U.S. patent application number 14/044460 was filed with the patent office on 2014-04-03 for gas turbine engine component.
The applicant listed for this patent is Rolls-Royce PLC. Invention is credited to Dougal Richard JACKSON, Ian TIBBOTT.
Application Number | 20140093379 14/044460 |
Document ID | / |
Family ID | 49263249 |
Filed Date | 2014-04-03 |
United States Patent
Application |
20140093379 |
Kind Code |
A1 |
TIBBOTT; Ian ; et
al. |
April 3, 2014 |
GAS TURBINE ENGINE COMPONENT
Abstract
A gas turbine engine component is described which (100),
comprises: a shell having an internal cavity for receiving a
multi-part insert; a multi-part insert located within the cavity,
wherein the multi-part insert comprises separate insert parts
assembled in an abutting relation with one another within the
cavity to provide the multi-part insert; an insertion aperture
within a wall of the shell which is sized to receive each of the
insert parts individually and wherein the multi-part insert cannot
be withdrawn from the cavity through the insertion aperture when
assembled.
Inventors: |
TIBBOTT; Ian; (Lichfield,
GB) ; JACKSON; Dougal Richard; (Stanton by Bridge,
GB) |
|
Applicant: |
Name |
City |
State |
Country |
Type |
Rolls-Royce PLC |
London |
|
GB |
|
|
Family ID: |
49263249 |
Appl. No.: |
14/044460 |
Filed: |
October 2, 2013 |
Current U.S.
Class: |
416/224 ;
29/889.7 |
Current CPC
Class: |
F05D 2260/30 20130101;
F01D 5/188 20130101; F01D 5/189 20130101; Y10T 29/49332 20150115;
F05D 2260/201 20130101; F05D 2260/202 20130101; Y10T 29/49336
20150115; B33Y 80/00 20141201 |
Class at
Publication: |
416/224 ;
29/889.7 |
International
Class: |
F01D 5/18 20060101
F01D005/18 |
Foreign Application Data
Date |
Code |
Application Number |
Oct 3, 2012 |
GB |
1217650.9 |
Oct 3, 2012 |
GB |
1217652.5 |
Claims
1. A gas turbine engine component (100), comprising a shell having
an internal cavity for receiving a multi-part insert; a multi-part
insert located within the cavity, wherein the multi-part insert
comprises separate insert parts assembled in an abutting relation
with one another within the cavity to provide the multi-part
insert; an insertion aperture within a wall of the shell which is
sized to receive each of the insert parts individually and wherein
the multi-part insert cannot be withdrawn from the cavity through
the insertion aperture when assembled.
2. A gas turbine component as claimed in claim 1, wherein the
cavity includes an insertion portion into which the insert parts
are inserted, and a receiving portion in which at least one of the
insert parts is located when the insert is assembled.
3. A gas turbine component as claimed in claim 2, wherein the
receiving portion is at least partially obscured by a wall or an
internal protuberant feature of the shell when viewed from the
insertion aperture.
4. A gas turbine component as claimed in claim 3, wherein the
obstruction of the receiving portion is caused by a twist along the
length of the cavity.
5. A gas turbine component as claimed in claim 3, wherein the
cavity includes a widened portion along the length thereof and the
receiving portion may be located within the widened portion of the
cavity.
6. A gas turbine engine as claimed in claim 1 where the assembled
insert includes at least one retention part, wherein the retention
part acts to engage with a portion of the cavity and the at least
one other insert part so as to retain the assembled insert within
the cavity.
7. A gas turbine component as claimed in claim 1, wherein the
retention part provides a resilient bias which acts to urge the
retention part and or another insert part against one or more walls
of the shell.
8. A gas turbine component as claimed in claim 1, wherein the
retention part may be oversized relative to the size required when
in situ such that inserting the retention part into the cavity
requires a deformation of the part and a resulting stressing to
provide the bias.
9. A gas turbine component as claimed in claim 1, wherein at least
one insert piece is made by additive layer manufacturing; and
wherein the insert includes formations which support the insert
within the shell and guide the cooling air around the inner surface
of the shell.
10. A gas turbine engine component as claimed in claim 1, wherein
the shell is a ceramic matrix composite shell.
11. A gas turbine engine component as claimed in claim 1, wherein
the formations include projections.
12. A gas turbine engine component according to claim 1, wherein at
least one of the insert parts predominantly includes trip strip
formations which lie along the inner surface of the shell when the
insert is assembled.
13. A gas turbine engine component as claimed in claim 1, wherein
the shell forms an aerofoil and includes a wall which divides the
shell into a front cavity at a leading edge region of the component
and a rear cavity at a trailing edge region of the component; and
wherein the multi-part insert is located in the front cavity or a
rear insert located in the rear cavity.
14. A gas turbine engine component as claimed in claim 13, wherein
the divider wall includes apertures which provide fluid
communication between the front and rear cavity, and at least one
part of the multiple insert parts includes a sealing plate to
restrict or prevent the flow of cooling air across the divider
wall.
15. A method of forming the gas turbine engine component according
to claim 1, the method including the steps of: providing the shell;
providing a plurality of insert parts which are configured to be
assembled in an abutting relation with one another within the
cavity to provide the multi-part insert; wherein the assembled
insert includes at least one retention part which engages with a
wall of the cavity and at least one other insert part so as to
retain the assembled insert within the cavity.
Description
FIELD OF THE INVENTION
[0001] The present invention relates to a gas turbine engine
component having a cavity defining shell which receives an insert
therein. The invention finds particular use in ceramic matrix
composite shells but also in more traditional metal shells.
BACKGROUND OF THE INVENTION
[0002] The performance of the simple gas turbine engine cycle,
whether measured in terms of efficiency or specific output, is
improved by increasing the turbine gas temperature. It is therefore
desirable to operate the turbine at the highest possible
temperature. For any engine cycle compression ratio or bypass
ratio, increasing the turbine entry gas temperature always produces
more specific thrust (e.g. engine thrust per unit of air mass
flow). However, as turbine entry temperatures increase, the life of
an uncooled turbine falls, necessitating the development of better
materials and the introduction of internal air cooling.
[0003] In modern engines, the high pressure (HP) turbine gas
temperatures are now much hotter than the melting point of the
blade materials used, and in some engine designs the intermediate
pressure (IP) and low pressure (LP) turbines are also cooled.
During its passage through the turbine; the mean temperature of the
gas stream decreases as power is extracted. Therefore the need to
cool the static and rotary parts of the engine structure decreases
as the gas moves from the HP stage(s) through the IP and LP stages
towards the exit nozzle.
[0004] Internal convection and external films are the main methods
of cooling the aerofoils. HP turbine nozzle guide vanes (NGVs)
consume the greatest amount of cooling air on high temperature
engines. HP blades typically use about half of the NGV cooling air
flow. The IP and LP stages downstream of the HP turbine use
progressively less cooling air.
[0005] FIG. 1 shows an isometric view of a conventional HP stage
cooled turbine. Block arrows indicate cooling air flows. The stage
has NGVs 100 with inner 102 and outer 104 platforms and HP rotor
blades 106 downstream of the NGVs, blade platform 112 and shroud
114. Cooling air can enter NGVs as a single end feed (i.e. in one
direction) or a dual end feed (i.e. an inboard and an outboard
feed). An aim of the dual feed is to ensure that adequate backflow
margin exists at all flight conditions.
[0006] The NGVs and HP blades are cooled using high pressure (HP)
air from the compressor that has by-passed the combustor and is
therefore relatively cool compared to the gas temperature. Typical
cooling air temperatures are between 800 and 1000K. Mainstream gas
temperatures can be in excess of 2100K.
[0007] The cooling air from the compressor that is used to cool the
hot turbine components is not used fully to extract work from the
turbine. Extracting coolant flow therefore has an adverse effect on
the engine operating efficiency. Thus, it is important that the
cooling air is used as effectively as possible.
[0008] Improvements in Ceramic Matrix Composite (CMC) technology
have resulted in its use in HP turbine components becoming more
common. CMC can be used to replace metal static components such as
high temperature seal segments, and also, more recently NGVs and
other aerofoil components.
[0009] CMC materials have a high temperature capability and low
thermal conductivity. Environmental barrier coatings (EBC) are
typically applied to the CMC material. It can be shown that using
coated CMC materials such as SiC--SiC, where long multi-strand
fibres of silicon carbide are integrated into a silicon carbide
matrix, cooling mass flows can be reduced by approximately 40%
relative to similar NGV designs made from single crystal nickel
alloys.
[0010] The introduction of CMCs does not eliminate the need for
cooling, although the quantity of coolant required to ensure
adequate durability reduces considerably. CMCs may be formed by a
laser sintering manufacturing process. However, this process can
only be used to produce relatively simple non-detailed structures
such as a hollow aerofoil shape with a centrally located divider
wall. A composite produced by laser sintering will generally be
porous, but the addition of a protective coating can help to
protect against environmental attack.
[0011] It is known that additional cooling of a hollow turbine
engine component can be achieved by providing sheet metal inserts
such as tubes or plates which provide impingement cooling by
directing cooling air onto the inside walls of the hollow
component. The sheet metal inserts may be adapted to provide
location supports in the form of pressed dimples.
[0012] With engine cycle gas temperatures rising and combustion
temperature profiles becoming flatter, as a consequence of the
drive to reduce NOx and CO.sub.2 emissions, there is an increasing
need to make better use of the cooling air in addition to utilising
the advantages provided by the CMC material.
[0013] Although the use of CMC material shells with the inserts of
the invention is particularly advantageous, the inserts can be used
with non-CMC materials, such as traditional metal shells which may
be cast as is known in the art. EP0392664 describes a blade for a
combined cycle turbine in which inserts are used to define conduits
for the transportation and recovery of steam for cooling purposes.
However, how the blades are constructed with the inserts is not
described.
[0014] The present invention seeks to provide inserts which may be
placed within shells having irregular cavities which may not
ordinarily be able to receive an insert.
SUMMARY OF THE INVENTION
[0015] The invention provides a gas turbine according to the
appended claims. In particular there is provided a gas turbine
engine component (100), comprising: a shell having an internal
cavity for receiving a multi-part insert; a multi-part insert
located within the cavity, wherein the multi-part insert comprises
separate insert parts assembled in an abutting relation with one
another within the cavity to provide the multi-part insert; an
insertion aperture within a wall of the shell which is sized to
receive each of the insert parts individually and wherein the
multi-part insert cannot be withdrawn from the cavity through the
insertion aperture when assembled.
[0016] The component may include an aerofoil for a gas turbine
engine. The component may be a blade or a vane. The blade or vane
may be for use in the turbine of the gas turbine engine.
[0017] The cavity may include an insertion aperture or portion into
which the insert parts are inserted, and a receiving portion in
which at least one of the insert parts is located when the insert
is assembled. The receiving portion may be at least partially
obscured by a wall or an internal protuberant feature of the shell
when viewed from the insertion aperture. The insertion aperture may
be defined by wall of a cavity, or may be defined as part of a
larger opening. The insertion aperture may be defined by a portion
of a larger opening through which an insertion part can be
inserted. The receiving portion may be different for each insert
part. The insertion aperture may be different for each insertion
part. An shell may have an insertion aperture in each end thereof,
each for a different insert part.
[0018] The obstruction of the receiving portion may be caused by a
twist along the length of the cavity. The obstructing wall may be
the leading or trailing edge of the aerofoil or the pressure or
suction surface wall. Alternatively or additionally, the
obstructing wall may be a dividing wall. The obstruction may be due
to a distortion in the shape of the cavity. The cavity may be
irregularly shaped along the length thereof. The cavity may be
twisted or bent along the length thereof. The twist may be chordal.
That is, the twist may be provided by an angular offset between a
first end and a second end of the aerofoil relative to the
longitudinal axis of the aerofoil. The cavity may include one or
more features around which the insert must be placed. The one or
more features may include cooling holes or projections.
[0019] The twisting may be due to the aerodynamic profiling of the
outer surface of the component. The cavity of the shell may be
provided by a wall of the shell. The internal surface of the shell
may be smooth. That is, the internal surface of the shell may be
devoid of surface features. Such features may include but are not
restricted to cooling and turbulating features such as pedestals
and trip strips.
[0020] The cavity may widen along the length thereof and the
receiving portion may be located within the widened portion of the
cavity. The cavity may include a recess. The recess may provide a
receiving portion for part of an insert. The recess may be towards
the trailing edge of the blade. The recess may be provided by
another part of the multi-part insert.
[0021] The maximum width of the assembled multi-part insert may be
greater than that of the maximum width of the insertion
aperture.
[0022] The assembled insert may include at least one retention
part, wherein the retention part acts to engage with a portion of
the cavity and the at least one other insert part so as to retain
the assembled insert within the cavity.
[0023] The retention part may provide an interference fit with
other insert parts and or a wall of the shell so as to provide a
chock. The retention part may provide a resilient bias.
[0024] The retention of the assembled insert with the retention
part may be for assembly purposes only. As such, the insertion
aperture may be partially or completely blocked after the insert is
located within the cavity. For example, the insertion aperture may
be covered with a cap or plate attached over the insertion
aperture.
[0025] The retention part may provide a resilient bias which acts
to urge the retention part and or another insert part against one
or more walls of the shell.
[0026] The retention piece may include two members joined at a
hinge portion. The hinge portion may be sprung loaded to provide
the resilient bias. The hinge portion may be plastically deformed
prior to assembly. The hinge portion may be connected to the
members so as to provide an angle of separation between the two
members. The angle of separation between the members may be greater
prior to assembly such that the arms need to be forcibly moved
together for insertion into the cavity. Forcing the arms of the
retention part together can elastically deform the hinge part such
that it is resiliently biased against a wall of a cavity or another
one of the insert parts when the retention part is placed in
situ.
[0027] The retention part may be oversized relative to the size
required when in situ such that inserting the retention part into
the cavity requires a deformation of the part and a resulting
stressing to provide the bias.
[0028] At least one insert piece may be made by additive layer
manufacturing. The insert may include formations which support the
insert within the shell and guide the cooling air around the inner
surface of the shell. The formations may include projections. The
shell may be a ceramic matrix composite shell.
[0029] The projections may be fins. The fins may be pin-fins. The
formations may form one or more chambers, between the insert and
the inner surface of the shell, the or each chamber being
configured so that, in use, the chamber receives cooling air from
the one or more flow channels, the cooling air pressure being lower
in the chamber than in the flow channels. The insert may form a
plurality of flow channels in fluid communication with one another
to define a multi-pass cooling arrangement.
[0030] At least one of the insert parts may predominantly include
trip strip formations which lie along the inner surface of the
shell when the insert is assembled. The separate insert parts may
include one or more support structures for engagement with the
insert and the trip strip formations. The strip trip formation and
or support structures may be elongate members in the form of bars
or rods. The strip trip insert part may have a ladder like
construction.
[0031] The shell forms an aerofoil and includes a divider wall
which divides the shell into a front cavity at a leading edge
region of the component and a rear cavity at a trailing edge region
of the component. The multi-part insert may be located in the front
cavity or a rear insert located in the rear cavity.
[0032] The divider wall may include apertures which provide fluid
communication between the front and rear cavity. At least one part
of the multiple insert parts may include a sealing plate to
restrict or prevent the flow of cooling air across the divider
wall. The sealing plate may be incorporated on the retaining part.
Alternatively or additionally, the sealing plate may be formed by
one or more insert parts.
[0033] In another aspect, the invention provides a method of
forming the gas turbine engine component according to any one of
the previous claims, the method including the steps of: providing
the shell; providing a plurality of insert parts which are
configured to be assembled in an abutting relation with one another
within the cavity to provide the multi-part insert; wherein the
assembled insert includes at least one retention part which engages
with a wall of the cavity and at least one other insert part so as
to retain the assembled insert within the cavity.
[0034] Other preferred features include, a gas turbine engine
component having a shell and an insert located inside the shell,
the insert forming one or more flow channels which, in use, receive
a flow of cooling air; wherein the insert is made by additive layer
manufacturing; and wherein the insert includes formations which
support the insert within the shell and guide the cooling air
around the inner surface of the shell.
[0035] The formations formed as part of an insert made by additive
layer manufacturing (ALM) may be intricate features which cannot be
formed as part of the shell and which cannot be formed with a high
level of dimensional accuracy on inserts produced from sheet metal.
By providing ALM inserts with supporting formations, the cooling
properties of the component can be greatly improved.
[0036] The use of ALM for the production of metal inserts can also
be advantageous in that the walls of the insert including the
impingement holes can be manufactured in one procedure without
requiring a separate tooling step to manufacture the holes.
Further, the inserts can be readily modified without a need for
expensive re-tooling and the time taken to manufacture inserts can
be reduced. Where the insert is metallic, the ALM process can be
direct laser deposition (DLD) (also known as direct metal
deposition (DMD)).
[0037] The shell may be a ceramic matrix composite shell. By
providing ALM inserts with supporting formations, the cooling
properties of a component having a CMC shell can be greatly
improved. More generally, it is possible to add to the benefits
provided by a CMC shell, such as its thermal properties, by
providing detailed structures that cannot be manufactured as part
of the CMC shell.
[0038] Alternatively, the shell may be a metal shell, such as
single crystal nickel alloy shell. The formations may include fins.
These fins may extend to the inner surface of the shell to support
the insert within the shell. The fins may be pin-fins which
advantageously enhance the heat transfer level by increasing the
turbulence of the cooling air flow and providing mixing of the
cooling air. The insert may also include impingement holes for
jetting cooling air from one or more flow channels onto the inner
surface of the shell.
[0039] The formations may form one or more chambers between the
insert and the inner surface of the shell, the or each chamber
being configured so that, in use, the chamber receives cooling air
from the one or more flow channels, the cooling air pressure being
lower in the chamber than in the flow channels. Each chamber can
contain cooling air at a different pressure. The or each chamber
can supply film cooling holes formed in the shell, the pressure of
cooling air at the film cooling holes being matched to the local
external pressure.
[0040] The insert may be tubular so that it forms a central flow
channel and fits inside the shell in a nested arrangement with
formations protruding outwardly from an outer wall of the insert
towards the inner surface of the shell. In this way, the chambers
can be located around the central flow channel. Another option is
for the insert to be a plate which extends from one part of the
inner surface of the shell to another part of the inner surface of
the shell to form a flow channel on at least one side of the
insert.
[0041] The insert may form a plurality of flow channels in fluid
communication with one another to define a multi-pass cooling
arrangement. In such a multi-pass cooling arrangement, the cooling
air can flow in opposite directions through successive channels.
Integral plates may be located at end walls of the component to
create suitable bend geometries between channels.
[0042] The insert may include trip strip formations which lie along
the inner surface of the shell. These formations can improve heat
transfer to the cooling air.
[0043] The gas turbine engine component may be an aerofoil. More
particularly, the gas turbine engine component may be a nozzle
guide vane (NGV) or a rotor blade. However, it is also possible
that the gas turbine engine component could be an NGV platform, a
shroud segment or a shroud liner.
[0044] When the component is an aerofoil, the shell can include a
divider wall which divides the shell into a front cavity at a
leading edge region of the component and a rear cavity at a
trailing edge region of the component. The divider wall can help to
prevent the aerofoil structure from rupturing under pressure loads
and also can help to prevent unwanted ballooning of the aerofoil
shape. The insert may be a front insert located in the front cavity
or a rear insert located in the rear cavity. Indeed, the aerofoil
may include respective inserts in both the front cavity and the
rear cavity.
[0045] When the insert is a rear insert, one or more chambers
defined by the insert can supply cooling air to trailing edge
discharge holes or slots, with the holes or slots receiving cooling
air at a pressure matched to the local external pressure.
[0046] The insert may include a sealing plate to prevent the flow
of cooling air across the divider wall. Such a sealing plate can
allow the divider wall to be discontinuous. In preventing the flow
of cooling fluid across the divider wall, the sealing plate can
help to reduce thermal induced stresses associated with hot
external walls and a cold divider.
[0047] The insert may be a unitary body, or may be formed from two
or more separately insertable insert parts. Forming the insert from
a plurality of insert parts can allow the insert to be fitted into
a shell which has, for example, a re-entrant cavity or is otherwise
configured in such a way as to prevent a unitary body from being
inserted.
[0048] A gas turbine engine component may be provided having a
shell and an insert located inside the shell, the insert may
include: a first wall containing first impingement holes which, in
use, jet cooling air onto a first region of the inner surface of
the shell; a second wall containing second impingement holes which,
in use, jet cooling air onto a second region of the inner surface
of the shell; and a fluid pathway formed between the two walls, the
pathway recycling the cooling air jetted onto the first region to
the inlets of the second impingement holes for jetting onto the
second region.
[0049] Advantageously, the insert allows jetted cooling air to be
used twice. In this way, film cooling effectiveness and film
coverage can be increased for a given quantity of cooling air mass
flow.
[0050] The shell may be a ceramic matrix composite shell.
Alternatively, the shell may be a metal shell, such as single
crystal nickel alloy shell.
[0051] The insert may be made by additive layer manufacturing (ALM)
or by casting. Where the insert is metallic, the ALM process can be
direct laser deposition (DLD) (also known as direct metal
deposition (DMD)). An insert made by ALM or casting can be produced
with a high level of intricacy and with high speed and
repeatability. For example, ALM facilitates the production of
features such as thin walls and internal cooling holes, as well as
internal heat transfer augmentation features like trip-strips,
pedestals, pin-fins etc.
[0052] The insert may include heat transfer formations at the first
and second regions which support the insert within the shell and
which guide the cooling air around the inner surface of the shell.
In this way, the cooling air can remove more heat from the walls of
the shell. In addition, as the insert supports itself, there may be
no need for extra support structures which can add to manufacturing
time and cost.
[0053] The geometry of the heat transfer formations at the first
region in particular may be chosen to restrict the flow rate of the
cooling air and to increase the pressure drop through the pathway.
The heat transfer formations may be pedestals or pin-fins, in which
case the flow rate of the cooling air may be controlled by the
number of pedestals/pin-fins, their density and their diameter.
Additionally or alternatively, the number of the impingement holes
and/or the diameter of the impingement holes can be used to control
the flow rate of the cooling air.
[0054] The shell may include exterior film cooling holes fed by
cooling air that has been jetted onto the second region of the
inner surface. This further recycling of the cooling air helps to
make even more effective use of the air.
[0055] The insert may include trip strip formations which lie along
the inner surface of the shell.
[0056] The gas turbine engine component may be an aerofoil. More
particularly, the gas turbine engine component may be a nozzle
guide vane (NGV) or a rotor blade. However, it is also possible
that the gas turbine engine component can be an NGV platform, a
shroud segment or a shroud liner.
[0057] Where the component is an aerofoil, the first and second
regions may be located at the suction side of the aerofoil.
[0058] The shell of the aerofoil may include a divider wall which
divides the shell into a front cavity at a leading edge region of
the aerofoil and a rear cavity at a trailing edge region of the
aerofoil. The insert can then be a front insert located in the
front cavity, or a rear insert located in the rear cavity. Indeed,
the aerofoil may have respective inserts in both the front cavity
and the rear cavity. The divider wall can help to prevent the
aerofoil structure from rupturing under pressure loads and also
helps to prevent unwanted ballooning of the aerofoil shape. The
insert may include a sealing plate to prevent a flow of cooling air
across the divider wall. In preventing such a flow, the sealing
plate can reduce thermal induced stresses associated with hot
external walls and a cold divider.
[0059] Where the insert of the aerofoil is a front insert, the
pathway may guide the recycled cooling air in an upstream direction
towards the leading edge. In this way, for the front cavity, the
first region of the inner surface of the shell may be located
further away from the leading edge of the aerofoil and the second
region of the inner surface of the shell may be located closer to
the leading edge. Any exterior film cooling holes fed by cooling
air that has been jetted onto the second region may therefore lie
at a position close to the leading edge, and can contribute to a
cooling film on the suction side of the aerofoil.
[0060] Where the insert is a rear aerofoil insert, the pathway may
guide the recycled cooling air in a downstream direction towards
the trailing edge. In this way, for the rear cavity, the first
region of the inner surface of the shell may be located further
away from the trailing edge of the aerofoil and the second region
of the inner surface of the shell may be located closer to the
trailing edge.
[0061] The insert may also include a bank of further heat transfer
formations, such as pedestals or pin fins, along the inner surface
of the shell to guide the cooling air along the inner surface of
the shell after it has been jetted onto the second region. In
respect of a rear aerofoil insert, the bank of further heat
transfer formations preferably guides the recycled cooling air in a
downstream direction towards the trailing edge to feed exit holes
or slots at the trailing edge.
[0062] The aerofoil insert may define one or more flow channels
which, in use, collect cooling air from one or both ends of the
aerofoil and distribute the cooling air through the shell, at least
a portion of the cooling air being distributed to the inlets of the
first impingement holes for jetting onto the first region.
[0063] The insert may be a unitary body, or may be formed from two
or more separately insertable insert parts. Forming the insert from
a plurality of insert parts can allow the insert to be fitted into
a shell which has, for example, a re-entrant cavity or is otherwise
configured in such a way as to prevent a unitary body from being
inserted.
[0064] Further optional features of the invention are set out
below.
BRIEF DESCRIPTION OF THE DRAWINGS
[0065] Embodiments of the invention will now be described by way of
example with reference to the accompanying drawings in which:
[0066] FIG. 1 shows an isometric view of a conventional HP stage
cooled turbine;
[0067] FIG. 2 shows a longitudinal cross-section through a ducted
fan gas turbine engine;
[0068] FIGS. 3(a) and (b) show cross sectional views of (a) a
ceramic matrix composite shell of a nozzle guide vane found in the
circled region labelled R in FIG. 2 and (b) front and rear inserts
to the shell;
[0069] FIG. 4 shows a cross-sectional view of the nozzle guide vane
of FIG. 3 with the inserts fitted inside the shell and cooling
flows indicated by arrows;
[0070] FIG. 5 shows a cross-sectional view of a second nozzle guide
vane;
[0071] FIG. 6 shows a cross-sectional view of the nozzle guide vane
of FIG. 5 with cooling flows indicated by arrows;
[0072] FIG. 7 shows a cross-sectional view of variant inserts for
the nozzle guide vane of FIGS. 5 and 6; and
[0073] FIGS. 8(a) and (b) show cross-sectional views of (a) a
ceramic matrix composite shell of a nozzle guide vane found in the
circled region labelled R in FIG. 2, and (b) front and rear inserts
to the shell;
[0074] FIG. 9 shows a cross-sectional view of the aerofoil of FIG.
8 with the inserts fitted inside the shell; and
[0075] FIGS. 10a to 12b show various aerofoil embodiments having
multi-part inserts according to the invention.
DETAILED DESCRIPTION AND FURTHER OPTIONAL FEATURES OF THE
INVENTION
[0076] With reference to FIG. 2, a ducted fan gas turbine engine
incorporating the invention is generally indicated at 10 and has a
principal and rotational axis X-X. The engine comprises, in axial
flow series, an air intake 11, a propulsive fan 12, an intermediate
pressure compressor 13, a high-pressure compressor 14, combustion
equipment 15, a high-pressure turbine 16, and intermediate pressure
turbine 17, a low-pressure turbine 18 and a core engine exhaust
nozzle 19. A nacelle 21 generally surrounds the engine 10 and
defines the intake 11, a bypass duct 22 and a bypass exhaust nozzle
23.
[0077] During operation, air entering the intake 11 is accelerated
by the fan 12 to produce two air flows: a first air flow A into the
intermediate pressure compressor 13 and a second air flow B which
passes through the bypass duct 22 to provide propulsive thrust. The
intermediate pressure compressor 13 compresses the air flow A
directed into it before delivering that air to the high pressure
compressor 14 where further compression takes place.
[0078] The compressed air exhausted from the high-pressure
compressor 14 is directed into the combustion equipment 15 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 16, 17, 18 before
being exhausted through the nozzle 19 to provide additional
propulsive thrust. The high, intermediate and low-pressure turbines
respectively drive the high and intermediate pressure compressors
14, 13 and the fan 12 by suitable interconnecting shafts.
[0079] A first example of a component having an insert will be
described with reference to FIGS. 3 and 4. FIG. 3 shows cross
sectional views of (a) a ceramic matrix composite (CMC) shell of a
gas turbine engine component in the form of a nozzle guide vane
(NGV) as found in the circled region labelled R in FIG. 2, and (b)
front and rear inserts to the shell. FIG. 4 shows a cross-sectional
view of the aerofoil of FIG. 3 with the inserts fitted inside the
shell and arrows indicating cooling air flows.
[0080] The NGV shell includes a divider wall 203 which divides the
shell into a front cavity 201 at a leading edge region of the
aerofoil and a rear cavity 202 at a trailing edge region of the
aerofoil. A front insert 210 made by direct laser deposition (OLD)
(a form of additive layer manufacturing) is located inside the
front cavity 210 and a rear insert 220 also made by DLD is located
inside the rear cavity 202.
[0081] The CMC shell includes film cooling holes 206 located at a
region of the suction side of the aerofoil closest to the leading
edge. Film cooling holes 206 are also located along the pressure
side of the aerofoil. A cooling flow outlet 207 is located at the
trailing edge of the CMC shell, in fluid communication with the
rear cavity, and may take the form of exit holes or slots.
[0082] Each of the DLD inserts 210, 220 of FIGS. 3 and 4 has a
tubular shape similar to the shape of the front and rear cavities
so that the front insert 210 is located inside the front cavity 201
in a nested arrangement, and the rear insert 220 is located inside
the rear cavity 202 in a nested arrangement. Each tubular insert
defines a central flow channel 211, 212, and cooling air is bled
out from each central flow channel to the inner surface of the
shell via impingement holes 216 formed in the walls of the
insert.
[0083] Each DLD insert 210, 220 includes formations 218, 219 which
extend outwards from an outer surface of the insert to an inner
surface of the shell to support the insert within the shell and
guide cooling air around the inner surface of the shell. The
formations include pin-fin formations 218 and chamber-forming
formations 219.
[0084] The rear insert 220 includes a sealing plate 270 located
along the divider wall 203 of the shell to help prevent the flow of
cooling air across the divider wall 203.
[0085] The flow of cooling air will now be described with reference
to FIG. 4. Large shaded arrows depict the flow of cooling air into
the aerofoil, inboard 311 and outboard 312 flows entering the front
cavity 201, and a single inboard flow of cooling air 313 entering
the rear cavity 202. Where the flow is a dual feed (an inboard and
outboard flow), the insert 210 preferably includes a baffle plate
(not shown). The baffle plate reduces differential pressures caused
by the dual feed, thereby reducing unwanted `blow through` effects.
The baffle plate can be formed as an integral part of the insert
210, which advantageously reduces the part count and cost, and
improves reliability.
[0086] The chamber-forming formations 219 form a plurality of
chambers 229 between each insert and the inner surface of the shell
200. Each chamber 229 is configured to receive cooling air from a
flow channel 211, 212 via impingement holes 216, the pressure of
the cooling air being lower in the chambers than in the flow
channel. Cooling air from the chambers 229 is used to supply film
cooling holes 206. The formations 219 of the front insert of the
aerofoil shown in FIG. 4 form four chambers between the insert 210
and the inner surface of the shell 200. A first chamber supplies
cooling air to film cooling holes 206 on the suction side, a second
chamber supplies cooling air to showerhead cooling holes 206 at the
leading edge region of the pressure side, and third and fourth
chambers supply cooling air to film cooling holes on the pressure
side further away from the leading edge.
[0087] The number of impingement holes 216 supplying a given
chamber and the number of film cooling holes 206 fed by that
chamber are selected so that each chamber is maintained at a
different pressure. Cooling air can therefore be supplied to the
film cooling holes 206 and the film cooling outlet 207 at pressures
which match the local external pressure. The front flow channel 211
has an internal pressure level which is controlled to ensure
adequate blowing rates through these film cooling holes, while
maintaining a safe backflow pressure margin to prevent hot gas
ingestion throughout the flight cycle.
[0088] A second example of a component having an insert will be
described with reference to FIGS. 5, 6 and 7. FIG. 5 shows a nozzle
guide vane 300 according to the second example, FIG. 6 shows the
cross-sectional view of FIG. 5 with arrows indicating cooling air
flows, and FIG. 7 shows a cross-sectional view of variant inserts
for the nozzle guide vane of FIGS. 5 and 6. The NGV has a CMC shell
400, including a divider wall 403 which divides the shell into a
front cavity at a leading edge region of the aerofoil and a rear
cavity at a trailing edge region of the aerofoil. A front insert
410 made by DLD is located inside the front cavity and a rear
insert 420, also made by DLD, is located inside the rear cavity.
Each insert 410, 420 includes a sealing plate 470 to prevent the
flow of cooling air across the divider wall 403.
[0089] The front insert 410 of the aerofoil shown in FIGS. 5 and 6
has formations, including a plate end 419 and pin-fins 418, which
extend from an insert plate 440 to the inner surface of the shell
to support the insert within the front cavity of the shell, and
thereby define a flow channel 411 at the pressure side of the front
cavity between the front insert and the inner surface of the shell,
and a chamber at the suction side of the front cavity between the
front insert and the inner surface of the shell. The chamber on the
suction side receives cooling air from the flow channel 411 via
impingement holes 416.
[0090] The rear insert 420 of the aerofoil shown in FIGS. 5 and 6
has formations in the form of sealing walls 475 which extend
outwardly from a central insert plate 430, to the inner surface of
the shell. There are four sealing walls 475 which, in addition to
the sealing plate 470, define a plurality of flow channels 441,
442, 443 in fluid communication with one another to form a
multi-pass cooling arrangement.
[0091] In FIG. 6, large straight arrows 512, 513 and 514 depict
flows of cooling air into the aerofoil 400. The multi-pass cooling
arrangement includes, in flow series, a pair of parallel first pass
chambers 441 (one on the pressure side and one on the suction side)
corresponding to a first pass flow channel, a pair of parallel
second pass chambers 442 (one on the pressure side and one on the
suction side) corresponding to a second pass flow channel and a
common third pass chamber 443 corresponding to a third pass flow
channel. The third chamber is located at a trailing edge region of
the rear cavity and feeds trailing edge discharge holes or slots
407. The first pass chamber 441 on the pressure side supplies film
cooling holes 406 on the pressure side of the NGV. Similarly, the
second pass chamber 442 on the pressure side supplies film cooling
holes 406 on the pressure side of the NGV.
[0092] Integral plates at end walls (not shown) create suitable
bend geometries to guide cooling air from the first pass chambers
441 to the second pass chambers 442 and from the second pass
chambers to the third chamber 443 in order that the chambers
operate as the rearward flowing, 3-pass cooling arrangement shown
by the curved arrows.
[0093] FIG. 7 shows variant front and rear inserts similar to those
of FIGS. 5 and 6 but having additional trip strip formations 460
which lie along the inner surface of the shell. The trip strip
formations are ladder-like in construction having a pair elongate
parallel rails which provide support for a linear array of equally
spaced trip strips or bars which run therebetween. The trip strips
are set at a compound angle to the rails, Trip strips are known in
the art and can locally enhance heat transfer to the cooling
air.
[0094] Although not shown in the above Figures, formations defining
a contra-flow cooling system can be incorporated into an insert, as
an alternative or an addition to the cooling structures described
above.
[0095] Any holes 216, 416 in the insert can be formed during the
DLD process so there is no need for subsequent machining of the
inserts.
[0096] In addition, the DLD process facilitates modification and
development of the insert design during the manufacturing process
as no tooling changes are required. For example, features such as
formations 218, 219, 418, 419, 460, 475 may be altered slightly
between the manufacture of different aerofoils 100, 400 of a single
engine 10 depending on the position of the respective aerofoils
within the engine to give a relative increase or decrease in the
cooling mass flow of the aerofoil.
[0097] FIG. 8 shows cross-sectional views of (a) a ceramic matrix
composite (CMC) shell 800 of a gas turbine engine component 100 in
the form of a nozzle guide vane (NGV) found in the circled region
labelled R in FIG. 2, and (b) front 210 and rear 820 inserts to the
shell.
[0098] FIG. 9 shows a cross-sectional view of the aerofoil of FIG.
8 with the inserts fitted inside the shell and arrows indicating
cooling air flows.
[0099] The shell 800 includes a divider wall 803 which divides the
shell into a front cavity 801 at a leading edge region of the
aerofoil and a rear cavity 802 at a trailing edge region of the
aerofoil. The front insert 810 is located inside the front cavity
801 and the rear insert 820 is located inside the rear cavity
802.
[0100] The CMC shell 800 includes exterior film cooling holes 806
located at the region of the suction side of the aerofoil closest
to the leading edge. More exterior film cooling holes 806 are
located along the pressure side of the aerofoil. The CMC shell 800
also includes exit holes or slots 807 at its trailing edge.
[0101] Each of the front and rear inserts includes a first wall
811, 821 having first impingement holes 813, 823 formed therein and
a second wall 812, 822 having second impingement holes 814, 824
formed therein. For each insert, a fluid pathway 815, 825 is formed
between the first wall 811, 821 and the second wall 812, 822.
[0102] The first impingement holes 813, 823 lie opposite a first
region 833, 843 of the inner surface of the shell and the second
impingement holes 814, 824 lie opposite a second region 834, 844 of
the inner surface of the shell. The first and second regions of the
aerofoil of FIGS. 3 and 4 are both located at the suction side of
the aerofoil. For each insert, the fluid pathway is formed between
the first region 833, 843 and the inlets of the second impingement
holes 814, 824 to recycle cooling air which has been jetted onto
the first region for jetting onto the second region.
[0103] The fluid pathway 815 of the front insert guides recycled
cooling air in an upstream direction towards the leading edge so
that, for the front cavity, the first region 833 is located further
away from the leading edge of the aerofoil and the second region
834 is located closer to the leading edge of the aerofoil. The
fluid pathway 825 of the rear insert guides recycled cooling air in
a downstream direction so that, for the rear cavity, the first
region 843 is located furthest away from the trailing edge of the
aerofoil and the second region 844 is located closest to the
trailing edge 807 of the aerofoil.
[0104] Heat transfer formations 853 are located at the first region
833, 843 and the second region 834, 844. The heat transfer
formations shown in FIGS. 3(b) and 4 are pin-fins.
[0105] In addition to the first wall 821 and second wall 822, the
rear insert 820 shown in FIGS. 3(b) and 4 includes a bank of
pin-fins 863 which extend along the inside surface of the shell
from the second region to the trailing edge. The rear insert also
defines a plurality of chambers 881, 882 at the pressure side of
the rear cavity. The chambers are interconnected via internal
passageways 829 so that they are in fluid communication with each
other. Two chambers 881, 882 are shown in the rear insert of the
aerofoil of FIGS. 3(b) and 4.
[0106] Each insert 810, 820 includes a sealing plate 870 which lies
along the divider wall 803 of the CMC shell 800 to prevent a flow
of cold air across the divider wall. The rear insert 820 also
includes trip strip formations 816 which lie along the inner
surface of the shell at the pressure side of the cavity to improve
heat transfer to the cooling air at this location.
[0107] The flow of cooling air will now be described with reference
to FIG. 9. Large shaded arrows depict the flow of cooling air into
the aerofoil: inboard 911 and outboard 912 flows entering the front
cavity 801, and a single inboard flow of cooling air 913 entering
the rear cavity 802. Where the flow is a dual feed (an inboard and
an outboard flow), the insert preferably includes a baffle plate
(not shown). The baffle plate reduces differential pressures caused
by the dual feed, therefore reducing unwanted `blow through`
effects. The baffle plate can be formed as an integral part of the
insert which advantageously reduces the part count and cost and
improves reliability.
[0108] In the front cavity 801, the first wall 811 defines a front
flow channel 860 at the pressure side of the cavity. Cooling air is
distributed from this front flow channel to the inlets of the first
impingement holes 813 for jetting onto the first region 833. The
front flow channel also supplies cooling air at a high pressure to
film cooling holes 806 on the pressure side in the form of a
leading edge showerhead cooling head arrangement. The front flow
channel has an internal pressure level which is controlled to
ensure adequate blowing rates through these cooling holes, while
maintaining a safe backflow pressure margin to prevent hot gas
ingestion throughout the flight cycle. Cooling air which has been
recycled and jetted onto the second region 834 will have a reduced
pressure compared to the cooling air supplied directly by the front
flow channel and can therefore be used to feed exterior film
cooling holes 806 on the suction side.
[0109] In the rear cavity 802, the plurality of chambers 881, 882
on the pressure side form a plurality of rear flow channels.
Cooling air enters the first chamber 881 and is distributed
therefrom to the inlets of the first impingement holes 823 for
jetting onto the first region 843. This first chamber also supplies
cooling at a high pressure to exterior film cooling holes 806 on
the pressure side of the aerofoil, as well as supplying cooling air
to the second chamber 882 via internal passageways 829. The second
chamber supplies cooling air to the bed of pin-fins 863 as well as
to further exterior film cooling holes 806 on the pressure side.
Both chambers have internal pressure levels which are controlled to
ensure adequate blowing rates through their cooling holes, while
maintaining a safe backflow pressure margin to prevent hot gas
ingestion throughout the flight cycle.
[0110] The CMC shell may be SiC--SiC and a protective coating may
be applied to the outside and/or inside surfaces of the shell 800
to prevent environmental attack. The inserts 810, 820 may be cast
(e.g. using the lost wax process) and then machined (e.g. for hole
drilling), or may be made using additive layer manufacturing such
as direct laser deposition (also known as direct metal deposition).
Additive layer manufacturing, and particularly direct laser
deposition, enables all of the detailed features of the inserts to
be manufactured in one procedure, including the impingement holes
813, 814, 823, 824. Further, it allows cooling schemes to be easily
changed, without the need for re-tooling.
[0111] The gas turbine component of the present invention can be an
NGV aerofoil, as described in detail in above, but can be any other
gas turbine aerofoil, including a rotor blade. The gas turbine
component may alternatively be an NGV platform, a shroud segment,
or a shroud liner.
[0112] The inserts described above can be used instead of, or in
combination with, sheet metal inserts.
[0113] Instead of forming each insert as a unitary body, as shown
in FIGS. 3 to 9, another option is to form the inserts from two or
more insert parts. This allows the inserts to be fitted into
cavities where a receiving portion in which part of the insert
would ideally be located is obstructed in some way such that a
complete insert cannot be directly inserted. The obstruction in
question may be provided by a wall of the cavity or by a
protuberant feature which extends from one or between two walls of
the cavity. An obstructed portion may be as viewed from outside the
shell through an insertion aperture, or by a part of the insert
having to enter the cavity along a first trajectory before being
located in a receiving portion along a second trajectory which is
different to the first trajectory. For example, an elongate insert
part having a longitudinal axis may be inserted into the cavity
with an axially extending trajectory, before being pushed laterally
into a recess or an otherwise obscured portion of the cavity.
[0114] FIGS. 10a, and 10b show a perspective view of an aerofoil
having a front insert 1010 which is a variant of the front insert
of FIGS. 5 and 6, and a rear insert 1020 which is a variant of the
rear insert of FIG. 7, the CMC shell 1000 being drawn as a
transparent body.
[0115] Thus, in FIGS. 10a and 10b there is shown an aerofoil in the
form of a vane similar to the NGV shown in FIG. 1. The aerofoil
includes an elongate shell 1000 having internal front 1001 and rear
1003 cavities. The outer surface of the shell has a predetermined
aerodynamic shape suitable for use as an NGV. As such, the aerofoil
is distorted from a straight radially extending form and includes a
chordal twist along its length. This distortion can be best seen in
FIG. 10b where the first end 1000a and second end 1000b of the
aerofoil are angularly offset from each other when viewed
approximately along the longitudinal axis of the aerofoil 1000.
This means that the front 1001 and rear cavities which extends
along the radial axis of the interior of the aerofoil 1000 have an
irregular shape with obstructed portions when viewed from the first
end along the longitudinal axis of the shell 1000.
[0116] It will be appreciated that the distortion of the cavities
is also affected by the internal profile of the shell walls which
may be varied but will typically be determined by the weight and
mechanical and thermal requirements of the aerofoil rather than the
fit of an insert. In the described example, the walls of the shell
have substantially uniform thickness.
[0117] The front cavity 1001 has a multi-part insert 1010 located
therein, which, in the described example, is made up from two
separate insert parts 1010a,b assembled in an abutting relation to
one another so as to provide the multi-part insert 1010. The rear
cavity 1003 also includes a multi-part insert 1020 having multiple
separate insert parts 1020a-f. The rear cavity insert 1020 is made
up from two main body parts 1020a,b and several trip-strip insert
parts 1020c-f which abut and engage the main body portions 1020b of
the rear insert 1020, and also the wall of the shell 1000. Thus,
the front insert 1010 is a multi-part insert formed from two insert
parts 1010a,b and the rear insert 1020 is formed from six insert
parts 1020a-f. In each cavity, the last insert park to be installed
locks the completed insert in place and ensures a tight fit between
the insert and the shell 1000 while accommodating manufacturing
tolerances.
[0118] To construct the vane with the assembled inserts 1010, 1020,
the insert parts 1010a,b, 1020a-f, are placed within the respective
cavities via an insertion aperture 1050. The insertion aperture
1050 may be any suitable entrance to the cavity and may be covered
and optionally sealed after the inserts 1010, 1020 have been
correctly located within the shell 1000. In the described example,
the insertion aperture 1050 is provided by the open end of the
aerofoil and is as large as can be accommodated by the walls of the
shell 1000. It will be appreciated that some constructions of the
component, particularly one which is cast for example, may only
include a partial opening in the end of the aerofoil. Further, the
insertion aperture may be defined by the walls of the shell, or a
particular portion or zone of a larger opening.
[0119] Although the insertion aperture 1050 of the rear cavity 1020
is as large as can be accommodated, the irregular shape of the rear
cavity 1020 means that the insertion of the assembled or unitary
insert 1010, 1020 into the cavity 1020 would not be possible. This
is because an insert which is shaped to match and abut the internal
walls of the cavity may be too large in parts to fit through the
insertion aperture 1050. Alternatively, the curvature or twist of
the insert may prevent it from being inserted along the length of
the cavity. Further, there may be features or recesses within the
cavity which the insert must either go around or be placed within
when being inserted. Thus, although the use of prior art inserts
has provided some benefits, applications have been limited due to
the restrictions placed on the inserts.
[0120] Providing a multi-part insert allows a first insert part to
be loaded into the cavity via an insertion aperture and
subsequently located into a receiving portion of the cavity.
Thereafter, the second insert part, or retaining part, is passed
into the cavity and engaged with the first insert part in an
abutting manner. The retaining part may provide a biasing force
which acts to urge the first insert part against a wall of the
cavity so as to retain it there, or may be manufactured to have an
interference fit with the first insert part so as to provide chock.
Thus, there is provided an assembled insert within the cavity which
cannot be withdrawn from the insertion aperture (or inserted if
assembled outside of the shell), but which can be located against
the wall of the shell.
[0121] In some embodiments, the resilient part may be the first or
an intermediate part loaded into the cavity. In this instance, the
loading of the resilient part will occur upon insertion of the last
part which will act to put the resilient part in a stressed
condition.
[0122] In the described example of FIGS. 10a and 10b, a receiving
portion 1060 can be taken to the rearmost portion of the rear
cavity 1003 in which the first insert part 1010a is located. The
insertion aperture 1050 can be taken to be at the first end 1000b
of the aerofoil toward the divider wall 1004. Thus, the first
insert part 1020a is inserted into the rear cavity 1003 through the
insertion aperture 1050 which is located at the wider end of the
open ended aerofoil towards the divider wall 1004 and with a
trajectory which is coincidental with the plane of the divider wall
1004. Once in place, the first insert part 1020a can be moved
toward the rear of the cavity until the distal ends of partitioning
walls 1021 abut the walls of the cavity. It will be appreciated
that the trip strip formations 1020c and 1020d can be mated to the
first insert part before or after the insertion depending on the
particular design, but it is envisaged that they are mated to the
first main body insert part 1020a prior to being loaded into the
rear cavity 1003. Next, the second main body insert part 1020b and
third trip strip 1020e can be placed within the rear cavity 1003
via the insertion aperture 1050 and pushed home to provide a chock
for retaining the first insert part 1020a in place. The final
insert part is the fourth trip strip 1020f formation which is slid
between a free end of a web of the first main insert part 1020a,
and a shoulder 1005 which protrudes into the rear cavity along the
length of the divider wall 1004 where the divider wall meets the
shell wall.
[0123] It will be noted from FIG. 10b, that the shape of the rear
cavity 1003 would prevent the insertion of the assembled insert
1020 into the cavity from the open end of the vane due to the
variance in amount of the chordal twist required between the front
and rear parts of the assembled insert.
[0124] The two insert parts 1010a,b of the front cavity 1001
include a curved member 1010a which sealably contacts the interior
of the leading edge of the aerofoil and extends around the suction
side toward the divider wall 1004. The second insert part 1010b is
in the form of a sealing plate 1014 which sealably abuts the
divider wall 1004. The sealing plate 1014 includes a short wall
along its length which includes a rebate for receiving the
corresponding free end of the first insert part 1010a.
[0125] The first insert part 1010a is made to be slightly flatter
than required when in situ such that the free end is closer to the
divider wall 1004 and inserting the second insert part 1010b urges
the first part 1010a towards the leading edge so as to provide the
biasing force for retaining the assembled insert 1010 in place.
[0126] In order to provide a correct fit, the insert parts 1010a,b
are arranged to be held in an abutting relation with a resilient
bias provided by one of the insert parts. The resilient bias in the
case of the front cavity is provided by the fore insert part 1010a
which is inserted after the sealing plate which is described above.
The fore insert part may be oversized slightly with respect to the
space in which it is designed to accommodate such that it must
elastically deform during insertion.
[0127] The elastic deformation is such that the part is
sufficiently stressed so as to provide the resilient bias between a
wall of the cavity and sealing plate. Alternatively, the insert
part may be made so as to be partly collapsible or compressible so
that the shape of the part is altered to allow it to be inserted.
In order to provide the collapsibility and compressibility, the
insert part may be made to size for the cavity before being
plastically deformed prior to insertion of the part.
[0128] The insert parts can incorporate rebates or other features
to allow them to be secured in an abutting relation and to provide
opposing surfaces for the retention of the parts via the resilient
bias. Hence, as seen in FIG. 10b, the sealing plate insert part
1010b in the front cavity 1001 and the free ends of the first and
second main body parts in the rear cavity 1003 include rebates for
receiving corresponding parts of abutting insert parts. Further,
the rails of trip-strip insert parts 1020c-f include protuberant
lips which engage with corresponding rebates in the main body
portions.
[0129] It will be noted that the shell is constructed from a CMC
material and as such has smooth outer and inner walls, principally
due to the difficulties of forming discrete features in a CMC
material. However, this may not always be the case, and the inserts
are applicable to other non-CMC constructed shells.
[0130] FIGS. 11a and 11b provide another example in which the rear
insert 1120 comprises three insert parts 1120 a-c. The first insert
part 1120b is V-shaped part having two plate-like members 1121a and
1121b which are joined at a hinged portion 1121c. The free ends of
the members 1121a,b (or arms) are tapered from the first end to the
second end so as to provide a smaller sectional area at the first
end so that it can be manoeuvred more readily into the insertion
aperture 1150, and to provide a generally wedge shaped insert part.
The second 1120b and third 1120c insert parts join along a mid-line
of the sealing plate and form a wedge shaped part in unison which
provides a chock for the first part 1120a when the insert parts
1120 are assembled into a complete insert. It will be appreciated
that the second and third parts are inserted from the opposite end
of the cavity through a second insertion aperture 1152.
[0131] The V-shaped first insert part 1120a is fabricated such that
the angle between the arms is greater than angle between the
corresponding portions of the rear cavity. Thus, to insert the
part, the arms are forceably moved together so as to elastically
stress the hinge portion as it is passed through the insertion
aperture. Once inside the cavity, the insert part can be pushed
into the receiving portion 1160 with the resilient bias of the arms
retaining the part in place.
[0132] The front cavity multi-part insert 1010 includes three parts
1010a-c. Here, the first insert part 1010a extends from the divider
wall 1004 toward the leading edge against the pressure surface of
the front cavity 1001. The second part 1010b abuts the free end of
the first insert part 1010a which is local to the leading edge and
extends around the suction surface toward the suction surface. The
third insert 1010c is generally L shaped with rebates provided on
the free ends of long and short members. The rebates provide a
flange which resides on the inside of the free ends of the
corresponding ends of the first and second insert parts. The arms
are joined at a hinge portion.
[0133] The first 1010a and second 1010b insert parts are made to
fit in a neutral or stress-free state within the front cavity 1001
whilst abutting the walls of the shell 1000. The third L-shaped
insert part is fabricated to have a larger angle than required such
that the hinge portion elastically deformed upon insertion so as to
provide a restoring force to bias against the free ends of the
first and second insert parts against the wall of the cavity via
the rebated portions.
[0134] A further example is shown in FIGS. 12a and 12b which
corresponds to the component described in FIGS. 8 and 9 above, but
with multiple insert parts in the front 1201 and rear cavities
1203. Hence, the front 1210 and rear 1220 inserts each include two
insert parts 1210a,b, 1220a,b, having similar features to those
described above in relation to FIGS. 10a to 11b. In this instance,
the front cavity 1201 has a first insert part 1210a which is
inserted first and provides the resilient bias once the sealing
part is inserted. The rear cavity 1203 has a first insert part
1220a which is inserted into the rear cavity via the insertion
aperture 1250 along a first trajectory before being pushed rearward
into the trailing edge which it is located in its corresponding
receiving portion 1260. The second insert 1220b provides the
sealing plate and a portion of wall which defines a cooling chamber
with the cavity wall. The wall is connected to the sealing plate
via a hinge portion which provides the resilient bias for retaining
the first insert part in place.
[0135] In addition to the above, it is possible in some embodiments
that multiple insert parts can be fitted inside one another so that
a single shell cavity includes an insert formed from two or more
nested insert parts. Each insert shown in FIGS. 3 and 4 seals its
cavity, as well as providing formations to support the insert and
guide cooling air around the inner surface of the shell. If two
nested insert parts are used in a cavity, the outer of the two
insert parts can provide the formations, and the inner of the two
insert parts can be configured to balloon under the pressure of the
inboard and/or outboard flows of cooling air to provide a sealing
load.
[0136] While the invention has been described in conjunction with
the exemplary embodiments described above, many equivalent
modifications and variations will be apparent to those skilled in
the art when given this disclosure. For example, the shell may be a
metal shell rather than a CMC shell. Accordingly, the exemplary
embodiments of the invention set forth above are considered to be
illustrative and not limiting. Various changes to the described
embodiments may be made without departing from the spirit and scope
of the invention.
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