U.S. patent application number 11/307607 was filed with the patent office on 2006-12-14 for bleed structure for a bleed passage in a gas turbine engine.
This patent application is currently assigned to VOLVO AERO CORPORATION. Invention is credited to Hans Martensson, Martin Nilsson.
Application Number | 20060277919 11/307607 |
Document ID | / |
Family ID | 36927683 |
Filed Date | 2006-12-14 |
United States Patent
Application |
20060277919 |
Kind Code |
A1 |
Martensson; Hans ; et
al. |
December 14, 2006 |
BLEED STRUCTURE FOR A BLEED PASSAGE IN A GAS TURBINE ENGINE
Abstract
Bleed structure (17) for a bleed passage in a gas turbine
engine. The structure includes a first wall portion (18) defining a
first side of an opening for the passage and a second wall portion
(19) defining a second side opposite the first side of the opening.
The first and second wall portions (18, 19) end at different
positions in an extension direction of the opening.
Inventors: |
Martensson; Hans;
(Trollhattan, SE) ; Nilsson; Martin; (Goteborg,
SE) |
Correspondence
Address: |
NOVAK DRUCE & QUIGG, LLP
1300 EYE STREET NW
400 EAST TOWER
WASHINGTON
DC
20005
US
|
Assignee: |
VOLVO AERO CORPORATION
SE-461 81
Trollhattan
SE
|
Family ID: |
36927683 |
Appl. No.: |
11/307607 |
Filed: |
February 14, 2006 |
Related U.S. Patent Documents
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Application
Number |
Filing Date |
Patent Number |
|
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PCT/SE05/00452 |
Mar 24, 2005 |
|
|
|
11307607 |
Feb 14, 2006 |
|
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60593941 |
Feb 25, 2005 |
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Current U.S.
Class: |
60/785 |
Current CPC
Class: |
F02K 3/075 20130101;
F05D 2250/52 20130101; F05D 2240/12 20130101; F05D 2250/314
20130101; F02C 6/08 20130101; F02C 9/18 20130101; F02C 3/13
20130101 |
Class at
Publication: |
060/785 |
International
Class: |
F02C 6/08 20060101
F02C006/08 |
Claims
1. A bleed structure (17,22,29,30,33,40,52,53) for a bleed passage
(16,55,56) in a gas turbine engine (1), said structure comprising:
a first wall portion (18,24,36,43) defining a first side of an
opening for the passage; and a second wall portion (19,26,37,44)
defining a second side opposite the first side of the opening and
wherein the first and second wall portions
(18,24,36,43;19,26,37,44) end at different positions in an
extension direction of the bleed passage opening.
2. The bleed structure as recited in claim 1, wherein one of the
first and second wall portions (18,44) is raised relative to the
adjacent surfaces of the structure.
3. The bleed structure as recited in claim 2, wherein the raised
wall portion forms an elongated projection (18,44) along said side
of the opening.
4. The bleed structure as recited in claim 1, wherein one of the
first and second wall portions (26,36) is lowered relative to the
adjacent surfaces of the structure.
5. The bleed structure as recited in claim 4, wherein the lowered
wall portion (26,36) is elongated and extends along said side of
the opening.
6. The bleed structure as recited in claim 2, wherein the other of
the first and second wall portions (19,24,37,43) is flush with the
adjacent surfaces of the structure.
7. The bleed structure as recited in claim 1, wherein at least one
airfoil (21,27,39,47) in said bleed passage opening for guiding a
gas flow in the passage.
8. The bleed structure as recited in claim 7, wherein a plurality
of airfoils (21,27,39,47) are arranged substantially in parallel to
each other in said bleed passage opening.
9. The bleed structure as recited in claim 1, wherein at least one
frame (28) surrounding the opening and the frame (28) comprises
said first and second wall portions (18,19).
10. The bleed structure as recited in claim 1, wherein said
structure forms an annular component (17) comprising a plurality of
circumferentially spaced bleed passage openings.
11. The bleed structure as recited in claim 1, wherein the passage
opening forms a continuous slot in a circumferential direction of
the structure.
12. The bleed structure as recited in claim 1, wherein the bleed
passage opening forms a bleed passage outlet.
13. The bleed structure as recited in claim 1, wherein the bleed
passage opening forms a bleed passage inlet.
14. The bleed structure as recited in claim 1, wherein a transition
from said first wall portion (18,24,36,43) to an adjacent gas duct
wall (23,25,34,41) is even so that any disturbance caused by bleed
on a passing gas flow is minimized.
15. The bleed structure as recited in claim 1, wherein a transition
from said second wall portion (19,26,37,44) to an adjacent gas duct
wall (23,25,34,41) is even so that any disturbance caused by bleed
on a passing gas flow is minimized.
16. The bleed structure as recited in claim 1, wherein the bleed
passage (16,55,56) defines a flow path for deflecting the gas with
a substantial inclination in relation to a passing gas flow.
17. A bleed arrangement for a gas turbine engine (1), said
arrangement comprising: a section of a primary gas duct (6) for the
engine, a section of a secondary gas duct (7) for the engine and at
least one bleed passage (16,55,56) connected to at least one of the
primary gas duct section and the secondary gas duct section; and a
bleed structure (17,22,29,30,33,40,52,53) for the bleed passage
(16,55,56) in the gas turbine engine (1), said structure comprising
a first wall portion (18,24,36,43) defining a first side of an
opening for the passage and a second wall portion (19,26,37,44)
defining a second side opposite the first side of the opening and
wherein the first and second wall portions
(18,24,36,43;19,26,37,44) end at different positions in an
extension direction of the bleed passage opening and wherein said
bleed structure (17,22,29,30,33,40,52,53) is arranged so that the
first wall portion (18,24,36,43) is located upstream of the opening
and the second wall portion (19,26,37,44) is located downstream of
the opening.
18. The bleed arrangement as recited in claim 17, wherein the bleed
passage (16,55,56) extends between the primary gas duct section and
the secondary gas duct section.
19. A gas turbine engine comprising a primary gas duct (6), a
secondary gas duct (7) and at least one bleed passage (16,55,56)
connected to at least one of the primary gas duct (6) and the
secondary gas duct (7); and a bleed structure
(17,22,29,30,33,40,52,53) for the bleed passage (16,55,56) in the
gas turbine engine (1), said structure comprising a first wall
portion (18,24,36,43) defining a first side of an opening for the
passage and a second wall portion (19,26,37,44) defining a second
side opposite the first side of the opening and wherein the first
and second wall portions (18,24,36,43;19,26,37,44) end at different
positions in an extension direction of the bleed passage opening
and wherein said bleed structure (17,22,29,30,33,40,52,53) is
arranged so that the first wall portion (18,24,36,43) is located
upstream of the opening and the second wall portion (19,26,37,44)
is located downstream of the opening.
20. The gas turbine engine a recited in claim 19, wherein the bleed
passage (17,22,29,30,52,53) is arranged to bleed a gas flow from
the primary gas duct (6) to the secondary gas duct (7).
21. The gas turbine engine a recited in claim 19, wherein the bleed
structure (17,22,29,30) is arranged at an outlet of the bleed
passage (16).
22. The gas turbine engine a recited in claim 20, wherein the bleed
structure (33,40) is arranged at an inlet of the bleed passage.
23. The gas turbine engine a recited in claim 21, wherein the bleed
structure (33,40) is arranged at an inlet of the bleed passage.
24. An aircraft engine comprising a primary gas duct (6), a
secondary gas duct (7) and at least one bleed passage (16)
extending between the primary gas duct (6) and the secondary gas
duct (17); and a bleed structure (17,22,29,30,33,40,52,53) for the
bleed passage (16,55,56) in the gas turbine engine (1), said
structure comprising a first wall portion (18,24,36,43) defining a
first side of an opening for the passage and a second wall portion
(19,26,37,44) defining a second side opposite the first side of the
opening and wherein the first and second wall portions
(18,24,36,43;19,26,37,44) end at different positions in an
extension direction of the bleed passage opening and wherein said
bleed structure (17,22,29,30,33,40,52,53) is arranged so that the
first wall portion (18,24,36,43) is located upstream of the opening
and the second wall portion (19,26,37,44) is located downstream of
the opening.
Description
CROSS-REFERENCE TO RELATED APPLICATIONS
[0001] The present application is a continuation patent application
of International Application No. PCT/SE2005/000452 filed 24 Mar.
2005. Further, the present application claims the benefit of U.S.
Provisional Application No. 60/593,941 filed 25 Feb. 2005. Said
applications are expressly incorporated herein by reference in
their entireties.
FIELD OF THE INVENTION
[0002] The present invention relates to a bleed structure for a
bleed passage in a gas turbine engine. The structure includes a
first wall portion defining a first side of an opening for the
passage and a second wall portion defining a second side, opposite
the first side of the opening. The bleed structure is intended to
be arranged in the gas turbine engine so that the first wall
portion is located upstream of the bleed passage opening and the
second wall portion is located downstream of the opening.
[0003] The bleed structure may be used in stationary gas turbine
engines, but is especially advantageous for aircraft jet engines.
Jet engine is meant to include various types of engines, which
admit air at relatively low velocity, heat it by combustion and
shoot it out at a much higher velocity. Accommodated within the
term jet engine are, for example, turbojet engines and turbo-fan
engines. The invention will below be described for a turbo-fan
engine, but may of course also be used for other engine types.
[0004] An aircraft gas turbine engine of the turbofan type
generally comprises a forward fan and booster compressor, a middle
core engine, and an aft low pressure power turbine. The core engine
comprises a high pressure compressor, a combustor and a high
pressure turbine in a serial relationship. The high pressure
compressor and high pressure turbine of the core engine are
interconnected by a high pressure shaft. The high-pressure
compressor, turbine and shaft essentially form a high pressure
rotor. The high-pressure compressor is rotatably driven to compress
air entering the core engine to a relatively high pressure. This
high pressure air is then mixed with fuel in the combustor and
ignited to form a high energy gas stream. The gas stream flows aft
and passes through the high-pressure turbine, rotatably driving it
and the high pressure shaft which, in turn, rotatably drives the
high pressure compressor.
[0005] The gas stream leaving the high pressure turbine is expanded
through a second or low pressure turbine. The low pressure turbine
rotatably drives the fan and booster compressor via a low pressure
shaft, all of which form the low pressure rotor. The low pressure
shaft extends through the high pressure rotor. Most of the thrust
produced is generated by the fan.
[0006] Part of the incoming air flow to the aircraft engine enters
an inner, primary gas duct, which guides the air to the combustor,
and part of the incoming air flow enters an outer, secondary gas
duct (fan duct) in which the engine bypass air flows.
[0007] In known aircraft engines, a bleed passage extends between
the primary gas duct and the secondary gas duct. According to a
known configuration, a variable bleed passage system is adapted to
bleed air from the primary gas duct to the secondary gas duct. In
certain operational conditions, compressed air is bled from the
primary gas duct via the bleed passage and introduced in a high
speed gas flow in the secondary gas duct.
[0008] There is a risk that the bleed air will negatively effect
the stability or efficiency of the engine or cause vibration
problems. A small air cushion is created when the bleed air meets
the gas flow in the fan duct, which locally increase the pressure
in the forward end of the outlet. This increased pressure creates a
non-uniform distribution of the bled gas flow, which leads to
losses. More specifically, for a set extension of the outlet in the
axial direction of the engine, the bleed gas will only flow into
the gas duct through a small part of the outlet at the downstream
end of the outlet.
SUMMARY OF THE INVENTION
[0009] The purpose of the invention is to achieve a bleed structure
for a gas turbine engine, which creates conditions for an effective
bleed while not negatively influencing the operation of the engine
or at least keep the negative effects to a minimum. More
specifically, the invention aims at improving the flow distribution
in the bleed passage with no substantial negative effects on the
gas flow in a gas duct from which the air is bled and/or in a gas
duct into which the bled air is introduced.
[0010] This purpose is achieved in that the first and second wall
portions end at different positions in an extension direction of
the bleed passage opening. Thus, the first and second wall portions
end at different positions in a direction of the bleed flow in the
bleed passage. In other words, the first and second wall portions
end at different positions in a direction perpendicular to a plane
in parallel to the walls defining the opening.
[0011] Such an opening configuration at a bleed passage outlet
creates conditions for a more favorable pressure distribution in a
gas flow in the bleed passage. Likewise, such an opening
configuration at a bleed passage inlet creates conditions for a
more favorable pressure distribution in the bleed passage.
[0012] The opening configuration is especially advantageous in
applications for bleed between a primary gas duct and a secondary
gas duct where a pressure difference is small between a compressor
portion and the secondary gas duct (fan duct) in order to secure
bleed through-flow to a sufficient extent and in the intended
direction. The opening configuration is further advantageous in
applications where there is a limited space available for the bleed
opening.
[0013] According to a preferred embodiment of the invention, for a
bleed passage outlet, an upstream wall portion ends at a position
closer to a wall defining the gas duct, which is opposite said
bleed passage opening, than the downstream wall portion. The speed
of the introduced bleed gas may then be leveled to some extent at
the outlet in the axial direction of the gas turbine and a larger
bleed flow may be introduced than according to presently employed
solutions. In other words, the bleed gas will flow into the gas
duct through a larger part of the outlet.
[0014] Thus, according to the preferred embodiment of the
invention, one of the first and second wall portions is raised
relative to the adjacent surfaces of the structure. This opening
configuration at the outlet creates conditions for introducing a
large bleed air flow into the gas duct.
[0015] According to a further preferred embodiment of the
invention, the other of the first and second wall portions is flush
with the adjacent surfaces of the structure. This opening
configuration at the outlet creates conditions for substantially
not negatively affecting the passing gas flow in the gas duct into
which the bleed air is introduced.
[0016] According to a further preferred embodiment of the
invention, one of the first and second wall portions is lowered
relative to the adjacent surfaces of the structure. This opening
configuration at the inlet creates conditions for substantially not
negatively affecting the passing gas flow in the gas duct from
which the bleed air is extracted.
[0017] According to a further preferred embodiment of the
invention, a transition from at least one of said first and second
wall portion to an adjacent gas duct wall is even so that any
disturbance caused by bleed on a passing gas flow is minimized. The
transition portion is preferably smooth, uninterrupted and
substantially flat.
[0018] According to a further preferred embodiment of the
invention, it comprises at least one airfoil in said bleed passage
opening for guiding a gas flow in the passage. By virtue of the
airfoils, the bleed air may be guided in a desired direction
to/from the bleed passage. Further, the airfoils create conditions
for a larger deflection of the bleed flow in a set axial
distance.
[0019] Further advantageous embodiments and further advantages of
the invention emerge from the detailed description below and the
claims.
BRIEF DESCRIPTION OF THE DRAWINGS
[0020] The invention will be described below, with reference to the
embodiments shown on the appended drawings, wherein:
[0021] FIG. 1 illustrates an aircraft engine in a schematic cut
side view;
[0022] FIG. 2 shows a cut side view of a first embodiment of a
bleed structure outlet configuration;
[0023] FIG. 3 shows a cut side view of a second embodiment of a
bleed structure outlet configuration;
[0024] FIG. 4 shows a schematic perspective view of a gas turbine
engine component comprising the bleed structure of FIG. 2;
[0025] FIG. 5 shows a perspective view of the bleed structure of
FIG. 2;
[0026] FIGS. 6, 7, 8 and 9 each shows a bleed outlet structure
according to an alternative embodiment;
[0027] FIG. 10 shows a cut portion of an aircraft engine according
to an alternative embodiment in a schematic side view;
[0028] FIG. 11 shows a cut side view of a first embodiment of a
bleed structure inlet configuration; and
[0029] FIG. 12 shows a cut side view of a second embodiment of a
bleed structure inlet configuration.
DETAILED DESCRIPTION
[0030] The invention will below be described for a turbofan gas
turbine aircraft engine 1, which in FIG. 1 is circumscribed about
an engine longitudinal central axis 2. The engine 1 comprises an
outer casing 3, or nacelle, an inner casing 4, and an intermediate
casing 5, which is concentric to the first two casings and divides
the gap between them into an inner primary gas duct 6 for the
compression of air and a secondary duct 7 in which the engine
bypass air flows. Thus, each of the gas ducts 6,7 is annular in a
cross section perpendicular to the engine longitudinal central axis
2.
[0031] The engine 1 comprises a fan 8 which receives ambient air 9,
a booster or low pressure compressor (LPC) 10 and a high pressure
compressor (HPC) 11 arranged in the primary gas duct 6, a combustor
12 which mixes fuel with the air pressurized by the high pressure
compressor 11 for generating combustion gases which flow downstream
through a high pressure turbine (HPT) 13 and a low pressure turbine
(LPT) 14 from which the combustion gases are discharged from the
engine.
[0032] A high pressure shaft joins the high pressure turbine 13 to
the high pressure compressor 11 to form a high pressure rotor. A
low pressure shaft joins the low pressure turbine 14 to the low
pressure compressor 10 to form a low pressure rotor. The high
pressure compressor 11, combustor 12 and high pressure turbine 13
are collectively referred to as a core engine. The low pressure
shaft is at least in part rotatably disposed co-axially with and
radially inwardly of the high pressure rotor.
[0033] A load carrying engine structure 15 is arranged between the
outer casing 3 and the inner casing 4.
[0034] A plurality of circumferentially spaced bleed passages 16
extend between the primary gas duct 6 and the secondary gas duct 7.
The bleed passages 16 define a flow path for routing air from the
primary gas duct 6 to the secondary gas duct 7 and more
specifically from an end region of the low pressure compressor 10.
A bleed passage inlet is arranged in a gap between an upstream
rotor and a downstream stator in the low pressure compressor
10.
[0035] FIG. 2 shows a cut side view of a first embodiment of a
bleed structure 17 forming a bleed passage outlet to the secondary
gas duct 7. The structure 17 comprises a first, upstream wall
portion 18 forming a leading edge of the outlet. The structure 17
further comprises a second, downstream wall portion 19 forming a
trailing edge of the outlet. The first and second wall portions
18,19 end at different distances in the extension direction of the
passage 16.
[0036] The wall 23 defining the gas duct 7 is substantially at the
same level across the bleed opening (outlet). Thus, the wall 23
extends along a substantially straight line across the opening.
[0037] More specifically, the upstream wall portion 18 is raised
relative to the adjacent surfaces of the structure and the gas duct
wall 23. Further, the upstream wall portion 18 is raised relative
to the downstream wall portion 19 so that a gas flow 107 in the
duct 7 is directed somewhat radially away from the outlet and
thereby creating a low pressure region outside the outlet. The
upstream wall portion 18 is smoothed and aerodynamically rounded
for reducing discontinuities in the fan gas duct flow. Further, a
transition from said upstream wall portion 18 to the adjacent gas
duct wall 23 is even so that any disturbance caused by bleed on the
passing gas flow 107 is minimized.
[0038] The raised upstream wall portion 18 forms an elongated
projection extending in the circumferential direction of the
structure along an upstream side of the outlet, see also FIG. 5.
The downstream wall portion 19 is substantially flush with the
adjacent surfaces of the structure and the gas duct wall 23. A
transition from the downstream wall portion 19 to the adjacent gas
duct wall 23 is even so that any disturbance caused by bleed on the
passing gas flow is minimized. Further, an end 20 of the downstream
wall portion 19 facing the outlet 17 is chamfered defining a flow
path for the bleed gas 116 from the bleed passage 16 to the gas
duct 7.
[0039] Four airfoils 21 (or stator vanes) are arranged
substantially in parallel to each other in the outlet, see also
FIG. 5, for guiding the bleed gas flow 116 to the secondary gas
duct 7. The airfoils 21 are arranged at a distance from each other
in the axial direction 2 of the engine 1.
[0040] The bleed passage 16 defines a flow path for deflecting the
gas with a substantial inclination in relation to the passing gas
flow 107 in the secondary gas duct 7. Preferably, the gas is
deflected at an angle of at least 45 degrees and especially at an
angle of at least 60 degrees in relation to the passing gas flow.
More specifically, in the shown embodiment, the gas is deflected at
substantially right angles with the passing gas flow.
[0041] FIG. 3 shows an alternative embodiment of a structure 22
forming a bleed passage outlet to the secondary gas duct 7. An
upstream wall portion 24 is substantially flush with the adjacent
surfaces of the structure and the gas duct wall 25. Further, a
transition from said upstream wall portion 24 to the adjacent gas
duct wall 25 is even so that any disturbance caused by bleed on the
passing gas flow 107 is minimized.
[0042] The downstream wall portion 26 is lowered relative to the
adjacent surfaces of the structure and the gas duct wall 25. More
specifically, the lowered wall portion 26 is elongated and extends
along the downstream side of the outlet. Further, the lowered wall
portion 26 has a contoured shape and shows a smooth, uninterrupted
surface facing the gas flow. A transition from the downstream wall
portion 26 to the adjacent gas duct wall 25 is even so that any
disturbance caused by bleed on the passing gas flow is minimized.
The downstream wall portion 26 is aerodynamically rounded for
reducing discontinuities in the fan gas duct flow 107. Four
airfoils 27 are arranged substantially in parallel to each other in
the outlet for guiding a bleed gas flow 116 to the secondary gas
duct 7.
[0043] FIG. 4 shows a perspective view of the outlet bleed
structure 17 in FIG. 1 and 2. The bleed structure 17 forms an
annular component comprising a plurality of circumferentially
spaced bleed passage outlets through the inner wall 23 of the
secondary gas duct 7. A rectangular frame 28 surrounds each outlet,
see also FIG. 5. The frames are joined to each other via flanges
31,32, see FIG. 5, to form said annular component. The structure 17
comprises means 50,51 for connection to an adjacent frame. The
connection means may for example comprise a bolt connection.
Through-holes 50,51 extend through each flange 31,32 for said
connection means.
[0044] The frame 28 comprises said upstream wall portion 18 and
downstream wall portion 19. Thus, the frames 28 are separate
pieces, which are positioned in a slot or aperture in the gas duct
wall 23. The frames 28 are arranged relative to the edges of the
gas duct wall defining the slot or aperture in such a manner that
the frames are substantially flush with the gas duct wall so that a
passing gas flow is not disturbed by the edges of the frames.
[0045] According to the embodiment shown in FIG. 4, the term "bleed
structure" comprises the plurality of frames forming the annular
component. According to an alternative, the bleed structure forms a
unison ring. According to a further alternative, the term "bleed
structure" comprises a single frame surrounding one or a plurality
of openings.
[0046] FIG. 5 illustrates the bleed structure 17 comprising a
rectangular frame with a rectangular opening and a grid of airfoils
21. The airfoils 21 extend between two opposite sides of the
rectangular frame and are fixedly attached to the frame. The
airfoils 21 are arranged in parallel to the upstream and downstream
wall portions 18,19. The elongated projection 18 has rounded edges
in the circumferential direction of the gas turbine. FIG. 6
illustrates an alternative bleed structure 52. The elongated
projection 18 extends at both ends a distance around the corner of
the opening. FIG. 7 illustrates a further alternative bleed
structure 29 comprising a rectangular frame with a substantially
circular opening. FIG. 8 illustrates a further alternative bleed
structure 30 comprising a rectangular frame with a substantially
elliptical opening. FIG. 9 illustrates a still further alternative
bleed structure 53. The elongated projection 18 extends at both
ends a distance around the curved periphery of the opening. Each of
the four alternative bleed structures 29,30,52,53 comprises a
raised upstream wall portion 18 and a grid of airfoils 21.
[0047] FIG. 10 illustrates a compressor portion of an aircraft
engine. More specifically, the region of the low pressure
compressor 10 and the high pressure compressor 11 is shown. A bleed
passage 55 is arranged to bleed air from the secondary gas duct 7
at a position upstream of the load carrying engine structure 15.
The bled air may be introduced into the gas flow of the primary gas
duct 6 or be used for cooling engine components or similar. An
inlet of the bleed passage 55 is arranged in the inner wall
defining the secondary gas duct 7.
[0048] A further bleed passage 56 is arranged to bleed air from the
secondary gas duct 7 at a position downstream of the load carrying
engine structure 15. The bled air is routed downstream for turbine
cooling, but may as an alternative be introduced into the gas flow
of the primary gas duct 6 or be used for cooling other engine
components. An inlet of the bleed passage 56 is arranged in the
inner wall defining the secondary gas duct 7.
[0049] FIG. 11 illustrates a first embodiment of a bleed passage
inlet structure 33. The inlet bleed structure 33 is arranged in a
wall 34 defining a gas duct from which gas is extracted. The gas
duct 35 may, according to one example, be formed by the secondary
gas duct 7 in FIG. 1, see FIG. 10.
[0050] The wall 34 defining the gas duct 7 is substantially at the
same level across the bleed opening (inlet). Thus, the wall 34
extends along a substantially straight line across the opening.
[0051] The bleed passage 55,56 defines a flow path for deflecting
the gas with a substantial inclination in relation to the passing
gas flow in the secondary gas duct 7. Preferably, the gas is
deflected at an angle of at least 45 degrees and especially at an
angle of at least 60 degrees in relation to the passing gas flow.
More specifically, in the shown embodiment, the gas is deflected at
substantially right angles with the passing gas flow.
[0052] An upstream wall portion 36 is lowered relative to the
adjacent surfaces of the structure and the gas duct wall 34.
Further, a transition from said upstream wall portion 36 to the
adjacent gas duct wall 34 is even so that any disturbance caused by
bleed on the passing gas flow is minimized. The downstream wall
portion 37 is substantially flush with the adjacent surfaces of the
structure and the gas duct wall 34. A transition from the
downstream wall portion 37 to the adjacent gas duct wall 34 is even
so that any disturbance caused by bleed on the passing gas flow is
minimized. More specifically, the lowered wall portion 36 is
elongated and extends along the upstream side of the inlet. The
lowered wall portion 36 extends away from the gas duct wall 34
defining a flow path for the bleed gas from the gas duct 35 to a
bleed passage 38. Further, the lowered wall portion 36 has a
contoured shape and shows a smooth, uninterrupted surface facing
the gas flow. The upstream wall portion 36 is aerodynamically
rounded for reducing discontinuities in the gas duct flow.
[0053] A plurality of airfoils 39 are arranged substantially in
parallel to each other in the inlet for guiding a bleed gas flow
from the gas duct 35.
[0054] FIG. 12 illustrates a second embodiment of a bleed passage
inlet structure 40. The inlet bleed structure 40 is arranged in a
wall 41 defining a gas duct 42 from which gas is extracted. The
structure 40 comprises a first, upstream wall portion 43 and a
second, downstream wall portion 44. The upstream wall portion 43 is
substantially flush with the adjacent surfaces of the structure and
the gas duct wall 34. Further, an end 45 of the upstream wall
portion 43 facing the inlet is chamfered defining a flow path for
the bleed gas from the gas duct 42 to a bleed passage 46. The
upstream wall portion 43 is smoothed and aerodynamically rounded
for reducing discontinuities in the gas duct flow. A plurality of
airfoils 47 are arranged in parallel to each other in the inlet for
guiding a bleed gas flow from the gas duct 42.
[0055] The downstream wall portion 44 is raised relative to the
adjacent surfaces of the structure and the gas duct wall 41.
Further, the downstream wall portion 44 is raised relative to the
upstream wall portion 43. The raised downstream wall portion 44
forms an elongated projection extending in the circumferential
direction of the structure along a downstream side of the inlet.
The downstream wall portion 44 has a substantially flat surface 48
facing the inlet and the surface 49 facing the gas duct 7 is
smoothed and aerodynamically rounded for reducing discontinuities
in the gas duct flow.
[0056] The bleed passage inlet structures of FIGS. 11 and 12 may
further have a similar frame configuration as shown in any of FIGS.
5-9.
[0057] The wall portions defining the bleed passage opening are
preferably stationary, i.e. non-variable with regard to each
other.
[0058] The invention is also related to an arrangement for a gas
turbine engine comprising the bleed structure described above. The
arrangement comprises a section of a primary gas duct 6 for the
engine, a section of a secondary gas duct 7 for the engine and said
at least one bleed passage 16 connected to at least one of the
primary gas duct section and the secondary gas duct section. Such
an arrangement may be fabricated to form a separate unit, which in
turn may be assembled to other units in order to build up an
engine.
[0059] The invention is not in anyway limited to the above
described embodiments, instead a number of alternatives and
modifications are possible without departing from the scope of the
following claims.
[0060] According to an alternative to the embodiment where the
bleed structure forms an annular component comprising a plurality
of circumferentially spaced bleed passage openings, it may form an
annular component comprising a continuous slot in a circumferential
direction of the structure.
[0061] According to a further alternative, the bleed passage is
arranged downstream of the combustor 12 for routing air from the
primary gas duct 6 to the secondary gas duct 7. More specifically,
it may be arranged between high pressure turbine 13 and the low
pressure turbine 14.
[0062] According to a further alternative, the outlet configuration
is not limited to be arranged through a radially inner wall of an
outer gas duct, but may also be arranged in a radially outer wall
of an inner gas duct, like the primary gas duct 6.
[0063] According to a further alternative, the inlet configuration
is not limited to be arranged through a radially inner wall of an
outer gas duct for extracting gas radially inwards, but may also be
arranged in a radially outer wall of an inner gas duct, like the
primary gas duct 6, for extracting gas radially outwards.
[0064] Further the inlet configuration is not limited to form an
inlet to a bleed passage between a primary and a secondary gas
duct. The inlet configuration may be used for a bleed passage from
a gas duct for routing air to secondary systems like turbine
cooling systems, aircraft systems etc.
[0065] Further, the number of airfoils in each bleed passage
opening may of course differ from the four airfoils shown in the
drawings.
[0066] Further, as an alternative to the embodiment where the bleed
structure forms an annular component comprising a plurality of
circumferentially spaced bleed passage openings, some of the
openings, for example every second opening in the circumferential
direction, is free from airfoils. According to a further
alternative embodiment, the bleed structure is free from any
airfoils.
[0067] The frames are preferably rounded in the circumferential
direction of the gas duct in order to form a circular, continuous,
uninterrupted ring, i.e. a ring free of any abrupt transitions
between adjacent frames.
[0068] The invention has been described above for a two shaft
engine, however, the invention may of course also be applied in a
one shaft engine or in a three shaft engine.
[0069] According to an alternative embodiment of the bleed
structure shown in FIG. 4 and 5, the frames are fastened to an
annular support member. Thus, in such a configuration, the frames
are not connected directly to each other, but instead to the
annular support member. Consequently, the flanges with holes for
connection means do not extend perpendicular to an opening plane of
the frame, but are instead arranged in line with the frame.
[0070] According to an alternative embodiment of the bleed
structure, there is no frame around the respective opening. Thus,
the opening ends directly in the gas duct wall.
* * * * *