GB2119452A - Shroud assemblies for axial flow turbomachine rotors - Google Patents
Shroud assemblies for axial flow turbomachine rotors Download PDFInfo
- Publication number
- GB2119452A GB2119452A GB08212167A GB8212167A GB2119452A GB 2119452 A GB2119452 A GB 2119452A GB 08212167 A GB08212167 A GB 08212167A GB 8212167 A GB8212167 A GB 8212167A GB 2119452 A GB2119452 A GB 2119452A
- Authority
- GB
- United Kingdom
- Prior art keywords
- shroud
- shroud segments
- segments
- supporting structure
- seal means
- 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.)
- Withdrawn
Links
- 230000000712 assembly Effects 0.000 title description 3
- 238000000429 assembly Methods 0.000 title description 3
- 238000005452 bending Methods 0.000 claims description 17
- 238000004519 manufacturing process Methods 0.000 claims description 16
- 239000011148 porous material Substances 0.000 claims description 2
- 238000001816 cooling Methods 0.000 description 21
- 238000007789 sealing Methods 0.000 description 17
- 239000007789 gas Substances 0.000 description 13
- 230000000694 effects Effects 0.000 description 9
- 238000000034 method Methods 0.000 description 6
- 238000005553 drilling Methods 0.000 description 4
- 230000008602 contraction Effects 0.000 description 3
- 238000010276 construction Methods 0.000 description 2
- 238000004513 sizing Methods 0.000 description 2
- 230000003068 static effect Effects 0.000 description 2
- 229910000990 Ni alloy Inorganic materials 0.000 description 1
- 229910045601 alloy Inorganic materials 0.000 description 1
- 239000000956 alloy Substances 0.000 description 1
- 238000005219 brazing Methods 0.000 description 1
- 230000000295 complement effect Effects 0.000 description 1
- 230000003247 decreasing effect Effects 0.000 description 1
- 230000002939 deleterious effect Effects 0.000 description 1
- 230000001419 dependent effect Effects 0.000 description 1
- 230000005489 elastic deformation Effects 0.000 description 1
- 238000002955 isolation Methods 0.000 description 1
- 239000000463 material Substances 0.000 description 1
- 229910001092 metal group alloy Inorganic materials 0.000 description 1
- 238000010008 shearing Methods 0.000 description 1
- 229910000601 superalloy Inorganic materials 0.000 description 1
- 238000003466 welding Methods 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
- 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
- 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
Landscapes
- Engineering & Computer Science (AREA)
- Mechanical Engineering (AREA)
- General Engineering & Computer Science (AREA)
- Turbine Rotor Nozzle Sealing (AREA)
Abstract
Supporting structure, in the form of carrier ring 55, circumferentially surrounds a shroud ring 41 comprising a number of shroud segments 43 connected to the carrier ring by lugs 57. To provide a good seal between the carrier ring and the shroud segments, resiliently deformable seal members 73 are held in a deformed condition between confronting front and rear edge portions of the carrier ring and the shroud segments. The seal members may be circumferentially extending bendable flanges which project from the carrier ring to seal against lands 75 on the shroud segments, or compressible porous rings, e.g. of woven wire. <IMAGE>
Description
SPECIFICATION
Shroud assemblies for turbomachine rotors
The present invention relates to shrouds and shroud assemblies for the high pressure stages of axial flow compressors and turbines such as are incorporated in gas turbine aeroengines.
Axial flow compressor or turbine rotor-blade stages operating at high gas temperatures in gas turbine aeroengines are now being provided with specially designed shroud rings for the purpose of maintaining more nearly optimum clearances between the tips of the rotor blades and the shrouds over as wide a range of rotor speeds and temperatures as possible. This is important because overlarge blade tip clearances reduce the efficiency of the compressor or turbine whilst clearances which are too small risk causing damage under some conditions due to interference between the blade tips and the shroud ring.
A known method of maintaining optimum blade tip clearances over a wide range of conditions involves matching the thermal response of the shroud ring and its supporting structure -- in terms of increase or decrease of diameter with operating temperature -- to the radial growth or shrinkage of the compressor or turbine rotor due to changing centrifugal forces and temperatures.
The better to achieve the required matching, the shroud rings can be composed of quite a large number of segments, each describing a relatively short arc length circumferentially of the rotor stage.
Such shroud segments must be individually connected to supporting structure circumferentially surrounding the shroud ring, and one problem which arises in such designs is excessive sealing clearances between the shroud segments and supporting structure. These excessive clearances can arise because of manufacturing tolerances in the production of the shroud segments and the supporting structure, and because of differing thermal expansion or expansion rates between the two types of components as the operating temperatures change. In the case of compressors the excessive clearances cause decreased efficiency because, for example, they allow air on the high pressure side of the rotor to leak back between the shroud segments and the supporting structure to the low pressure side of the rotor.
In the case of turbines, the excessive clearances increase the consumption of the high pressure cooling air which is fed to the shroud segments and the adjacent components in order to cool them. This reduces the efficiency of the engine.
Large clearances also decrease the effictiveness of the cooling air in cooling the shroud segments by allowing cooling air to escape which would otherwise pass through small cooling air passages in the shroud segments.
The present invention seeks to alleviate the above problems by providing means whereby the manufacturing tolerances, and even the differing thermal characteristics, can be accommodated.
Accordingly, the present invention provides a shroud assembly for an axial flow turbomachine rotor in which supporting structure circumferentially surrounds a shroud ring comprising a plurality of shroud segments connected to the supporting structure, andresiliently deformable seal means are provided on the supporting structure near the axially opposite edges of the shroud ring to seal between confronting portions of the supporting structure and the shroud segments, the seal means being held between said confronting portions in a deformed condition. The deformation of the seal means when held in the assembly is effective to accommodate predetermined manufacturing tolerances in the shroud segments and the supporting structure which could otherwise cause unwanted clearance between the seal means and the shroud segments.In addition, or alternatively, the deformation of the seal means when held in the assembly is sufficient to accommodate at least most of any reduction in deformation of the seal means due to predetermined differential thermal expansion between the shroud segments and the supporting structure during operation of the compressor rotor at elevated temperatures, which differential expansion could otherwise cause unwanted clearances between the seal means and the shroud segments.
The resiliently deformable seal means may comprise circumferentially extending bendable flanges which project from the supporting structure to span the gap between the confronting portions of the supporting structure and the shroud segments, the deformation of the seal means comprising bending of said flanges. In cases where the shroud segments are adapted to be cooled by means of compressed air supplied to chamber means between the support structure and the shroud segments, the bendable flanges may be adapted to allow a controlled amount of said compressed air to exit from said chamber means. Alternatively, the bendable flanges may be replaced by circumferentially extending pads of bulk-compressible material, such as woven wire pads, which again allow a controlled amount of said compressed air to exit from said chamber means.
Further aspects of the invention will be apparent from the following description of preferred embodiments and the appended claims.
Embodiments of the invention will now be described by way of example only and with reference to the accompanying drawings, in which:
Figure 1 is a sectional side elevation of part of a high pressure compressor in a gas turbine aeroengine, incorporating a shroud assembly according to the invention;
Figure 2 is a front elevation of a pair of shroud segments seen in isolation from other structure;
Figure 3 is a sectional side elevation of part of a high pressure turbine in a gas turbine aeroengine, incorporating a shroud assembly according to the invention; and
Figure 4 is a view similar to that within area B in Figure 3, but showing a variant form of the invention at an enlarged scale.
In Figure 1, a high pressure compressor 1 comprises a plurality of rotor blade stages, only one of which is shown, at reference 3. Rotor blade stage 3 is preceding by-stator blades 5, which direct compressor gases 7 onto the rotor blades 9 of rotor stage 3 at a suitable angle. Rotor blades 9 do work on the compressor gases 7, compressing them to a higher pressure and passing them onto succeeding stator blades 1 which inturn deflect the gases onto a further rotor stage. Only the
radially outer portions of the blades are shown.
Rotor stage 3 is circumferentially surrounded
by a shroud ring 13 comprising a number of
shroud segments 1 5. The shroud segments 1 5 are
connected to supporting structure in the form of a
carrier ring 17, which in turn is attached to-a
compressor casing member 1 9 by means of a
number of circumferentially spaced bolts 21
passing through holes 23. Casing member 19 is
itself supported from an outer casing (not shown)
of the engine via frusto-conical panel 25 and
bolted flanges 27.
The function of the shroud ring 1 3 and carrier
ring 1 7 acting together is to maintain an optimum clearance between the tips of the rotor blades 9 and the inner surfaces of the shroud segments 1 5.
The smaller the gap which can be safely
maintained between the blade tips and the shroud segments, the less leakage there will be over the tips of the blades between their high pressure and
low pressure flanks, and the greater will be the
efficiency of the compressor. A known method of
maintaining an optimum clearance is so to design
the panel 25, casing member 19, carrier ring 17,
and shroud ring 13, that as they expand and
contract due to changing temperatures during
operation of the compressor, their collective
influence on the diameter of the shroud ring 1 3 causes the increase or decrease of that diameter
to match, at least approximately, the radial growth
or shrinkage of the compressor rotor 3 due to
centrifugal forces and temperature changes.Exact
matching of expansion or contraction under all
conditions of operation is not possible using this
method and so in order to allow the blade tips and
shroud segments to come into contact with each
other without damage which is significantly
deleterious to compressor performance, the inner
surfaces of the shroud segments 1 5 are provided
with abradeable iinings 29, which can be worn
away by the blade tips without structural damage
to either the blades or the shroud segments.
In the present design the shroud ring 13 is
composed of a number of segments 1 5 to allow for better control of the necessary changes in its
diameter, each segment describing a relatively
short arc length. The carrier ring 1 7 is also segmented, the carrier segments being of greater arc length than the shroud segments, e.g. such that each carrier segment carries three or four shroud segments.
Each carrier segment is held to casing member
19 by means of one bolt 21 positioned centrally of
the circumferential span of each carrier segment
so that circumferential expansion and contraction
of the segments relative to the casing does not
impose shearing forces on the bolts. Differential
expansion and contraction between the carrier
segments and the casing in the radial sense is
accommodated by elastic deformation of the bolt
and threads.
The shroud segments 1 5 are connected to
carrier ring If by means of circumferentially - extending discontinuous "hook" flanges or rows of
lugs 31 (see also Figure 2) which project from the
outel sides of the shroud segments near their
axially opposed (i.e. front and rear) edges, portions
33 of the lugs 31 being oriented to extend parallel
to the inner surface of the shroud ring for
reception in annular grooves 35 in the carrier ring.
Discrete lugs 31 are used rather than continuous flanges in order to cut down conductive heat
transfer between the shroud segments 1 5 and the
carrier ring 17 and also to reduce warping of the
shroud segments due to differential expansion
between the hook flanges and the inner parts of
the segments.
In accordance with the invention, carrier ring 1 7 is provided with resilientiy deformable seal
members 37, portions 38 of the outer sides of the
shroud segments near their axially opposed (i.e.
front and rear) edges being seated against them.
In the Figure 1 embodiment seal members 37
comprise circumferentially extending hairpin flanges or lips on the front and rear portions of
carrier ring 17, these lips being thin enough to be
"springy". This is best seen in the inset to Figure
1, which is an enlarged view of the circular area A
in Figure 1.Lugs 31 are dimensioned so that the
hairpin lips 37 are sprung against sealing lands 38
forming edge portions of the shroud segments as
shown so that they are effective to seal between
the shroud segments and the carrier ring, i.e. lugs
31 hold the shroud segments 1 5 in the assembly
so that the hairpin lips 37 are in a bent condition
due to the edge portions 38 of the shroud
segments being pulled against them, the bending
occurring chiefly in their frusto-conical "arm"
portions 39 (see inset).
This bending is desirable in order to ensure that
the seal between the shroud segments and the
carrier ring, which helps to prevent leakage of
compressor gases from the high pressure side of
compressor rotor stage 3 back to the low pressure side via the spaces between the outer side of the
shroud segments and the carrier ring 17, is
maintained against the effect of manufacturing
tolerances in the shroud segments and the
supporting structure.
The effect of manufacturing tolerances can, for
example, be to increase or decrease the radii of
curvature of the front and rear edge portions of the
shroud segments, or the radii of curvature of the
sealing lips 37, from their nominal value. Similarly,
critical dimensions of lugs 31 or grooves 35 can
also vary from their nominal values. Hence within allowed tolerances, and in the absence of an amount of bending in the sealing members 37 adequate to accommodate them, it would be possible for over or under-sizing of critical dimensions in the carrier ring, or consistent (nonrandom) over- or under-sizing of critical dimensions of the shroud segments, to cause clearance of large circumferential extent between the shroud segments 15 and the sealing members 37, thus allowing compressor gases to escape and reduce the efficiency of the engine.In designing the assembly, therefore, a nominal degree of bending is specified for the sealing lips 37 with the shroud segments assembled to the carrier ring, this nominal degree of bending being sufficient to accommodate variations in the actual degree of bending due to the manufacturing tolerances. The actual degree of bending is of course the nominal degree of bending, with a deviation from nominal caused by actual variations in dimensions within the tolerances.
It is pointed out that because sealing lips 37 in
Figure 1 are continuous over the circumferential extent of the carrier ring 1 7, or at least, are continuous over the circumferential extent of.any segments thereof, whereas the shroud segments describe shorter areas in the circumferential direction, any segment-to-segment (nonconsistent or random) variations in critical dimensions of the shroud segments will result in clearances between some of the shroud segments and the sealing lips 37, because the sealing lips will only bear against those shroud segments whose dimensions place their sealing surfaces at the most radially outward positions.However, because random variations in dimensions within tolerances will in general be smaller than consistent variations due to trends within tolerances, the clearances caused by the random variations will also be small, and hence preferably to the clearances which would arise if the sealing lips 37 were not resiliently bendable.
Besides making allowance for manufacturing tolerances, it is probably even more advantageous to ensure that the seal between shroud segments and carrier ring is maintained against the effect of differential thermal expansion between the shroud segments and the carrier ring. Differential expansion arises because:
a) the shroud segments and the carrier ring may be made from different alloys, and so may have different coefficients of thermal expansion;
b) the shroud segments and the carrier ring will be at somewhat different temperatures during operation of the compressor; and
c) the shroud segments and the carrier ring may expand or contract at differing rates due to the face that they have different thermal masses, i.e.
different thermal inertias.
such differential expansion, it will be seen, could well open up clearances between the shroud segments and the sealing lips 37, and in order to accommodate them, it can be arranged that the nominal degree of bending in the lips 37 when the shroud segments are assembled to the carrier ring
(i.e. in the cold-assembled condition) is sufficient
to accommodate not only the-manufacturing
tolerances, but also as much as possible
(preferably all) of any reduction in bending due to
known differential expansion effects caused by
operation of the compressor. It is, of course, also
necessary to ensure that the nominal degree of
bending is such as to accommodate any increase
ih bending due to differential expansion effects,
without overstraining the lips 37.Again, the actual
degree of bending in the lips when the shroud
segments are held in the assembly at room
temperature will be the nominal degree of
bending, with a deviation from-that nominal
degree due to actual variations in dimensions
within manufacturing tolerances.
In Figure 1, the shroud segments and stator
blades 5 and 11 are shown as being supported by
a carrier ring, which in turn is suspended from a
casing member 1 9. However, it would be possible
in suitable cases to dispense with the carrier ring as a separate component and support the shroud segments and stator blades directly from a thicker casing member having suitable location, retaining and sealing features machined into it.
In Figure 3, the invention is shown as applied to
a high pressure turbine rotor shroud ring 41 comprising a large number of shroud segments 43. The turbine rotor 45 has rotor blades 47, the high temperature turbine gases 49 being directed onto them by nozzle guide vanes 51, and being received by stator blades 53 after energy has been extracted by the rotor blades. As was the case in the embodiment illustrated in Figures 1 and 2, the shroud segments 43 are connected to a carrier ring 55 (which is also segmented) by means of "hook" lugs 57 whose axially oriented portions 59 engage complementary grooves 61 provided on the-carrier ring 55.The segmented construction of the carrier ring 55 allows it to be moved radially inwards or outwards by a limited amount so as to maintain an optimum clearance between the tips of the turbine rotor blades 47 and the inside surfaces of the shroud segments 43, the radial movement being controlled by known means (not shown) to match the radial growth or shrinkage of the rotor blades 47 due to changes in temperature and centrifugal force during operation of the turbine. The carrier ring 55 has side walls 63 which extend radially outwards, adjacent static structure being sealed against these side walls by means of spring-loaded annular "piston-ring" type seals 65.
To survive in the extreme conditions of a high pressure turbine stage in an aeroengine, the shroud segments 43 are made from a superalloy and are actively cooled by means of a supply of pressurised cooling air 67 which is fed to the interior of the carrier ring 55 and thence through apertures 69 in the radially inner side of the carrier ring to annular space 71 between the carrier ring and the shroud segments.
As was the case in Figure 1, resiliently deformable seal members 73, in the form of "hairpin"-section flanges or lips, are sprung against lands 75 near the front and rear edges of the shroud segments, the degree of bending in members 73 being determined by the radial dimensions of the lugs 57, the manufacturing tolerances, and differential thermal expansion as already explained in relation to the previous embodiment.Due to the evident similarity between the compressor and turbine embodiments of the invention as shown in Figures 1 and 2 and Figure 3, no further explanation of the nature and effectiveness of the seal between the carrier ring and the shroud segments is required, except in relation to the cooling of the shroud segments 43 by the cooling air 67, and in relation to the circumferentially extending row of small cooling air holes 77 in the seal members 73 of
Figure 3, which are not present in Figure 1.
The mode of cooling the shroud segments is not specifically detailed in Figure 3, but could be, for example, by allowing the cooling air to pass into the turbine passage via a large number of small diameter drillings (not shown) through the thickness of the shroud segments, the drillings being appropriately distributed over the circumferential and axial extent of the segments and being effective to provide an insulating film of
relatively cool air over their radially inner surfaces.
The invention would in such a case be employed for the purpose of ensuring that the required amount of cooling air finds its way into the turbine passage through the drillings, thereby efficiently cooling the shroud segments, rather than through excessive clearances between sealing surfaces of the shroud segments and the carrier ring, thereby depriving the drillings of some of their cooling air, raising the temperature of the shroud segments-to unacceptable levels and reducing the efficiency of the engine.However, the invention could equally well be employed to ensure the effectiveness of
known impingement cooling techniques in which
a thin perforated plate is placed in close proximity to the radially outer surface of the shroud segments and pressure differences across the plate causes the cooling air to impinge on the radially outer surface in the form of small jets issuing from the perforations; again, excessive clearances between sealing surfaces of the shroud
segments and the carrier ring could cause a
reduction in the pressure difference across the
perforated plate and reduce the effectiveness of the impingement cooling process whilst reducing the efficiency of the engine.
The function of holes 77 in sealing members 73 is two-fold. Firstly, they allow a controlled amount of the cooling air 67, which is at a higher pressure in chamber 71 than the turbine gases in the turbine passage, to pressurise the small chambers 79 and 81 defined between the shroud structure and adjacent static structure and thus prevent undesirable ingress of high temperature turbine gases through the gaps 80, 82 between the front and rear edges of the shroud segments and the adjacent edges of the vanes 51 and stator blades 53 respectively. Secondly, the flow of cooling air through holes 77 tends to cool seal members 73 and prevent conduction of some of the heat from the shroud segments into the carrier ring.
Figure 4 shows an alternative embodiment of the invention in which the "hairpin"-section flanges 73 of Figure 3 are replaced by resilient circumferentially extending pads 83 of bulkcompressible porous material such as woven wire, which are brazed or welded to the carrier ring 55' and which are held compressed between the carrier ring and the lands 75' on the shroud segments, the amount of deformation of pads 83 being once again dependent on the dimensions of lugs 57', manufacturing tolerances, and differential expansion, and being sufficient in the cold-assembled condition to accommodate the latter two.
The porosity of pads 83 is such as to allow controlled escape of the cooling air through them for the same reasons as outlined in relation to
Figure 3. If the pads are of woven wire construction, the density of the weave must be high and the wire must be particularly springy in nature. This can be achieved with nickel alloy wires, especially since the ring experiences a flow of cooling air through it which helps retain its resilience against the high temperatures experienced by the shroud segments.
No matter whether segmented shroud rings for turbines or compressors are considered, it is necessary to seal between adjacent shroud segments, as well as between the carrier ring and the shroud segments. Figures 1 and 2 show one known method of achieving this, in which thin springy strips 85 of a suitable metallic alloy, conforming to the shape of the outer surface of the shroud segments, have one side fixed by brazing or welding to the outer surface of one of each pair of adjacent shroud segments and bridge the gaps 87 (Figure 2) between segments so that their free (unfixed) sides are in sliding contact with - the outer surface of the other segment in each pair. It will be noted that the so-called "strip seals" 85 extend over substantially the whole axial spans of the segments between the sealing lands 38. A more effective type of strip-seal can be provided by making the shroud segments radially thick enough to have matching slots machined in the confronting edges of the shroud segments, the metallic strips being a clearance (sliding) fit in the slots so as to bridge the gaps 87.
Although the embodiments of the invention have been described as accommodating the effects of manufacturing tolerances, or the effects of manufacturing tolerances plus the effects of differential expansion, it will be apparent that the embodiments could be designed to accommodate the effects of differential expansion only, e.g., in a case in which losses of efficiency caused by clearances due to manufacturing tolerances are negligible or unimportant.
Claims (9)
1. A shroud assembly for an axial flow turbomachine rotor in which supporting structure circumferentially surrounds a shroud ring comprising a plurality of shroud segments connected to the supporting structure, and resiliently deformable seal means are provided on the supporting structure near the axially opposite edges of the shroud ring to seal between confronting portions of the supporting structure and the shroud segments, the seal means being held between said confronting portions in a deformed condition.
2. A shroud assembly according to claim 1 in which the deformation of the seal means when held in the assembly is effective to accommodate predetermined manufacturing tolerances in the shroud segments and the supporting structure which could otherwise cause unwanted clearances between the seal means and the shroud segments.
3. A shroud assembly according to claim 1 or claim 2 in which the deformation of the seal means when held in the assembly is sufficient to accommodate at least most of any reduction in deformation of the seal means due to predetermined differential thermal expansion between the shroud segments and the supporting structure during operation of the compressor rotor at elevated temperatures, which differential expansion could otherwise cause unwanted clearances between the seal means and the shroud segments.
4. A shroud assembly according to any one of claims 1 to 3 in which connection of the shroud segments to the carrier ring is achieved by circumferentially extending interengaging means provided on the supporting structure and the shroud segments, said interengaging means being effective to pull the confronting portions towards each other against the resilience of the seal means.
5. A shroud assembly according to any one of claims 1 to 4 in which the resiliently deformable seal means comprise circumferential extending bendable flanges which project from the supporting structure to span the gap between the confronting portions of the supporting structure and the shroud segments, the deformation of the seal means comprising bending of said flanges.
6. A shroud assembly according to claim 5 in which the shroud segments are adapted to be cooled by means of compressed air supplied to - chamber means between the support structure and the shroud segments; and the bendable flanges are adapted to allow a controlled amount of said compressed air to exit from said chamber means.
7. A shroud assembly according to any one of claims 1 to 4 in which the shroud segments are adapted to be cooled by means of compressed air supplied to chamber means between the support structure and the shroud segments, and the resiliently deformable seal means comprise circumferentially extending pads of bulkcompressible porous material which allow a controlled amount of said compressed air to exit from said chamber means.
8. A shroud assembly substantially as described in this specificiation with reference to and as illustrated in Figures 1 and2 of the accompanying drawings.
9. A shroud assembly substantially as described
in this specification with reference to and as
illustrated by Figure 3 or Figure 4 of the
accompanying drawings.
Priority Applications (1)
| Application Number | Priority Date | Filing Date | Title |
|---|---|---|---|
| GB08212167A GB2119452A (en) | 1982-04-27 | 1982-04-27 | Shroud assemblies for axial flow turbomachine rotors |
Applications Claiming Priority (1)
| Application Number | Priority Date | Filing Date | Title |
|---|---|---|---|
| GB08212167A GB2119452A (en) | 1982-04-27 | 1982-04-27 | Shroud assemblies for axial flow turbomachine rotors |
Publications (1)
| Publication Number | Publication Date |
|---|---|
| GB2119452A true GB2119452A (en) | 1983-11-16 |
Family
ID=10529998
Family Applications (1)
| Application Number | Title | Priority Date | Filing Date |
|---|---|---|---|
| GB08212167A Withdrawn GB2119452A (en) | 1982-04-27 | 1982-04-27 | Shroud assemblies for axial flow turbomachine rotors |
Country Status (1)
| Country | Link |
|---|---|
| GB (1) | GB2119452A (en) |
Cited By (16)
| Publication number | Priority date | Publication date | Assignee | Title |
|---|---|---|---|---|
| GB2244523A (en) * | 1990-05-31 | 1991-12-04 | Gen Electric | Gas turbine shroud assembly |
| GB2245316A (en) * | 1990-06-21 | 1992-01-02 | Rolls Royce Plc | Improvements in shroud assemblies for turbine rotors |
| EP0417958A3 (en) * | 1989-09-15 | 1992-10-28 | Rolls-Royce Plc | Improvements in or relating to shroud rings |
| GB2249356B (en) * | 1990-11-01 | 1995-01-18 | Rolls Royce Plc | Shroud liners |
| WO1995012056A1 (en) * | 1993-10-27 | 1995-05-04 | United Technologies Corporation | Mounting and sealing arrangement for a turbine shroud segment |
| WO1995013456A1 (en) * | 1993-11-08 | 1995-05-18 | United Technologies Corporation | Turbine shroud segment |
| WO1995030072A1 (en) * | 1994-04-28 | 1995-11-09 | United Technologies Corporation | Shroud segment having a cut-back retaining hook |
| US6877952B2 (en) * | 2002-09-09 | 2005-04-12 | Florida Turbine Technologies, Inc | Passive clearance control |
| EP1178182A4 (en) * | 2000-03-07 | 2005-09-07 | Mitsubishi Heavy Ind Ltd | Gas turbine split ring |
| WO2006043987A3 (en) * | 2004-05-17 | 2006-09-14 | Carlton Forge Works | Method and system for improved blade tip clearance in a gas turbine jet engine |
| JP2009133308A (en) * | 2007-11-13 | 2009-06-18 | Snecma | Turbine or compressor stage for turbomachine |
| EP2218882A1 (en) * | 2009-02-16 | 2010-08-18 | Siemens Aktiengesellschaft | Stator vane carrier system |
| FR2955359A1 (en) * | 2010-01-21 | 2011-07-22 | Snecma | Turbine stage for turbine engine e.g. turbojet or turbo propeller of airplane, has casing whose annular attachment radially extends towards upstream end of sectorized ring, where radial clearance is determined between casing and ring |
| DE102013205883A1 (en) * | 2013-04-03 | 2014-10-09 | MTU Aero Engines AG | Guide vane segment with integrated heat insulation |
| US9617822B2 (en) | 2013-12-03 | 2017-04-11 | Baker Hughes Incorporated | Compliant seal for irregular casing |
| CN118257747A (en) * | 2024-04-18 | 2024-06-28 | 中国航发湖南动力机械研究所 | Impeller cover and compressor |
Citations (6)
| Publication number | Priority date | Publication date | Assignee | Title |
|---|---|---|---|---|
| GB1199974A (en) * | 1966-11-02 | 1970-07-22 | United Aircraft Corp | Turbine Blade Seal Assembly |
| GB1406098A (en) * | 1971-11-10 | 1975-09-17 | Bbc Brown Boveri & Cie | Fluid flow machines |
| US4053254A (en) * | 1976-03-26 | 1977-10-11 | United Technologies Corporation | Turbine case cooling system |
| GB2037902A (en) * | 1978-12-20 | 1980-07-16 | United Technologies Corp | Turbine seal and vibration damper |
| GB1574981A (en) * | 1976-11-22 | 1980-09-17 | Gen Electric | Ceramic turbine shroud assemblies |
| GB2080439A (en) * | 1980-07-18 | 1982-02-03 | United Technologies Corp | An axially flexible radially stiff retaining ring for sealing in a gas turbine engine |
-
1982
- 1982-04-27 GB GB08212167A patent/GB2119452A/en not_active Withdrawn
Patent Citations (6)
| Publication number | Priority date | Publication date | Assignee | Title |
|---|---|---|---|---|
| GB1199974A (en) * | 1966-11-02 | 1970-07-22 | United Aircraft Corp | Turbine Blade Seal Assembly |
| GB1406098A (en) * | 1971-11-10 | 1975-09-17 | Bbc Brown Boveri & Cie | Fluid flow machines |
| US4053254A (en) * | 1976-03-26 | 1977-10-11 | United Technologies Corporation | Turbine case cooling system |
| GB1574981A (en) * | 1976-11-22 | 1980-09-17 | Gen Electric | Ceramic turbine shroud assemblies |
| GB2037902A (en) * | 1978-12-20 | 1980-07-16 | United Technologies Corp | Turbine seal and vibration damper |
| GB2080439A (en) * | 1980-07-18 | 1982-02-03 | United Technologies Corp | An axially flexible radially stiff retaining ring for sealing in a gas turbine engine |
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| Publication number | Priority date | Publication date | Assignee | Title |
|---|---|---|---|---|
| EP0417958A3 (en) * | 1989-09-15 | 1992-10-28 | Rolls-Royce Plc | Improvements in or relating to shroud rings |
| US5127793A (en) * | 1990-05-31 | 1992-07-07 | General Electric Company | Turbine shroud clearance control assembly |
| GB2244523B (en) * | 1990-05-31 | 1993-09-08 | Gen Electric | Turbine shroud assembly |
| GB2244523A (en) * | 1990-05-31 | 1991-12-04 | Gen Electric | Gas turbine shroud assembly |
| GB2245316A (en) * | 1990-06-21 | 1992-01-02 | Rolls Royce Plc | Improvements in shroud assemblies for turbine rotors |
| US5161944A (en) * | 1990-06-21 | 1992-11-10 | Rolls-Royce Plc | Shroud assemblies for turbine rotors |
| GB2245316B (en) * | 1990-06-21 | 1993-12-15 | Rolls Royce Plc | Improvements in shroud assemblies for turbine rotors |
| GB2249356B (en) * | 1990-11-01 | 1995-01-18 | Rolls Royce Plc | Shroud liners |
| US5927942A (en) * | 1993-10-27 | 1999-07-27 | United Technologies Corporation | Mounting and sealing arrangement for a turbine shroud segment |
| WO1995012056A1 (en) * | 1993-10-27 | 1995-05-04 | United Technologies Corporation | Mounting and sealing arrangement for a turbine shroud segment |
| WO1995013456A1 (en) * | 1993-11-08 | 1995-05-18 | United Technologies Corporation | Turbine shroud segment |
| WO1995030072A1 (en) * | 1994-04-28 | 1995-11-09 | United Technologies Corporation | Shroud segment having a cut-back retaining hook |
| EP1178182A4 (en) * | 2000-03-07 | 2005-09-07 | Mitsubishi Heavy Ind Ltd | Gas turbine split ring |
| US6877952B2 (en) * | 2002-09-09 | 2005-04-12 | Florida Turbine Technologies, Inc | Passive clearance control |
| WO2006043987A3 (en) * | 2004-05-17 | 2006-09-14 | Carlton Forge Works | Method and system for improved blade tip clearance in a gas turbine jet engine |
| JP2009133308A (en) * | 2007-11-13 | 2009-06-18 | Snecma | Turbine or compressor stage for turbomachine |
| EP2218882A1 (en) * | 2009-02-16 | 2010-08-18 | Siemens Aktiengesellschaft | Stator vane carrier system |
| FR2955359A1 (en) * | 2010-01-21 | 2011-07-22 | Snecma | Turbine stage for turbine engine e.g. turbojet or turbo propeller of airplane, has casing whose annular attachment radially extends towards upstream end of sectorized ring, where radial clearance is determined between casing and ring |
| DE102013205883A1 (en) * | 2013-04-03 | 2014-10-09 | MTU Aero Engines AG | Guide vane segment with integrated heat insulation |
| DE102013205883B4 (en) * | 2013-04-03 | 2020-04-23 | MTU Aero Engines AG | Arrangement of guide vane segments and method for producing such an arrangement |
| US9617822B2 (en) | 2013-12-03 | 2017-04-11 | Baker Hughes Incorporated | Compliant seal for irregular casing |
| CN118257747A (en) * | 2024-04-18 | 2024-06-28 | 中国航发湖南动力机械研究所 | Impeller cover and compressor |
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Legal Events
| Date | Code | Title | Description |
|---|---|---|---|
| WAP | Application withdrawn, taken to be withdrawn or refused ** after publication under section 16(1) |