CA2031084A1 - Contoured rotor blade shroud segment - Google Patents
Contoured rotor blade shroud segmentInfo
- Publication number
- CA2031084A1 CA2031084A1 CA 2031084 CA2031084A CA2031084A1 CA 2031084 A1 CA2031084 A1 CA 2031084A1 CA 2031084 CA2031084 CA 2031084 CA 2031084 A CA2031084 A CA 2031084A CA 2031084 A1 CA2031084 A1 CA 2031084A1
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- Canada
- Prior art keywords
- shroud
- inner face
- curvature
- radius
- gas turbine
- 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.)
- Abandoned
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- Turbine Rotor Nozzle Sealing (AREA)
Abstract
CONTOURED ROTOR BLADE SHROUD SEGMENT
Abstract of the Disclosure A shroud segment adapted to mount to fixed support structure of the casing of the turbine or compressor of a gas turbine engine has an outer face and an inner face formed with end portions having a first radius of curvature which are located on either side of a center portion having a second radius of curvature. A temperature differential between the inner and outer faces of the shroud segment during operation of the gas turbine engine causes the shroud segment to chord or straighten out circumferentially so that a combined, substantially uniform radius of curvature is formed on the inner face which closely conforms to the path of motion of the rotor blades of the gas turbine engine.
Abstract of the Disclosure A shroud segment adapted to mount to fixed support structure of the casing of the turbine or compressor of a gas turbine engine has an outer face and an inner face formed with end portions having a first radius of curvature which are located on either side of a center portion having a second radius of curvature. A temperature differential between the inner and outer faces of the shroud segment during operation of the gas turbine engine causes the shroud segment to chord or straighten out circumferentially so that a combined, substantially uniform radius of curvature is formed on the inner face which closely conforms to the path of motion of the rotor blades of the gas turbine engine.
Description
CONTOURED ROTOR BLADE SHROUD SEGMENT
. _ Field of the Invention This invention relates to gas turbine engines, and, more particularly, to improved shroud segments mounted to the casing of the high pressure turbine or compressor of a gas turbine engine such as a jet engine.
Backqround of the Invention The turbines and compressors of gas turbine engines such as jet engines each include one or more circumferentially extending rows or stages of rotating rotor blades which are axially spaced between rows or stages of fixed stator vanes. Each rotor blade has a blade root mounted to the rotor disk, and an air foil extending radially outwardly from the root which terminates at a blade tip. In many gas turbine engine designs, a number of abutting, circumferentially extending shroud segments are carried by the turbine or compressor case to form an essentially continuous cylindrical-shaped surface along which the tips of the rotor blades tangentially pass. Each of these shroud ~ 8 ~ 13DV-8814 segments includes an outer face, and an inner, arcuate-shaped face along which the blade tips pass, opposite end portions which abut with adjacent shrouds and opposed side mounting rails which mount to sta-tionary hangers on the casing of the turbine and/orcompressors.
The shroud segments, particularly those located in the turbine of a jet engine, are subjected to high temperatures at their inner face along which the rotor blades pass. In an effort to lower the temperature of the shroud segments and increase their durability, cooling air from an intermediate stage of the compressor is often directed onto the outer face of the shrouds. This cooling air is intended to raduce the overall temperature of the entire shroud without directly contacting the inner face and dis-rupting the air flow through the turbine or compres-sor.
A major design consideration in any jet engine is the reduction of specific fuel consumption.
one source of decreased specific fuel consumption in many jet engine designs is pressure losses resulting from the creation of a relatively large radial tip clearance between the tip of the rotor blades and the inner face of the shroud segments. It is believed that one source of increased radial tip clearance is attributable to a problem known as "chording".
Chording results from the temperature differential between the high temperature inner face and the cooler outer face of the shroud segments. Impingement of cooling air on the outer face of the shroud segments while the inner face is subjected to high temperatures causes the shrouds to chord or "straighten out"
circumferentially, i.e., the end portions of the inner face of the shroud tend to move radially outwardly relative to the center portion of the inner face of the shroud. While the interconnection of the side mounting rails of the shroud segments with the sta-tionary hooks on the case of the compressor or turbine is intended to resist chording or straightening-out of the shroud segments, such resistance is overcome by the temperature gradient between the outer and inner faces thereof.
Because the inner face of these shroud segments is formed with a substantially constant radius of curvature from end~to-end, chording has the effect of creating a wedge-shaped space or gap between the tip of the rotor blades and each end portion of the inner face of the shroud segments. Such chording can also cause additional blade tip rubs in the central portion of the shroud segment inner surface.
These rubs produce friction which further increases the radial temperature gradient, thereby causing even further chording and rubs. This increase in radial ~ ~ 3 ~ 13DV-8814 tip clearance at the end portions of each shroud segment has been found to be equivalent to a uniform tip clearance increase of about 0.00~ inches in some types of jet engines, resulting in pressure losses which reduce specific fuel consumption by a signifi-cant amount, e.g., about 0.4%.
One attempt to reduce chording, or the straightening out of the shroud segments, has been to form one or more radially extending notches or grooves in each of the side mounting rails of the shroud segments which mount to stationary structure of the turbine or compressor casing. These radial grooves are intended to reduce or eliminate the "beam strength" of the shroud segments by making them lS discontinuous along the length of their side mounting rails.
It has been found that the presence of radial notches or grooves in the shroud segments creates high stress concentrations at the inner end of such grooves. These stress concentrations can create cracking or fracturing of the shroud segments which can propagate from the groove and result in premature failure of the shroud segment.
SummarY of the Invention It is therefore among the objectives of this invention to provide shroud segments adapted to mount to the casing of the turbine or compressor of a gas $~ 13DV-8814 turbine engine, such as a jet engine, which impro~e specific fuel consumption of the jet engine and which exhibit improved durability.
These objecti~es are accomplished in a shroud segment having a shroud body including opposed ends, an outer face, an inner face and forward and aft side mounting rails which are adapted to mount to fix~d support structure on the casing of a turbine or compressor of a jet engine. The inner face of each shroud segment has a center portion located between a pair of end portions, each of which extend from the opposite ends of the shroud body toward the center portion of the inner face. The inner face is formed in a generally concavely arcuate shape in which each lS f the end portions are formed with a first radius of curvature, and the center portion is formed with a second, smaller radius of curvature. This configura-tion of the inner face of the shroud body is intended to take into account "chording" of the shroud segment during operation of the gas turbine engine so as to avoid the formation of a large radial ~ip clearance between the rotor blade tips and the shroud segments.
This invention is predicated upon the concept of forming a shroud segment with an inner face which, during operation of the turbine or compressor of a gas turbine engine, undergoes chording and is deflected into a position which closely conforms to ~ ~ 3 ~ 13DV-8814 the path of motion of the blade tips of the rotor blades of such engine. As mentioned above, "chording"
refers to beam-bending of the shroud segment resulting from the creation of a temperature gradient or differ-ential between the inner and outer surfaces of theshroud segment. Cooling air is directed onto the outer face of the shroud segment while its inner face is subjected to the relatively high temperatures of the turbine or compressor of the gas turbine engine.
This temperature differential creates bending forces which are transmitted along the length of the shroud segment and which force the outer ends of the inner face to move radially outwardly with respect to the center of the inner face. In other shroud segment desiyns, this radially outward movement of the end portions of the inner face of the shroud segment creates a wedge-shaped gap or clearance between the rotor blade tips and shroud segments during operation of the gas turbine engine.
Using numerically controlled grinding equipment, the inner face of the shroud segment of this invention is formed with two different radii of curvature to account for chording of the shroud segment. The end portions of the inner face are ground with a first radius of curvature while the center portion of the inner face is ground with a second, smaller radius of curvature. ~hen the shroud 13DV-~814 . .
segments become hot during operation of the gas turbine engine, "chording" or bending of the outer portions of the inner face with respect to its center portion results in the formation of an inner face having a combined or deflected radius of curvature which is approximately equal to the path of travel of the rotor blade tips of the gas turbine engine. In other words, as the end portions of the inner face of the shroud segment are deflected radially outwardly with respect to the center portion ther~of, a new radius of curvature is formed with the end portions in this deflected position which closely conforms to the path of travel of the rotor blade tips.
The configuration oE the inner face of the shroud segments of this invention therefore provides a substantially constant and relatively small radial tip clearance during operation of the gas turbine enyine.
This reduces pressure losses in the turbine or com-pressor and thus improves the specific fuel consump-tion of the gas turbine engine. No attempt is made toresist chording of the shroud segments in this inven-tian, and therefore radial slots used in some other shroud segment designs to reduce chording can be eliminated which reduces or eliminates the formation of stress concentrations in the shroud segments ~hich had been a problem in other designs.
`g ~
Descri~tion of the Drawings The structure, operation and advantages of the presently preferred embodiment of this invention will become further apparent upon consideration of the following description, taken in conjunction with the accompanying drawings, wherein:
Fig. 1 is a schematic cross sectional view of a portion of the turbine of a jet engine illustrat-ing the mounting of the shroud segment to the turbine case;
Fig. 2 is a perspective view of a shroud segment of this invention;
Fig. 3 ls a side cross sectional view of a group of three shroud segments in an undeflected position with a pair of turbine blades moving there-pas~; and Fig. 4 is a view similar to Fig. 3 except with the shroud segments in a deflected position.
Detailed Descript~on of the Invention Referring now to Fig. 1, a shroud segment 10 in accordance with this invention is shown in position within the turbine 12 of a gas turbine engine such as shown, for example, in United States Patent No. 4,177,004 issued December 4, 1979, Riedmiller et al. The turbine 12 is shown for purposes of illustrating the positioning of shroud segment 10, and it should be understood that the ~3~
shroud segment 10 could be utilized in turbines of other designs and/or within the high pressure com-pressor of a gas turbine engine. The detailed con-struction of the turbine 12 forms no part of this invention per se and is thus described briefly herein.
A first stage stator vane 14 is bolted at its inner band 16 to a first stage support 18 which provides both radial and axial support for the stator vane 14. An outer band 20 carried by the outside diameter of the stator vane 14 is mounted by a ring 22 to a vane support 24. As used herein, the term "outer" refers to a direction toward the top of Fig.
1, and the term "inner" refers to the opposite direc-tion.
The stator vane 14 is cooled by compressor discharge air which enters a forward plenum 26 Aefined on its outer side by a compressor rear frame 28 and on its inner side by an impingement plate 30. The term "forward" as used herein refers to the lefthand side of Yig. 1, and "aft" refers to the righthand side of Fig. 1. The impingement plate 30 is secured by a plurality of bolts 32 to a seal 34, and this seal 34 is secured to the vane support 24 by fastener 36. The seal 34 is annular in shape and extends radially outwardly from the vane support 24 to a pad 38 on the compressor rear frame 2~ to isolate the forward plenum 26 from an aft plenum 40 which contains cooling air at a lower pressure and temperature from that of forward plenum 26.
A first stage of rotor blades 42, tangen-tially rotatable on a rotor disk 44 are located aft of the stator vanes 14 and each have a blade tip 46 immediately adjacent the shroud segments 10. The shroud segments 10 are supported on the stationary structure of the turbine 12 such that a relatively small radial tip clearance 48 is maintained between the inner face of the shroud segments 10 and the blade tip 46 of the rotor blades 42, as described below.
Support for the shroud segments 10 is provided on the forward end by a plurality of shroud support plates 50 which are connected to the vane support 24 by the fastener 36. Each shroud support plate 50 is formed with an a-xial flange 52 for mount-ing the forward side of the shroud segment 10, as described below. Structure for supporting the aft side of shroud segment 10 includes a rim 54 integrally formed with the vane support 24 having an outer flange 56 and an inner flange 58. This rim 54 mounts a C-clamp 60 havinq an outer flange 62 which engages the outer flange 56 of rim 54, and an inner flange 64 which mounts the shroud segment 10 as described below.
The shroud segment 10 is cooled by cooling air discharged from the compressor (not shown) of the gas turbine engine which enters a cavity 66 formed by ~3~
the vane support 24, rim 54 and shroud support plates 50. The cooling air enters the cavity 66 through apertures 31 formed in the impingement plate 30, through apertures 23 of ring 22 and an opening 68 in the shroud support plates 50. The cooling air impinges upon the outer surface or face of the shroud segment 10 to reduce the overall temperature of the shroud segment 10.
Referring now to Fig. 2, one of the shroud segments 10 is illustrated in detail. Each shroud segment 10 includes a shroud body formed with a forward side mounting rail 72, an aft side mounting rail 74, a center stiffener 75 extending between the side rails 72, 74, opposed end plates 76, 78, an outer face 80 and an inner Pace 82 described in detail below.
The forward and aft side mounting rails 72, 74 each comprise an outer arm 84 and an inner arm 86 which are ~oth connected at one edge to a side plate 88. The arms 84, 86 of the forward side mounting rail 72 are spaced from one another to form a forward slot 90 which is adapted to receive the axial flange 52 of shroud support segment 50. Similarly, the arms 84, 86 of the aft side mounting rail 74 are formed with an aft slot 92 therebetween which is adapted to receive the inner flange 6~ of the C-clamp 60. The shroud se~ment 10 is thus mounted to the stationary structure of the turbine 12 and a plurality of circumferentially extending shroud segments 10 a~ut one another at their end plates 76, 78 to form a substantially continuous cylindrical-shaped surface consisting of adjacent S inner ~aces 82 of abutting shroud segments lo.
As best shown in Figs. 3 and 4, the inner face 82 of each shroud segment 10 is formed with a center portion 94 and two end portions 96, 98 which extend toward the center portion 94 from the end plates 76, 78, respectively, of the shroud body. As shown in Fig. 3, wherein the inner face 82 is in an undeflected position, the end portions 96, 98 of the inner face 82 are formed with a first radius of curvature which is substantially coincident with the path 100 of movement of the blade tips 46 of rotor blades 42. The center portion 94 of the inner face 82 of shroud segment 10 is formed with a second radius of curvature, which is smaller than the first radius of curvature of end portions 96, 98, so that a gap or space 102 of about .005 to .olo inches is formed between the center portion 94 and the rotor blade path 100. It is contemplated that these grinding radiuses could be formed in the shroud segment 10 by commer-cially available computer numerically controlled 2S grinding equipment or any other machining technique.
The purpose o~ forming the inner face 82 o f the shroud segment 10 with two radii of curvature is to account for "chording", i.e., the radially outward deflection of the end portions 96, 98 of inner face 82, and the radially inward deflection of the center portion 94 of inner face 82, which results from the temperature differential between the cooled outer face 80 and the hot inner face 82 of shroud segment 10. As discussed above, the inner face 82 undergoes deflec-tion in response to this temperature differential because the shroud segment 10 acts as a beam and transmits bending forces along the length thereof.
As illustrated in Fig. 4, this bending or chording of the shroud segment 10 is accommodated by the configuration of the inner face 82. With the shroud segment 10 in a heated, operating condition, the end portions 96, 98 of the inner face 82 deflect radially outwardly and the center portion 94 thereof deflects radially inwardly such that the combined, deflected radius of curvature of the inner face 82 of shroud segment 10 is substantially equal to the path ~o 100 of motion of the rotor blade kips 46. In this deflected position of the inner face 82, a substan-tially uniform and relatively small gap 104 is created between the rotor blade tips 46 and inner face 82 which substantially reduces pressure losses in the turbine or compressor of the gas turbine engine.
While the invention has been described with reference to a preferred embodiment, it will be understood by those skilled in the art that various changes may be made and equivalents may be substituted for elements thereof without departing from the scope of the invention. In addition, many modifications may be made to adapt a particular situation or material to the teachings of the invention without departing from the essential scope thereof. Therefore, it is intended that the invention not be limited to the particular embodiment disclosed as the best mode contemplated for carrying out this invention, but that the invention will include all embodiments falling within the scope of the appended claims.
. _ Field of the Invention This invention relates to gas turbine engines, and, more particularly, to improved shroud segments mounted to the casing of the high pressure turbine or compressor of a gas turbine engine such as a jet engine.
Backqround of the Invention The turbines and compressors of gas turbine engines such as jet engines each include one or more circumferentially extending rows or stages of rotating rotor blades which are axially spaced between rows or stages of fixed stator vanes. Each rotor blade has a blade root mounted to the rotor disk, and an air foil extending radially outwardly from the root which terminates at a blade tip. In many gas turbine engine designs, a number of abutting, circumferentially extending shroud segments are carried by the turbine or compressor case to form an essentially continuous cylindrical-shaped surface along which the tips of the rotor blades tangentially pass. Each of these shroud ~ 8 ~ 13DV-8814 segments includes an outer face, and an inner, arcuate-shaped face along which the blade tips pass, opposite end portions which abut with adjacent shrouds and opposed side mounting rails which mount to sta-tionary hangers on the casing of the turbine and/orcompressors.
The shroud segments, particularly those located in the turbine of a jet engine, are subjected to high temperatures at their inner face along which the rotor blades pass. In an effort to lower the temperature of the shroud segments and increase their durability, cooling air from an intermediate stage of the compressor is often directed onto the outer face of the shrouds. This cooling air is intended to raduce the overall temperature of the entire shroud without directly contacting the inner face and dis-rupting the air flow through the turbine or compres-sor.
A major design consideration in any jet engine is the reduction of specific fuel consumption.
one source of decreased specific fuel consumption in many jet engine designs is pressure losses resulting from the creation of a relatively large radial tip clearance between the tip of the rotor blades and the inner face of the shroud segments. It is believed that one source of increased radial tip clearance is attributable to a problem known as "chording".
Chording results from the temperature differential between the high temperature inner face and the cooler outer face of the shroud segments. Impingement of cooling air on the outer face of the shroud segments while the inner face is subjected to high temperatures causes the shrouds to chord or "straighten out"
circumferentially, i.e., the end portions of the inner face of the shroud tend to move radially outwardly relative to the center portion of the inner face of the shroud. While the interconnection of the side mounting rails of the shroud segments with the sta-tionary hooks on the case of the compressor or turbine is intended to resist chording or straightening-out of the shroud segments, such resistance is overcome by the temperature gradient between the outer and inner faces thereof.
Because the inner face of these shroud segments is formed with a substantially constant radius of curvature from end~to-end, chording has the effect of creating a wedge-shaped space or gap between the tip of the rotor blades and each end portion of the inner face of the shroud segments. Such chording can also cause additional blade tip rubs in the central portion of the shroud segment inner surface.
These rubs produce friction which further increases the radial temperature gradient, thereby causing even further chording and rubs. This increase in radial ~ ~ 3 ~ 13DV-8814 tip clearance at the end portions of each shroud segment has been found to be equivalent to a uniform tip clearance increase of about 0.00~ inches in some types of jet engines, resulting in pressure losses which reduce specific fuel consumption by a signifi-cant amount, e.g., about 0.4%.
One attempt to reduce chording, or the straightening out of the shroud segments, has been to form one or more radially extending notches or grooves in each of the side mounting rails of the shroud segments which mount to stationary structure of the turbine or compressor casing. These radial grooves are intended to reduce or eliminate the "beam strength" of the shroud segments by making them lS discontinuous along the length of their side mounting rails.
It has been found that the presence of radial notches or grooves in the shroud segments creates high stress concentrations at the inner end of such grooves. These stress concentrations can create cracking or fracturing of the shroud segments which can propagate from the groove and result in premature failure of the shroud segment.
SummarY of the Invention It is therefore among the objectives of this invention to provide shroud segments adapted to mount to the casing of the turbine or compressor of a gas $~ 13DV-8814 turbine engine, such as a jet engine, which impro~e specific fuel consumption of the jet engine and which exhibit improved durability.
These objecti~es are accomplished in a shroud segment having a shroud body including opposed ends, an outer face, an inner face and forward and aft side mounting rails which are adapted to mount to fix~d support structure on the casing of a turbine or compressor of a jet engine. The inner face of each shroud segment has a center portion located between a pair of end portions, each of which extend from the opposite ends of the shroud body toward the center portion of the inner face. The inner face is formed in a generally concavely arcuate shape in which each lS f the end portions are formed with a first radius of curvature, and the center portion is formed with a second, smaller radius of curvature. This configura-tion of the inner face of the shroud body is intended to take into account "chording" of the shroud segment during operation of the gas turbine engine so as to avoid the formation of a large radial ~ip clearance between the rotor blade tips and the shroud segments.
This invention is predicated upon the concept of forming a shroud segment with an inner face which, during operation of the turbine or compressor of a gas turbine engine, undergoes chording and is deflected into a position which closely conforms to ~ ~ 3 ~ 13DV-8814 the path of motion of the blade tips of the rotor blades of such engine. As mentioned above, "chording"
refers to beam-bending of the shroud segment resulting from the creation of a temperature gradient or differ-ential between the inner and outer surfaces of theshroud segment. Cooling air is directed onto the outer face of the shroud segment while its inner face is subjected to the relatively high temperatures of the turbine or compressor of the gas turbine engine.
This temperature differential creates bending forces which are transmitted along the length of the shroud segment and which force the outer ends of the inner face to move radially outwardly with respect to the center of the inner face. In other shroud segment desiyns, this radially outward movement of the end portions of the inner face of the shroud segment creates a wedge-shaped gap or clearance between the rotor blade tips and shroud segments during operation of the gas turbine engine.
Using numerically controlled grinding equipment, the inner face of the shroud segment of this invention is formed with two different radii of curvature to account for chording of the shroud segment. The end portions of the inner face are ground with a first radius of curvature while the center portion of the inner face is ground with a second, smaller radius of curvature. ~hen the shroud 13DV-~814 . .
segments become hot during operation of the gas turbine engine, "chording" or bending of the outer portions of the inner face with respect to its center portion results in the formation of an inner face having a combined or deflected radius of curvature which is approximately equal to the path of travel of the rotor blade tips of the gas turbine engine. In other words, as the end portions of the inner face of the shroud segment are deflected radially outwardly with respect to the center portion ther~of, a new radius of curvature is formed with the end portions in this deflected position which closely conforms to the path of travel of the rotor blade tips.
The configuration oE the inner face of the shroud segments of this invention therefore provides a substantially constant and relatively small radial tip clearance during operation of the gas turbine enyine.
This reduces pressure losses in the turbine or com-pressor and thus improves the specific fuel consump-tion of the gas turbine engine. No attempt is made toresist chording of the shroud segments in this inven-tian, and therefore radial slots used in some other shroud segment designs to reduce chording can be eliminated which reduces or eliminates the formation of stress concentrations in the shroud segments ~hich had been a problem in other designs.
`g ~
Descri~tion of the Drawings The structure, operation and advantages of the presently preferred embodiment of this invention will become further apparent upon consideration of the following description, taken in conjunction with the accompanying drawings, wherein:
Fig. 1 is a schematic cross sectional view of a portion of the turbine of a jet engine illustrat-ing the mounting of the shroud segment to the turbine case;
Fig. 2 is a perspective view of a shroud segment of this invention;
Fig. 3 ls a side cross sectional view of a group of three shroud segments in an undeflected position with a pair of turbine blades moving there-pas~; and Fig. 4 is a view similar to Fig. 3 except with the shroud segments in a deflected position.
Detailed Descript~on of the Invention Referring now to Fig. 1, a shroud segment 10 in accordance with this invention is shown in position within the turbine 12 of a gas turbine engine such as shown, for example, in United States Patent No. 4,177,004 issued December 4, 1979, Riedmiller et al. The turbine 12 is shown for purposes of illustrating the positioning of shroud segment 10, and it should be understood that the ~3~
shroud segment 10 could be utilized in turbines of other designs and/or within the high pressure com-pressor of a gas turbine engine. The detailed con-struction of the turbine 12 forms no part of this invention per se and is thus described briefly herein.
A first stage stator vane 14 is bolted at its inner band 16 to a first stage support 18 which provides both radial and axial support for the stator vane 14. An outer band 20 carried by the outside diameter of the stator vane 14 is mounted by a ring 22 to a vane support 24. As used herein, the term "outer" refers to a direction toward the top of Fig.
1, and the term "inner" refers to the opposite direc-tion.
The stator vane 14 is cooled by compressor discharge air which enters a forward plenum 26 Aefined on its outer side by a compressor rear frame 28 and on its inner side by an impingement plate 30. The term "forward" as used herein refers to the lefthand side of Yig. 1, and "aft" refers to the righthand side of Fig. 1. The impingement plate 30 is secured by a plurality of bolts 32 to a seal 34, and this seal 34 is secured to the vane support 24 by fastener 36. The seal 34 is annular in shape and extends radially outwardly from the vane support 24 to a pad 38 on the compressor rear frame 2~ to isolate the forward plenum 26 from an aft plenum 40 which contains cooling air at a lower pressure and temperature from that of forward plenum 26.
A first stage of rotor blades 42, tangen-tially rotatable on a rotor disk 44 are located aft of the stator vanes 14 and each have a blade tip 46 immediately adjacent the shroud segments 10. The shroud segments 10 are supported on the stationary structure of the turbine 12 such that a relatively small radial tip clearance 48 is maintained between the inner face of the shroud segments 10 and the blade tip 46 of the rotor blades 42, as described below.
Support for the shroud segments 10 is provided on the forward end by a plurality of shroud support plates 50 which are connected to the vane support 24 by the fastener 36. Each shroud support plate 50 is formed with an a-xial flange 52 for mount-ing the forward side of the shroud segment 10, as described below. Structure for supporting the aft side of shroud segment 10 includes a rim 54 integrally formed with the vane support 24 having an outer flange 56 and an inner flange 58. This rim 54 mounts a C-clamp 60 havinq an outer flange 62 which engages the outer flange 56 of rim 54, and an inner flange 64 which mounts the shroud segment 10 as described below.
The shroud segment 10 is cooled by cooling air discharged from the compressor (not shown) of the gas turbine engine which enters a cavity 66 formed by ~3~
the vane support 24, rim 54 and shroud support plates 50. The cooling air enters the cavity 66 through apertures 31 formed in the impingement plate 30, through apertures 23 of ring 22 and an opening 68 in the shroud support plates 50. The cooling air impinges upon the outer surface or face of the shroud segment 10 to reduce the overall temperature of the shroud segment 10.
Referring now to Fig. 2, one of the shroud segments 10 is illustrated in detail. Each shroud segment 10 includes a shroud body formed with a forward side mounting rail 72, an aft side mounting rail 74, a center stiffener 75 extending between the side rails 72, 74, opposed end plates 76, 78, an outer face 80 and an inner Pace 82 described in detail below.
The forward and aft side mounting rails 72, 74 each comprise an outer arm 84 and an inner arm 86 which are ~oth connected at one edge to a side plate 88. The arms 84, 86 of the forward side mounting rail 72 are spaced from one another to form a forward slot 90 which is adapted to receive the axial flange 52 of shroud support segment 50. Similarly, the arms 84, 86 of the aft side mounting rail 74 are formed with an aft slot 92 therebetween which is adapted to receive the inner flange 6~ of the C-clamp 60. The shroud se~ment 10 is thus mounted to the stationary structure of the turbine 12 and a plurality of circumferentially extending shroud segments 10 a~ut one another at their end plates 76, 78 to form a substantially continuous cylindrical-shaped surface consisting of adjacent S inner ~aces 82 of abutting shroud segments lo.
As best shown in Figs. 3 and 4, the inner face 82 of each shroud segment 10 is formed with a center portion 94 and two end portions 96, 98 which extend toward the center portion 94 from the end plates 76, 78, respectively, of the shroud body. As shown in Fig. 3, wherein the inner face 82 is in an undeflected position, the end portions 96, 98 of the inner face 82 are formed with a first radius of curvature which is substantially coincident with the path 100 of movement of the blade tips 46 of rotor blades 42. The center portion 94 of the inner face 82 of shroud segment 10 is formed with a second radius of curvature, which is smaller than the first radius of curvature of end portions 96, 98, so that a gap or space 102 of about .005 to .olo inches is formed between the center portion 94 and the rotor blade path 100. It is contemplated that these grinding radiuses could be formed in the shroud segment 10 by commer-cially available computer numerically controlled 2S grinding equipment or any other machining technique.
The purpose o~ forming the inner face 82 o f the shroud segment 10 with two radii of curvature is to account for "chording", i.e., the radially outward deflection of the end portions 96, 98 of inner face 82, and the radially inward deflection of the center portion 94 of inner face 82, which results from the temperature differential between the cooled outer face 80 and the hot inner face 82 of shroud segment 10. As discussed above, the inner face 82 undergoes deflec-tion in response to this temperature differential because the shroud segment 10 acts as a beam and transmits bending forces along the length thereof.
As illustrated in Fig. 4, this bending or chording of the shroud segment 10 is accommodated by the configuration of the inner face 82. With the shroud segment 10 in a heated, operating condition, the end portions 96, 98 of the inner face 82 deflect radially outwardly and the center portion 94 thereof deflects radially inwardly such that the combined, deflected radius of curvature of the inner face 82 of shroud segment 10 is substantially equal to the path ~o 100 of motion of the rotor blade kips 46. In this deflected position of the inner face 82, a substan-tially uniform and relatively small gap 104 is created between the rotor blade tips 46 and inner face 82 which substantially reduces pressure losses in the turbine or compressor of the gas turbine engine.
While the invention has been described with reference to a preferred embodiment, it will be understood by those skilled in the art that various changes may be made and equivalents may be substituted for elements thereof without departing from the scope of the invention. In addition, many modifications may be made to adapt a particular situation or material to the teachings of the invention without departing from the essential scope thereof. Therefore, it is intended that the invention not be limited to the particular embodiment disclosed as the best mode contemplated for carrying out this invention, but that the invention will include all embodiments falling within the scope of the appended claims.
Claims (6)
1. A shroud segment for the rotor blades of a gas turbine engine having shroud mounting structure and shroud cooling structure, comprising:
a shroud body having opposed first and second ends, a forward mounting rail, an aft mounting rail, an outer face, and an inner face;
said forward and aft mounting rails being adapted to mount to the shroud mounting structure associated with the gas turbine engine so that said inner face of said shroud body faces the rotor blades of the gas turbine engine and is exposed to high temperatures, and so that said outer face of said shroud body is impinged with cooling air from the shroud cooling structure of the gas turbine engine, whereby a temperature differential is created between said inner and outer faces of said shroud body which causes said inner face to move between an undeflected position and a deflected position;
said inner face of said shroud body being formed in a non-uniform, arcuate shape in said undeflected position wherein at least one portion of said inner face has a first radius of curvature and another portion of said inner face has a second radius of curvature, said portions of said inner face having said first and second radii of curvature forming a combined, substantially uniform radius of curvature in said deflected position of said inner face which closely conforms to the path of motion of the rotor blades of the gas turbine engine.
a shroud body having opposed first and second ends, a forward mounting rail, an aft mounting rail, an outer face, and an inner face;
said forward and aft mounting rails being adapted to mount to the shroud mounting structure associated with the gas turbine engine so that said inner face of said shroud body faces the rotor blades of the gas turbine engine and is exposed to high temperatures, and so that said outer face of said shroud body is impinged with cooling air from the shroud cooling structure of the gas turbine engine, whereby a temperature differential is created between said inner and outer faces of said shroud body which causes said inner face to move between an undeflected position and a deflected position;
said inner face of said shroud body being formed in a non-uniform, arcuate shape in said undeflected position wherein at least one portion of said inner face has a first radius of curvature and another portion of said inner face has a second radius of curvature, said portions of said inner face having said first and second radii of curvature forming a combined, substantially uniform radius of curvature in said deflected position of said inner face which closely conforms to the path of motion of the rotor blades of the gas turbine engine.
2. The shroud segment of claim 1 in which said at least one portion of said inner face of said shroud body includes a pair of end portions of said inner face, each formed with said first radius of curvature, and said another portion of said inner face includes a center portion formed with said second radius of curvature, said center portion being located between said end portions.
3. The shroud segment of claim 2 in which said second radius of curvature of said center portion of said inner face is smaller than said first radius of curvature of said end portions of said inner face.
4. A shroud segment for the rotor blades of a gas turbine engine having shroud mounting structure and shroud cooling structure, comprising:
a shroud body having opposed first and second ends, a forward mounting rail, an aft mounting rail, an outer face, and an inner face;
said forward and aft mounting rails being adapted to mount to the shroud mounting structure associated with the gas turbine engine so that said inner face of said shroud body faces the rotor blades of the gas turbine engine and is exposed to high temperatures, and so that said outer face of said shroud body is impinged with cooling air from the shroud cooling structure of the gas turbine engine, whereby a temperature differential is created between said inner and outer faces of said shroud body;
said inner face of said shroud body having a center portion located between a pair of end portions, each of said end portions being formed with a first radius of curvature and said center portion being formed with a second radius of curvature, whereby said inner face of said shroud body undergoes deflection as a result of said temperature differential between said inner and outer faces of said shroud body so that said first and second radii of curvature of said inner face form a combined, substantially uniform radius of curvature which closely conforms to the path of motion of the rotor blades of the gas turbine engine.
a shroud body having opposed first and second ends, a forward mounting rail, an aft mounting rail, an outer face, and an inner face;
said forward and aft mounting rails being adapted to mount to the shroud mounting structure associated with the gas turbine engine so that said inner face of said shroud body faces the rotor blades of the gas turbine engine and is exposed to high temperatures, and so that said outer face of said shroud body is impinged with cooling air from the shroud cooling structure of the gas turbine engine, whereby a temperature differential is created between said inner and outer faces of said shroud body;
said inner face of said shroud body having a center portion located between a pair of end portions, each of said end portions being formed with a first radius of curvature and said center portion being formed with a second radius of curvature, whereby said inner face of said shroud body undergoes deflection as a result of said temperature differential between said inner and outer faces of said shroud body so that said first and second radii of curvature of said inner face form a combined, substantially uniform radius of curvature which closely conforms to the path of motion of the rotor blades of the gas turbine engine.
5. The shroud segment of claim 1 in which said second radius of curvature of said center portion of said inner face is smaller than said first radius of curvature of said end portions of said inner face.
6. The invention as defined in any of the preceding claims including any further features of novelty disclosed.
Applications Claiming Priority (2)
| Application Number | Priority Date | Filing Date | Title |
|---|---|---|---|
| US46584590A | 1990-01-16 | 1990-01-16 | |
| US465,845 | 1990-01-16 |
Publications (1)
| Publication Number | Publication Date |
|---|---|
| CA2031084A1 true CA2031084A1 (en) | 1991-07-17 |
Family
ID=23849397
Family Applications (1)
| Application Number | Title | Priority Date | Filing Date |
|---|---|---|---|
| CA 2031084 Abandoned CA2031084A1 (en) | 1990-01-16 | 1990-11-29 | Contoured rotor blade shroud segment |
Country Status (1)
| Country | Link |
|---|---|
| CA (1) | CA2031084A1 (en) |
Cited By (1)
| Publication number | Priority date | Publication date | Assignee | Title |
|---|---|---|---|---|
| US20230080931A1 (en) * | 2021-09-15 | 2023-03-16 | Toshiba Energy Systems & Solutions Corporation | Turbine stage sealing mechanism |
-
1990
- 1990-11-29 CA CA 2031084 patent/CA2031084A1/en not_active Abandoned
Cited By (1)
| Publication number | Priority date | Publication date | Assignee | Title |
|---|---|---|---|---|
| US20230080931A1 (en) * | 2021-09-15 | 2023-03-16 | Toshiba Energy Systems & Solutions Corporation | Turbine stage sealing mechanism |
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