CA1117429A - Support member and a component supported thereby - Google Patents
Support member and a component supported therebyInfo
- Publication number
- CA1117429A CA1117429A CA000352259A CA352259A CA1117429A CA 1117429 A CA1117429 A CA 1117429A CA 000352259 A CA000352259 A CA 000352259A CA 352259 A CA352259 A CA 352259A CA 1117429 A CA1117429 A CA 1117429A
- Authority
- CA
- Canada
- Prior art keywords
- annular
- shroud ring
- turbine
- filaments
- support member
- 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
- HQVNEWCFYHHQES-UHFFFAOYSA-N silicon nitride Chemical compound N12[Si]34N5[Si]62N3[Si]51N64 HQVNEWCFYHHQES-UHFFFAOYSA-N 0.000 claims abstract description 13
- 229910052581 Si3N4 Inorganic materials 0.000 claims abstract description 10
- 229910010293 ceramic material Inorganic materials 0.000 claims description 13
- 239000000919 ceramic Substances 0.000 claims description 9
- PXHVJJICTQNCMI-UHFFFAOYSA-N Nickel Chemical compound [Ni] PXHVJJICTQNCMI-UHFFFAOYSA-N 0.000 claims description 6
- 238000003491 array Methods 0.000 claims description 4
- 239000000956 alloy Substances 0.000 claims description 3
- 229910045601 alloy Inorganic materials 0.000 claims description 3
- 229910052759 nickel Inorganic materials 0.000 claims description 3
- 239000007789 gas Substances 0.000 description 23
- 230000008602 contraction Effects 0.000 description 5
- 239000000463 material Substances 0.000 description 3
- 230000004075 alteration Effects 0.000 description 2
- 238000002485 combustion reaction Methods 0.000 description 2
- 230000004323 axial length Effects 0.000 description 1
- 230000009977 dual effect Effects 0.000 description 1
- 239000004744 fabric Substances 0.000 description 1
- 238000009413 insulation Methods 0.000 description 1
- 150000002500 ions Chemical class 0.000 description 1
- 238000011068 loading method Methods 0.000 description 1
- 238000012423 maintenance Methods 0.000 description 1
- 238000004519 manufacturing process Methods 0.000 description 1
- 150000004767 nitrides Chemical class 0.000 description 1
- 230000002093 peripheral effect Effects 0.000 description 1
- HBMJWWWQQXIZIP-UHFFFAOYSA-N silicon carbide Chemical group [Si+]#[C-] HBMJWWWQQXIZIP-UHFFFAOYSA-N 0.000 description 1
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
- F01D25/00—Component parts, details, or accessories, not provided for in, or of interest apart from, other groups
- F01D25/24—Casings; Casing parts, e.g. diaphragms, casing fastenings
- F01D25/246—Fastening of diaphragms or stator-rings
-
- 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
- F01D11/00—Preventing or minimising internal leakage of working-fluid, e.g. between stages
- F01D11/08—Preventing or minimising internal leakage of working-fluid, e.g. between stages for sealing space between rotor blade tips and stator
-
- F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
- F16—ENGINEERING ELEMENTS AND UNITS; GENERAL MEASURES FOR PRODUCING AND MAINTAINING EFFECTIVE FUNCTIONING OF MACHINES OR INSTALLATIONS; THERMAL INSULATION IN GENERAL
- F16J—PISTONS; CYLINDERS; SEALINGS
- F16J15/00—Sealings
- F16J15/16—Sealings between relatively-moving surfaces
- F16J15/32—Sealings between relatively-moving surfaces with elastic sealings, e.g. O-rings
- F16J15/3284—Sealings between relatively-moving surfaces with elastic sealings, e.g. O-rings characterised by their structure; Selection of materials
- F16J15/3288—Filamentary structures, e.g. brush seals
-
- F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
- F16—ENGINEERING ELEMENTS AND UNITS; GENERAL MEASURES FOR PRODUCING AND MAINTAINING EFFECTIVE FUNCTIONING OF MACHINES OR INSTALLATIONS; THERMAL INSULATION IN GENERAL
- F16M—FRAMES, CASINGS OR BEDS OF ENGINES, MACHINES OR APPARATUS, NOT SPECIFIC TO ENGINES, MACHINES OR APPARATUS PROVIDED FOR ELSEWHERE; STANDS; SUPPORTS
- F16M1/00—Frames or casings of engines, machines or apparatus; Frames serving as machinery beds
- F16M1/04—Frames or casings of engines, machines or apparatus; Frames serving as machinery beds for rotary engines or similar machines
-
- F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
- F05—INDEXING SCHEMES RELATING TO ENGINES OR PUMPS IN VARIOUS SUBCLASSES OF CLASSES F01-F04
- F05D—INDEXING SCHEME FOR ASPECTS RELATING TO NON-POSITIVE-DISPLACEMENT MACHINES OR ENGINES, GAS-TURBINES OR JET-PROPULSION PLANTS
- F05D2230/00—Manufacture
- F05D2230/60—Assembly methods
- F05D2230/64—Assembly methods using positioning or alignment devices for aligning or centring, e.g. pins
- F05D2230/642—Assembly methods using positioning or alignment devices for aligning or centring, e.g. pins using maintaining alignment while permitting differential dilatation
Abstract
ABSTRACT OF DISCLOSURE
A turbine suitable for a gas turbine engine is provided with a stage of rotary aerofoil blades which is surrounded by a shroud ring made of silicon nitride. In order to minimise load transfer between the shroud ring and the turbine casing the shroud ring is radially supported by and radially spaced apart from the turbine casing by an annular brush seal.
A turbine suitable for a gas turbine engine is provided with a stage of rotary aerofoil blades which is surrounded by a shroud ring made of silicon nitride. In order to minimise load transfer between the shroud ring and the turbine casing the shroud ring is radially supported by and radially spaced apart from the turbine casing by an annular brush seal.
Description
This invention relates to a structure comprising a support member and a component supported thereby.
It has long been a problem to support components which are subject to thermal e~pansion and contraction with support members which are also subject to such thermal e~pansion and contraction but at a different rate. If the two are rigidly connected, each will be subject to stresses which may eventually lead to their mechanical failure.
This is particularly so in the case when either or both of the support member and component are made from a brittle material such as a ceramic.
This i~ a problem which can arise in gas turbine engines and in particular in the combustion and turbine regions of such engines.
Turbines suitable for gas turbine engines con~entionally comprise a casing enclosing alternate stages of rotary and stationary aerofoil blades positioned in an annular gas passage. In order to ensure the efficient operation of such turbine~, it is important that the clearances between the tips of the rotary aerofoil blades and the radially outer wall of the gas passage are as small as possible.
If the clearances are too great, excessive gas lea~age occurs across the blade tips, thereby reducing turbine efficiency. There i8 a danger however that if clearances are reduced so as to reduce leakage, it is likely that under certain turbine operating conditions, the tips of the rotary blades will make contact with the gas passage wall, thereby causing both blade and wall damage.
In an attempt to ensure that optimum blade tip clearances are achieved and maintained with minimal gas leakage across them, it ha~
been suggested to surround a stagc of rotary aerofoil blades with a shroud ring. The shroud ring is conventionally attached to ~e turbine casi~g in such a man~er that it provides a radially inner sur~ace which defines a portion of the radially outer wall of the turbine annular gas passage. Since the shroud ring is an item which is comparatively simple to manufacture, it may be closely toleranced so as to ensure that rotary aerofoil blade tip clearances are as near to the optimum as is possible. ~owever, shroud rin~s still present problems in ensuring that optimum tip clearances are maintained during turbine operation. These problems are associated mainly with the differing rstes of thermal e~pan~ion of the turbine casing, the shroud ring and the rotary aerofoil blade assembly. Thus, for instance, although the turbine casing and ~hroud ring may be formed from materials having the same or similar rates of thermal expansion, the difference in their masses and the temperatures to which they are e~ psed during turbine operation ensures that they usually expand and contract at differing rates. Consequently there is a danger of the shroud ring and possibly the turbine casing being distorted. Similarly the shroud ring and rotary aerofoil blade stage are li~ely to radiaIly e~pand and contract at differing rates, thereby causing variations in the tip clearances of the rotary aerofoil blades.
It is an ob~ect of the present invention to provide a structure comprising a support member and a component supported thereby in which loadings between them are minimised.
It is a further object of the present invention to provide a turbine which includes a turbine casing, shroud ring and rotary aerofoil blade stage which is so adapted as to minimise variations in the clearances between the tips of the rotary aerofoil blades and the shroud rinB during turbine operation.
~ccording to the present invention, a structure comprises a support member and a component supported thereby, one of said supp rt member and said component being provided with an array of upstanding filaments 90 arranged as to define a brush seal, said component being surrounded by said brush seal in such a manner that said component is both supported from and spac0d apart from said member by Qaid brush seal.
Said component may be of circular cross-section and said brush ~eal comprise an annular array of upstanding filaments, the arrangement being such that said brush seal constitutes the sole means of radial supporb for said component~
Said upstanding filaments are preferably mounted on an annular radially inwardly facing surface of said support member.
~5 According to a further aspect of the present invention a turbine suitable for a gas turbine engine comprises a turbine casing enclosing means adapted to cooperate with said casing to define an annular gas passage, a ~tage of rotary aerofoil blades positioned within said annular gas passage and a shroud ring surrounding but not engaging said rotary aerofoil blades, said shroud ring comprising a ceramic material, adapted to constitute a portion of the radially outer wall of said annular gas passage and both radialIy supported from and radially spaced apart from said turbine casing by an annular array of upstanding filaments mounted on said turbine casing and so arranged as to define an annular brush seal.
Annular brush seals are known in the art and conventionally comprise an annular array of upstanding generally radially extending resilient filaments which are anchored at either of their radially irner or outer ends by a support member. The free ends of the filaments engage the psripheral surface of a me~ber so that a seal is provided between the peripheral surface of the member and tXe filament support.
The upstanding filaments may be anchored by clampirg or alternatively by constituting part of a woven structure such as a velvet-like fabric, Since the shroud ring i9 radially supported from and radialIy spaced apart from the turbine casing by a brush seal comprising a plurality of resilient filaments, it is free to move relative to the casing over a restricted range without the seal between the casing and shroud ring being broken, In particular the shroud ring and casing may e~pand or contract at differing rates without the seal between them being broken and also without any significant load transfer taking place between them.
The lack of any significant load transfer between the shroud 3 ring and casing under a large range of thermal conditions means that the shroud ring may comprise a ceramic material which, under normal circumstances would not tolerate direct attachment to the casing, Since oeramics generally have low rates of thermal expansion, the use of a shroud ring which comprises a ceramic material is highly advantageous in the maintenance of small blade tip clearances which vary little during turbine operation. ~hus whilst the rotary 174~
aerofoil blade stage may expand and contract radiaIIy during turbine operation, the tip clearances between the rotary aerofoil blades and shroud ring vary over a smaller range than is the cass when conventional metallic shroud rings are utilised.
The turbine casing is preferably axially divided radially outwardly of the rotary aerofoil blade stage 90 as to define a circumferentially e~tending housing adapted to accomodate said shroud ring comprising a ceramic material and is additionally adapted to define an annular chamber radially outwardly of said shroud ring housing, said annular brush seal being located within said annular chamber between the radially outer wall of said chamber and the radially outer surface of said shroud ring.
The annular brush seal preferably comprises a support member carrying at least one annular array of upstanding filaments which are inclined to the radii of said shroud ring.
The shroud ring may comprise a metallic ringff haped support member which is adapted to carry the ceramic portion of the shroud ring and also engage the annular brush seal.
~he shroud ring may be supported by an annular brush seal comprising two or more annular arrays of filaments which are coaxially mounted, a~ially spaced apart and carried by the same support member.
If two or more annular arrays of filaments are utilised, the filaments of each annular array are preferably inclined to the radii of said shroud ring in a direction which is opposite to that of the filaments of its adjacent array.
The support member carrying the or each annular array of filaments is preferably mounted on the radialIy outer wall of said annular chamber defined by said turbine casing 90 that the free ends of said filament engage and support said shroud ring.
Said annular brush seal filaments are preferably formed from a nickel base alloy, Said shroud ring preferably comprises an annular silicon nitride portion.
~il7~
Said ~hroud ring may additionally co~prise a further ceramic material interposed between said qilicon nitride portion and said ring shaped support member.
Said further ceramic material may be so adapted that insulating air gaps are defined between said further ceramic material and each of ~aid silicon nitride Fortion and said ring-shaped support member.
The invention will now be described by way of example with reference to the accompanying drawings in which:-Figure 1 is a sectioned side view of a portion of a gas turbine engine incorporating a turbine in accordance with the present invention.
Figure 2 i9 a view on section line A-A of Figure 1, Figure 3 is a sectioned side view of an alternative form of the present invention.
Figure 4 is a vie~r on section line B-B of Figure 3.
~ith reference to Figure 1 a gas turbine engine portio~ generally indicated at 10 comprise~ a combustion chamber 11 and a turbine 12.
The turbine 12 in turn comprises a casing 13 which defines the radially outer wall of an annular gas passage 14. The passage 14 ~ contains, in flow 3eries, stages of stationary nozzle guide vanes 15, rotary high pressure aerofoil blades 16, low pressure stator vanes 17 and rotary low pressure aerofoil blades 18. The stages of rotary aerofoil blades 16 and 18 are mounted for rotation on discs 19 and 20 respectively. The nozzle guide vanes 15 and rotary aerofoil ~5 blades 16 constitute the high pressure section of the turbine 10 and the stator vanes 17 and rotary aerofoil blades 18 the low pressure se~tion. The platforms 21, 22,23 and 24 of the nozzle guide vanes 15, rotary aerofoil blades 16, stator vanes 17 and rotary aerofoil blades 18 respectively define the radially inner wall of the gas pas3age 14.
The turbine casing 13 is axial~y divided radially outwardly of the rotary high pressure aerofoil blade array 16 to provide a circumferentially extending housing 25 for a silicon nitride shroud ring 26. The housing 25 is of sufficient axial length to permit the i5 shroud ring 26 to float radialIy with respect to the axis of rotation of the turbine 10. The wall~ of the housing 25 extend radia11y outwardly to coopeFate ~ith a generally T-shaped cross-section 11~'7 ring 27 so that together they define an annular chamber 28. The radially outer wall 29 of the chamber 28 is provided with a rece3s 30 which accommodates a support ring 31 carrying two annular arrays of upstanding generally radially inwardly extending nickel ba~e alloy filaments 33 and 34. The free ends of the filaments 33 and 34 engage and support the radially outer surface of the shroud ring 26 so that the shroud ring 26 i9 radially spaced apart from the turbine casing 13 but is located axially by the walls of the housing 25.
Thus ths filaments 33 and 34 and support ring 31 constitute a br~sh seal which provides the sole radial support for the shroud ring 26.
The filaments 33 whilst being generally radially extending, are inclined to the radii of the shroud ring 26 as can be seen in Figure
It has long been a problem to support components which are subject to thermal e~pansion and contraction with support members which are also subject to such thermal e~pansion and contraction but at a different rate. If the two are rigidly connected, each will be subject to stresses which may eventually lead to their mechanical failure.
This is particularly so in the case when either or both of the support member and component are made from a brittle material such as a ceramic.
This i~ a problem which can arise in gas turbine engines and in particular in the combustion and turbine regions of such engines.
Turbines suitable for gas turbine engines con~entionally comprise a casing enclosing alternate stages of rotary and stationary aerofoil blades positioned in an annular gas passage. In order to ensure the efficient operation of such turbine~, it is important that the clearances between the tips of the rotary aerofoil blades and the radially outer wall of the gas passage are as small as possible.
If the clearances are too great, excessive gas lea~age occurs across the blade tips, thereby reducing turbine efficiency. There i8 a danger however that if clearances are reduced so as to reduce leakage, it is likely that under certain turbine operating conditions, the tips of the rotary blades will make contact with the gas passage wall, thereby causing both blade and wall damage.
In an attempt to ensure that optimum blade tip clearances are achieved and maintained with minimal gas leakage across them, it ha~
been suggested to surround a stagc of rotary aerofoil blades with a shroud ring. The shroud ring is conventionally attached to ~e turbine casi~g in such a man~er that it provides a radially inner sur~ace which defines a portion of the radially outer wall of the turbine annular gas passage. Since the shroud ring is an item which is comparatively simple to manufacture, it may be closely toleranced so as to ensure that rotary aerofoil blade tip clearances are as near to the optimum as is possible. ~owever, shroud rin~s still present problems in ensuring that optimum tip clearances are maintained during turbine operation. These problems are associated mainly with the differing rstes of thermal e~pan~ion of the turbine casing, the shroud ring and the rotary aerofoil blade assembly. Thus, for instance, although the turbine casing and ~hroud ring may be formed from materials having the same or similar rates of thermal expansion, the difference in their masses and the temperatures to which they are e~ psed during turbine operation ensures that they usually expand and contract at differing rates. Consequently there is a danger of the shroud ring and possibly the turbine casing being distorted. Similarly the shroud ring and rotary aerofoil blade stage are li~ely to radiaIly e~pand and contract at differing rates, thereby causing variations in the tip clearances of the rotary aerofoil blades.
It is an ob~ect of the present invention to provide a structure comprising a support member and a component supported thereby in which loadings between them are minimised.
It is a further object of the present invention to provide a turbine which includes a turbine casing, shroud ring and rotary aerofoil blade stage which is so adapted as to minimise variations in the clearances between the tips of the rotary aerofoil blades and the shroud rinB during turbine operation.
~ccording to the present invention, a structure comprises a support member and a component supported thereby, one of said supp rt member and said component being provided with an array of upstanding filaments 90 arranged as to define a brush seal, said component being surrounded by said brush seal in such a manner that said component is both supported from and spac0d apart from said member by Qaid brush seal.
Said component may be of circular cross-section and said brush ~eal comprise an annular array of upstanding filaments, the arrangement being such that said brush seal constitutes the sole means of radial supporb for said component~
Said upstanding filaments are preferably mounted on an annular radially inwardly facing surface of said support member.
~5 According to a further aspect of the present invention a turbine suitable for a gas turbine engine comprises a turbine casing enclosing means adapted to cooperate with said casing to define an annular gas passage, a ~tage of rotary aerofoil blades positioned within said annular gas passage and a shroud ring surrounding but not engaging said rotary aerofoil blades, said shroud ring comprising a ceramic material, adapted to constitute a portion of the radially outer wall of said annular gas passage and both radialIy supported from and radially spaced apart from said turbine casing by an annular array of upstanding filaments mounted on said turbine casing and so arranged as to define an annular brush seal.
Annular brush seals are known in the art and conventionally comprise an annular array of upstanding generally radially extending resilient filaments which are anchored at either of their radially irner or outer ends by a support member. The free ends of the filaments engage the psripheral surface of a me~ber so that a seal is provided between the peripheral surface of the member and tXe filament support.
The upstanding filaments may be anchored by clampirg or alternatively by constituting part of a woven structure such as a velvet-like fabric, Since the shroud ring i9 radially supported from and radialIy spaced apart from the turbine casing by a brush seal comprising a plurality of resilient filaments, it is free to move relative to the casing over a restricted range without the seal between the casing and shroud ring being broken, In particular the shroud ring and casing may e~pand or contract at differing rates without the seal between them being broken and also without any significant load transfer taking place between them.
The lack of any significant load transfer between the shroud 3 ring and casing under a large range of thermal conditions means that the shroud ring may comprise a ceramic material which, under normal circumstances would not tolerate direct attachment to the casing, Since oeramics generally have low rates of thermal expansion, the use of a shroud ring which comprises a ceramic material is highly advantageous in the maintenance of small blade tip clearances which vary little during turbine operation. ~hus whilst the rotary 174~
aerofoil blade stage may expand and contract radiaIIy during turbine operation, the tip clearances between the rotary aerofoil blades and shroud ring vary over a smaller range than is the cass when conventional metallic shroud rings are utilised.
The turbine casing is preferably axially divided radially outwardly of the rotary aerofoil blade stage 90 as to define a circumferentially e~tending housing adapted to accomodate said shroud ring comprising a ceramic material and is additionally adapted to define an annular chamber radially outwardly of said shroud ring housing, said annular brush seal being located within said annular chamber between the radially outer wall of said chamber and the radially outer surface of said shroud ring.
The annular brush seal preferably comprises a support member carrying at least one annular array of upstanding filaments which are inclined to the radii of said shroud ring.
The shroud ring may comprise a metallic ringff haped support member which is adapted to carry the ceramic portion of the shroud ring and also engage the annular brush seal.
~he shroud ring may be supported by an annular brush seal comprising two or more annular arrays of filaments which are coaxially mounted, a~ially spaced apart and carried by the same support member.
If two or more annular arrays of filaments are utilised, the filaments of each annular array are preferably inclined to the radii of said shroud ring in a direction which is opposite to that of the filaments of its adjacent array.
The support member carrying the or each annular array of filaments is preferably mounted on the radialIy outer wall of said annular chamber defined by said turbine casing 90 that the free ends of said filament engage and support said shroud ring.
Said annular brush seal filaments are preferably formed from a nickel base alloy, Said shroud ring preferably comprises an annular silicon nitride portion.
~il7~
Said ~hroud ring may additionally co~prise a further ceramic material interposed between said qilicon nitride portion and said ring shaped support member.
Said further ceramic material may be so adapted that insulating air gaps are defined between said further ceramic material and each of ~aid silicon nitride Fortion and said ring-shaped support member.
The invention will now be described by way of example with reference to the accompanying drawings in which:-Figure 1 is a sectioned side view of a portion of a gas turbine engine incorporating a turbine in accordance with the present invention.
Figure 2 i9 a view on section line A-A of Figure 1, Figure 3 is a sectioned side view of an alternative form of the present invention.
Figure 4 is a vie~r on section line B-B of Figure 3.
~ith reference to Figure 1 a gas turbine engine portio~ generally indicated at 10 comprise~ a combustion chamber 11 and a turbine 12.
The turbine 12 in turn comprises a casing 13 which defines the radially outer wall of an annular gas passage 14. The passage 14 ~ contains, in flow 3eries, stages of stationary nozzle guide vanes 15, rotary high pressure aerofoil blades 16, low pressure stator vanes 17 and rotary low pressure aerofoil blades 18. The stages of rotary aerofoil blades 16 and 18 are mounted for rotation on discs 19 and 20 respectively. The nozzle guide vanes 15 and rotary aerofoil ~5 blades 16 constitute the high pressure section of the turbine 10 and the stator vanes 17 and rotary aerofoil blades 18 the low pressure se~tion. The platforms 21, 22,23 and 24 of the nozzle guide vanes 15, rotary aerofoil blades 16, stator vanes 17 and rotary aerofoil blades 18 respectively define the radially inner wall of the gas pas3age 14.
The turbine casing 13 is axial~y divided radially outwardly of the rotary high pressure aerofoil blade array 16 to provide a circumferentially extending housing 25 for a silicon nitride shroud ring 26. The housing 25 is of sufficient axial length to permit the i5 shroud ring 26 to float radialIy with respect to the axis of rotation of the turbine 10. The wall~ of the housing 25 extend radia11y outwardly to coopeFate ~ith a generally T-shaped cross-section 11~'7 ring 27 so that together they define an annular chamber 28. The radially outer wall 29 of the chamber 28 is provided with a rece3s 30 which accommodates a support ring 31 carrying two annular arrays of upstanding generally radially inwardly extending nickel ba~e alloy filaments 33 and 34. The free ends of the filaments 33 and 34 engage and support the radially outer surface of the shroud ring 26 so that the shroud ring 26 i9 radially spaced apart from the turbine casing 13 but is located axially by the walls of the housing 25.
Thus ths filaments 33 and 34 and support ring 31 constitute a br~sh seal which provides the sole radial support for the shroud ring 26.
The filaments 33 whilst being generally radially extending, are inclined to the radii of the shroud ring 26 as can be seen in Figure
2. The filaments 34 are also inclined to the radii of the shroud ring 26 but in the opposite direction. Thus together the fila~ents 33 and 34 oppose any tendency for the shroud ring 26 to rotate in either a clockwise or anti-clockwise direction.
The filaments 33 and 34 serve a dual role. They firstly support the shroud ring 26 from the turbine casing 13 in such a manner that any radial growth or contraction of the turbine casing 13 due to thermal expansion or contraction iq not transmitted to the shroud ring 26. Thus any alterations in the radial distance between the turbine casing 13 and the shroud ring 26 arising from relative radial expansion or contraction results in the fil~ments 33 and 34 fle~ing in the manner of springs 90 as to accommodate those alterations.
Consequently little load transfer occurs between the turbine casing 13 and the qhroud ring 26, thereby permitting the shroud ring 26 to be formed from a brittle material such as silicon nitride. It will be appreciated, however, that the present invention is ~snerally applicable to shroud rings comprising any convenient ceramic material.
Since ceramics in general and silicon nitride in particular have low coefficients of thermal expansion, they can be e~cpected to dimensionally alter very little during turbine operation. It follows from this that during turbine operation, the clearance between 1`~ 17~
the tips of the rotary aerofoil blades 16 and the shroud ring 26 effectively only vary b~ the amount that the blades 16 and their associated disc 19 thermally expand and contract in a radial direction. Thus tip clearances are unaffected by the amount that the turbine casing 13 may thermally expand or contract during turbine operation.
The second role served by the filaments 33 and 34 is in providing an a~ial gas seal across the shroud ring 26. Thus during the operation of the turbine 12 some of the hot exhaust gases directed by the stage of no~zle guide vanes 15 onto the stage of rotary aerofoil blades 16 escape through the housing 25 and into the annular chsmber 28, The filaments 33 and 34 prevent these gases from passing across the annular chamber 28 and re-entering the annular gas passage downstream of the rotary aerofoil blade stage 16.
Consequently the only gas leakage across the rotary aerofoil blade stage 16 is across the blade tips.
In certain instances, the temperatures which are encountered in a gas turbine engine turbine are 90 high that the silicon nitride heats up to such an extent that the filaments 33 and 34 may be in danger of heat damage. In such circumstances it is preferred to utilise a shroud ring which has improved heat insulation properties, Such a shroud ring 26~ is shoNn in Fieures 3 and 4.
The shroud ring 26a comprises a silicon nitride ring portion 36 which is similar to the previously described shroud ring 26.
~Iowever the radially outer surface of the silicon nitride ring portion 36 is provided with an annular array of ceramic blocks 37.
~he annular array of ceramic blocks 37 is surrounded in turn by a metallic ring-shaped support member 38 which serves to retain the ceramic blocks 37 in position around the silicon nitride ring portion 36, ~ he ceramic blocks 37 are pravided with cut-out portions 39 and 40 on their radially inner and outer surfaces respectively, These cut-out portions 39 and 40 cooperate with the silicon carbide ring portion 36 and the ring shaped support member 38 respectively to '4~
define insulating air gaps 41 and 42. Thus the air gaps 41 and 42 together with the ceramic blocks 37 ensure that the filaments ~3 and ~4 do not overheat.
Although the present invention has been described with reference to the high pressure stage of a turbine, it will be appreciated that the invention is in fact applicable to any turbine stage.
It will also be appreciated that whilst the pressnt invention has been described with reference to the mounting of a shroud ring within the turbine of a gas turbine engine, it does have broader applications, Thus in its broadest aspect, the present invention relates generally to the mounting of circular cross-section components by means of an array of upstanding filaments which are so arranged as to define a brush seal. Moreover the array of upstan~ing filaments could be mounted in a support member or alternatively on the component itself so that the free ends of the upstanding filaments engage the support member.
3o ~5
The filaments 33 and 34 serve a dual role. They firstly support the shroud ring 26 from the turbine casing 13 in such a manner that any radial growth or contraction of the turbine casing 13 due to thermal expansion or contraction iq not transmitted to the shroud ring 26. Thus any alterations in the radial distance between the turbine casing 13 and the shroud ring 26 arising from relative radial expansion or contraction results in the fil~ments 33 and 34 fle~ing in the manner of springs 90 as to accommodate those alterations.
Consequently little load transfer occurs between the turbine casing 13 and the qhroud ring 26, thereby permitting the shroud ring 26 to be formed from a brittle material such as silicon nitride. It will be appreciated, however, that the present invention is ~snerally applicable to shroud rings comprising any convenient ceramic material.
Since ceramics in general and silicon nitride in particular have low coefficients of thermal expansion, they can be e~cpected to dimensionally alter very little during turbine operation. It follows from this that during turbine operation, the clearance between 1`~ 17~
the tips of the rotary aerofoil blades 16 and the shroud ring 26 effectively only vary b~ the amount that the blades 16 and their associated disc 19 thermally expand and contract in a radial direction. Thus tip clearances are unaffected by the amount that the turbine casing 13 may thermally expand or contract during turbine operation.
The second role served by the filaments 33 and 34 is in providing an a~ial gas seal across the shroud ring 26. Thus during the operation of the turbine 12 some of the hot exhaust gases directed by the stage of no~zle guide vanes 15 onto the stage of rotary aerofoil blades 16 escape through the housing 25 and into the annular chsmber 28, The filaments 33 and 34 prevent these gases from passing across the annular chamber 28 and re-entering the annular gas passage downstream of the rotary aerofoil blade stage 16.
Consequently the only gas leakage across the rotary aerofoil blade stage 16 is across the blade tips.
In certain instances, the temperatures which are encountered in a gas turbine engine turbine are 90 high that the silicon nitride heats up to such an extent that the filaments 33 and 34 may be in danger of heat damage. In such circumstances it is preferred to utilise a shroud ring which has improved heat insulation properties, Such a shroud ring 26~ is shoNn in Fieures 3 and 4.
The shroud ring 26a comprises a silicon nitride ring portion 36 which is similar to the previously described shroud ring 26.
~Iowever the radially outer surface of the silicon nitride ring portion 36 is provided with an annular array of ceramic blocks 37.
~he annular array of ceramic blocks 37 is surrounded in turn by a metallic ring-shaped support member 38 which serves to retain the ceramic blocks 37 in position around the silicon nitride ring portion 36, ~ he ceramic blocks 37 are pravided with cut-out portions 39 and 40 on their radially inner and outer surfaces respectively, These cut-out portions 39 and 40 cooperate with the silicon carbide ring portion 36 and the ring shaped support member 38 respectively to '4~
define insulating air gaps 41 and 42. Thus the air gaps 41 and 42 together with the ceramic blocks 37 ensure that the filaments ~3 and ~4 do not overheat.
Although the present invention has been described with reference to the high pressure stage of a turbine, it will be appreciated that the invention is in fact applicable to any turbine stage.
It will also be appreciated that whilst the pressnt invention has been described with reference to the mounting of a shroud ring within the turbine of a gas turbine engine, it does have broader applications, Thus in its broadest aspect, the present invention relates generally to the mounting of circular cross-section components by means of an array of upstanding filaments which are so arranged as to define a brush seal. Moreover the array of upstan~ing filaments could be mounted in a support member or alternatively on the component itself so that the free ends of the upstanding filaments engage the support member.
3o ~5
Claims (14)
1. A structure comprising a support member, a component supported thereby and an array of upstanding filaments mounted on one of said support member and said component and so arranged as to define a brush seal, said component being surrounded by said brush seal in such a manner that said component is both supported from and spaced apart from said support member by said brush seal.
2. A structure as claimed in claim 1 wherin said component is of substantially circular cross-section and said brush seal comprises an annular array of upstanding filaments, the arrangement being such that said brush seal constitutes the sole means of radial support for said component.
3. A structure as claimed in claim 2 wherein said upstanding filaments are mounted on an annular radially inwardly facing surface of said support member.
4. A turbine suitable for a gas turbine engine comprising a turbine casing enclosing means adapted to cooperate with said casing to define an annular gas passage, a stage of rotary aerofoil blades positioned within said annular gas passage and a shroud ring surrounding but not engaging said aerofoil blades, said shroud ring comprising a ceramic material, adapted to constitute a portion of the radially outer wall of said annular gas passage and both radially supported by-and radially spaced apart from said turbine casing by an annular array of upstanding filaments mounted on said turbine casing and so arranged as to define an annular brush seal.
5. A turbine casing suitable for a gas turbine engine as claimed in claim 4 wherein said turbine casing is axially divided radially outwardly of said rotary aerofoil blade stage so as to define a circumferentially extending housing adapted to accommodate said shroud ring comprising a ceramic material and is additionally adapted to define an annular chamber radially outwardly of said shroud ring housing, said annular brush seal being located within said annular chamber between the radially outer wall of said chamber and the radially outer surface of said shroud ring.
6. A turbine suitable for a gas turbine engine as claimed in claim 4 wherein said shroud ring comprises a ring-shaped support member which is adapted to carry the ceramic portion of the shroud ring and engage the annular brush seal.
7. A turbine suitable for a gas turbine engine as claimed in claim 4 wherein said annular brush seal comprises a support member carrying at least one annular array of upstanding filaments which are inclined to the radii of said shroud ring.
8. A turbine suitable for a gas turbine engine as claimed in claim wherein said annular brush seal comprises two or more annular arrays of filaments which are coaxially mounted, axially spaced apart and carried by the same support member.
9. A turbine suitable for a gas turbine engine as claimed in claim 8 wherein the filaments of each annular array are inclined to the radii of said shroud ring in a direction which is opposite to that of the filaments of its adjacent array.
10. A turbine suitable for a gas turbine engine as claimed in claim 7 wherein the support member carrying the annular array of filaments is mounted on the radially outer wall of said annular chamber defined by said turbine casing so that the free ends of said filaments engage and support said shroud ring.
11. A turbine suitable for a gas turbine engine as claimed in claim 4 wherein said annular brush seal filaments are formed from a nickel base alloy.
12. A turbine suitable for a gas turbine engine as claimed in claim 4 wherein said shroud ring comprises an annular silicon nitride portion.
13. A turbine suitable for a gas turbine engine as claimed in claim 12 wherein said shroud ring additionally comprises a further ceramic material interposed between said silicon nitride portion and said ring shaped support member.
14. A turbine suitable for a gas turbine engine as claimed in claim 13 wherein said further ceramic material is so adapted that insulating air gaps are defined between said further ceramic material and each of said silicon nitride portion and said ring shaped support member.
Applications Claiming Priority (2)
| Application Number | Priority Date | Filing Date | Title |
|---|---|---|---|
| GB7922802 | 1979-06-30 | ||
| GB7922802 | 1979-06-30 |
Publications (1)
| Publication Number | Publication Date |
|---|---|
| CA1117429A true CA1117429A (en) | 1982-02-02 |
Family
ID=10506210
Family Applications (1)
| Application Number | Title | Priority Date | Filing Date |
|---|---|---|---|
| CA000352259A Expired CA1117429A (en) | 1979-06-30 | 1980-05-20 | Support member and a component supported thereby |
Country Status (6)
| Country | Link |
|---|---|
| JP (1) | JPS5612020A (en) |
| CA (1) | CA1117429A (en) |
| DE (1) | DE3023609C2 (en) |
| FR (1) | FR2465874B1 (en) |
| GB (1) | GB2051962B (en) |
| IT (1) | IT1132095B (en) |
Families Citing this family (15)
| Publication number | Priority date | Publication date | Assignee | Title |
|---|---|---|---|---|
| GB2070700B (en) * | 1980-03-01 | 1983-10-05 | Rolls Royce | Gas turbine seals |
| US4398866A (en) * | 1981-06-24 | 1983-08-16 | Avco Corporation | Composite ceramic/metal cylinder for gas turbine engine |
| GB2397102B (en) * | 1981-12-30 | 2004-11-03 | Rolls Royce | Turbine shroud assembly |
| GB2254378B (en) * | 1981-12-30 | 1993-03-31 | Rolls Royce | Gas turbine engine ring shroud ring mounting |
| JPS6117402U (en) * | 1984-07-09 | 1986-01-31 | トヨタ自動車株式会社 | gas turbine engine |
| DE3514377A1 (en) * | 1985-04-20 | 1986-10-23 | MTU Motoren- und Turbinen-Union München GmbH, 8000 München | HEAT EXCHANGER |
| DE3514382C1 (en) * | 1985-04-20 | 1986-06-12 | Motoren Turbinen Union | Brush seal |
| JPS61250984A (en) * | 1985-04-30 | 1986-11-08 | 株式会社山武 | Manufacture of airtight terminal |
| DE3535106A1 (en) * | 1985-10-02 | 1987-04-16 | Mtu Muenchen Gmbh | DEVICE FOR THE EXTERNAL SHEATHING OF THE BLADES OF AXIAL GAS TURBINES |
| JPH0194906A (en) * | 1987-10-02 | 1989-04-13 | Gokou Seisakusho:Kk | Filter in toilet device |
| US6217277B1 (en) * | 1999-10-05 | 2001-04-17 | Pratt & Whitney Canada Corp. | Turbofan engine including improved fan blade lining |
| DE19962316C2 (en) * | 1999-12-23 | 2002-07-18 | Mtu Aero Engines Gmbh | brush seal |
| DE102004025142B4 (en) * | 2004-05-21 | 2007-08-02 | Mtu Aero Engines Gmbh | sealing arrangement |
| US7771160B2 (en) * | 2006-08-10 | 2010-08-10 | United Technologies Corporation | Ceramic shroud assembly |
| CN116201635A (en) * | 2023-05-05 | 2023-06-02 | 中国航发沈阳发动机研究所 | Core machine for controlling stability of rotor shafting based on runner matching |
Family Cites Families (7)
| Publication number | Priority date | Publication date | Assignee | Title |
|---|---|---|---|---|
| FR1219504A (en) * | 1958-03-25 | 1960-05-18 | Zd Y V I | Sealing Ring for Gas Turbine Impeller |
| FR1339482A (en) * | 1961-11-28 | 1963-10-04 | Licentia Gmbh | Rotor seal with radially movable sealing rings, especially for turbo-engines |
| BE756582A (en) * | 1969-10-02 | 1971-03-01 | Gen Electric | CIRCULAR SCREEN AND SCREEN HOLDER WITH TEMPERATURE ADJUSTMENT FOR TURBOMACHINE |
| GB1335145A (en) * | 1972-01-12 | 1973-10-24 | Rolls Royce | Turbine casing for a gas turbine engine |
| DE2366059C3 (en) * | 1973-03-16 | 1981-08-27 | Skf Kugellagerfabriken Gmbh, 8720 Schweinfurt | Seal for sealing a shaft against a bearing housing |
| GB1450553A (en) * | 1973-11-23 | 1976-09-22 | Rolls Royce | Seals and a method of manufacture thereof |
| GB1483661A (en) * | 1974-12-27 | 1977-08-24 | Lucas Industries Ltd | Gas turbine engines |
-
1980
- 1980-04-28 GB GB8013985A patent/GB2051962B/en not_active Expired
- 1980-05-20 CA CA000352259A patent/CA1117429A/en not_active Expired
- 1980-06-05 FR FR8012541A patent/FR2465874B1/en not_active Expired
- 1980-06-11 IT IT22725/80A patent/IT1132095B/en active
- 1980-06-17 JP JP8205780A patent/JPS5612020A/en active Granted
- 1980-06-24 DE DE3023609A patent/DE3023609C2/en not_active Expired
Also Published As
| Publication number | Publication date |
|---|---|
| GB2051962B (en) | 1982-12-15 |
| IT1132095B (en) | 1986-06-25 |
| DE3023609C2 (en) | 1984-08-09 |
| JPS6147290B2 (en) | 1986-10-18 |
| JPS5612020A (en) | 1981-02-05 |
| FR2465874A1 (en) | 1981-03-27 |
| GB2051962A (en) | 1981-01-21 |
| DE3023609A1 (en) | 1981-01-15 |
| FR2465874B1 (en) | 1986-06-06 |
| IT8022725A0 (en) | 1980-06-11 |
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| MKEX | Expiry |