U.S. patent application number 15/517262 was filed with the patent office on 2017-08-31 for centrifugal compressor diffuser passage boundary layer control.
The applicant listed for this patent is GENERAL ELECTRIC COMPANY. Invention is credited to David Vickery PARKER, Caitlin Jeanne SMYTHE, James Richard WILSON.
Application Number | 20170248155 15/517262 |
Document ID | / |
Family ID | 53872192 |
Filed Date | 2017-08-31 |
United States Patent
Application |
20170248155 |
Kind Code |
A1 |
PARKER; David Vickery ; et
al. |
August 31, 2017 |
CENTRIFUGAL COMPRESSOR DIFFUSER PASSAGE BOUNDARY LAYER CONTROL
Abstract
A centrifugal compressor diffuser (42) includes a plurality of
diffuser flow passages (22) extending through an annular diffuser
housing (20) and circumferentially bounded by diffuser vanes (23)
and axially bounded by forward and aft walls (101, 100). A diffuser
boundary layer bleed (96) for the passages may include boundary
layer bleed apertures (106) or slots (130) disposed through the
forward wall (101) and a downstream facing wall (142) canted at an
acute cant angle to a downstream diffuser airflow direction (103)
in the passages. Diffuser bleed flow (112) is bled from a diffuser
boundary layer. Boundary layer bleed apertures can be located
downstream of throat sections (28) of the flow passages near
pressure sides of the vanes. A centrifugal compressor (18) may
include the diffuser surrounding an annular centrifugal compressor
impeller (32) and apparatus for flowing impeller bleed flow (102)
from a radial clearance between an impeller tip (36) and a diffuser
annular inlet (27) with diffuser bleed flow either mixed or
separately to cool a turbine (16).
Inventors: |
PARKER; David Vickery;
(Middleton, MA) ; SMYTHE; Caitlin Jeanne;
(Cambridge, MA) ; WILSON; James Richard; (Melrose,
MA) |
|
Applicant: |
Name |
City |
State |
Country |
Type |
GENERAL ELECTRIC COMPANY |
Schenectady |
NY |
US |
|
|
Family ID: |
53872192 |
Appl. No.: |
15/517262 |
Filed: |
August 11, 2015 |
PCT Filed: |
August 11, 2015 |
PCT NO: |
PCT/US2015/044673 |
371 Date: |
April 6, 2017 |
Related U.S. Patent Documents
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Application
Number |
Filing Date |
Patent Number |
|
|
62060991 |
Oct 7, 2014 |
|
|
|
Current U.S.
Class: |
1/1 |
Current CPC
Class: |
F02C 3/08 20130101; F05D
2250/324 20130101; F04D 29/682 20130101; F02C 6/08 20130101; F01D
9/045 20130101; F04D 17/10 20130101; F05D 2270/101 20130101; F01D
25/12 20130101; F05D 2270/17 20130101; F02C 7/18 20130101; F05D
2220/32 20130101; F04D 29/284 20130101; F05D 2240/12 20130101; F04D
29/444 20130101; F04D 27/009 20130101 |
International
Class: |
F04D 29/44 20060101
F04D029/44; F04D 29/28 20060101 F04D029/28; F04D 27/00 20060101
F04D027/00; F04D 17/10 20060101 F04D017/10 |
Goverment Interests
GOVERNMENT INTERESTS
[0001] This invention was made with government support under
government contract No. W911-W6-11-2-0009 by the Department of
Defense. The government has certain rights to this invention.
Claims
1. A gas turbine engine centrifugal compressor diffuser comprising:
an annular diffuser housing, diffuser vanes axially extending
between a forward wall and an aft wall of he diffuser housing, a
plurality of diffuser flow passages extending through the housing
and spaced about a circumference of the housing, the diffuser flow
passages bounded by the diffuser vanes and the forward and aft
walls, and a diffuser boundary layer bleed for bleeding diffuser
bleed flow from a diffuser boundary layer in each of the diffuser
flow passages.
2. The diffuser according to claim 1 further comprising the
diffuser boundary layer bleed configured for bleeding the diffuser
bleed flow from the diffuser boundary layer at a position located
in a region of flow weakness in each of the diffuser flow
passages.
3. The diffuser according to claim 1 further comprising the
diffuser boundary layer bleed including boundary layer bleed
apertures disposed through the forward wall.
4. The diffuser according to claim 3 further comprising each of the
boundary layer bleed apertures being a slot including a downstream
facing wall angled or canted at an acute cant angle with respect to
a downstream diffuser airflow direction in each of the diffuser
flow passages respectively.
5. The diffuser according to claim 3 further comprising the
boundary layer bleed apertures positioned or located downstream of
throat sections of the diffuser flow passages near pressure sides
of the diffuser vanes.
6. The diffuser according to claim 5 further comprising each of the
boundary layer bleed apertures being a slot including a downstream
facing wall angled or canted at an acute cant angle with respect to
a downstream diffuser airflow direction in each of the diffuser
flow passages respectively.
7. A gas turbine engine centrifugal compressor comprising: an
annular centrifugal compressor impeller, a diffuser annularly
surrounding the impeller, a plurality of diffuser flow passages
extending through a housing of the diffuser and spaced about a
circumference of the housing, each of the passages including a
throat section and a diffusing section downstream of the throat
section, the diffuser flow passages circumferentially bounded by
diffuser vanes extending axially between forward and aft walls of
the diffuser, and a diffuser boundary layer bleed for bleeding
diffuser bleed flow from a diffuser boundary layer in each of the
diffuser flow passages.
8. The centrifugal compressor according to claim 7 further
comprising the diffuser boundary layer bleed configured for
bleeding the diffuser bleed flow from the diffuser boundary layer
at a position located in a region of flow weakness in each of the
diffuser flow passages.
9. The diffuser according to claim 7 further comprising the
diffuser boundary layer bleed including boundary layer bleed
apertures disposed through the forward wall.
10. The centrifugal compressor according to claim 9 further
comprising each of the boundary layer bleed apertures being a slot
including a downstream facing wall angled or canted at an acute
cant angle with respect to a downstream diffuser airflow direction
in each of the diffuser flow passages respectively.
11. The centrifugal compressor according to claim 10 further
comprising the boundary layer bleed apertures positioned or located
downstream of throat sections of the diffuser flow passages near
pressure sides of the diffuser vanes.
12. The centrifugal compressor according to claim 11 further
comprising each of the boundary layer bleed apertures being a slot
including a downstream facing wall angled or canted at an acute
cant angle with respect to a downstream diffuser airflow direction
in each of the diffuser flow passages respectively.
13. The centrifugal compressor according to claim 9 further
comprising: a radial clearance between an impeller tip of the
impeller and an annular inlet of the diffuser, a means for mixing
impeller bleed flow from the radial clearance with the diffuser
bleed flow from the boundary layer bleed apertures for providing
turbine cooling air and flowing the turbine cooling air to a
turbine, or a means for flowing the impeller bleed flow and the
diffuser bleed flow separately to the turbine.
14. The centrifugal compressor according to claim 13 further
comprising each of the boundary layer bleed apertures being a slot
including a downstream facing wall angled or canted at an acute
cant angle with respect to a downstream diffuser airflow direction
in each of the diffuser flow passages respectively.
15. The centrifugal compressor according to claim 13 further
comprising the boundary layer bleed apertures positioned or located
downstream of throat sections of the diffuser flow passages near
pressure sides of the diffuser vanes.
16. The centrifugal compressor according to claim 15 further
comprising each of the boundary layer bleed apertures being a slot
including a downstream facing wall angled or canted at an acute
cant angle with respect to a downstream diffuser airflow direction
in each of the diffuser flow passages respectively.
17. The centrifugal compressor according to claim 9 further
comprising: a radial clearance between an impeller tip of the
impeller and an annular inlet of the diffuser, the radial clearance
in fluid communication with a radially inner manifold, the boundary
layer bleed apertures in flow communication with a radially outer
manifold, the radially inner manifold in fluid communication with
the radially outer manifold such that the impeller bleed flow flows
into the radially outer manifold and mixes with the diffuser bleed
flow to form turbine cooling air, and means for flowing turbine
cooling air out of the radially outer manifold.
18. The centrifugal compressor according to claim 17 further
comprising each of the boundary layer bleed apertures being a slot
including a downstream facing wall angled or canted at an acute
cant angle with respect to a downstream diffuser airflow direction
in each of the diffuser flow passages respectively.
19. The centrifugal compressor according to claim 18 further
comprising the boundary layer bleed apertures positioned or located
downstream of throat sections of the diffuser flow passages near
pressure sides of the diffuser vanes.
20. The centrifugal compressor according to claim 19 further
comprising each of the boundary layer bleed apertures being a slat
including a downstream facing wall angled or canted at an acute
cant angle with respect to a downstream diffuser airflow direction
in each of the diffuser flow passages respectively.
21. The centrifugal compressor according to claim 9 further
comprising: a radial clearance between an impeller tip of the
impeller and an annular inlet of the diffuser, the radial clearance
in fluid communication with a radially inner annular manifold,
inter-manifold apertures disposed between the inner annular
manifold and a plurality of radially outer annular manifolds, a
means for porting and flowing the impeller bleed flow from the
radial clearance through a plurality of circumferentially
distributed impeller bleed flow manifold ports in and through an
diffuser forward casing surrounding the centrifugal compressor to
the high pressure turbine for turbine cooling, the diffuser
boundary layer bleed in fluid flow communication with and operable
for bleeding the diffuser bleed flow into an annular diffuser bleed
manifold, and a means for porting and flowing the diffuser bleed
flow through a plurality of circumferentially distributed diffuser
bleed manifold ports in and through the diffuser forward casing to
the high pressure turbine for turbine cooling.
22. The centrifugal compressor according to claim 21 further
comprising each of the boundary layer bleed apertures being a slot
including a downstream facing wall angled or canted at an acute
cant angle with respect to a downstream diffuser airflow direction
in each of the diffuser flow passages respectively.
23. The centrifugal compressor according to claim 22 further
comprising the boundary layer bleed apertures positioned or located
downstream of throat sections of the diffuser flow passages near
pressure sides of the diffuser vanes.
24. The centrifugal compressor according to claim 23 further
comprising each of the boundary layer bleed apertures being a slot
including a downstream facing wall angled or canted at an acute
cant angle with respect to a downstream diffuser airflow direction
in each of the diffuser flow passages respectively.
Description
BACKGROUND OF THE INVENTION
Technical Field
[0002] The present invention relates to bleed air from gas turbine
engine centrifugal compressors.
[0003] One type of gas turbine engine includes a centrifugal
compressor having a rotatable impeller to accelerate and, thereby,
increase the kinetic energy of air flowing therethrough. A diffuser
is generally located immediately downstream of and surrounding the
impeller. The diffuser operates to decrease the velocity of the air
flow leaving the impeller and transform the energy thereof to an
increase in static pressure, thus, pressurizing the air.
[0004] A conventional gas turbine engine typically includes a
compressor, combustor, and turbine, both rotating turbine
components such as blades, disks and retainers, and stationary
turbine components, such as vanes, shrouds, and frames routinely
require cooling due to heating thereof by hot combustion gases.
Cooling of the turbine, especially the rotating components, is
important to the proper function and safe operation of the engine.
It is known to bleed cooling air from the centrifugal compressor to
help cool the turbine.
[0005] Failure to adequately cool a turbine disk and its blading,
for example, by providing cooling air deficient in supply pressure,
volumetric flow rate or temperature margin, may be detrimental to
the life and mechanical integrity of the turbine. Depending on the
nature and extent of the cooling deficiency, the impact on engine
operation may range from relatively benign blade tip distress,
resulting in a reduction in engine power and useable blade life, to
a rupture of a turbine disk, resulting in an unscheduled engine
shutdown.
[0006] Balanced with the need to adequately cool the turbine is the
desire for higher levels of engine operating efficiency which
translate into lower fuel consumption and lower operating costs.
Since turbine cooling air is typically drawn from one or more
stages of the compressor and channelled by various means, such as
pipes, ducts, and internal passageways to the desired components,
such air is not available to be mixed with fuel, ignited in the
combustor and undergo work extraction in the primary gas flowpath
of the turbine.
[0007] Total cooling flow bled from the compressor is a loss in the
engine operating cycle and it is desirable to keep such losses to a
minimum.
[0008] Some conventional engines employ clean air bleed systems to
cool turbine components in gas turbines using an axi-centrifugal
compressor as is done in the General Electric CFE738 engine. The
turbine cooling supply air exits the centrifugal diffuser through a
small gap between the diffuser exit and deswirler inner shroud.
Other turbine cooling air methods include extracting cooling from
the impeller or from a gap between the impeller and the diffuser
exit.
[0009] U.S. Pat. No. 5,555,7211 to Bourneuf, et al, which issued on
Sep. 17, 1996 and is entitled AGas Turbine Engine Cooling Supply
Circuit@, discloses using bleed air from an impeller stage of a
centrifugal compressor in a turbine cooling supply circuit for a
gas turbine. U.S. Pat. No. 5,555,721 discloses impeller tip forward
bleed flow and impeller tip aft bleed flow for cooling turbine
components. U.S. Pat. No. 5,555,721 is assigned to the General
Electric Co., the same assignee as this patent and is incorporated
herein by reference.
[0010] U.S. Pat. No. 8,087,249 to Ottaviano, et al. which issued
Jan. 3, 2012, and is entitled ATurbine Cooling Air From A
Centrifugal Compressor@ discloses a gas turbine engine turbine
cooling system including an impeller and a diffuser directly
downstream of the impeller and a bleed for bleeding clean cooling
air from downstream of the diffuser. U.S. Pat. No. 8,087,249 is
assigned to the General Electric Co., the same assignee as this
patent and is incorporated herein by reference.
[0011] Thus, there continues to be a demand for advancements in
diffuser design and geometry that improves aerodynamic performance
and reduces the overall engine radial envelope.
BRIEF DESCRIPTION OF THE INVENTION
[0012] A diffuser for a centrifugal compressor includes an annular
diffuser housing, diffuser vanes axially extending between a
forward wall and an aft wall of the diffuser housing, a plurality
of diffuser flow passages extending through the housing and spaced
about a circumference of the housing. The diffuser flow passages
are bounded by the diffuser vanes and the forward and aft walls. A
diffuser boundary layer bleed is provided for bleeding diffuser
bleed flow from a diffuser boundary layer in each of the diffuser
flow passages.
[0013] The diffuser boundary layer bleed may be configured for
bleeding the diffuser bleed flow from the diffuser boundary layer
at a position located in a region of flow weakness in each of the
diffuser flow passages.
[0014] The diffuser boundary layer bleed may include boundary layer
bleed apertures disposed through the forward wall. Each of the
boundary layer bleed apertures may be a slot including a downstream
facing wall angled or canted at an acute cant angle with respect to
a downstream diffuser airflow direction in each of the diffuser
flow passages respectively.
[0015] The boundary layer bleed apertures may be positioned or
located downstream of throat sections of the diffuser flow passages
near pressure sides of the diffuser vanes.
[0016] A centrifugal compressor including an annular centrifugal
compressor impeller, a diffuser annularly surrounding the impeller,
and a plurality of diffuser flow passages extending through a
housing of the diffuser and spaced about a circumference of the
housing. Each of the passages includes a throat section and a
diffusing section downstream of the throat section. The diffuser
flow passages are circumferentially bounded by diffuser vanes
extending axially between forward and aft walls of the diffuser and
a diffuser boundary layer bleed is provided for bleeding diffuser
bleed flow from a diffuser boundary layer in each of the diffuser
flow passages.
[0017] The centrifugal compressor may also include a radial
clearance between an impeller tip of the impeller and an annular
inlet of the diffuser, a means for mixing impeller bleed flow from
the radial clearance with diffuser bleed flow from the boundary
layer bleed apertures for providing turbine cooling air, and a
means for flowing the turbine cooling air to a turbine or a means
for flowing impeller bleed flow and the diffuser bleed flow
separately to the turbine.
BRIEF DESCRIPTION OF THE DRAWINGS
[0018] FIG. 1 is a sectional view illustration of a gas turbine
engine with a centrifugal compressor for mixing impeller tip bleed
flow and diffuser bleed flow in the compressor section before using
the flows for cooling turbine components.
[0019] FIG. 2 is an enlarged sectional view illustration of the
centrifugal compressor and a diffuser with diffuser bleed holes
illustrated in FIG. 1.
[0020] FIG. 3 is an aft looking forward perspective view
illustration of the diffuser and the diffuser bleed holes through
3-3 in FIG. 2.
[0021] FIG. 4 is an enlarged perspective view illustration of the
bleed holes illustrated in FIG. 3.
[0022] FIG. 5 is a perspective view illustration of a portion of
the diffuser and the diffuser bleed holes illustrated in FIG.
2.
[0023] FIG. 6 is an enlarged sectional view illustration of the
centrifugal compressor tip and the diffuser bleed holes illustrated
in FIG. 2.
[0024] FIG. 7 is a sectional view illustration of a gas turbine
engine centrifugal compressor with an alternative arrangement for
separately flowing impeller tip bleed for cooling turbine
components.
[0025] FIG. 8 is a sectional view illustration of the gas turbine
engine illustrated in FIG. 7 with an arrangement for separately
flowing diffuser bleed flow for cooling turbine components.
[0026] FIG. 9 is an enlarged perspective view illustration of one
of the impeller bleed flow ports illustrated in FIG. 7 and as taken
through 9-9 in FIG. 10.
[0027] FIG. 10 is a forward looking aft perspective view
illustration of an aft casing surrounding the centrifugal
compressor and including the impeller and bleed flow ports
illustrated in FIGS. 7 and 8 respectively.
[0028] FIG. 11 is cutaway perspective view illustration of impeller
bleed flowpaths for one of the impeller bleed flow ports
illustrated in FIGS. 7 and 9.
[0029] FIG. 12 is an enlarged perspective view illustration of one
of the diffuser bleed flow ports illustrated in FIG. 8 and as taken
through 12-12 in FIG. 10.
[0030] FIG. 13 is cutaway perspective view illustration of a
diffuser bleed flowpath through one of the diffuser bleed flow
ports illustrated in FIG. 8 and as taken through 12-12 in FIG.
10.
DETAILED DESCRIPTION OF THE INVENTION
[0031] Illustrated in FIG. 1 is a gas turbine engine high pressure
centrifugal compressor 18 in a high pressure gas generator 10 of a
gas turbine engine 8. The high pressure centrifugal compressor 18
is a final compressor stage of a high pressure compressor 4. The
high pressure gas generator 10 has a high pressure rotor 12
including, in downstream serial or flow relationship, the high
pressure compressor 14, a combustor 52, and a high pressure turbine
16. The rotor 12 is rotatably supported about an engine axis 25 by
bearings in engine frames not illustrated herein.
[0032] The exemplary embodiment of the high pressure compressor 14
illustrated herein includes a five stage axial compressor 30
followed by the centrifugal compressor 18 having an annular
centrifugal compressor impeller 32. Outlet guide vanes 40 are
disposed between the five stage axial compressor 30 and the single
stage centrifugal compressor 18. Compressor discharge pressure
(CDP) air 76 exits the impeller 32 and passes through a diffuser 42
annularly surrounding the impeller 32 and then through a deswirl
cascade 44 into a combustion chamber 45 within the combustor 52.
The combustion chamber 45 is surrounded by annular radially outer
and inner combustor casings 46, 47. Air 76 is conventionally mixed
with fuel provided by a plurality of fuel nozzles 48 and ignited
and combusted in an annular combustion zone 50 bounded by annular
radially outer and inner combustion liners 72, 73.
[0033] The combustion produces hot combustion gases 54 which flow
through the high pressure turbine 16 causing rotation of the high
pressure rotor 12 and continue downstream for further work
extraction in a low pressure turbine 78 and final exhaust as is
conventionally known. In the exemplary embodiment depicted herein,
the high pressure turbine 16 includes, in downstream serial flow
relationship, first and second high pressure turbine stages 55, 56
having first and second stage disks 60, 62. A high pressure shaft
64 of the high pressure rotor 12 connects the high pressure turbine
16 in rotational driving engagement to the impeller 32. A first
stage nozzle 66 is directly upstream of the first high pressure
turbine stage 55 and a second stage nozzle 68 is directly upstream
of the second high pressure turbine stage.
[0034] Referring to FIG. 1, the compressor discharge pressure (CDP)
air 76 is discharged from the impeller 32 of the centrifugal
compressor 18, used to combust fuel in the combustor 52, and to
cool components of turbine 16 subjected to the hot combustion gases
54; such as, the first stage nozzle 66, first and second stage
shrouds 71, 69 surrounding the first and second high pressure
turbine stages 55, 56 respectively. The high pressure compressor 14
includes a compressor aft casing 110 and a diffuser forward casing
114 as more fully illustrated in FIGS. 1 and 2. The compressor aft
casing 110 generally surrounds the axial compressor 30 and the
diffuser forward casing 114 generally surrounds the centrifugal
compressor 18 and supports the diffuser 42 directly downstream of
the centrifugal compressor 18. The compressor discharge pressure
(CDP) air 76 is discharged from the impeller 32 of the centrifugal
compressor 18 directly into the diffuser 42.
[0035] Referring to FIGS. 2 and 3, the impeller 32 includes a
plurality of centrifugal compressor blades 84 radially extending
from a rotor disc portion 82. Opposite and axially forward of the
compressor blades 84 is an annular blade tip shroud 90. The shroud
90 is adjacent to blade tips 86 of the compressor blades 84
defining a blade tip clearance 80 therebetween. The diffuser 42
includes an annular diffuser housing 20 having a plurality of
tangentially disposed diffuser flow passages 22 extending radially
therethrough, spaced about a circumference 26 of the housing 20,
and through which diffuser airflow 103 flows in a downstream
direction. Diffuser vanes 23 axially extend between a forward wall
101 and the aft wall 100 of the diffuser 42.
[0036] Referring to FIGS. 2 and 3, the diffuser vanes 23
circumferentially extend between adjacent ones of the diffuser flow
passages 22. The diffuser flow passages 22 are partly defined and
circumferentially bounded by the circumferentially spaced apart
diffuser vanes 23. Adjacent ones of the passages 22 intersect with
each other at radially inner inlet sections 24 of the passages 22
that define a quasi-vaneless annular inlet 27 of the diffuser 42.
Each passage 22 further includes a throat section 28 downstream of
and integral with the inner inlet section 24. Each passage 22
further includes a diffusing section 99 immediately downstream of
the throat section 28.
[0037] Referring to FIGS. 2 and 6, a centrifugal compressor first
cooling air source 92 for turbine cooling air 88 is a small
predetermined radial clearance (C) located between an impeller tip
36 of the rotating impeller 32 and the annular inlet 27 of the
static diffuser 42. Impeller bleed flow 102 from the radial
clearance (C) is collected in a radially inner manifold 104. The
predetermined. radial clearance (C) is designed to accommodate
thermal and mechanical growth of the impeller 32 and is open to or
in fluid communication with the radially inner manifold 104.
[0038] Referring to FIGS. 3-6, we have found that the diffuser
airflow 103 on one side of the passage (such as passage 22) in
multi-passage diffusers (such as the diffuser 42) that follow or
are downstream of centrifugal impellers (such as the impeller 32)
is often weak and may be subject to separation. Separation in the
passage can generate high losses that lowers engine specific fuel
consumption (SFC). This area or region of weak flow 127 is also
believed to be a contributor to surge that limits flow range of the
compressor.
[0039] A centrifugal compressor stage second cooling air source 94
for turbine cooling air 88 includes a diffuser boundary layer bleed
96 for bleeding diffuser bleed flow 112 from a diffuser boundary
layer 113 in each of the diffuser flow passages 22 of the diffuser
42, illustrated herein as plurality of boundary layer bleed
apertures 106. The diffuser boundary layer bleed 96, also referred
to as fluidic bleed, helps reduce the weak flow and limit or
prevent the unwanted flow separation. The diffuser boundary layer
bleed 96 bleeds diffuser bleed flow 112 from the diffuser boundary
layer 113 into a radially outer manifold 116.
[0040] The radially inner and outer manifolds 104, 116 are in fluid
communication such that the impeller bleed flow 102 from the
radially inner manifold 104 flows into the radially outer manifold
116. The impeller and diffuser bleed flows 102, 112 are mixed in
the radially outer manifold 116 to provide the turbine cooling air
88 which is then ported or otherwise flowed from radially outer
manifold 116 through a plurality of circumferentially distributed
manifold ports 117 to the high pressure turbine 16. The turbine
cooling air 88 may be channelled or flowed therefrom by external
piping (not shown) to cool the first and second stage shrouds 71,
69 (illustrated in FIG. 1).
[0041] Substantially axially extending beams or struts 122 separate
the radially inner and outer manifolds 104, 116 and the impeller
bleed flow 102 passes between the struts 122 as it flows from the
radially inner manifold 104 into the radially outer manifold 116.
The fluidic bleed flow illustrated herein as the diffuser boundary
layer bleed 96 represents a small amount of flow, less than 1% of
the engine core flow. The fluidic bleed is strategically removed
near the inception of the weak flow to improve the overall
performance of the diffuser.
[0042] Referring to FIGS. 3-5, the boundary layer bleed apertures
106 may be holes or slots 130 through the forward wall 101 of the
diffuser 42 as illustrated herein. The boundary layer bleed
apertures 106 or slots 130 lead into and are in flow communication
with the radially outer manifold 116. The slot 130 is positioned or
located downstream of the throat section 28 near a pressure side
126 of the diffuser vane 23 at a position where the flow would
begin to show weakness or instability in a diffuser without the
diffuser boundary layer bleed 96. This position is located in what
is referred to as a region of flow weakness 127. A slot width W may
be sized with manufacturing constraints such as a minimum tool
size. A slot length L may be selected to enable up to 3% of the
engine core flow to be used.
[0043] The slot 130 should ideally be angled such that the diffuser
bleed flow 112 exits the slot perpendicular to a forward surface
105 of the forward wall 101 of the diffuser 42 in a radial plane
132 passing through the engine centerline or axis 25 as illustrated
in FIG. 5. However, because of constraints such as the slot
extending through or very near a bend 134 in the forward wall 101
of the diffuser 42 this angle may be different. The slot 130 has
radially outer and inner walls 136, 138, as illustrated in FIG. 6,
and upstream and downstream facing walls 140, 142, as illustrated
in FIGS. 4 and 5 respectively, extending through the forward wall
101. The downstream facing wall 142 is designed to scoop boundary
layer air 144 in the diffuser boundary layer 113 only. Thus, the
downstream facing wall 142 is angled or canted at an acute cant
angle B of less than 90 degrees with respect to the diffuser
airflow 103 (parallel to the direction boundary layer air 144 in
the downstream direction in the diffuser flow passages 22 of the
diffuser 42. It appears that an acute cant angle B of 45 degrees is
desirable. However, the acute cant angle B is limited by geometry
and manufacturing constraints on the outside of the diffuser so
that an acute cant angle, for example about 22.5 degrees, is more
practical.
[0044] Illustrated in FIGS. 7-13 is a gas turbine engine with a
centrifugal compressor similar to the one illustrated in FIGS. 1-3
but with an alternative arrangement or design for separately
gathering and flowing the impeller tip bleed and diffuser bleed
flow for cooling turbine components. The impeller bleed flow 102
front the radial clearance (C), illustrated in FIG. 9, is flowed
into and collected in a radially inner annular manifold 154
illustrated in FIGS. 7 and 9. Inter-manifold apertures 160 are
disposed between the inner annular manifold 154 and a plurality of
radially outer annular manifolds 156 illustrated in FIGS. 7, 9, and
13. The inter-manifold apertures 160 allow the impeller bleed flow
102 to flow front the inner annular manifold 154 into the outer
annular manifolds 156. The impeller bleed flow 102 from the outer
annular manifolds 156 is then ported or otherwise flowed through a
plurality of circumferentially distributed impeller bleed flow
manifold ports 157, illustrated in FIG. 10, to the high pressure
turbine 16 for turbine cooling.
[0045] Referring to FIGS. 8, 10, and 11-13, the diffuser boundary
layer bleed 96 bleeds diffuser bleed flow 112 from the diffuser
boundary layer 113 into an annular diffuser bleed manifold 158 from
where the diffuser bleed flow 112 is then ported or otherwise
flowed through a plurality of circumferentially distributed
diffuser bleed manifold ports 159 to the high pressure turbine 16
for turbine cooling. FIG. 10 illustrates the relative
circumferential and axial locations of the impeller bleed flow
manifold ports 157 and the diffuser bleed manifold ports 159 on and
through the diffuser forward casing 114.
[0046] While there have been described herein what are considered
to be preferred and exemplary embodiments of the present invention,
other modifications of the invention shall be apparent to those
skilled in the art from the teachings herein and, it is therefore,
desired to be secured in the appended claims all such modifications
as fall within the true spirit and scope of the invention.
Accordingly, what is desired to be secured by Letters Patent of the
United States is the invention as defined and differentiated in the
following claims.
* * * * *