Apparatus and method for cooling a turbine shroud segment and vane outer shroud

Description

TECHNICAL FIELD

[0001] The invention relates generally to gas turbine engines and more particularly to turbine shroud and downstream vane outer shroud cooling.

BACKGROUND OF THE ART

[0002] A gas turbine engine usually includes a hot section, i.e., a turbine section which includes at least one rotor stage, for example, having a plurality of shroud segments disposed circumferentially one adjacent to another to form a shroud ring surrounding a turbine rotor, and at least one stator vane stage disposed immediately downstream and/or upstream of the rotor stage, formed with outer and inner shrouds and a plurality of radial stator vanes extending therebetween. Being exposed to very hot gases, the rotor stage and the stator vane stage need to be cooled. Since flowing coolant through the rotor and stator vane stages diminishes overall engine performance, it is typically desirable to minimize cooling flow consumption without degrading durability of components of the turbine section. Hereintofore, in various approaches for development of adequate cooling arrangements, efforts have been directed to reuse cooling air used for cooling of an upstream component of the turbine section. For example, gas turbine engine designers have been continuously seeking improved turbine section cooling arrangements for cooling turbine shroud segments, and for further reuse of the turbine shroud segment cooling air in an effective manner to cool a downstream stator vane stage.

[0003] Accordingly, there is a need to provide an improved cooling arrangement in a gas turbine engine for effectively cooling in a serial flow relationship, both turbine shroud segments and a stator vane stage thereof.

[0004] A turbine shroud segment having the features of the preamble of claim 1 is disclosed in US 6899518 B2.

SUMMARY OF THE INVENTION

[0005] It is therefore an object of the present invention to provide a cooling arrangement in a gas turbine engine for cooling a portion of an outer wall of an annular gas path of a turbine section.

[0006] The present invention therefore provides a turbine shroud segment as set forth in claim 1.

[0007] The present invention also provides gas turbine engine structure as set forth in claim 6.

[0008] The present invention also provides a method of reusing turbine shroud cooling air for impingement cooling on a downstream turbine vane outer shroud as set forth in claim 10.

[0009] These and other aspects of the present invention will be better understood with reference to preferred embodiments described hereinafter.

DESCRIPTION OF THE DRAWINGS

[0010] Reference is now made to the accompanying figures depicting aspects of the present invention, in which:

Figure 1 is a schematic cross-sectional view of a gas turbine engine;

Figure 2 is an axial cross-sectional view of a turbine shroud assembly used in the gas turbine engine of Figure 1, in accordance with one embodiment of the present invention; and

Figure 3 is a perspective view of a shroud segment used in the turbine shroud assembly of Figure 2.

DETAILED DESCRIPTION OF THE PREFERRED EMBODIMENTS

[0011] Referring to Figure 1, a turbofan gas turbine engine incorporates an embodiment of the present invention, presented as an example of the application of the present invention, and includes a housing or a nacelle 10, a core casing 13, a low pressure spool assembly seen generally at 12 which includes a fan 14, low pressure compressor 16 and low pressure turbine 18, and a high pressure spool assembly seen generally at 20 which includes a high pressure compressor 22 and a high pressure turbine 24. There is provided a burner 25 for generating combustion gases. The low pressure turbine 18 and high pressure turbine 24 include a plurality of rotor stages 28 and stator vane stages 30.

[0012] Referring to Figures 1-3, each of the rotor stages 28 has a plurality of rotor blades 33 encircled by a turbine shroud assembly 32 and each of the stator vane stages 30 includes a stator vane assembly 34 which is positioned upstream and/or downstream of a rotor stage 31, for directing combustion gases into or out of an annular gas path 36 within a corresponding turbine shroud assembly 32, and through the corresponding rotor stage 31.

[0013] The stator vane assembly 34, for example a first stage of a low pressure turbine (LPT) vane assembly, is disposed, for example, downstream of the shroud assembly 32 of one rotor stage 28, and includes, for example a plurality of stator vane segments (not indicated) joined one to another in a circumferential direction to form a turbine vane outer shroud 38 which comprises a plurality of axial stator vanes 40 (only a portion of one is shown) which divide a downstream section of the annular gas path 36 relative to the rotor stage 28, into sectoral gas passages for directing combustion gas flow out of the rotor stage 28.

[0014] The shroud assembly 32 in the rotor stage 28 includes a plurality of shroud segments 42 (only one shown) each of which includes a platform 44 having front and rear radial legs 46, 48 with respective hooks (not indicated). The shroud segments 42 are joined one to another in a circumferential direction and thereby form the shroud assembly 32.

[0015] The platform 44 of each shroud segment 42 has outer and inner surfaces 50, 52 and is defined axially between leading and trailing ends 54, 56, and circumferentially between opposite sides 58, 60 thereof. The platforms 44 of the segments collectively form a turbine shroud ring (not indicated) which encircles the rotor blades 33 and in combination with the rotor stage 28, defines a section of the annular gas path 36. The turbine shroud ring is disposed immediately upstream of and abuts the turbine vane outer shroud 38, to thereby form a portion of an outer wall (not indicated) of the annular gas path 36.

[0016] The front and rear radial legs 46, 48 are axially spaced apart and integrally extend from the outer surface 50 radially and outwardly such that the hooks of the front a rear radial legs 46, 48 are conventionally connected with an annular shroud support structure 62 which is formed with a plurality of shroud support segments (not indicated) and is in turn supported within the core casing 13. An annular cavity 64 is thus defined axially between the front and rear legs 46, 48 and radially between the platforms 44 of the shroud segments 42 and the annular shroud support structure 62. The annular middle cavity is in fluid communication with a cooling air source, for example bleed air from the low or high pressure compressors 16, 22 and thus the cooling air under pressure is introduced into and accommodated within the annular cavity 64.

[0017] The platform 44 of each shroud segment 42 preferably includes a passage, for example a plurality of holes 66 extending axially within the platform 44 for directing cooling air therethrough for transpiration cooling of the platform 44. For convenience of the hole drilling, a groove 68 extending in a circumferential direction with opposite ends closed is provided, for example, on the outer surface 50 of the platform 44 such that holes 66 can be drilled from the trailing end 56 of the platform straightly and axially towards and terminate at the groove 68. Thus, the groove 68 forms a common inlet of the holes 66 for intake of cooling air accommodated within the cavity 64. However, other types of outlets can be made to achieve the convenience of the hole drilling process. It is also preferable to provide one or more outlets of the holes 66 in order to adequately discharge the cooling air from the holes 66 and reduce the contact surface of the trailing end 56 of the platform 44 of the shroud segments 42 with respect to the turbine vane outer shroud 38. For example, an elongate recess 70 is provided in the trailing end 56 of the platform 44 with an opening on the inner surface 52 of the platform 44, thereby forming a common outlet of the holes 66 to discharge the cooling air, for example to the gas path 36. Other types of outlets can be used for adequately discharging the cooling air from the holes 66.

[0018] The groove 68 is in fluid communication with the middle cavity 64 and thus cooling air introduced into the cavity 64 is directed into and through the axial holes 66 for effectively cooling the platform 44 of the shroud segments 42. The cooling air is then discharged through the elongate recess 70 at the trailing end 56 of the platform 42, impinging on a downstream engine part such as the turbine vane outer shroud 38, before entering the gas path 36.

[0019] The groove 68 which functions as the common inlet of the holes 66 is preferably located close to the front leg 46 such that the holes 66 extend through a major section of the entire axial length of the platform 44 of the shroud segment 42, thereby efficiently cooling the platform 44 of the shroud segment 42.

[0020] The holes 66 are preferably substantially evenly spaced apart in a circumferential direction and are preferably aligned with the turbine vane outer shroud. Thus, the cooling air impinges on the leading end of the turbine vane outer shroud 38. The number of holes 66 in each shroud segment 42 is determined such that the cooling air discharged from the holes 66 effectively cools the entire circumference of the leading end of the turbine vane outer shroud 38.

[0021] During engine operation, cooling air introduced into the cavity 64 is directed within and through the platforms 44 of the shroud segments 42 via the holes 66, in order to cool the turbine shroud ring. Simultaneous with cooling of the turbine shroud ring, cooling air flowing through the substantially axial straight holes 66 forms a plurality of substantially straight axial cooling air streams directed towards the leading end of the turbine vane outer shroud 38, preferably with a high velocity thereof, for impingement cooling on the turbine vane outer shroud 38. In contrast to conventional impingement holes defined in a plate, the substantially axial straight holes 66 direct the cooling air through the entire length thereof, thereby forming substantially straight cooling air streams with relatively high directionality. In other words, the substantially straight cooling air streams are more individually focused and interfere less with adjacent air streams when approaching the leading end of the turbine vane outer shroud 38, which results in greater impingement effects on the leading end of the turbine vane outer shroud 38. Due to the restriction of the elongate recess 70 in the trailing end 56 of the platforms 44, the cooling air upon impingement on the leading end of the turbine vane outer shroud 38, is then directed radially, inwardly and rearwardly, thereby further film cooling a front portion of the inner surface of the turbine vane outer shroud 38 and a portion of the axial stator vanes 40, prior to being discharged into the annular gas path 36.

[0022] The above description is meant to be exemplary only, and one skilled in the art will recognize that changes may be made to the embodiments described without departure from the scope of the invention disclosed. For example, the present invention can be applicable in any type of gas turbine engine other than the described turbofan gas turbine engine. Other modifications which fall within the scope of the present invention will be apparent to those skilled in the art, in light of a review of this disclosure, and such modifications are intended to fall within the appended claims.

Claims

1. A turbine shroud segment (42) for a cooling arrangement in a turbine section of a gas turbine engine for cooling in a serial flow relationship, both the turbine shroud segment (42) and a turbine vane outer shroud (38) disposed downstream of the turbine shroud segment (42), the turbine shroud segment comprising a plurality of passages (66) extending in a shroud platform (44) for directing cooling air therethrough to cool the turbine shroud segment (42) and to discharge the cooling air at a trailing end (56) of the shroud platform (44) for impingement on the turbine vane outer shroud (38), the shroud platform (44) comprising a common outlet (70) for said plurality of passages (66) on the trailing end of said platform (44); characterised in that said common outlet (70) has an opening on the radially inner surface (52) of the platform (44).

2. The shroud segment as claimed in claim 1 wherein the plurality of passages (66) extend through a major section of an entire axial length of the shroud platform (44).

3. The shroud segment as claimed in claim 2 wherein the plurality of passages (66) are in fluid communication with a cavity (64) defined between front and rear legs (46, 48) of the turbine shroud segment (42).

4. The shroud segment as claimed in any preceding claim wherein the plurality of passages (66) comprises a plurality of substantially axial straight holes.

5. A gas turbine engine structure for defining a portion of an outer wall of an annular gas path of a turbine section, comprising a turbine shroud (32) comprising a plurality of turbine shroud segments (42) as claimed in claim 1, and a turbine vane outer shroud (38) with a plurality of vanes (40) disposed immediately downstream of the turbine shroud, the passages (66) substantially aligning with the turbine vane outer shroud (38), whereby cooling air delivered from said passages impinges over substantially the entire extent of a circumference of a leading end of the turbine vane outer shroud (38).

6. The gas turbine engine structure as claimed in claim 5 wherein the passages (66) comprise a plurality of holes having at least one inlet (68) thereof on an outer surface (50) of the platform (44).

7. The gas turbine engine structure as claimed in claim 6 wherein the at least one inlet (68) of the holes (66) is located in a position close to and downstream a front leg (46) of the shroud.

8. The gas turbine engine structure as claimed in claim 6 or 7 wherein the holes (66) extend in a substantially straight direction over a major section of the entire axial length of the platform (44).

9. A method of reusing turbine shroud cooling air for impingement cooling on a downstream turbine vane outer shroud (38), the method comprising steps of:

(a) directing cooling air within and through a platform (44) of a shroud segment (42) as claimed in claim 4 of a turbine shroud for cooling the turbine shroud; and

(b) using the cooling air of step (a) to form a plurality of substantially straight cooling air streams axially towards a leading end of the turbine vane outer shroud (38) for impingement cooling on the turbine vane outer shroud.

10. The method as claimed in claim 9 wherein steps (a) and (b) are conducted substantially simultaneously.

11. The method as claimed in claim 10 wherein steps (a) and (b) are practiced by directing the cooling air through a plurality of substantially axial and straight passages (66) extending within the platform (44) of the shroud segment (42) of the turbine shroud to deliver the substantially straight cooling air streams with a high velocity thereof.

12. The method as claimed in claim 11 wherein the cooling air is directed from a cavity (64) defined between front and rear legs of the shroud segment (42), into the substantially axial and straight passages (66).

13. The method as claimed in any of claims 9 to 12 further comprising a step (c) of discharging the cooling air into a gas path upon the impingement cooling thereof on the turbine vane outer shroud (38).

14. The method as claimed in any of claims 9 to 13 comprising directing the substantially straight cooling air streams in a manner for impingement cooling on a substantially entire circumference of the leading end of the turbine vane outer shroud (38).

Ansprüche

1. Turbinen-Mantelringsegment (42) für eine Kühlanordnung in einem Turbinenabschnitt eines Gasturbinenmotors zum Kühlen, in einem Reihenströmungsverhältnis, sowohl des Turbinen-Mantelringsegments (42) als auch eines Deckbands einer Turbinenleitschaufel (38), das stromabwärts von dem Turbinen-Mantelringsegment (42) angeordnet ist, wobei das Turbinen-Mantelringsegment eine Mehrzahl von Durchlässen (66) umfasst, die sich in einer Mantelringplattform (44) erstrecken, um Kühlluft dadurch zu lenken, um das Turbinen-Mantelringsegment (42) zu kühlen und die Kühlluft an einem hinteren Ende (56) der Mantelringplattform (44) abzulassen, damit sie auf das Deckband der Turbinenleitschaufel (38) trifft, wobei die Mantelringplattform (44) einen gemeinsamen Auslass (70) für die Mehrzahl von Durchlässen (66) am hinteren Ende der Plattform (44) umfasst; dadurch gekennzeichnet, dass der gemeinsame Auslass (70) eine Öffnung an der radial inneren Fläche (52) der Plattform (44) aufweist.

2. Mantelringsegment nach Anspruch 1, wobei die Mehrzahl von Durchlässen (66) sich durch einen Hauptabschnitt einer gesamten axialen Länge der Mantelringplattform (44) erstreckt.

3. Mantelringsegment nach Anspruch 2, wobei die Mehrzahl von Durchlässen (66) in Fluidverbindung mit einem Hohlraum (64) steht, der zwischen einem vorderen und hinteren Schenkel (46, 48) des Turbinen-Mantelringsegments (42) definiert ist.

4. Mehrzahl von Durchlässen (66) nach einem der vorangehenden Ansprüche, wobei die Mehrzahl von Durchlässen (66) eine Mehrzahl von im Wesentlichen axialen graden Löchern umfasst.

5. Gasturbinenmotorstruktur zum Definieren eines Abschnitts einer Außenwand eines ringförmigen Gaswegs eines Turbinenabschnitts, umfassend einen Turbinenmantelring (32), der eine Mehrzahl von Mantelringsegmenten (42) nach Anspruch 1 umfasst, und ein Deckband der Turbinenleitschaufel (38) mit einer Mehrzahl von Schaufeln (40), die unmittelbar stromabwärts von dem Turbinenmantelring angeordnet sind, wobei die Durchlässe (66) im Wesentlichen an dem Deckband der Turbinenleitschaufel (38) ausgerichtet sind, wodurch Kühlluft, die von den Durchlässen abgegeben wird, im Wesentlichen die gesamte Erstreckung eines Umfangs einer vorderen Endes des Deckbands der Turbinenleitschaufel (38) anbläst.

6. Gasturbinenmotorstruktur nach Anspruch 5, wobei die Durchlässe (66) eine Mehrzahl von Löchern umfassen, die wenigstens einen Einlass (68) davon an einer Außenfläche (50) der Plattform (44) aufweisen.

7. Gasturbinenmotorstruktur nach Anspruch 6, wobei der wenigstens eine Einlass (68) der Löcher (66) an einer Position nahe einem vorderen Schenkel (46) des Mantelrings und stromabwärts davon angeordnet ist.

8. Gasturbinenmotorstruktur nach Anspruch 6 oder 7, wobei die Löcher (66) sich in einer im Wesentlichen geraden Richtung über einen Hauptabschnitt der gesamten axialen Länge der Plattform (44) erstrecken.

9. Verfahren zum Wiederverwenden von Turbinen-Mantelringkühlluft zur Anblaskühlung an einem stromabwärtigen Deckband einer Turbinenleitschaufel (38), wobei das Verfahren folgende Schritte umfasst:

(a) Lenken von Kühlluft in und durch eine Plattform (44) eines Mantelringsegments (42) nach Anspruch 4 eines Turbinenmantelrings zum Kühlen des Turbinenmantelrings; und

(b) Verwenden der Kühlluft aus Schritt (a) zum Bilden einer Mehrzahl von im Wesentlichen geraden Kühlluftströmen axial zu einem vorderen Ende des Deckbands der Turbinenleitschaufel (38) hin zum Anblaskühlen des Deckbands der Turbinenleitschaufel.

10. Verfahren nach Anspruch 9, wobei die Schritte (a) und (b) im Wesentlichen gleichzeitig durchgeführt werden.

11. Verfahren nach Anspruch 10, wobei die Schritte (a) und (b) ausgeführt werden, indem die Kühlluft durch eine Mehrzahl von im Wesentlichen axialen und geraden Durchlässen (66) gelenkt wird, die sich in der Plattform (44) des Mantelsegments (42) des Turbinenmantelrings erstrecken, um die im Wesentlichen geraden Kühlluftströme mit einer hohen Geschwindigkeit derselben abzugeben.

12. Verfahren nach Anspruch 11, wobei die Kühlluft von einem Hohlraum (64), der zwischen dem vorderen und hinteren Schenkel des Mantelringsegments (42) definiert ist, in die im Wesentlichen axialen und geraden Durchlässe (66) gelenkt wird.

13. Verfahren nach einem der Ansprüche 9 bis 12, ferner umfassend einen Schritt (c) des Ablassens der Kühlluft in einen Gasweg während des Anblaskühlens derselben an dem Deckband der Turbinenleitschaufel (38).

14. Verfahren nach einem der Ansprüche 9 bis 13, umfassend Lenken der im Wesentlichen geraden Kühlluftströme in einer Weise zum Anblaskühlen an einem im Wesentlichen gesamten Umfang des vorderen Endes des Deckbands der Turbinenleitschaufel (38).

Revendications

1. Segment d'enveloppe de turbine (42) pour un agencement de refroidissement dans une section de turbine d'un moteur à turbine à gaz pour un refroidissement dans une relation d'écoulement en série, à la fois le segment d'enveloppe de turbine (42) et une enveloppe externe d'aube de turbine (38) étant disposée en aval du segment d'enveloppe de turbine (42), le segment d'enveloppe de turbine comprenant une pluralité de passages (66) s'étendant dans une plate-forme d'enveloppe (44) pour diriger l'air de refroidissement au travers pour refroidir le segment d'enveloppe de turbine (42) et pour évacuer l'air de refroidissement au niveau d'une extrémité de fuite (56) de la plate-forme d'enveloppe (44) pour impact sur l'enveloppe externe d'aube de turbine (38), la plate-forme d'enveloppe (44) comprenant un refoulement commun (70) pour ladite pluralité de passages (66) sur l'extrémité de fuite de ladite plate-forme (44) ; caractérisé en ce que ledit refoulement commun (70) a une ouverture sur la surface radialement interne (52) de la plate-forme (44).

2. Segment d'enveloppe selon la revendication 1, dans lequel la pluralité de passages (66) s'étend à travers une section majeure d'une longueur axiale entière de la plate-forme d'enveloppe (44).

3. Segment d'enveloppe selon la revendication 2, dans lequel la pluralité de passages (66) est en communication fluidique avec une cavité (64) définie entre des pattes avant et arrière (46, 48) du segment d'enveloppe de turbine (42).

4. Pluralité de passages (66) selon une quelconque revendication précédente, dans laquelle les passages (66) comprennent une pluralité de trous droits sensiblement axiaux.

5. Structure de moteur de turbine à gaz destinée à définir une portion d'une paroi externe d'un trajet de gaz annulaire d'une section de turbine, comprenant une enveloppe de turbine (32) comprenant une pluralité de segments d'enveloppe de turbine (42) tels que revendiqués dans la revendication 1, et une enveloppe externe d'aube de turbine (38) dotée d'une pluralité d'aubes (40) disposée immédiatement en aval de l'enveloppe de turbine, les passages (66) s'alignant sensiblement avec l'enveloppe externe d'aube de turbine (38), moyennant quoi de l'air de refroidissement délivré desdits passages vient impacter sensiblement l'étendue entière d'une circonférence d'une extrémité d'attaque de l'enveloppe externe d'aube de turbine (38).

6. Structure de moteur de turbine à gaz selon la revendication 5, dans laquelle les passages (66) comprennent une pluralité de trous ayant au moins une admission (68) de ceux-ci sur une surface externe (50) de la plate-forme (44).

7. Structure de moteur de turbine à gaz selon la revendication 6, dans laquelle la au moins une admission (68) des trous (66) est située dans une position proche et en aval d'une patte avant (46) de l'enveloppe.

8. Structure de moteur de turbine à gaz selon la revendication 6 ou 7, dans laquelle les trous (66) s'étendent dans une direction sensiblement droite sur une section majeure de longueur axiale entière de la plate-forme (44).

9. Procédé de réutilisation d'air de refroidissement d'enveloppe de turbine pour un refroidissement par impact sur une enveloppe externe d'aube de turbine aval (38), le procédé comprenant les étapes de :

(a) direction de l'air de refroidissement à l'intérieur de et à travers une plate-forme (44) d'un segment d'enveloppe (42) tel que revendiqué dans la revendication 4 d'une enveloppe de turbine pour refroidir l'enveloppe de turbine ; et

(b) utilisation de l'air de refroidissement de l'étape (a) pour former une pluralité de courants d'air de refroidissement sensiblement droits axialement vers une extrémité d'attaque de l'enveloppe externe d'aube de turbine (38) pour un refroidissement par impact sur l'enveloppe externe d'aube de turbine.

10. Procédé selon la revendication 9, dans lequel les étapes (a) et (b) sont conduites sensiblement simultanément.

11. Procédé selon la revendication 10, dans lequel les étapes (a) et (b) sont pratiquées en dirigeant l'air de refroidissement à travers une pluralité de passages sensiblement axiaux et droits (66) s'étendant à l'intérieur de la plate-forme (44) du segment d'enveloppe (42) de l'enveloppe de turbine pour délivrer les courants d'air de refroidissement sensiblement droits à une vitesse élevée.

12. Procédé selon la revendication 11, dans lequel l'air de refroidissement est dirigé à partir d'une cavité (64) définie entre des pattes avant et arrière du segment d'enveloppe (42), dans les passages sensiblement axiaux et droits (66).

13. Procédé selon l'une quelconque des revendications 9 à 12, comprenant en outre une étape (c) d'évacuation de l'air de refroidissement dans un trajet de gaz lors du refroidissement par impact de celui-ci sur l'enveloppe externe d'aube de turbine (38).

14. Procédé selon l'une quelconque des revendications 9 à 13, comprenant la direction des courants d'air de refroidissement sensiblement droits d'une façon permettant un refroidissement par impact sur une circonférence sensiblement entière de l'extrémité d'attaque de l'enveloppe externe d'aube de turbine (38).

Drawing