficiency.

[0029] In some embodiments, the downstream end of the ring includes discharge ports suitable for fluidly connecting the second internal cavity with the gas passage of the turbomachine.

[0030] In certain embodiments, the diameter of the discharge ports is larger than the diameter of the ports in the second set of ports. This limits the effects of air blockage in the second internal cavity, which would increase the pressure within it. Larger ports also simplify the manufacturing of these discharge ports, as downstream cooling, and therefore fluid acceleration, is not required.

[0031] In some embodiments, a section of the evacuation orifices is larger than a section of the orifices 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 drainage orifices have a diameter between 1 and 1.5 mm.

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

[0035] The present exposition also relates to an assembly comprising a turbomachine housing and a ring according to any of the preceding embodiments, mounted on the housing, the housing surrounding the ring and delimiting an annular cooling chamber arranged radially outside the ring.

[0036] The present description 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 below of various embodiments of the invention given to Title of non-exhaustive examples. This description refers to the attached figure pages, on which:

[0038] [Fig. 1] Fig. 1 schematically represents 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 cross-sectional view of a ring sector according to one embodiment of the invention,

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

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

[0042] [Fig.5] Figure 5 is a perspective and cross-sectional view of the ring sector of [Fig.3] along a cutting plane B. Description of the implementation methods

[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 hot gas flow through a turbomachine, indicated by 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. 1 schematically represents part of a high-pressure turbine 1 of a turbomachine.

[0046] The high-pressure turbine 1 includes, in particular, a fixed annular housing 2 forming a turbomachine casing. A fixed turbine ring 4 is attached to this housing 2 and surrounds a plurality of movable turbine blades 6. These movable blades 6 are arranged downstream of fixed blades 8 with respect to the flow direction F of the hot gases from a combustion chamber 12 of the turbomachine and passing through the turbine. Thus, the turbine ring 4 surrounds a passage 14 (or stream) of hot gas flow.

[0047] 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 continuous circular surface. However, it is also conceivable that the turbine ring consists of a single continuous piece. The present invention applies equally to a single turbine ring and to a turbine ring sector.

[0048] Referring to [Fig. 2], illustrating in cross-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 annular wall 17 (hereinafter "internal wall 17") and a wall radially external 18 (hereinafter "external wall 18") annular offset radially outwards relative to the internal wall 17.

[0049] The inner 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 hooks 22 and downstream hooks 24 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 through 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 blower 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 outer 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 outer wall 18. The second internal cavity 20 also communicates fluidly with the passage 14 via drainage 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 in a 6*6 matrix, allowing for uniform 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 can 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 radially offset 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.

[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 specifically, the external wall 18 comprises an upstream portion of external wall 181 and a downstream portion of external wall 182, connected to each other by an external wall shoulder 183, such that the upstream portion of external wall 181 is offset radially inwards relative to the downstream portion of external wall 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 in substantially 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 located 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 includes the entry of cooling air into the annular cooling chamber 30 through 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 through 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 and then into the downstream portion 13 of the first internal cavity 10.

[0062] The cooling air, thus homogenized 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 through 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 to the passage 14, through 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 between 0.5 and 1 mm, the orifices 52 of the second set 50 of orifices have a diameter between 0.3 and 0.5 mm, and the discharge orifices 62 have a diameter 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 embodiments, it is evident that modifications and changes can 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 the radial walls circumferentially around the central axis X. This would limit parasitic effects and thus further improve thermal performance. Moreover, individual features of the various embodiments illustrated / mentioned can be combined in additional embodiments. Therefore, The description and drawings should be considered in an illustrative rather than restrictive sense.

Claims

Demands

1. A fixed ring (4) extending about a central axis (X), adapted to be mounted on a turbomachine housing (2) to surround a gas passage (14) of the turbomachine (1), the housing (2) defining an annular cooling chamber (30) disposed radially outside the ring (4), said ring (4) comprising a plurality of ring sectors (16) arranged circumferentially end-to-end about the central axis (X), each ring sector (16) comprising a radially inner wall (17) and a radially outer wall (18) defining at least partially a first internal cooling cavity (10) adapted to communicate fluidly with the annular cooling chamber (30) via a first set of orifices (40), and an intermediate wall (19) disposed 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) 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 each ring sector (16), the first set of orifices (40) being formed in the radially external 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 disposed in the same radial position.

2. 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 radially offset inwards relative to the downstream portion (13).

3. Ring (4) according to claim 2, wherein 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, wherein the upstream portion (11) of the first internal cavity (10) is able to communicate fluidly with the annular cooling chamber (30) via the first assembly of orifices (40), and the downstream portion (13) of the first internal cavity (10) communicates fluidly 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, wherein the downstream end (162) of the ring (4) includes discharge ports (62) suitable for fluidly communicating the second internal cavity (20) with the gas passage (14) of the turbomachine (1).

7. Ring (4) according to claim 6, wherein a diameter of the discharge orifices (62) is greater than a diameter of the orifices (52) of the second set of orifices (50).

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

9. Assembly comprising a turbomachine housing (2) and a ring (4) according to any one of claims 1 to 8 mounted on the housing (2), the housing (2) surrounding the ring (4) and defining an annular cooling chamber (30) arranged radially outside the ring (4).

10. High pressure turbine (1) comprising an assembly according to claim 9.

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