JP2005030313A - Shroud segment - Google Patents

Abstract

<P>PROBLEM TO BE SOLVED: To suppress the lowering of the mechanical strength of an overall low-pressure turbine case 3 by minimizing the number of positions on a low-pressure turbine case where stress concentration occurs. <P>SOLUTION: An arc-shaped bump 51 having a bump plane 51fa and a bump arc face 51fb is formed on the surface of a back plate 37 at the rear end side, and a honeycomb cell 59 allowing the contact of the tip part of a turbine blade 19 therewith is integrally formed with the rear surface of the back plate 37. The bump plane 51fa is surface-finished by machining so that a distance between the bump plane 51fa and a case hook plane 49fa when the thermal displacement of the back plate 37 in the shroud axial direction in operating an engine is estimated becomes a specified distance m setting the flow of cooling air CA flowing rearwards between the turbine case 3 and the back plate 37. <P>COPYRIGHT: (C)2005,JPO&NCIPI

Description

  The present invention relates to the technical field of gas turbine engines such as aircraft engines, and more particularly to a shroud segment of a turbine shroud used for a turbine in a gas turbine engine.

  A turbine shroud is used for a turbine in a gas turbine engine such as an aircraft engine, and the turbine shroud is disposed so as to surround a plurality of turbine blades in the turbine case and is heated by the influence of combustion gas. It suppresses. The turbine shroud is usually divided into a plurality of shroud segments so that excessive thermal stress is not generated in the turbine shroud during engine operation.

  Furthermore, a general configuration of the shroud segment will be described as follows. That is, a general shroud segment has an arc-shaped back plate as a segment body, and the front end side of the back plate is held so as to be movable in the shroud axial direction by its own thermal displacement with respect to the turbine case. The rear end of the back plate is held so as not to move in the shroud axial direction with respect to the case hook in the turbine case. A blade contact member that allows contact with the tip of the turbine blade is integrally provided on the back surface of the back plate.

  Further, on the front stage side of the turbine, the cooling air is configured to flow backward between the back plate and the turbine case, and the cooling air passes through the case hook in the turbine case. A plurality of possible through holes for cooling air are formed at appropriate intervals in the circumferential direction. Here, the flow rate of the cooling air flowing between the turbine case and the back plate is set according to the size and number of the through holes for cooling air.

  Therefore, by shielding the turbine case from the combustion gas by the plurality of shroud segments (in other words, by the turbine shroud), the temperature of the turbine case due to the influence of the combustion gas can be suppressed. Furthermore, under the state where the flow rate of the cooling air is adjusted by the plurality of cooling air through holes, the cooling air flows from the front direction between the turbine case and the back plate, so that the cooling action by the cooling air works. Thus, the high temperature of the turbine case can be reliably and sufficiently suppressed.

In addition, there exists a thing shown to patent document 1 as a prior art relevant to this invention.
JP-A-11-62509

  By the way, as described above, a plurality of cooling air through holes are necessary so that a flow rate of cooling air flowing between the turbine case and the back plate can be set, while a plurality of the cooling holes are formed in the case hook. Is processed (formed), the number of places where stress concentration occurs in the turbine case increases, the mechanical strength of the turbine case as a whole decreases, and it is difficult to extend the life of the turbine case. is there.

  Therefore, the present invention is a novel device capable of setting the flow rate of cooling air flowing between the turbine case and the back plate without machining a cooling air through hole in the rear case hook in the turbine case. An object is to provide a shroud segment of construction.

According to the first aspect of the present invention, the shroud segment is formed by dividing the turbine shroud that is disposed so as to surround the plurality of turbine blades in the turbine case and suppresses the high temperature of the turbine case due to the influence of combustion gas. In
An arcuate back plate whose front end side is held immovably in the shroud axial direction with respect to the turbine case, and whose rear end side is held movable with respect to the turbine case in the shroud axial direction by its own thermal displacement;
A bump plane formed on the rear end surface of the back plate, fitable to a rear case hook in the turbine case, parallel to the shroud radial direction and opposed to the case hook plane of the rear case hook; An arc-shaped bump having a bump arc surface orthogonal to the bump plane and capable of facing the case hook arc surface of the rear case hook;
A blade contact member provided integrally with the back surface of the back plate and allowing contact with the tip of the turbine blade;
The flow rate of the cooling air flowing backward between the turbine case and the back plate is a gap between the bump plane and the case hook plane when assuming the thermal displacement of the back plate in the shroud axial direction when the engine is operating. The bump plane is surface-finished by machining so that a predetermined gap is set.

  Note that the front end side of the back plate may be directly held so as not to move in the shroud axial direction with respect to the turbine case, or may be held indirectly. Similarly, even if the rear end side of the back plate is directly held movably in the shroud axial direction by its own thermal displacement (displacement of the back plate) with respect to the turbine case, it is held indirectly. There is no problem.

  According to the invention specific matter of the first aspect, the gap between the bump plane and the case hook plane when the thermal displacement in the shroud axial direction of the back plate during engine operation is assumed becomes the predetermined gap. Further, since the bump plane is surface-finished by machining, the cooling air flowing between the turbine case and the back plate is not processed in the rear case hook in the turbine case without machining the cooling air through hole. The flow rate can be set.

  In addition to the above-described operation, the turbine case is shielded from the combustion gas by a plurality of the shroud segments (in other words, by the turbine shroud), thereby suppressing an increase in the temperature of the turbine case due to the influence of the combustion gas. it can. Further, under the condition that the flow rate of the cooling air is adjusted by the gap between the bump plane and the case hook plane, the cooling air flows backward between the turbine case and the back plate, thereby cooling the cooling air. Therefore, it is possible to reliably and sufficiently suppress the high temperature of the turbine case.

In the invention according to claim 2, in addition to the specific matters of the invention according to claim 1, it is formed on the surface of the front end side of the back plate and is engaged with the peripheral groove of the front case hook in the turbine case. And an arcuate segment hook that is held so as not to move in the shroud axial direction with respect to the circumferential groove of the front case hook.
Here, the segment hook is held immovably in the shroud axial direction with respect to the circumferential groove of the front case hook, so that the front end side of the back plate is held immovable in the shroud axial direction with respect to the turbine case. Will be.

  According to invention of Claim 2, there exists an effect | action similar to the effect | action by the invention specific matter of Claim 1.

  In the invention of claim 3, in addition to the invention specific matter of claim 1 or 2, it is provided on the front end surface of the back plate, and between the back plate and the turbine nozzle of the previous stage. A front seal that suppresses leakage of combustion gas to the turbine case side and suppresses leakage of cooling air from the space between the back plate and the preceding turbine nozzle to the mainstream side is provided.

  According to the invention specific matter of the third aspect, in addition to the action of the invention specific matter of the first or second aspect, the front seal causes the turbine case side to pass from between the back plate and the preceding turbine nozzle. Therefore, the shielding action by the plurality of shroud segments (in other words, the shielding action by the turbine shroud) can be promoted.

  Further, since the front seal can suppress the leakage of cooling air from between the back plate and the preceding turbine nozzle to the mainstream side, the cooling action by the cooling air can be promoted.

  The invention according to claim 4 is configured such that, in addition to the matters specifying the invention according to claim 3, the front end portion of the front seal can be brought into contact with the outer band of the preceding turbine nozzle by its own elastic force. It is characterized by that.

  According to the invention specific matter of claim 4, the same effect as the effect of the invention specific matter of claim 3 is obtained.

  In the invention according to claim 5, in addition to the invention specific matter according to any one of claims 1 to 4, in the rear-stage turbine nozzle at the rear end of the back plate. A rotation notch that can be engaged with a part of the outer band is formed.

  According to the invention specific matter of claim 5, in addition to the action of the invention specific matter of any one of claims 1 to 4, the turbine nozzle at the rear stage is connected to the turbine case. In other words, the turbine notch is not attached to the turbine case by engaging the rotation notch with a part of the outer band of the turbine nozzle in the subsequent stage under the state where the turbine is stopped. The shroud segment can be stopped with respect to the turbine case without processing a fixing pin mounting hole in the case.

  According to the first aspect of the present invention, the flow rate of the cooling air flowing between the turbine case and the back plate can be set without forming a cooling air through hole in the rear case hook in the turbine case. It is possible to extend the life of the turbine case by reducing the number of places where stress concentration occurs in the turbine case and suppressing the decrease in mechanical strength of the entire turbine case.

  According to invention of Claim 2, there exists an effect similar to the effect of invention of Claim 1.

  According to the invention described in claim 3, in addition to the effects of the invention described in claim 1 or 2, it is possible to promote the shielding action by the plurality of shroud segments and the cooling action by the cooling air. The turbine shroud can efficiently suppress the temperature increase of the turbine case due to the influence of combustion gas, and the life of the turbine case can be sufficiently extended.

  According to invention of Claim 4, there exists an effect similar to the effect of invention of Claim 4.

  According to the invention described in claim 5, in addition to the effects of the invention described in any one of claims 1 to 4, machining of the mounting hole for the rotation pin in the turbine case. The shroud segment can be stopped with respect to the turbine case, so the number of mounting holes formed in the turbine case is reduced, and the number of places where stress concentration occurs in the turbine case is further reduced. The life of the turbine case can be further extended.

  The low-pressure turbine in the aircraft engine according to the best mode and the shroud segment according to the best mode will be described with reference to FIGS.

  Here, FIG. 1 is an enlarged view of a portion I in FIG. 5, FIG. 2 is an enlarged view of a portion II in FIG. 1, and FIG. 3 is a front view of the shroud segment according to the best mode. FIG. 4 is a plan view of the shroud segment according to the best mode, and FIG. 5 is a partial cross-sectional view of the low-pressure turbine in the aircraft engine according to the best mode.

  Further, “front and back” means the front and back in FIG. 1, FIG. 2, FIG. 4, FIG. The orientation of each drawing will be described based on the state at the time of publication of the patent publication.

  As shown in FIG. 5, the low-pressure turbine 1 in the aircraft engine according to the best mode includes a low-pressure turbine case 3 as a low-pressure turbine body. The low-pressure turbine case 3 includes a main turbine case 5 and a rear turbine case 7 integrally provided on the rear end side of the main turbine case 5, and the front end side of the main turbine case 5 is a high-pressure turbine in a high-pressure turbine. It is connected to the case 9.

  In the main turbine case 5, a plurality of stages of turbine nozzles 11, 13, 15, and 17 that rectify the combustion gas are arranged at appropriate intervals in the case axial direction (front-rear direction). , 17 are segmented. In the main turbine case 5, a plurality of rotatable turbine disks (not shown) are alternately arranged with the turbine nozzles 11, 13, 15, and 17 at appropriate intervals in the case axial direction. A plurality of turbine blades 19, 21, 23, 25 are provided on the outer periphery of the turbine disk (only one is shown in FIG. 5 for each stage). Here, the multi-stage turbine disk is integrally connected, and the multi-stage turbine disk is integrally connected to the low-pressure compressor rotor (not shown) of the low-pressure compressor and the fan rotor (not shown) of the fan. Has been. Furthermore, in the main turbine case 5, a plurality of turbine shrouds 27, 29, 31, and 33 that suppress high temperature due to the influence of combustion gas surround a plurality of corresponding turbine nozzles 11, 13, 15, and 17. The turbine shrouds 27, 29, 31, and 33 are each segmented.

  Accordingly, the plurality of stages of the turbine disks are integrally rotated by the expansion of the combustion gas from the combustor (not shown), so that the low pressure turbine 1 can obtain a driving force and the plurality of stages of the low pressure compressors. The rotor and the fan rotor can be integrally rotated to drive the low-pressure compressor and the fan in conjunction with each other.

  As shown in FIGS. 1 to 3, the shroud segment 35 according to the best mode is obtained by dividing the first stage turbine shroud 27 in the low-pressure turbine 1, and an arc-shaped back plate 37 is used as a segment body. It has. The back plate 37 is formed integrally with the first plate portion 37a extending in the shroud axial direction (case axial direction, front-rear direction), and the front end edge of the first plate portion 37a, and the shroud axial center side (from FIG. 1). And a second plate portion 37b protruding downward (in FIG. 3).

  An arc-shaped segment hook 43 that can be engaged with the circumferential groove 41 of the front case hook 39 in the main turbine case 5 is integrally formed on the surface of the front end side of the back plate 37. The U-shaped C clip 45 is held so as not to move in the shroud axial direction with respect to the circumferential groove 41 of the front case hook 39. Here, since the segment hook 43 is held so as not to move in the shroud axial direction with respect to the circumferential groove 41 of the front case hook 39, the front end side of the back plate 37 cannot move in the shroud axial direction with respect to the main turbine case 5. Will be held.

  Further, the rear end side of the back plate 37 has its own thermal displacement (with respect to the back plate 37) with respect to the main turbine case 5 by the cooperation of the outer band 47 in the rear turbine nozzle 13 and the rear case hook 49 in the main turbine case 5. It is held so as to be movable in the shroud axis direction by thermal displacement.

  As shown in FIGS. 1 and 2, an arc-shaped bump 51 that can be fitted to a rear case hook 49 in the main turbine case 5 is formed on the surface of the rear end side of the back plate 37. A bump plane 51fa that is parallel to the shroud radial direction (vertical direction in FIGS. 1 and 2) and that can face the case hook plane 49fa of the rear case hook 49, and a case orthogonal to the bump plane 51fa and the case of the rear case hook 49 And a bump arc surface 51fa that can be opposed to the hook arc surface 49fa. Further, on the bump arc surface 51fb of the bump 51, an arc-shaped projection row 53 that can contact the case hook arc surface 49fb of the rear case hook 49 is formed.

  Here, the gap between the bump plane 51fa and the case hook plane 49fa assuming the thermal displacement in the shroud axial direction of the back plate 37 during engine operation flows backward between the main turbine case 5 and the back plate 37. The bump plane 51fa is surface-finished by machining in the same manner as the case hook plane 49fa so as to be a predetermined gap m that sets the flow rate of the cooling air CA. When the bump 51 is fitted to the rear case hook 49, the gap between the bump arc surface 51fb and the case hook arc surface 49fb is always larger than the gap between the bump plane 51fa and the case hook plane 49fa. It is set to be large. The cooling air CA is air compressed by the compressor and is introduced between the main turbine case 5 and the back plate 37 through the introduction hole 55h of the outer band 55 in the turbine nozzle 11 at the preceding stage.

   As shown in FIGS. 1 and 4, a rotation stop notch 57 that can engage with a part (tab) 47 a of the outer band 47 in the turbine nozzle 13 at the rear stage is formed at the rear end of the back plate 37.

  As shown in FIG. 1, a honeycomb cell (an example of a turbine contact member) 59 that allows contact with the tip fin (tip portion) 19 a of the turbine blade 19 is integrally provided on the back surface of the back plate 37. Note that another turbine contact member may be used instead of the honeycomb cell 59.

  An arc-shaped front seal 61 is integrally provided on the front end surface of the back plate 37, and this front seal is provided between the back plate 37 and the turbine nozzle 11 in the preceding stage to the low pressure turbine case 3 side. While suppressing a leak, the leak of the cooling air CA from the back plate 37 and the turbine nozzle 11 of the front | former stage to the mainstream side (lower side in FIG. 1) is suppressed. Further, the front seal 61 is configured such that the front end portion thereof can come into contact with the outer band 55 of the preceding turbine nozzle 11 by its own elastic force (elastic force of the front seal 61).

  A pair of first seal grooves 65 (only one is shown in FIG. 1) into which side portions of the first spline seal plate 63 can be inserted are formed on both side end surfaces of the back plate 37, respectively. The seal grooves 65 are each configured to extend in the shroud axial direction from the vicinity of the front end to the vicinity of the rear end of the first plate portion 37a. Here, the first spline seal plate 63 suppresses leakage of combustion gas from the gap between the first plate portions 37a of the pair of adjacent shroud segments 35 to the low-pressure turbine case 3 side, and also prevents the pair of adjacent shroud segments 35 from adjoining. The leakage of the cooling air CA from the gap of the first plate portion 37a to the mainstream side is suppressed.

  Further, a pair of second seal grooves 69 (only one is shown in FIG. 1) into which side portions of the second spline seal plate 67 can be inserted are formed on both side end surfaces of the back plate 37, respectively. The second seal grooves 69 are each configured to extend in the shroud radial direction from the proximal end to the vicinity of the distal end of the second plate portion 37b. The second spline seal plate 67 suppresses the leakage of combustion gas from the gap between the second plate portions 37b of the pair of adjacent shroud segments 35 to the low pressure turbine case 3 side, and the second spline seal plate 67 The leakage of the cooling air CA from the gap between the two plate portions 37b to the mainstream side is suppressed.

  Further, as shown in FIGS. 1, 3, and 4, a pair of seal gaps 71 are formed at the front portions of both end surfaces of the back plate 37, and each seal gap 71 has a first seal groove 65. And the base end side of the second seal groove 69 respectively.

  Next, the operation of the best mode will be described.

  The bump flat surface 51fa is surface-finished by machining so that the gap between the bump flat surface 51fa and the case hook flat surface 49fa when the engine is operated is assumed to be thermally displaced in the shroud axial direction. Therefore, the flow rate of the cooling air CA flowing between the main turbine case 5 and the back plate 37 can be set without processing the through hole for cooling air in the rear case hook 49 in the main turbine case 5.

  Further, under the state in which the rear stage turbine nozzle 13 is rotated with respect to the low pressure turbine case 3, the rotation notch 57 is engaged with a portion 47 a of the outer band 47 in the rear stage turbine nozzle 13, whereby the low pressure turbine case. The shroud segment 35 can be rotated with respect to the low-pressure turbine case 3 without attaching a rotation-stopping pin to the low-pressure turbine case 3, in other words, without forming a rotation-pin mounting hole in the low-pressure turbine case 3.

  In addition to the above-described operation, the first spline seal plate 63, the second spline seal plate 67 and the like suppress the leakage of the combustion gas, while the plurality of shroud segments 35 (in other words, by the turbine shroud 27) reduce the combustion gas from the low pressure turbine. By shielding the case 3, the high temperature of the low-pressure turbine case 3 due to the influence of the combustion gas can be suppressed. Further, under the state where the flow rate of the cooling air CA is adjusted by the gap between the bump plane 51fa and the case hook plane 49fa while suppressing the leakage of the cooling air CA by the first spline seal plate 63, the second spline seal plate 67, and the like. As the cooling air CA flows backward between the main turbine case 5 and the back plate 37, the cooling action by the cooling air CA works, and the high temperature of the low-pressure turbine case 3 can be reliably and sufficiently suppressed.

  Here, in addition to the first spline seal plate 63 and the second spline seal plate 67, the front seal 61 suppresses the leakage of combustion gas from between the back plate 37 and the preceding turbine nozzle 11 to the low-pressure turbine case 3 side. Therefore, the shielding action by the plurality of shroud segments 35 (in other words, the shielding action by the turbine shroud 27) can be promoted. In addition to the first spline seal plate 63 and the second spline seal plate 67, the front seal 61 can suppress the leakage of the cooling air CA from the space between the back plate 37 and the preceding turbine nozzle 11 to the mainstream side. The cooling action by the cooling air CA can be promoted.

  As described above, according to the best mode, the flow rate of the cooling air CA flowing between the main turbine case 5 and the back plate 37 is reduced without machining the cooling air through hole in the rear case hook 49 of the main turbine case 5. Since the shroud segment 35 can be rotated with respect to the low pressure turbine case 3 without machining the mounting hole for the rotation stop pin in the low pressure turbine case 3, the stress concentration occurs in the low pressure turbine case 3. As much as possible, the decrease in mechanical strength of the low-pressure turbine case 3 as a whole can be suppressed, and the life of the low-pressure turbine case 3 can be extended.

  Particularly, since the shielding action by the plurality of shroud segments 35 and the cooling action by the cooling air CA can be promoted, the turbine shroud 27 can efficiently suppress the high temperature of the low-pressure turbine case 3 due to the influence of the combustion gas. The life of the low-pressure turbine case 3 can be made sufficiently longer.

  The present invention is not limited to the description of the above-described best mode, and can be implemented in various other modes by making appropriate modifications.

It is an enlarged view of the I section in FIG. It is an enlarged view of the II section in FIG. It is a front view of the shroud segment concerning the best form. It is a top view of the shroud segment concerning the best form. 1 is a partial cross-sectional view of a low-pressure turbine in an aircraft engine according to the best mode.

Explanation of symbols

DESCRIPTION OF SYMBOLS 1 Low pressure turbine 3 Low pressure turbine case 5 Main turbine case 7 Rear turbine case 35 Shroud segment 37 Back plate 39 Front case hook 41 Circumferential groove 49 Rear case hook 49fa Case hook plane 49fb Case hook arc surface 51 Bump 51fa Bump plane 51fb Bump arc surface 59 Honeycomb cell 61 Front seal

Claims (5)

In a shroud segment formed by dividing a turbine shroud that is arranged so as to surround a plurality of turbine blades in a turbine case and suppresses a high temperature of the turbine case due to the influence of combustion gas,
An arcuate back plate whose front end side is held immovably in the shroud axial direction with respect to the turbine case, and whose rear end side is held movable with respect to the turbine case in the shroud axial direction by its own thermal displacement;
A bump plane formed on the rear end surface of the back plate, fitable to a rear case hook in the turbine case, parallel to the shroud radial direction and opposed to the case hook plane of the rear case hook; An arc-shaped bump having a bump arc surface orthogonal to the bump plane and capable of facing the case hook arc surface of the rear case hook;
A blade contact member provided integrally with the back surface of the back plate and allowing contact with the tip of the turbine blade;
The flow rate of the cooling air flowing backward between the turbine case and the back plate is a gap between the bump plane and the case hook plane when assuming the thermal displacement of the back plate in the shroud axial direction when the engine is operating. The shroud segment is characterized in that the bump plane is surface-finished by machining so that a predetermined gap is set. It is formed on the front end surface of the back plate, can be engaged with the circumferential groove of the front case hook in the turbine case, and is held immovable in the shroud axial direction with respect to the circumferential groove of the front case hook. The shroud segment according to claim 1, comprising an arc segment hook. Provided on the front end surface of the back plate, suppresses leakage of combustion gas from between the back plate and the preceding turbine nozzle to the turbine case side, and from the back plate to the main stream side from the preceding turbine nozzle. The shroud segment according to claim 1, further comprising: a front seal that suppresses leakage of cooling air. The shroud segment according to claim 3, wherein a front end portion of the front seal is configured to come into contact with an outer band of a turbine nozzle at a front stage by its own elastic force. The rotation stop notch which can be engaged with a part of outer band in the turbine nozzle of the back | latter stage is formed in the rear end of the said backplate, The claim in any one of Claims 1-4 characterized by the above-mentioned. The shroud segment described.

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