U.S. patent application number 15/995525 was filed with the patent office on 2018-12-06 for clearance control arrangement.
This patent application is currently assigned to ROLLS-ROYCE plc. The applicant listed for this patent is ROLLS-ROYCE plc. Invention is credited to Leo V. LEWIS, Simon PITT.
Application Number | 20180347391 15/995525 |
Document ID | / |
Family ID | 59349941 |
Filed Date | 2018-12-06 |
United States Patent
Application |
20180347391 |
Kind Code |
A1 |
LEWIS; Leo V. ; et
al. |
December 6, 2018 |
CLEARANCE CONTROL ARRANGEMENT
Abstract
A clearance control arrangement (26) for a rotor (28), the
arrangement comprising a rotor and a casing (32) radially outside
the rotor. An annular array of segment assemblies (33) mounted to
the casing and radially spaced from the rotor by a clearance (42).
Each segment assembly comprising a heat transfer cavity (48)
radially adjacent to the casing. A birdmouth cavity (66) towards
the rear of the segment assembly. A bypass hole (68) configured to
deliver air to the birdmouth cavity to reduce the amount of air
which leaks from the heat transfer cavity to the birdmouth
cavity.
Inventors: |
LEWIS; Leo V.; (Kenilworth,
GB) ; PITT; Simon; (Derby, GB) |
|
Applicant: |
Name |
City |
State |
Country |
Type |
ROLLS-ROYCE plc |
London |
|
GB |
|
|
Assignee: |
ROLLS-ROYCE plc
London
GB
|
Family ID: |
59349941 |
Appl. No.: |
15/995525 |
Filed: |
June 1, 2018 |
Current U.S.
Class: |
1/1 |
Current CPC
Class: |
F05D 2260/201 20130101;
F04D 29/526 20130101; F05D 2240/55 20130101; F01D 25/14 20130101;
F01D 25/246 20130101; F05D 2240/14 20130101; F05D 2240/11 20130101;
F01D 11/24 20130101; F01D 11/10 20130101; F01D 11/005 20130101 |
International
Class: |
F01D 11/24 20060101
F01D011/24; F01D 25/14 20060101 F01D025/14 |
Foreign Application Data
Date |
Code |
Application Number |
Jun 1, 2017 |
GB |
1708744.6 |
Claims
1. A clearance control arrangement for a rotor, the arrangement
comprising: a rotor; a casing radially outside the rotor; an
annular array of segment assemblies mounted to the casing and
radially spaced from the rotor by a clearance; each segment
assembly comprising: a heat transfer cavity radially adjacent to
the casing; a birdmouth cavity towards the rear of the segment
assembly; a bypass hole configured to deliver air to the birdmouth
cavity to reduce the amount of air which leaks from the heat
transfer cavity to the birdmouth cavity; and a birdmouth seal
defined at the radially outer extent of a rear segment carrier.
2. An arrangement as claimed in claim 1 wherein the birdmouth
cavity is downstream of the birdmouth seal.
3. An arrangement as claimed in claim 2 further comprising a rear
hook which supports the rear segment carrier, the birdmouth cavity
formed between the rear hook and the rear segment carrier.
4. An arrangement as claimed in claim 1 wherein the birdmouth
cavity is upstream of the birdmouth seal.
5. An arrangement as claimed in claim 4 wherein the birdmouth
cavity is separated from the heat transfer cavity by a rib.
6. An arrangement as claimed in claim 1 further comprising a
segment cooling cavity at the radially inner extent of the segment
assembly.
7. An arrangement as claimed in claim 6 wherein the bypass hole is
configured to receive air from the segment cooling cavity.
8. An arrangement as claimed in claim 7 further comprising a supply
cavity radially between the heat transfer cavity and the segment
cooling cavity.
9. An arrangement as claimed in claim 8 wherein the bypass hole is
configured to receive air from the supply cavity
10. An arrangement as claimed in claim 6 further comprising a
supply cavity radially between the heat transfer cavity and the
segment cooling cavity.
11. An arrangement as claimed in claim 10 wherein the bypass hole
is configured to receive air from the supply cavity.
12. An arrangement as claimed in claim 1 comprising an array of
bypass holes.
13. An arrangement as claimed in claim 1 further comprising a first
supply hole configured to allow ingress of air to the heat transfer
cavity.
14. An arrangement as claimed in claim 1 further comprising a front
hook which supports a front segment carrier, the front hook
configured to allow ingress of air to the heat transfer cavity.
15. An arrangement as claimed in claim 1 further comprising an
array of controlled entry holes configured to allow ingress of air
to the heat transfer cavity.
16. An arrangement as claimed in claim 1 wherein the segment
assembly includes cooling air delivery holes through its radially
inner wall.
17. A gas turbine engine comprising an arrangement as claimed in
claim 1.
Description
CROSS-REFERENCE TO RELATED APPLICATION
[0001] This application is based upon and claims the benefit from
British Patent Application No. GB 1708744.6, filed on 1 Jun. 2017,
the entire contents of which are hereby incorporated by
reference.
BACKGROUND
Field
[0002] The present disclosure concerns a clearance control
arrangement for a rotor. It finds utility for a rotor stage of a
gas turbine engine.
Description of Related Art
[0003] A gas turbine engine rotor stage typically has a rotor with
a casing radially outside it. Mounted radially inside the casing is
an array of segments. There is a small clearance between the
segments and the tips of the rotor blades. Cooling air may be
directed into the segment assemblies and directed towards the rotor
blades to cool the segment. Cool air may also be impinged on the
outside of the casing to change the rate at which it expands or
contracts thermally to maintain the clearance at a preferred
level.
SUMMARY
[0004] According to a first aspect there is provided a clearance
control arrangement for a rotor, the arrangement comprising: [0005]
a rotor; [0006] a casing radially outside the rotor; [0007] an
annular array of segment assemblies mounted to the casing and
radially spaced from the rotor by a clearance; each segment
assembly comprising: [0008] a heat transfer cavity radially
adjacent to the casing; [0009] a birdmouth cavity towards the rear
of the segment assembly; and [0010] a bypass hole configured to
deliver air to the birdmouth cavity to reduce the amount of air
which leaks from the heat transfer cavity to the birdmouth
cavity.
[0011] Advantageously the birdmouth cavity is independently fed by
the bypass hole so that leakage from the heat transfer cavity is
reduced. Advantageously the mass flow into the heat transfer cavity
can therefore be reduced to a level which is suitable for its
primary purpose of controlling the clearance.
[0012] The arrangement may further comprise a birdmouth seal
defined at the radially outer extent of a rear segment carrier.
Advantageously the pressure differential across the birdmouth seal
can be reduced by supplying air to it through the bypass hole.
[0013] The birdmouth cavity may be downstream of the birdmouth
seal. The arrangement may further comprise a rear hook which
supports the rear segment carrier. The birdmouth cavity may be
formed between the rear hook and the rear segment carrier. The
birdmouth cavity may be formed at an extant junction between the
rear hook and the rear segment carrier, for example by providing a
radius or chamfer on one or both components. Alternatively the
birdmouth cavity may be formed by providing an additional flange on
the rear hook or rear segment carrier to form a new cavity.
[0014] Alternatively the birdmouth cavity may be upstream of the
birdmouth seal. The birdmouth cavity may be separated from the heat
transfer cavity by a rib. The rib may extend towards the casing.
Advantageously the pressure differential across the rib may be
substantially equalised so no applied sealing is required.
Advantageously the birdmouth cavity is contained within the space
envelope of the segment assembly.
[0015] The arrangement may further comprise a segment cooling
cavity at the radially inner extent of the segment assembly. The
segment cooling cavity may supply air into the clearance. The
bypass hole may be configured to receive air from the segment
cooling cavity. Advantageously the bypass hole is supplied from a
substantially unmetered cavity.
[0016] The arrangement may further comprise a supply cavity
radially between the heat transfer cavity and the segment cooling
cavity. The bypass hole may be configured to receive air from the
supply cavity. Advantageously the birdmouth cavity is supplied by
air which is independent of that used for the segment cooling or
for affecting the temperature of the casing via the heat transfer
cavity.
[0017] The arrangement may comprise an array of bypass holes. The
bypass holes may be regularly spaced or irregularly spaced in the
circumferential direction. There may be one bypass hole in each
segment assembly. Alternatively there may be more than one bypass
hole in each segment assembly. In a further alternative there may
be one bypass hole in one or more of the segment assemblies and
more than one bypass hole in one or more others of the segment
assemblies.
[0018] The arrangement may further comprise a first supply hole
configured to allow ingress of air to the heat transfer cavity.
There may be an array of first supply holes. The array may extend
in the radial and/or circumferential directions.
[0019] The arrangement may further comprise a front hook which
supports a front segment carrier. The front hook may be configured
to allow ingress of air to the heat transfer cavity. The front hook
may be intermittent in the circumferential direction. Alternatively
it may include one or more slots, holes or apertures.
[0020] The arrangement may comprise an array of controlled entry
holes configured to allow ingress of air to the heat transfer
cavity. The controlled entry holes may be supplied from the segment
cooling cavity or the supply cavity. Alternatively they may be
supplied from outside, upstream of, the segment assembly.
[0021] The segment assembly may include cooling air delivery holes
through its radially inner wall. The cooling air delivery holes may
be supplied from the segment cooling cavity. The cooling air
delivery holes may be angled or positioned to preferentially cool
portions of the rotor blade tip across the clearance.
[0022] According to a second aspect there is provided a gas turbine
engine comprising an arrangement as described.
[0023] The skilled person will appreciate that except where
mutually exclusive, a feature described in relation to any one of
the above aspects may be applied mutatis mutandis to any other
aspect. Furthermore except where mutually exclusive any feature
described herein may be applied to any aspect and/or combined with
any other feature described herein.
BRIEF DESCRIPTION OF THE SEVERAL VIEWS OF THE DRAWINGS
[0024] Embodiments will now be described by way of example only,
with reference to the Figures, in which:
[0025] FIG. 1 is a sectional side view of a gas turbine engine;
[0026] FIG. 2 is a schematic illustration of a clearance control
arrangement; AND
[0027] FIGS. 3-9 are schematic illustrations of other clearance
control arrangements.
DETAILED DESCRIPTION
[0028] With reference to FIG. 1, a gas turbine engine is generally
indicated at 10, having a principal and rotational axis 11. The
engine 10 comprises, in axial flow series, an air intake 12, a
propulsive fan 13, an intermediate pressure compressor 14, a
high-pressure compressor 15, combustion equipment 16, a
high-pressure turbine 17, an intermediate pressure turbine 18, a
low-pressure turbine 19 and an exhaust nozzle 20. A nacelle 21
generally surrounds the engine 10 and defines both the intake 12
and the exhaust nozzle 20.
[0029] The gas turbine engine 10 works in the conventional manner
so that air entering the intake 12 is accelerated by the fan 13 to
produce two air flows: a first air flow into the intermediate
pressure compressor 14 and a second air flow which passes through a
bypass duct 22 to provide propulsive thrust. The intermediate
pressure compressor 14 compresses the air flow directed into it
before delivering that air to the high pressure compressor 15 where
further compression takes place.
[0030] The compressed air exhausted from the high-pressure
compressor 15 is directed into the combustion equipment 16 where it
is mixed with fuel and the mixture combusted. The resultant hot
combustion products then expand through, and thereby drive the
high, intermediate and low-pressure turbines 17, 18, 19 before
being exhausted through the nozzle 20 to provide additional
propulsive thrust. The high 17, intermediate 18 and low 19 pressure
turbines drive respectively the high pressure compressor 15,
intermediate pressure compressor 14 and fan 13, each by suitable
interconnecting shaft.
[0031] Other gas turbine engines to which the present disclosure
may be applied may have alternative configurations. By way of
example such engines may have an alternative number of
interconnecting shafts (e.g. two) and/or an alternative number of
compressors and/or turbines. Further the engine may comprise a
gearbox provided in the drive train from a turbine to a compressor
and/or fan. FIG. 2 shows a clearance control arrangement 26. The
clearance control arrangement 26 includes a rotor blade 28 which is
one of an annular array of rotor blades 28. The rotor blades 28 may
form any one of the rotor stages of the intermediate pressure
compressor 14, high pressure compressor 15, high pressure turbine
17, intermediate pressure turbine 18 or low pressure turbine 19.
Each rotor blade 28 includes a tip 30 at its radially outer end.
The tip 30 may be parallel to the engine axis 11 or may be angled,
curved or another more complex shape as known to the skilled
reader. The tip 30 may include fences, shrouds or other
features.
[0032] The clearance control arrangement 26 also includes a casing
32 which is annular and is arranged radially outside the rotor
blades 28. The casing 32 may extend axially parallel to the engine
axis 11 or may have a conical shape or other more complex shape.
Typically the shape of the casing 32 radially outside the rotor
blades 28 approximately matches the shape inscribed by the rotor
blade tips 30.
[0033] A plurality of segments 34 are mounted radially inside the
casing 32. The segments 34 form an annular array. There may be the
same number of segments 34 as there are rotor blades 28, or there
may be more segments 34 or fewer segments 34. Each segment 34
extends circumferentially so that the radially inner surfaces of
all the segments 34 form a substantially continuous fluid-washed
surface over which working fluid of the gas turbine engine 10 flows
as it passes between and over the tips 30 of the rotor blades
28.
[0034] The segments 34 are each mounted to a segment carrier 35.
The segment 34 and segment carrier 35 together are referred to as
the segment assembly 33. Alternatively the segments 34 may be
integrally formed with or coupled to the segment carrier 35 so that
the segment assembly 33 forms a single part.
[0035] Radially outside the casing 32 there may be one or more cool
air chambers 36 having an array of cooling holes 38 through its
radially inner surface. The cool air chamber 36 is selectively
filled with cool air, for example by opening or closing a valve 40.
The cool air is delivered from the cool air chamber 36 through the
cooling holes 38 to impinge against the casing 32 in the axial
vicinity of the rotor blades 28. The cool air acts to retard the
thermal growth of the casing 32 and therefore causes the radial
clearance 42 between the rotor tips 30 and the segments 34 to be
held small.
[0036] Each segment assembly 33 includes a number of cavities and
chambers. At the radially inner extent of the segment assembly 33
may be a segment cooling cavity 44. The segment cooling cavity 44
includes cooling air delivery holes 46 through its radially inner
wall, which is formed by the segment 34. The holes 46 form an array
arranged in any suitable pattern in order to cool the segment 34.
They may also form vortices or other fluid forms to aerodynamically
reduce the clearance 42 perceived by the working fluid.
[0037] Each segment assembly 33 includes a heat transfer cavity 48
at the radially outer extent of the segment assembly 33, radially
proximal the casing 32. The inner extent of the heat transfer
cavity 48 may be defined by a plate 50. The outer extent is defined
by the casing 32. The upstream extent is defined by the front
segment carrier 35 and front hook 52. The downstream extent is
defined by the rear segment carrier 35 and rear hook 54. The front
hook 52 is configured to support the front segment carrier 35
whilst the rear hook 54 is configured to support the rear segment
carrier 35. The front hook 52 may be a fully annular ring or may be
intermittent in the circumferential direction. The rear hook 54 is
a fully annular ring.
[0038] The heat transfer cavity 48 may be supplied with air through
the intermittent gaps in the front hook 52. Alternatively there may
be a first supply hole or array of first supply holes 56 which
allows air ingress to the heat transfer cavity 48. The first supply
hole 56 may be provided through the front segment carrier 35 in
some arrangements. Additionally or alternatively the heat transfer
cavity 48 may be supplied with air via controlled entry holes 64
through the plate 50, as shown in FIG. 4.
[0039] The heat transfer cavity 48 may be supplied with relatively
hot air from a chamber upstream of the segment assemblies 33 and
radially inside the casing 32, or from some other source.
Alternatively relatively cool air may be supplied from a chamber
upstream of the segment assemblies 33, for example air which has
been pre-cooled through a heat exchanger or similar. The air
supplied to the heat transfer cavity 48 acts to control the heat
transfer coefficient across the casing 32. Advantageously the air
impinged on the casing 32 from the heat transfer cavity 48 may be
hotter than the casing 32 itself and so may heat up the casing 32,
at least in the axial vicinity of the rotor blades 28. For engine
transients, where the rotor disc grows thermally more quickly than
the casing 32 would otherwise expand, such impingement heating can
maintain the size of the clearance 42, and therefore prevent the
clearance 42 reducing to levels where the blade tips 30 may rub the
radially inner surface of the segments 34 causing permanent damage,
by heating the casing 32 at a similar rate to the thermal growth of
the rotor disc. The amount of air needed to be delivered to the
heat transfer cavity 48 in order to control the clearance 42 is
small. Alternatively the air supplied through the controlled entry
holes 64 may reduce the heat transfer across the casing 32.
[0040] There is a birdmouth seal 58 at the rear of the heat
transfer cavity 48. The birdmouth seal 58 may be formed between the
radially outer end of the rear segment carrier 35 and the casing
32. Alternatively it may be formed between an axially extending
portion of the rear hook 54 and the rear segment carrier 35.
Alternatively the birdmouth seal 58 may be between a different
portion of the rear hook 54 and the rear segment carrier 35. Air
from the heat transfer cavity 48 leaks through the birdmouth seal
58 to an area axially downstream of the segment assembly 33. The
leakage through the birdmouth seal 58 is governed by the pressure
differential across it, which is generally large. This means that
the mass flow pulled from the heat transfer cavity 48 across the
birdmouth seal 58 to the downstream area is large and may become
the governing factor for the amount of air supplied to the heat
transfer cavity 48 in known segment assemblies 33.
[0041] Air may be supplied to the segment cooling cavity 44 via a
second supply hole or array of second supply holes 60. There may
also be a metering hole or array of metering holes 62, shown in
FIG. 3, which deliver air from the heat transfer cavity 48 to the
segment cooling cavity 44. The metering hole 62 can be sized to
control the amount of air drawn into and through the heat transfer
cavity 48 for rotor tip clearance purposes.
[0042] The clearance control arrangement 26 includes a birdmouth
cavity 66. The birdmouth cavity 66 is situated towards the rear of
the segment assembly 33. The birdmouth cavity 66 may be upstream or
downstream of the birdmouth seal 58 as will be described. The
birdmouth cavity 66 is supplied with air through a bypass hole or
an array of bypass holes 68. The bypass holes 68 deliver air to the
birdmouth cavity 66 in order to reduce the pressure differential
across the birdmouth seal 58. Consequently the leakage mass flow
from the heat transfer cavity 48 reduces and so the amount of air
drawn into the heat transfer cavity 48 also reduces.
[0043] In a first aspect of the clearance control arrangement 26,
as shown in FIG. 2, the birdmouth cavity 66 is formed downstream of
the birdmouth seal 58, between the rear hook 54 and the rear
segment carrier 35. The rear hook 54 is shaped to have two axially
extending portions with a radially extending wall joining their
downstream ends. The axially extending portion which is closer to
the casing 32 forms the hook which supports the rear segment
carrier 35. The other axially extending portion of the rear hook 54
closes the birdmouth cavity 66. The bypass hole 68 is arranged to
supply air from the segment cooling cavity 44 to the birdmouth
cavity 66.
[0044] Advantageously the bypass hole 68 supplies the air which
then leaks past the inner axially extending portion. Advantageously
the birdmouth cavity 66 is pressurised enough to reduce the
pressure differential, and thus leakage across, the birdmouth seal
58. The birdmouth cavity 66 may be sufficiently pressurised so that
there is substantially no pressure differential, and consequently
no air flow, across the birdmouth seal 58.
[0045] Optionally the segment assembly 33 may include a supply
cavity 70 radially between the heat transfer cavity 48 and the
segment cooling cavity 44, as shown in FIG. 5. The radially outer
extent of the supply cavity 70 may be defined by the plate 50 and
the radially inner extent may be defined by a second plate 72. The
supply cavity 70 may receive air through the second supply hole 60.
It may be supplied with relatively hot air from a chamber upstream
of the segment assemblies 33 and radially inside the casing 32, or
from some other source. Alternatively relatively cool air may be
supplied from a chamber upstream of the segment assemblies 33, for
example air which has been pre-cooled through a heat exchanger or
similar.
[0046] The supply chamber 70 may deliver air to the segment cooling
cavity 44 through a delivery hole or array of delivery holes 74 in
the second plate 72. The supply cavity 70 may also be the source
for the air which is delivered to the heat transfer cavity 48
through the optional controlled entry holes 64. The bypass hole 68
may receive air from the supply cavity 70 for delivery to the
birdmouth cavity 66.
[0047] Advantageously the pressure in the birdmouth cavity 66 may
be high enough to set the leakage across the birdmouth seal 58 to
zero, or close to zero. This is because the flow supplied to the
birdmouth cavity 66 does not affect the pressure of the flow in the
segment cooling cavity 44. Advantageously the tip clearance 42 is
wholly controlled by sized holes and not by leakage flows.
[0048] In a second aspect of the segment assembly 33, shown in FIG.
6 and FIG. 7, the birdmouth cavity 66 is again formed between the
rear hook 54 and the rear segment carrier 35. However, unlike in
the first aspect the rear hook 54 only includes one axially
extending portion. The rear hook 54 includes a chamfer, radius or
other cut away on the radially outer surface of the axially
extending portion which leaves a space adjacent to the rear segment
carrier 35 to form the birdmouth cavity 66. Thus in the second
aspect the birdmouth cavity 66 is smaller than in the first aspect.
Nevertheless it is sufficiently large to be pressurised and
therefore to reduce the pressure differential across the birdmouth
seal 58 so that the leakage mass flow is reduced.
[0049] Advantageously existing segment assemblies 33 may be adapted
to provide the birdmouth cavity 66 of the second aspect by
machining away a part of the rear hook 54 and drilling the bypass
hole 68. Thus the benefit of the birdmouth cavity 66 an be realised
via a modification of existing hardware.
[0050] The birdmouth cavity 66 is again fed with air through the
bypass hole 68. The bypass hole 68 may be supplied from the segment
cooling cavity 44, FIG. 6 or from the supply cavity 70, FIG. 7. The
bypass hole 68 may be wholly within the rear segment carrier 35 or
may partially pass through the plate 50.
[0051] In the second aspect the plate 50 may include the metering
hole 62 and/or may include the controlled entry holes 64.
Alternatively the heat transfer cavity 48 may be supplied solely by
the first supply hole 56 either through the intermittent gaps in
the front hook 52 or through the front segment carrier 35.
[0052] A third aspect is shown in FIG. 8 and FIG. 9. Each includes
the optional controlled entry holes 64 which may alternatively be
omitted. In the third aspect the rear hook 54 does not include a
chamfer, radius or cut away. Instead the birdmouth cavity 66 is
provided upstream of the birdmouth seal 58. A radially extending
rib 76 is provided to truncate the heat transfer cavity 48 in the
axial direction. The rib 76 may be mounted to or integrally formed
with the plate 50 and extends towards the casing 32. The radially
outer end of the rib 76 is close to, but not attached to, the
casing 32 and does not require any applied sealing. The pressure
across the rib 76 is balanced by supplying air from the segment
cooling cavity 44 to the birdmouth cavity 66 via the bypass hole
68. Because the pressure is balanced there is little or no leakage
of air from the heat transfer cavity 48 into the birdmouth cavity
66. The metering hole 62 may therefore be beneficial to maintain
flow through the heat transfer cavity 48. Furthermore the leakage
across the birdmouth seal 58 downstream of the birdmouth cavity 66
is supplied by the birdmouth cavity 66, which itself is supplied
from the segment cooling cavity 44. The bypass hole 68 may have
approximately twice the flow area of the metering hole 62. Thus any
variation in the amount of flow in the heat transfer cavity 48 will
be approximately one third of the variation in the amount of flow
across the birdmouth seal 58. This compares favourably with known
arrangements where the variation of flow in the heat transfer
cavity 48 matched the variation in flow across the birdmouth seal
58.
[0053] FIG. 9 is similar to FIG. 8 but includes the optional supply
cavity 70. The bypass hole 68 is illustrated to be configured to
receive air from the segment cooling cavity 44. However, it may
alternatively be supplied from the supply cavity 70. In this
alternative the bypass hole 68 and metering hole 62 may be mutually
offset circumferentially.
[0054] Advantageously the birdmouth cavity 66 supplied by the
bypass hole 68 allows the air flow requirement for the heat
transfer across the casing 32, in the heat transfer cavity 48, to
be independent of the leakage across the birdmouth seal 58.
Advantageously the mass flow to be delivered into the heat transfer
cavity 48 can be reduced relative to known arrangements without a
separately supplied birdmouth cavity 66. This improves the
transient rotor tip clearance control.
[0055] Advantageously, deterioration through life, variation
between turbine stages of a gas turbine engine 10 and variation
between the turbines of different gas turbine engines 10 can be
better accommodated since the segment assembly 33 is less sensitive
to changes or differences in the birdmouth leakage. That is, if the
birdmouth seal 58 deteriorates through life or is less effective
(within its defined tolerance limits) the flow across the birdmouth
seal 58 will be larger than intended. However, the required
increase in air flow will be predominantly or wholly sourced from
the segment cooling cavity 44 or supply cavity 70 and not from the
heat transfer cavity 48 so the effect on the tip clearance control
is minimal.
[0056] The clearance control arrangement 26 finds particular
utility for a rotor in a gas turbine engine 10. Such a gas turbine
engine 10 may be used to power an aircraft or a marine vessel. The
arrangement 26 may be used on one or more than one rotor stage. For
example it may be used for a rotor stage of the high pressure
turbine 17, the intermediate pressure turbine 18 or the low
pressure turbine 19. It may be used on each of several rotor stages
of one of the turbines 17, 18, 19 whether the stages are
consecutive or separated by other rotor stages. The arrangement 26
may also be used for rotor stages of the compressors, 14, 15.
[0057] It will be understood that the invention is not limited to
the embodiments above-described and various modifications and
improvements can be made without departing from the concepts
described herein. Except where mutually exclusive, any of the
features may be employed separately or in combination with any
other features and the disclosure extends to and includes all
combinations and sub-combinations of one or more features described
herein.
* * * * *