CA1135195A - Partially segmented supporting and sealing structure for a guide vane array of a gas turbine engine - Google Patents
Partially segmented supporting and sealing structure for a guide vane array of a gas turbine engineInfo
- Publication number
- CA1135195A CA1135195A CA000373476A CA373476A CA1135195A CA 1135195 A CA1135195 A CA 1135195A CA 000373476 A CA000373476 A CA 000373476A CA 373476 A CA373476 A CA 373476A CA 1135195 A CA1135195 A CA 1135195A
- Authority
- CA
- Canada
- Prior art keywords
- radially
- gas turbine
- turbine engine
- nozzle
- ring
- Prior art date
- Legal status (The legal status is an assumption and is not a legal conclusion. Google has not performed a legal analysis and makes no representation as to the accuracy of the status listed.)
- Expired
Links
Classifications
-
- F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
- F01—MACHINES OR ENGINES IN GENERAL; ENGINE PLANTS IN GENERAL; STEAM ENGINES
- F01D—NON-POSITIVE DISPLACEMENT MACHINES OR ENGINES, e.g. STEAM TURBINES
- F01D9/00—Stators
- F01D9/02—Nozzles; Nozzle boxes; Stator blades; Guide conduits, e.g. individual nozzles
- F01D9/04—Nozzles; Nozzle boxes; Stator blades; Guide conduits, e.g. individual nozzles forming ring or sector
- F01D9/041—Nozzles; Nozzle boxes; Stator blades; Guide conduits, e.g. individual nozzles forming ring or sector using blades
Landscapes
- Engineering & Computer Science (AREA)
- Mechanical Engineering (AREA)
- General Engineering & Computer Science (AREA)
- Turbine Rotor Nozzle Sealing (AREA)
Abstract
ABSTRACT OF THE DISCLOSURE
A nozzle of a gas turbine engine includes an array of radially extending guide vanes fixedly supported at their tip portions to an outer shroud, while the root portion of each vane is connected to a segment of a radially inner, segmented shroud. Each segment is, in turn, connected via a slip fit inter-connection to a radially inner ring sup-port structure of generally L-shaped configuration. A
spring extends between the segment and the base of the L-shaped ring support for forming a flexible coupling be-tween the inner ring and the segment, while the remaining portion of the ring support structure defines a pressure dam to reduce leakage through the flexible coupling connec-tion. The segmented nozzle construction effectively re-duces stress levels resulting from differential thermal excursions of the components of the nozzle under transient and steady state operation of the engine, while also mini-mizing leakage through the flexible coupling.
A nozzle of a gas turbine engine includes an array of radially extending guide vanes fixedly supported at their tip portions to an outer shroud, while the root portion of each vane is connected to a segment of a radially inner, segmented shroud. Each segment is, in turn, connected via a slip fit inter-connection to a radially inner ring sup-port structure of generally L-shaped configuration. A
spring extends between the segment and the base of the L-shaped ring support for forming a flexible coupling be-tween the inner ring and the segment, while the remaining portion of the ring support structure defines a pressure dam to reduce leakage through the flexible coupling connec-tion. The segmented nozzle construction effectively re-duces stress levels resulting from differential thermal excursions of the components of the nozzle under transient and steady state operation of the engine, while also mini-mizing leakage through the flexible coupling.
Description
11351~ t PARTIALLY SEGMENTED SUPPORTING AND SEALING
STRUCTURE FOR A GUID~ VANE ARRAY OF A GAS
TURBINE ENG_NE
Background of the Invention The present invention relates to a turbine nozzle as employed in a multi-stage turbine of a gas turbine engine, and more partlcularly, a supporting and sealing structure for an array of radially extending guide vanes of a tur-bine nozzle wherein the root ends of the guide vanes are flexibly connected by a segmented inner shroud, and which supporting structure includes a pressure dam to minimize the amount of leakage introduced by providing flexibility in the nozzle assembly.
In a multi-stage turbine of a gas turbine engine, stationary Vane assemblies are inserted between the rotor wheels, as well as at the entrance and exit of the turbine unit. In the operation of the gas turbine engine, the sta-tionary vane assemblies function to alter the static pres-sure and change the velocity of the high pressure, high temperature gases flowing through the turbine. Heretofore, in order to insure the structural integrity of a vane as-sembly ~s it is subjected to thermal excursions of the com-ponents of the assembly during transient and steady state operating conditions of the gas turbine engine, it has been common to cast the entire nozzle ass~mbly in one piece. The one piece assembly included an outer, unitary shroud, an inner, unitary shroud, and the array of radially extending guide vanes. With this prior art construction, it has been found that during transient and steady state operation of the gas turbine engine, the temperature differentials be-tween the thin, fast responding vanes and the slower, more massive shroud rings, causes a differential thermal growth 1~3519S
STRUCTURE FOR A GUID~ VANE ARRAY OF A GAS
TURBINE ENG_NE
Background of the Invention The present invention relates to a turbine nozzle as employed in a multi-stage turbine of a gas turbine engine, and more partlcularly, a supporting and sealing structure for an array of radially extending guide vanes of a tur-bine nozzle wherein the root ends of the guide vanes are flexibly connected by a segmented inner shroud, and which supporting structure includes a pressure dam to minimize the amount of leakage introduced by providing flexibility in the nozzle assembly.
In a multi-stage turbine of a gas turbine engine, stationary Vane assemblies are inserted between the rotor wheels, as well as at the entrance and exit of the turbine unit. In the operation of the gas turbine engine, the sta-tionary vane assemblies function to alter the static pres-sure and change the velocity of the high pressure, high temperature gases flowing through the turbine. Heretofore, in order to insure the structural integrity of a vane as-sembly ~s it is subjected to thermal excursions of the com-ponents of the assembly during transient and steady state operating conditions of the gas turbine engine, it has been common to cast the entire nozzle ass~mbly in one piece. The one piece assembly included an outer, unitary shroud, an inner, unitary shroud, and the array of radially extending guide vanes. With this prior art construction, it has been found that during transient and steady state operation of the gas turbine engine, the temperature differentials be-tween the thin, fast responding vanes and the slower, more massive shroud rings, causes a differential thermal growth 1~3519S
-2-or thermal gradient to develop within the nozzle assembly, as well as different temperature levels throughout the noz-zle assembly. The result of the differential thermal grad-ients causes differential thermal excursions of the parts of the nozzle assembly, thereby leading to the development of local stresses and cracks in the interconnections be-tween the vanes and the shrouds. In addition, the inner shroud of a stationary turbine nozzle is usually sealed by ; a sheet metal member which is usually brazed to the inner shroud, and it has been found that the thermal excursions of the parts of the turbine nozzle have caused distortion and separation of the brazed connections due to the ther-mal loading on the sheet metal pieces, thereby resulting in pressure leakage through the Vane assembly.
Accordingly, it is an object of the subject invention to overcome the shortcomings of the prior art turbine nozzle assemblies and to provide a new and improved supporting and sealing structure for an array of radially extending guide vanes of a nozzle of a gas turbine engine, which supporting and sealing structure provides a flexible coup-ling between the individual vanes and the inner shroud.
It is another object of the present invention to pro-vide a new and improved supportlng and sealing structure for an array of radially extending guide vanes of a gas turbine engine wherein the flexible coupling between the root ends of the vanes and the inner segmented shroud is sealed by a flexible, pressure dam to minimize leakage through the flexible coupling.
It is a further object of the present invention to provide a new and improved supporting and sealing structure for an array of radially extending guide vanes of a nozzle of a gas turbine engine including means for maintaining the radial and axial alignment of the vanes under transient and steady state operating conditions of the gas turbine engine.
Summary of the Invention The nozzle of the subject invention is embodied in a gas turbine engine, and includes a radially inner shroud ring, a radially outer shroud ring, and a plurality of rad-ially extending Vane structures respectively disposed be-t~een the xadially lnner and the radially outer shroud rings. Each vane is firmly secured at its tip end to the radially outer shroud ring, while the root end of each vane is secured to the inner shroud ring by an inner support and sealing structure. The latter includes a radially inner ring structure of generally L-shaped cross-section includ-ing a generally cylindrical ~ase, and a radially outwardly extending disc. The root end of each vane is connected to a structural segment which forms a portion of a plurality of segments defining a segmented ring. Each structural segment includes a radially inwardly extending lug which is adapted to engage a cooperating slot which extends in two mutually perpendicular directions on the radially outwardly extending disc of the inner ring structure to define a slip fit connection. The latter functions to retain the inner support ~nd sealing structure concentric to the outer shroud ring. The sllp fit connection between the struc-tural segments and the disc also functions to define a pressure dam for minimizing pressure leakage through the flexible coupling of the vanes to the inner shroud. A
spring of generally C-shaped cross-section preferably ex-tends between each segment and thè base of the L-shaped in-ner ring structure, thereby providing a flexible restrain-; ing interconnection between the inner shroud and the vanes.
The new and improved sealin~ and supporting structure ofthe subject invention provides flexibility in the nozzle assembly, thereby eliminating the development of local stresses within the nozzle assembly, while minimizing the amount of leakage introduced by providing flexibility in the nozzle assembly. The flexibility of the sub~ect in-vention is obtained by the provision of the segmented in-ner shroud and the springs. The pressure dam is effective to reduce leakage, and by virtue of the slip fit intercon-nection between the structural segments and the disc por-tion of the inner ring structure, the pressure dam is main-tained during thermal excursions of the components of the nozzle assembly, during both transient and steady state operating conditions of the turbine engine.
Description of the Drawings Other objects and advantages of the invention will be-come apparent from a reading of the following detailed de-scription taken in conjunction wlth the drawings in which:
FIG. 1 is a front elevational view of the new and im-proved nozzle assembly of the subject invention;
FIG. 2 is a cross-sectional view taken along line 2-2 in FIG. l;
FIG. 3 is a rear elevational view of the new and im-proved nozzle assembly of the subject invention; and FIG. 4 is a cross-sectional view taken along line 4-4 in FIG. 2.
Detailed Descri tion of the Preferred Embodiment p Referring to FIGS. 1, 2, and 3, the stationary turbine nozzle assembly of the subject invention is generally des-ignated by the numeral 10 and basically comprises a radi-ally outer shroud ring 12, a radially inner shroud ring 14, and an array of radially extending guide vanes 16 disposed between rings 12 and 14. The radially outer tip portions 18 of each guide vane 16 is secured to the inner surface of the outer shroud ring 12 by a rigid connection, such as by brazing or castin~. On the other hand, the root portion 20 of each guide vane 16 is flexibly connected to the inner shroud by means of the supporting and sealing str~cture of the subject invention. The supporting and sealing struc-ture enables the guide vanes 16 to undergo thermal excur-sions during transient and steady state operation of the gas turbine engine, without resulting in distortion or the development of local stresses on the assembly 10 which could lead to the development of local cracks in the assem-bly.
The supporting and sealing structure of the subject invention includes a segmented ring 30 which is defined by a plurality of individual segments 32 arranged concentri-cally with the radially outer shroud ring 12. Each segment32 is connected to the root end 20 of a radial~y extending guide Vane 16. As illustrated in FIGS. 2, 3, and 4, depend-ing from each segment 32 and extending radially inward of the segment 32, is a T-shaped lug portion 34. Each T---5--shaped lug 34 includes a leg portion 36 which is aligned with the longitudinal axis of the gas turbine engine, and a transverse bar segment 38 extending orthagonal to the longitudinal axis of the engine. The supporting and seal-ing structure 40 further includes a radially inner ringsupport structure 40 which is generally L-shaped in cross-section (see FIG. 2) and includes a gènerally cylindrical base 42 and a radially outward extending disc portion 44.
Secured to the disc portion 44 is an angled ring member 50 which includes an array of radially extending cut-outs 52 so as to define a generally scalloped configuration, as viewed from the rear of the assembly 10 (see FIG. 3). The angled cross-section of the ring 50 (see FIGS. 2 and 4) re-sults in a circumferential space or slot 60 extending about lS the radially outer diameter of the disc portion 44 of the ring support structure 40. As illustrated, the circumfer-ential slot 60 is downstream of the disc portion 44.
The leg portions 36 of the T-shaped lugs 34 are re-spectively slidably mounted in the cut-outs 52, while the transverse bar segment 38 of each lug 34 is slidably mount-ed in the space 60 defined between the disc 44 and the angled ring 50 (see FIGS. 2 and 4). By this arrangement, a slip fit interconnection is defined between each segment 32 and the inner ring support structure 40, with the slip fit connection effectively maintaining the continuity be-tween the segmented ring 30 and the inner support ring 40 so as to define a pressure dam for minimizing pressure leak-age through the flexible coupling of the supporting and sealing structure. It is noted that the pressure dam is maintained throughout the various transient and steady state operating conditions of the gas turbine engine, dur-ing which time the thermal excursions of the vanes cause the segments 32 to move relative to the inner support ring 40.
Disposed at the upstream end of each segment 32 and ex-tending between said segment 32 and the upstream end of the base 42 is a spring means in the form of a C-shaped, flex-ible spring 70. As shown in FIG. 1, a plurality of springs 70 are provided preferably corresponding to the number of 11;~51~S
segments 32 of the segmented r~ng 30. Each spring 70 is connected at its opposite ends to a segment 32 and to the base 42 of the inner support ring 40. By this arrangement, the springs 70 provide a constant biasing force for main-taining the guide vanes 16 in axial and radial alignmentduring both transient and steady state operating conditions of the gas turbine engine when the stationary vane assembly 10 and the components thereof are subjected to thermal ex-cursions. Accordingly, the arrangement of springs 70, seg-mented ring 30, and the inner ring support 40 effectivelydefines a flexible coupling as part of the supporting and sealing structure of the subject invention. Furthermore, axial positioning of the guide vanes 16 is assured by vir-tue of the slip fit interconnection between the T-shaped lugs 34 and the inner ring support structure 40, and in particular, the interconnection between the transverse bar segments 38 of the lugs 34 and the circumferential slot 60 defined between the disc 44 and angled ring 50.
In operation, the supporting and sealing structure 30 insures that the required sealing of the pressure upstream of the vane assembly is maintained relative to the differ-ential pressure downstream of the vane assembly, and by virtue of the flexible coupling interconnection, differen-tial thermal expansion and excursions of the shrouds and the guide vanes is readily accommodated without the devel-opment of local stresses which could lead to cracks in the assembly 10.
Accordingly, the subject invention provides a par-tially segmented turbine nozzle having a flexible support and sealing inner shroud mem~er which is effective to ac-commodate and neutralize thermal excursions of components of the nozzle during transient and steady state operating conditions of the gas turbine engine. The flexible coup-ling at the inner shroud of the subject nozzle assembly in-sures that the structural integrity of the fixed, usuallybrazed, connections of the vane tips to the outer shroud is maintained. Furthermore, the subject construction elimin-ates local stress problems brought about by differential thermal expansions of the components of the assembly.
113Sl~
Flexibility of the subject turbine nozzle is achieved by the arrangement of segmenting the inner shroud and the pro-vision of the springs which maintain the radial positions of the inner ring structure 40, while providing flexibility of the guide vanes in the radial direction. The pressure dam forming a portion of the supporting and sealing inner shroud construction is effective to reduce leakage through the segmented inner shroud, and the pressure dam includes the slip fit construction so as to maintain the pressure dam during various operating conditions of the gas turbine engine, while enabling free movement of the guide vanes in the radial direction. Still further, the specific con-struction of the pressure dam of the subject invention func-tions to maintain and locate the axial position of the seg-mented ring, and the slip fit construction further aids inmaintaining concentricity of the inner shroud.
Although the invention has been described with respect to a preferred embodiment, it is readily apparent that those skilled in the art will be able to make numerous mod-ifications of the exemplary embodiments without departingfrom the spirit and scope of the invention. All such modi-fications are intended to be included within the spirit and scope of the invention as def~ned by the appended claims.
Accordingly, it is an object of the subject invention to overcome the shortcomings of the prior art turbine nozzle assemblies and to provide a new and improved supporting and sealing structure for an array of radially extending guide vanes of a nozzle of a gas turbine engine, which supporting and sealing structure provides a flexible coup-ling between the individual vanes and the inner shroud.
It is another object of the present invention to pro-vide a new and improved supportlng and sealing structure for an array of radially extending guide vanes of a gas turbine engine wherein the flexible coupling between the root ends of the vanes and the inner segmented shroud is sealed by a flexible, pressure dam to minimize leakage through the flexible coupling.
It is a further object of the present invention to provide a new and improved supporting and sealing structure for an array of radially extending guide vanes of a nozzle of a gas turbine engine including means for maintaining the radial and axial alignment of the vanes under transient and steady state operating conditions of the gas turbine engine.
Summary of the Invention The nozzle of the subject invention is embodied in a gas turbine engine, and includes a radially inner shroud ring, a radially outer shroud ring, and a plurality of rad-ially extending Vane structures respectively disposed be-t~een the xadially lnner and the radially outer shroud rings. Each vane is firmly secured at its tip end to the radially outer shroud ring, while the root end of each vane is secured to the inner shroud ring by an inner support and sealing structure. The latter includes a radially inner ring structure of generally L-shaped cross-section includ-ing a generally cylindrical ~ase, and a radially outwardly extending disc. The root end of each vane is connected to a structural segment which forms a portion of a plurality of segments defining a segmented ring. Each structural segment includes a radially inwardly extending lug which is adapted to engage a cooperating slot which extends in two mutually perpendicular directions on the radially outwardly extending disc of the inner ring structure to define a slip fit connection. The latter functions to retain the inner support ~nd sealing structure concentric to the outer shroud ring. The sllp fit connection between the struc-tural segments and the disc also functions to define a pressure dam for minimizing pressure leakage through the flexible coupling of the vanes to the inner shroud. A
spring of generally C-shaped cross-section preferably ex-tends between each segment and thè base of the L-shaped in-ner ring structure, thereby providing a flexible restrain-; ing interconnection between the inner shroud and the vanes.
The new and improved sealin~ and supporting structure ofthe subject invention provides flexibility in the nozzle assembly, thereby eliminating the development of local stresses within the nozzle assembly, while minimizing the amount of leakage introduced by providing flexibility in the nozzle assembly. The flexibility of the sub~ect in-vention is obtained by the provision of the segmented in-ner shroud and the springs. The pressure dam is effective to reduce leakage, and by virtue of the slip fit intercon-nection between the structural segments and the disc por-tion of the inner ring structure, the pressure dam is main-tained during thermal excursions of the components of the nozzle assembly, during both transient and steady state operating conditions of the turbine engine.
Description of the Drawings Other objects and advantages of the invention will be-come apparent from a reading of the following detailed de-scription taken in conjunction wlth the drawings in which:
FIG. 1 is a front elevational view of the new and im-proved nozzle assembly of the subject invention;
FIG. 2 is a cross-sectional view taken along line 2-2 in FIG. l;
FIG. 3 is a rear elevational view of the new and im-proved nozzle assembly of the subject invention; and FIG. 4 is a cross-sectional view taken along line 4-4 in FIG. 2.
Detailed Descri tion of the Preferred Embodiment p Referring to FIGS. 1, 2, and 3, the stationary turbine nozzle assembly of the subject invention is generally des-ignated by the numeral 10 and basically comprises a radi-ally outer shroud ring 12, a radially inner shroud ring 14, and an array of radially extending guide vanes 16 disposed between rings 12 and 14. The radially outer tip portions 18 of each guide vane 16 is secured to the inner surface of the outer shroud ring 12 by a rigid connection, such as by brazing or castin~. On the other hand, the root portion 20 of each guide vane 16 is flexibly connected to the inner shroud by means of the supporting and sealing str~cture of the subject invention. The supporting and sealing struc-ture enables the guide vanes 16 to undergo thermal excur-sions during transient and steady state operation of the gas turbine engine, without resulting in distortion or the development of local stresses on the assembly 10 which could lead to the development of local cracks in the assem-bly.
The supporting and sealing structure of the subject invention includes a segmented ring 30 which is defined by a plurality of individual segments 32 arranged concentri-cally with the radially outer shroud ring 12. Each segment32 is connected to the root end 20 of a radial~y extending guide Vane 16. As illustrated in FIGS. 2, 3, and 4, depend-ing from each segment 32 and extending radially inward of the segment 32, is a T-shaped lug portion 34. Each T---5--shaped lug 34 includes a leg portion 36 which is aligned with the longitudinal axis of the gas turbine engine, and a transverse bar segment 38 extending orthagonal to the longitudinal axis of the engine. The supporting and seal-ing structure 40 further includes a radially inner ringsupport structure 40 which is generally L-shaped in cross-section (see FIG. 2) and includes a gènerally cylindrical base 42 and a radially outward extending disc portion 44.
Secured to the disc portion 44 is an angled ring member 50 which includes an array of radially extending cut-outs 52 so as to define a generally scalloped configuration, as viewed from the rear of the assembly 10 (see FIG. 3). The angled cross-section of the ring 50 (see FIGS. 2 and 4) re-sults in a circumferential space or slot 60 extending about lS the radially outer diameter of the disc portion 44 of the ring support structure 40. As illustrated, the circumfer-ential slot 60 is downstream of the disc portion 44.
The leg portions 36 of the T-shaped lugs 34 are re-spectively slidably mounted in the cut-outs 52, while the transverse bar segment 38 of each lug 34 is slidably mount-ed in the space 60 defined between the disc 44 and the angled ring 50 (see FIGS. 2 and 4). By this arrangement, a slip fit interconnection is defined between each segment 32 and the inner ring support structure 40, with the slip fit connection effectively maintaining the continuity be-tween the segmented ring 30 and the inner support ring 40 so as to define a pressure dam for minimizing pressure leak-age through the flexible coupling of the supporting and sealing structure. It is noted that the pressure dam is maintained throughout the various transient and steady state operating conditions of the gas turbine engine, dur-ing which time the thermal excursions of the vanes cause the segments 32 to move relative to the inner support ring 40.
Disposed at the upstream end of each segment 32 and ex-tending between said segment 32 and the upstream end of the base 42 is a spring means in the form of a C-shaped, flex-ible spring 70. As shown in FIG. 1, a plurality of springs 70 are provided preferably corresponding to the number of 11;~51~S
segments 32 of the segmented r~ng 30. Each spring 70 is connected at its opposite ends to a segment 32 and to the base 42 of the inner support ring 40. By this arrangement, the springs 70 provide a constant biasing force for main-taining the guide vanes 16 in axial and radial alignmentduring both transient and steady state operating conditions of the gas turbine engine when the stationary vane assembly 10 and the components thereof are subjected to thermal ex-cursions. Accordingly, the arrangement of springs 70, seg-mented ring 30, and the inner ring support 40 effectivelydefines a flexible coupling as part of the supporting and sealing structure of the subject invention. Furthermore, axial positioning of the guide vanes 16 is assured by vir-tue of the slip fit interconnection between the T-shaped lugs 34 and the inner ring support structure 40, and in particular, the interconnection between the transverse bar segments 38 of the lugs 34 and the circumferential slot 60 defined between the disc 44 and angled ring 50.
In operation, the supporting and sealing structure 30 insures that the required sealing of the pressure upstream of the vane assembly is maintained relative to the differ-ential pressure downstream of the vane assembly, and by virtue of the flexible coupling interconnection, differen-tial thermal expansion and excursions of the shrouds and the guide vanes is readily accommodated without the devel-opment of local stresses which could lead to cracks in the assembly 10.
Accordingly, the subject invention provides a par-tially segmented turbine nozzle having a flexible support and sealing inner shroud mem~er which is effective to ac-commodate and neutralize thermal excursions of components of the nozzle during transient and steady state operating conditions of the gas turbine engine. The flexible coup-ling at the inner shroud of the subject nozzle assembly in-sures that the structural integrity of the fixed, usuallybrazed, connections of the vane tips to the outer shroud is maintained. Furthermore, the subject construction elimin-ates local stress problems brought about by differential thermal expansions of the components of the assembly.
113Sl~
Flexibility of the subject turbine nozzle is achieved by the arrangement of segmenting the inner shroud and the pro-vision of the springs which maintain the radial positions of the inner ring structure 40, while providing flexibility of the guide vanes in the radial direction. The pressure dam forming a portion of the supporting and sealing inner shroud construction is effective to reduce leakage through the segmented inner shroud, and the pressure dam includes the slip fit construction so as to maintain the pressure dam during various operating conditions of the gas turbine engine, while enabling free movement of the guide vanes in the radial direction. Still further, the specific con-struction of the pressure dam of the subject invention func-tions to maintain and locate the axial position of the seg-mented ring, and the slip fit construction further aids inmaintaining concentricity of the inner shroud.
Although the invention has been described with respect to a preferred embodiment, it is readily apparent that those skilled in the art will be able to make numerous mod-ifications of the exemplary embodiments without departingfrom the spirit and scope of the invention. All such modi-fications are intended to be included within the spirit and scope of the invention as def~ned by the appended claims.
Claims (14)
PROPERTY OR PRIVILEGE IS CLAIMED ARE DEFINED AS FOLLOWS:
1. A supporting and sealing structure for an array of radially extending guide vanes of a nozzle of a gas turbine engine, said supporting and sealing structure comprising:
a radially outer shroud means rigidly connected to the radially outer tip portions of the vanes; and radially inner shroud means flexibly connected to the radially inner root portions of the vanes to de-fine a flexible coupling to accommodate radial ex-cursions and reduce the stress levels of the vanes during thermal loading, said inner shroud means in-cluding a pressure dam to reduce leakage through said flexible coupling connection.
a radially outer shroud means rigidly connected to the radially outer tip portions of the vanes; and radially inner shroud means flexibly connected to the radially inner root portions of the vanes to de-fine a flexible coupling to accommodate radial ex-cursions and reduce the stress levels of the vanes during thermal loading, said inner shroud means in-cluding a pressure dam to reduce leakage through said flexible coupling connection.
2. A supporting and sealing structure for an array of radially extending guide vanes of a nozzle of a gas turbine engine as in claim 1 wherein said radially inner shroud means comprises:
a radially inner ring support structure of a gen-erally L-shaped cross-section including a cylindrical base and a radially outward extending disc;
a segmented ring disposed radially outward of said inner ring support structure and defined by a plural-ity of segments, each segment being respectively con-nected to the root portion of a vane and being slid-ably engagable with the disc portion of the inner ring support structure to define the pressure dam; and spring means forming part of said flexible coupl-ing and extending between said cylindrical base and the outer segmented ring.
a radially inner ring support structure of a gen-erally L-shaped cross-section including a cylindrical base and a radially outward extending disc;
a segmented ring disposed radially outward of said inner ring support structure and defined by a plural-ity of segments, each segment being respectively con-nected to the root portion of a vane and being slid-ably engagable with the disc portion of the inner ring support structure to define the pressure dam; and spring means forming part of said flexible coupl-ing and extending between said cylindrical base and the outer segmented ring.
3. A supporting and sealing structure for an array of radially extending guide vanes of a nozzle of a gas turbine engine as in claim 2 wherein said spring means comprises a plurality of individual springs extending between said cyl-indrical base and the individual segments of the outer seg-mented ring.
4. A supporting and sealing structure for an array of radially extending guide vanes of a nozzle of a gas turbine engine as in claim 3 wherein each spring is of generally C-shaped configuration in cross-section.
5. A supporting and sealing structure for an array of radially extending guide vanes of a nozzle of a gas turbine engine as in claim 1 wherein the radially outer tip portion of each vane is brazed to the radially outer shroud means.
6. A supporting and sealing structure for an array of radially extending guide vanes of a nozzle of a gas turbine engine as in claim 2 wherein said radially inner ring sup-port structure is of unitary construction.
7. A supporting and sealing structure for an array of radially extending guide vanes of a nozzle of a gas turbine engine as in claim 2 wherein each of the plurality of seg-ments of the segmented ring includes a depending, radially inward extending lug, and wherein the peripheral edge of the radially outward extending disc includes a correspond-ing plurality of radially-extending slots, each lug being respectively engaged with a slot to define a slip fit con-nection for retaining the inner ring support structure con-centric with the outer shroud and maintaining the pressure dam upon radial excursions of the guide vanes.
8. A supporting and sealing structure for an array of radially extending guide vanes of a nozzle of a gas turbine engine as in claim 7 wherein each lug is T-shaped in cross-section, and each slot is of a corresponding configuration.
9. A supporting and sealing structure for an array of radially extending guide vanes of a nozzle of a gas turbine engine as in claim 4 wherein the width of each spring in-creases radially outward as viewed along the longitudinal axis of the gas turbine engine.
10. In a nozzle for a gas turbine engine including a radial inner shroud ring, a radial outer shroud ring co-axial with said radial inner shroud ring, and a plurality of radially extending vane structures disposed between said radially inner and said radially outer shroud rings, each of said vanes being firmly secured at one end to the radi-ally outer shroud ring, each of said vanes being secured at its other end to said inner shroud ring by an inner support and sealing structure, said inner support structure com-prising:
a radial inner ring structure of generally L-shaped cross-section including a cylindrical base and a radially outward extending disc;
a segmented ring disposed radially outward of said inner ring support structure and defined by a plural-ity of segments each of which is respectively connect-ed to said other end of each vane structure, said out-er segmented ring being slidably engaged with the rad-ially outwardly extending disc of the inner ring sup-port structure to retain the vane in radial alignment and to define a pressure dam to minimize leakage through said inner support structure; and spring means extending between said cylindrical base and the segmented ring to define a flexible coupl-ing between the pressure dam and the segmented ring whereby said inner support structure defines a flexible connection between the other end of each vane structure and said inner shroud ring for accommodating radial ex-cursions, reducing the stress level of each vane struc-ture during thermal loading, and concentrically lo-cating the inner support and sealing structure.
a radial inner ring structure of generally L-shaped cross-section including a cylindrical base and a radially outward extending disc;
a segmented ring disposed radially outward of said inner ring support structure and defined by a plural-ity of segments each of which is respectively connect-ed to said other end of each vane structure, said out-er segmented ring being slidably engaged with the rad-ially outwardly extending disc of the inner ring sup-port structure to retain the vane in radial alignment and to define a pressure dam to minimize leakage through said inner support structure; and spring means extending between said cylindrical base and the segmented ring to define a flexible coupl-ing between the pressure dam and the segmented ring whereby said inner support structure defines a flexible connection between the other end of each vane structure and said inner shroud ring for accommodating radial ex-cursions, reducing the stress level of each vane struc-ture during thermal loading, and concentrically lo-cating the inner support and sealing structure.
11. In a nozzle for a gas turbine engine as in claim 10 wherein said spring means comprises a plurality of in-dividual springs of generally C-shaped configuration in cross-section, and extending between said cylindrical base and the individual segments of the outer segmented ring.
12. In a nozzle for a gas turbine engine as recited in claim 10 wherein each of the plurality of segments of the segmented ring includes a depending, radially inward ex-tending lug, and wherein the peripheral edge of the radi-ally outward extending disc includes a corresponding plur-ality of peripheral slots, each lug being respectively en-gaged with a slot to define a slip fit connection for re-taining the inner ring structures in alignment and main-taining the pressure dam upon radial excursions of the vane structures.
13. In a nozzle for a gas turbine engine as recited in claim 12 wherein each lug is T-shaped in cross-section, and each slot is of a corresponding configuration.
14. In a nozzle for a gas turbine engine as in claim 11 wherein the width of each spring increases radially out-ward as viewed along the longitudinal axis of the gas tur-bine engine.
Applications Claiming Priority (2)
| Application Number | Priority Date | Filing Date | Title |
|---|---|---|---|
| US15104780A | 1980-05-19 | 1980-05-19 | |
| US151,047 | 1980-05-19 |
Publications (1)
| Publication Number | Publication Date |
|---|---|
| CA1135195A true CA1135195A (en) | 1982-11-09 |
Family
ID=22537110
Family Applications (1)
| Application Number | Title | Priority Date | Filing Date |
|---|---|---|---|
| CA000373476A Expired CA1135195A (en) | 1980-05-19 | 1981-03-20 | Partially segmented supporting and sealing structure for a guide vane array of a gas turbine engine |
Country Status (8)
| Country | Link |
|---|---|
| JP (1) | JPS5716205A (en) |
| BR (1) | BR8102869A (en) |
| CA (1) | CA1135195A (en) |
| DE (1) | DE3108319C2 (en) |
| FR (1) | FR2482657A1 (en) |
| GB (1) | GB2076069B (en) |
| IT (1) | IT1137478B (en) |
| SE (1) | SE448757B (en) |
Families Citing this family (2)
| Publication number | Priority date | Publication date | Assignee | Title |
|---|---|---|---|---|
| JP3631271B2 (en) * | 1993-11-19 | 2005-03-23 | ユナイテッド テクノロジーズ コーポレイション | Inner shroud integrated stator vane structure |
| GB2434182A (en) * | 2006-01-11 | 2007-07-18 | Rolls Royce Plc | Guide vane arrangement for a gas turbine engine |
Family Cites Families (9)
| Publication number | Priority date | Publication date | Assignee | Title |
|---|---|---|---|---|
| GB589541A (en) * | 1941-09-22 | 1947-06-24 | Hayne Constant | Improvements in axial flow turbines, compressors and the like |
| FR954577A (en) * | 1947-09-02 | 1950-01-03 | ||
| GB816288A (en) * | 1957-03-25 | 1959-07-08 | Gen Motors Corp | Improvements relating to labyrinth seals in turbines or compressors |
| US3552753A (en) * | 1968-06-26 | 1971-01-05 | Westinghouse Electric Corp | High efficiency static seal assembly |
| US3529906A (en) * | 1968-10-30 | 1970-09-22 | Westinghouse Electric Corp | Static seal structure |
| US3647311A (en) * | 1970-04-23 | 1972-03-07 | Westinghouse Electric Corp | Turbine interstage seal assembly |
| US3829233A (en) * | 1973-06-27 | 1974-08-13 | Westinghouse Electric Corp | Turbine diaphragm seal structure |
| US4011718A (en) * | 1975-08-01 | 1977-03-15 | United Technologies Corporation | Gas turbine construction |
| SE398659B (en) * | 1976-05-05 | 1978-01-09 | Stal Laval Turbin Ab | SEALING DEVICE IN A GAS TURBINE |
-
1981
- 1981-02-25 SE SE8101236A patent/SE448757B/en not_active IP Right Cessation
- 1981-02-26 GB GB8106159A patent/GB2076069B/en not_active Expired
- 1981-03-02 DE DE3108319A patent/DE3108319C2/en not_active Expired
- 1981-03-20 CA CA000373476A patent/CA1135195A/en not_active Expired
- 1981-03-24 FR FR8105817A patent/FR2482657A1/en active Granted
- 1981-05-08 BR BR8102869A patent/BR8102869A/en unknown
- 1981-05-14 IT IT21699/81A patent/IT1137478B/en active
- 1981-05-19 JP JP7433881A patent/JPS5716205A/en active Granted
Also Published As
| Publication number | Publication date |
|---|---|
| GB2076069B (en) | 1983-12-21 |
| IT1137478B (en) | 1986-09-10 |
| IT8121699A0 (en) | 1981-05-14 |
| BR8102869A (en) | 1982-02-02 |
| JPS6153521B2 (en) | 1986-11-18 |
| SE8101236L (en) | 1981-11-20 |
| SE448757B (en) | 1987-03-16 |
| FR2482657B1 (en) | 1985-03-22 |
| DE3108319C2 (en) | 1986-12-18 |
| GB2076069A (en) | 1981-11-25 |
| DE3108319A1 (en) | 1982-01-28 |
| FR2482657A1 (en) | 1981-11-20 |
| JPS5716205A (en) | 1982-01-27 |
Similar Documents
| Publication | Publication Date | Title |
|---|---|---|
| CA2513047C (en) | Duct with integrated baffle | |
| EP0578460B1 (en) | Gas turbine engine | |
| EP0578461B1 (en) | Turbine nozzle support arrangement | |
| EP0161203B1 (en) | First stage turbine vane support structure | |
| US7229249B2 (en) | Lightweight annular interturbine duct | |
| JP2828651B2 (en) | Stator assembly of axial rotating machine | |
| US4815933A (en) | Nozzle flange attachment and sealing arrangement | |
| US3938906A (en) | Slidable stator seal | |
| CA2853959C (en) | Integrated strut and vane arrangements | |
| US5104286A (en) | Recirculation seal for a gas turbine exhaust diffuser | |
| EP0844369A1 (en) | A bladed rotor and surround assembly | |
| EP0616113B1 (en) | Gas turbine engine and method of assembling a seal in said gas turbine engine | |
| CA2076083C (en) | Flow activated flowpath liner seal | |
| US5176496A (en) | Mounting arrangements for turbine nozzles | |
| GB1600721A (en) | Turbine shroud support | |
| EP0469826A2 (en) | Seal assembly for segmented turbine engine structures | |
| EP3085904A1 (en) | Shroud assembly and shroud for gas turbine engine | |
| GB2103294A (en) | Shroud assembly for a gas turbine engine | |
| GB1600722A (en) | Combined turbine shroud and vane support structure | |
| JP5551165B2 (en) | Improved high-pressure turbine for turbomachinery, associated guide vane sector, and associated aircraft engine | |
| GB2198489A (en) | Gas turbine engine seal assembly | |
| CA2598435C (en) | Interturbine duct with integrated baffle and seal | |
| EP2715069B1 (en) | Piston seal ring | |
| JPH01247702A (en) | Labylinth-seal | |
| JPH0520570B2 (en) |
Legal Events
| Date | Code | Title | Description |
|---|---|---|---|
| MKEX | Expiry |