U.S. patent application number 13/408251 was filed with the patent office on 2012-10-04 for gas turbine engine component.
This patent application is currently assigned to ROLLS-ROYCE PLC. Invention is credited to Peter T. IRELAND, Anthony J. RAWLINSON, Ian TIBBOTT, Lynne H. TURNER.
Application Number | 20120251295 13/408251 |
Document ID | / |
Family ID | 44067422 |
Filed Date | 2012-10-04 |
United States Patent
Application |
20120251295 |
Kind Code |
A1 |
TURNER; Lynne H. ; et
al. |
October 4, 2012 |
GAS TURBINE ENGINE COMPONENT
Abstract
A component of a gas turbine engine is provided. The component
includes an external wall which, in use, is exposed on one surface
thereof to working gas flowing through the engine. The component
further includes effusion cooling holes formed in the external
wall. In use, cooling air blows through the cooling holes to form a
cooling film on the surface of the external wall exposed to the
working gas. The component further includes an air inlet
arrangement which receives the cooling air for distribution to the
cooling holes. The component further includes a plurality of
metering feeds and a plurality of supply plena. The metering feeds
meter the cooling air from the air inlet arrangement to respective
of the supply plena, which in turn supply the metered cooling air
to respective portions of the cooling holes.
Inventors: |
TURNER; Lynne H.; (Bristol,
GB) ; IRELAND; Peter T.; (Derby, GB) ;
TIBBOTT; Ian; (Lichfield, GB) ; RAWLINSON; Anthony
J.; (Derby, GB) |
Assignee: |
ROLLS-ROYCE PLC
LONDON
GB
|
Family ID: |
44067422 |
Appl. No.: |
13/408251 |
Filed: |
February 29, 2012 |
Current U.S.
Class: |
415/115 |
Current CPC
Class: |
F05D 2240/11 20130101;
F05D 2260/205 20130101; F23R 2900/03041 20130101; F05D 2260/202
20130101; F23R 3/002 20130101; F05D 2260/201 20130101; F01D 11/24
20130101 |
Class at
Publication: |
415/115 |
International
Class: |
F01D 5/18 20060101
F01D005/18 |
Foreign Application Data
Date |
Code |
Application Number |
Mar 28, 2011 |
GB |
1105105.9 |
Claims
1. A component of a gas turbine engine, the component including: an
external wall which, in use, is exposed on one surface thereof to
working gas flowing through the engine, effusion cooling holes
formed in the external wall, in use, cooling air blowing through
the cooling holes to form a cooling film on the surface of the
external wall exposed to the working gas, and an air inlet
arrangement which receives the cooling air for distribution to the
cooling holes; wherein the component further includes: a plurality
of metering feeds and a plurality of supply plena, the metering
feeds metering the cooling air from the air inlet arrangement to
respective of the supply plena, and the supply plena supplying the
metered cooling air to respective portions of the cooling
holes.
2. A component according to claim 1, wherein the supply plena are
partially defined by the other surface of the external wall.
3. A component according to claim 2, wherein the metering feeds are
configured to form impingements jets from the cooling air metered
therethrough, the impingements jets impinging on said other
surface.
4. A component according to claim 1, further including a plurality
of secondary plena which are partially defined by the other surface
of the external wall, and are each in flow series between a
respective supply plenum and its respective portion of the cooling
holes.
5. A component according to claim 4, further including a plurality
of jet-forming passages which deliver the cooling air from the
supply plena to the secondary plena, the jet-forming passages being
configured to form impingements jets from the cooling air delivered
therethrough, and the impingements jets impinging on said other
surface.
6. A component according to claim 1 which further includes an inlet
plenum in flow series between the air inlet arrangement and the
supply plena, the metering feeds feeding the cooling air from the
inlet plenum to respective of the supply plena.
7. A component according to claim 1, wherein, in use, the pressure
of the working gas to which the external wall is exposed varies
from a leading edge to a trailing edge of the external wall, the
supply plena, metering feeds and cooling holes being configured
such that the pressure of the cooling air blowing through each
cooling hole matches the local pressure of the working gas at that
cooling hole.
8. A component according to claim 7, wherein the working gas varies
from a higher pressure to a lower pressure from the leading edge to
the trailing edge of the external wall.
9. A component according to claim 1 which provides an endwall to
the working gas annulus of the engine, the external wall being the
endwall.
10. A component according to claim 9 which is a shroud segment.
11. A component according to claim 9 which is a turbine blade or a
vane, a platform of the blade or vane forming the endwall.
12. A component according to claim 1 which is a combustor, the
external wall at least partially defining a combustion chamber of
the combustor.
13. A gas turbine engine having one or more components according to
claim 1.
Description
[0001] The present invention relates to a cooled component for use
in gas turbine engines.
[0002] With reference to FIG. 1, a ducted fan gas turbine engine
generally indicated at 10 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.
[0003] The gas turbine engine 10 works in a conventional manner so
that 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 14 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.
[0004] 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.
[0005] The performance of gas turbine engines, whether measured in
terms of efficiency or specific output, is improved by increasing
the turbine gas temperature. It is therefore desirable to operate
the turbines at the highest possible temperatures. For any engine
cycle compression ratio or bypass ratio, increasing the turbine
entry gas temperature produces more specific thrust (e.g. engine
thrust per unit of air mass flow). However as turbine entry
temperatures increase, the life of an un-cooled turbine falls,
necessitating the development of better materials and the
introduction of internal air cooling.
[0006] In modern engines, the high-pressure turbine gas
temperatures are hotter than the melting point of the material of
the blades and vanes, necessitating internal air cooling of these
airfoil components. During its passage through the engine, 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 high-pressure
stage(s), through the intermediate-pressure and low-pressure
stages, and towards the exit nozzle.
[0007] FIG. 2 shows an isometric view of a typical single stage
cooled turbine. Cooling air flows are indicated by arrows.
[0008] Internal convection and external films are the prime methods
of cooling the gas path components--airfoils, platforms, shrouds
and shroud segments etc. High-pressure turbine nozzle guide vanes
31 (NGVs) consume the greatest amount of cooling air on high
temperature engines. High-pressure blades 32 typically use about
half of the NGV flow. The intermediate-pressure and low-pressure
stages downstream of the HP turbine use progressively less cooling
air.
[0009] The high-pressure turbine airfoils are cooled by using high
pressure 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 1000 K, while
gas temperatures can be in excess of 2100 K.
[0010] 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. Therefore, as extracting coolant flow has an adverse
effect on the engine operating efficiency, it is important to use
the cooling air effectively.
[0011] Ever increasing gas temperature levels combined with a drive
towards flatter combustion radial profiles, in the interests of
reduced combustor emissions, have resulted in an increase in local
gas temperature experienced by the extremities of the blades and
vanes, and the working gas annulus endwalls formed e.g. by the NGV
31 inner and outer platforms 33, the blade 32 platform 34, and the
blade shroud 35. Effusion cooling holes 36 can be formed in the
endwalls of such components, so that cooling air can blow
therethrough to form a cooling film over the surface of the endwall
exposed to the working gas.
[0012] The pressure field into which the cooling film is introduced
typically decreases from the leading edge to the trailing edge of
the endwall of the component. Thus when the film is supplied from a
single pressure source, the film blowing rate and effectiveness at
the trailing edge is determined by a need to provide a safe
pressure margin at the leading edge. This leads to higher blowing
rates at the trailing edge than required, which compromises the
balance between film effectiveness and system mass flow rate.
[0013] Accordingly, a first aspect of the present invention
provides a component of a gas turbine engine, the component
including:
[0014] an external wall which, in use, is exposed on one surface
thereof to working gas flowing through the engine,
[0015] effusion cooling holes formed in the external wall, in use
cooling air blowing through the cooling holes to form a cooling
film on the surface of the external wall exposed to the working
gas, and
[0016] an air inlet arrangement which receives the cooling air for
distribution to the cooling holes;
[0017] wherein the component further includes:
[0018] a plurality of metering feeds and a plurality of supply
plena, the metering feeds metering the cooling air from the air
inlet arrangement to respective of the supply plena, and the supply
plena supplying the metered cooling air to respective portions of
the cooling holes.
[0019] Advantageously, the metering feeds can be configured to
provide different cooling air pressures in the supply plena. In
this way, the supply plena and their metering feeds allow the
cooling air blown through the cooling holes to be driven by
different source pressures. The film blowing rate at different
positions on the external wall can thus be matched to the working
gas pressure field, leading to enhanced film effectiveness and
reduced aerodynamic losses.
[0020] The component may have any one or, to the extent that they
are compatible, any combination of the following optional
features.
[0021] Different flow cross-sectional areas of the metering feeds
can provide the different cooling air pressures in the supply
plena.
[0022] The supply plena may be partially defined by the other
surface of the external wall. The metering feeds can then be
configured to form impingements jets from the cooling air metered
therethrough, the impingements jets impinging on said other
surface. The jets can provide further cooling of the external
wall.
[0023] Alternatively, the component may further include a plurality
of secondary plena which are partially defined by the other surface
of the external wall, and are each in flow series between a
respective supply plenum and its respective portion of the cooling
holes. In such an arrangement, the supply plena can still provide
different source pressures for the cooling holes. However, the
component may also include a plurality of jet-forming passages
which deliver the cooling air from the supply plena to the
secondary plena, the jet-forming passages being configured to form
impingements jets from the cooling air delivered therethrough, and
the impingements jets impinging on said other surface. Thus the
metering feeds do not then have to perform a jet-forming
function.
[0024] The component may further include an inlet plenum in flow
series between the air inlet arrangement and the supply plena, the
metering feeds feeding the cooling air from the inlet plenum to
respective of the supply plena. The inlet plenum can help to ensure
an even distribution of cooling air into the supply plena. However,
a different option is to arrange for the entrances to the metering
feeds to form directly the cooling air inlet arrangement, e.g. by
extending the metering feeds to the rear of the segment.
[0025] Typically, in use, the pressure of the working gas to which
the external wall is exposed varies from a leading edge to a
trailing edge of the external wall, the supply plena, metering
feeds and cooling holes being configured such that the pressure of
the cooling air blowing through each cooling hole matches the local
pressure of the working gas at that cooling hole. For example, the
working gas can vary from a higher pressure to a lower pressure
from the leading edge to the trailing edge of the external
wall.
[0026] The cooling holes may be angled in the external wall to
further reduce aerodynamic losses. The holes may have fan-shaped or
conical exit geometries, e.g. to improve spreading of the cooling
film and to reduce the exit velocity of the blown cooling air.
[0027] The cooling holes also effect cooling of the external wall
by heat transfer from the walls of the holes to the air blowing
therethrough. Increasing the cooling hole internal surface
roughness can thus enhance cooling effectiveness, as can increasing
the lengths of the cooling holes (e.g. by angling the holes and/or
increasing the external wall thickness).
[0028] The component may provide an endwall to the working gas
annulus of the engine, the external wall being the endwall. For
example, the component can be a shroud segment, and in particular a
high-pressure or intermediate-pressure shroud segment.
Alternatively, the component can be a turbine blade or a vane, a
platform of the blade or vane forming the endwall, and in
particular a high-pressure turbine blade, or a high-pressure or
intermediate-pressure nozzle guide vane.
[0029] Alternatively, the component can be a combustor, the
external wall at least partially defining a combustion chamber of
the combustor.
[0030] A second aspect of the present invention provides gas
turbine engine having one or more components according to the
previous aspect.
[0031] Embodiments of the invention will now be described by way of
example with reference to the accompanying drawings in which:
[0032] FIG. 1 shows a schematic longitudinal cross-section through
a ducted fan gas turbine engine;
[0033] FIG. 2 shows an isometric view of a typical single stage
cooled turbine;
[0034] FIG. 3 shows a schematic longitudinal cross-section through
a first embodiment of a shroud segment for a high-pressure or
intermediate-pressure turbine stage of a gas turbine engine;
and
[0035] FIG. 4 shows (a) a second embodiment of a schematic
longitudinal cross-section through another shroud segment, and, and
(b) a plot of axial variation of the gas path static pressure
adjacent the endwall of the segment of the first or second
embodiment.
[0036] FIG. 3 shows a first embodiment of a schematic longitudinal
cross-section through a shroud segment for a high-pressure or
intermediate-pressure turbine stage of a gas turbine engine. The
segment provides an endwall 40 to the working gas annulus with an
external gas washed surface 41 that is exposed to the working gas
flowing through the engine. Cooling air is blown through a
plurality of effusion cooling holes 42 formed in the endwall to
form a cooling film over the gas washed surface that protects the
endwall from the working gas. Heat transfer from the walls of the
holes to the air blowing therethrough also cools the endwall. To
prevent working gas being ingested into the segment through the
holes, the source pressure of the cooling air must exceed that of
working gas.
[0037] The cooling air is typically compressed air bled from the
compressor section of the engine and bypassing the combustor. The
direction of flow of the cooling air is indicated by arrows in FIG.
3. The air enters the segment through an air inlet aperture 43 at
the rear of the segment and fills an inlet plenum 44. A plurality
of supply plena 45 are arranged between the inlet plenum and the
endwall 40, with the inner surface 46 of the endwall partially
defining the supply plena, and the individual supply plena being
separated from each other by internal walls 50. The cooling air is
fed into the supply plena from the inlet plenum via respective
metering feeds 47. The inlet plenum evens the distribution of the
air flow to the supply plena, the pressure (P1-P8) of the cooling
air in each supply plenum being a function of at least the flow
cross-sectional area of the respective feed into that plenum.
[0038] Each supply plenum 45 then supplies the cooling air for a
respective portion of the cooling holes 42. In particular, as shown
in FIG. 3, the supply plena are arranged in a line from the leading
edge 48 to the trailing edge 49 of the endwall 40, with each supply
plena supplying a respective row of cooling holes. The row by row
hole diameter and the number of holes in each row can vary
depending on the cooling duty. Each supply plena operates at a
different pressure as determined by its metering feed 47. This
allows local target pressure margins to be maintained above the
working gas path pressure seen by the gas washed surface 41. In
particular, axial variation of the gas path static pressure
distribution can be accommodated to provide a more uniform blowing
rate and mass flow into the cooling film.
[0039] Optionally, the metering feeds 47 can form impingement jets
from the metered cooling air. These jets then impinge on the inner
surface 46 of the endwall to enhance heat transfer from the
endwall.
[0040] However, another option is to separate the metering and jet
forming functionality. Accordingly, FIG. 4(a) shows a second
embodiment of a schematic longitudinal cross-section through
another shroud segment. Corresponding features have the same
reference numbers in FIGS. 3 and 4(a). FIG. 4(b) shows a plot of
axial variation of the gas path static pressure adjacent the
endwall 40 of the segment of the first or second embodiment. The
position of the tips of the turbine blades which sweep across the
endwall is indicated.
[0041] In the second embodiment, each supply plenum 45 is fed by a
respective metering feed 47, as in the first embodiment. The
positions of the metering feeds are only schematically indicated in
FIG. 4(a). However, the entrances of the feeds, which are at the
rear side of the segment, directly form an air inlet arrangement
into the segment, i.e. a separate air inlet aperture and
intermediate inlet plenum are not needed. Further, a row of
secondary plena 51 are provided between the supply plena and the
endwall 40, with the inner surface 46 of the endwall now partially
defining the secondary plena. The secondary plena, like the supply
plena, are separated from each other by internal walls 52. Each
supply plenum has a respective secondary plenum, with jet-forming
passages 53 delivering the cooling air from the supply plena to the
secondary plena. The jets of cooling air produced by these passages
impinge on the inner surface 46 of the endwall to enhance heat
transfer from the endwall. Thus the metering feeds 47 ultimately
determine the cooling air pressure in the secondary plena, but do
not form the impingement jets.
[0042] In both embodiments, the provision of supply plena 45 and
metering feeds 47 allows better control of mass flow rate through
individual rows of cooling holes 42. In this way, a shroud segment
can achieve a high level of cooling film effectiveness, as the film
can be introduced onto the gas washed surface 41 with a momentum
which matches the pressure of the gas in contact with the wall. By
matching the cooling air blowing rate to the local gas pressure,
engine specific fuel consumption can be reduced as less cooling
flow is required. Further, improved cooling film effectiveness
allows higher turbine entry temperatures to be achieved.
[0043] Although described above in relation to shroud segments, the
present invention may also be applied to e.g. a platform of a
high-pressure turbine blade, or the platforms of a high-pressure or
intermediate-pressure nozzle guide vane. The present invention may
additionally be applied to combustor chamber wall cooling.
[0044] 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. 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.
* * * * *