Note: Descriptions are shown in the official language in which they were submitted.
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.
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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
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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
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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
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