Note: Descriptions are shown in the official language in which they were submitted.
COMBUSTOR FOR GAS TURBINE ENGINE
TECHNICAL FIELD
[0001] The
present application relates to gas turbine engines and to a combustor
thereof.
BACKGROUND OF THE ART
[0002] In
combustors of gas turbine engines, an efficient use of primary zone volume in
annular combustor is desired. An important component in improving the mixing
within the
primary zone of the combustor is creating high swirl, while minimizing the
amount of
components. It has been found however that high velocity outer annulus flow
produces low
local static pressure drop, and the inability to turn the flow to feed a row
of large dilution holes
at the inner and outer diameters of an annular combustor may result in poor
hole discharge
coefficient and low penetration angle of the air jets.
SUMMARY OF THE INVENTION
[0003] In
one aspect, the present invention provides at least an annular scoop ring on a
combustor liner defining a combustion chamber; the ring including a solid
radial inner portion
provided with bores defined in the ring and communicating with the combustion
chamber to
form air dilution inlets, and a radial outer portion in the form of a C-shaped
scoop open to
receive high velocity, annular air flow. The bores communicate with the scoop
to direct the air
into the combustion chamber wherein the bores form air jet nozzles to generate
jet
penetration and trajectory within the combustor.
[0004] In a
more specific embodiment the radial thickness of the inner portion of the
scoop ring must meet a minimum ratio of UD=1 where L is the axial length of
the bore and D
is the diameter of the bore.
[0005] In a
still more specific embodiment, the combustor is an annular combustor with
inner and outer liners and there is at least an annular scoop ring on each
inner and outer
liner.
[0006]
Further details of these arid other aspects of the present invention will be
apparent
from the detailed description and figures included below.
DESCRIPTION OF THE DRAWINGS
[0007]
Reference is now made to the accompanying figures depicting embodiments of
the present invention, in which:
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[0008] Fig. 1 is a schematic cross-sectional view of a turbofan gas
turbine engine;
[0009] Fig.
2 is a side cross-sectional view of a combustor assembly in accordance with
one embodiment;
[0010] Fig. 3 is a fragmentary perspective view of a detail shown in Fig. 2;
[0011] Fig. 4 is a fragmentary perspective view of another detail shown in
Fig. 2;
[0012] Fig. 5 is a schematic section view showing an axial length to diameter
ratio of a bore
of a scoop ring of the combustor of Fig. 2;
[0013] Figs. 6A and 6B are respectively outer radial and section views of a
scoop ring of
the combustor, with internal guide vanes; and
[0014] Figs. 7A and 7B are respectively outer radial and section views of a
scoop ring of
the combustor, with directional inlet holes.
DETAILED DESCRIPTION OF THE PREFERRED EMBODIMENTS
[0015] Fig.
1 illustrates a gas turbine engine 10 of a type preferably provided for use in
subsonic flight, comprising in serial flow communication a fan 12 through
which ambient air is
propelled, a compressor section 14 for pressurizing the air, a combustor 16 in
which the
compressed air is mixed with fuel and ignited for generating an annular stream
of hot
combustion gases, and a turbine section 18 for extracting energy from the
combustion gases.
[0016] The
combustor 16 is illustrated in Fig. 1 as being of the reverse-flow type;
however the skilled reader will appreciate that the description herein may be
applied to many
combustor types, such as straight-flow combustors, radial combustors, lean
combustors, and
other suitable annular combustor configurations. The combustor 16 has an
annual geometry
with an inner liner 20 and an outer liner 30 defining therebetween an annular
combustor
chamber in which fuel and air mix and combustion occurs. As shown in figure 2,
the upstream
end A of the combustor 16 may contain a manifold, fuel and air nozzles.
Downstream, is the
mixing channel B which includes channel walls 50 and 60 providing a narrow,
annular throat
favoring complete mixing of the fuel and air. The inner and outer liners 20
and 30 flare out,
downstream of the mixing channel B into the dilution zone C, within the
combustion zone.
[0017] The
liners 20 and 30 are provided with various patterns of cooling inlets
represented by the 27 in liner 20, for instance. Annular scoop rings 70 and 80
are provided as
integral to the liners 20 and 30 respectively. The scoop rings 70, 80 may also
be separately
fabricated and welded to the liners. Associated with annular rings 70 and 80
are patterns of
air diluting inlets 26, 36, respectively.
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[0018] Annular ring 80 will now be described in detail. Annular ring
70 is similar to
annular ring 80. Annular ring 80 includes a radially inner portion 82 in the
form of an annular,
solid block, i.e., having a greater thickness than the surrounding liner. A C-
Shaped or U-
shaped appendage extends radially outwardly from the inner block forming an
air scoop 84,
open to receive the annular flow air. The dilution air inlets 36 and cooling
inlets 37 are in the
form of bores extending through the solid block of the inner portion 82 and
communicating
with the combustion chamber. As described in the above mentioned US Patent
Application
Serial No. 13/795,089, the bores forming the inlets 36 and 37 will be oriented
individually at
predetermined directions, either at an angle to the radial axis, such as
tangential, acute or
obtuse depending on the penetration or swirl required of the air jets formed
by the bores
making up the inlets 36 and 37.
[0019] In order to ensure the formation of air jets by means of the
bores making up inlets
36, the radial thickness of the inner block portion 82 must be sufficient to
meet a minimum
ratio of L/D=1 where L is the axial length of the bore and D is the diameter
of the bore (as
shown in Fig. 5). The thickness of the inner block portion may be greater,
thus increasing the
bore length. The block portions may be integrally formed with the liner, or
attached thereto
(e.g., welding, etc).
[0020] The provision of the scoop portion 84 immediately adjacent
the inlets 36 captures
the dynamic head in the outer air flow to increase the inlet feed static
pressure and for a
better right angle turn into the inlets 36. The jet flow formed by the bores,
defining the inlets
36, result in improved discharge coefficient, higher pressure drop and deeper
jet penetration.
[0021] Referring to Fig. 4, dilution air inlets 36 are
circumferentially distributed on the
respective scoop ring 80, in the dilution zone C of the combustor 16.
According to an
embodiment, the dilution air inlets 26 and 36 are equidistantly distributed,
and opposite one
another across the combustion chamber. It is observed that the central axis of
one or more of
the bores forming the dilution air inlets 26 and 36, generally shown as D, may
have an axial
component and/or a tangential component, as opposed to being strictly radial.
Referring to
Fig. 4, the central axis D is oblique relative to a radial axis R of the
combustor 16, in a plane
in which lies a longitudinal axis X of the combustor 16. Hence, the axial
component DX of the
central axis D is oriented downstream, i.e., in the same direction as that of
the flow of the fuel
and air, whereby the central axis D leans towards a direction of flow (for
instance generally
parallel to the longitudinal axis X). In an embodiment, the central axis D
could lean against a
direction of the flow.
[0022] It should however be understood that the inlets 26 and 36 may
have both the
axial component DX and the tangential component DZ. The tangential component
DZ is
oblique relative to radial axis R in an axial plane, i.e., the axial plane
being defined as having
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the longitudinal axis X of the combustor 16 being normal to the axial plane.
In Fig. 4, the
tangential component DZ is in a counter clockwise direction.
[0023] Referring to Fig. 4, the plurality of cooling air inlets 27
may be defined in the inner
liner 20 and at least cooling air inlets 37 in the scoop ring 80 relative to
liner 30. The scoop
ring 80 has a set of dilution air inlets 37 in an alternating sequence with
the set of dilution air
inlets 36. The dilution air inlets 37 have a smaller diameter than that of the
dilution air inlets
36. This alternating sequence is a configuration considered to maximize the
volume of
dilution in a single circumferential ring.
[0024] The scoop portion 84, of the scoop ring 80, is open upstream
to the direction of
annular airflow, in other words, downstream relative to the direction of flow
within the
combustion chamber, while the scoop 74 of scoop ring 70 is open upstream to
the reverse
direction of annular airflow adjacent the liner 20, but upstream to the
direction of flow of fuel
and air within the combustion chamber. Hence, the scoop rings 70 and 80 face
opposite
directions, although they could face a similar direction as well. The shape of
the scoop
portion 74, 84 of the scoop ring 70, 80 may be of various open configurations
such as U-
shaped, C-shaped or other open shapes. The scoop portion 84 includes a forward
extending
lip 84a which may be designed at a selected angle and extension length to
optimize the air
entrance trajectory and the feed static pressure. For the purposes of this
description, the term
C-shape is meant to cover the various shapes. Slots 85 may be provided in the
scoop portion
84 to relieve any hoop stresses. Like slots may also be provided in the scoop
ring 70.
[0025] The openings to the diluting air inlets 26, 36 are located on
the inner surface of
the scoop portion 74, 84, near the bight of the C-shaped portion. The figures
show a single
row of inlets 26, 27, 36, 37, but multiple rows are considered as well.
Sectional dimensions
for the inlets 26, 27, 36, 37 may also vary. Referring to Fig. 5, one of the
scoop rings 70 and
80 is illustrated as having dimensions d, I and h, and angles a and 13 that
can be adjusted in
order to obtain the desired effect, for instance to optimize the entrance
trajectory and feed
static pressure in the case of angle 13-
[0026] Referring to Figs. 6A and 6B, internal guide vanes 90 may be
provided in the
scoop rings 70 and/or 80, to give tangential direction to the incoming flow,
hence providing
control of the tangential component of the air jet entering the combustor.
Alternatively, or
additionally, referring to Figs. 7A and 7B, directional inlet holes 100 may be
provided in the
scoop rings 70 and/or 80, for the same tangential component purpose. In the
case of
directional inlet holes 100, they are defined in a radial block 101 added in
the scoop rings 70
and/or 80.
[0027] The above description is meant to be exemplary only, and one
skilled in the art
will recognize that changes may be made to the embodiments described without
departing
from the scope of the invention disclosed. For instance, the annular scoop
rings 70, 80 may
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be present on the outer liner, on the inner liner, or in tandem, so as to
obtain the desired
mass flow rate and flow feature. Other modifications which fall within the
scope of the
present invention will be apparent to those skilled in the art, in light of a
review of this
disclosure, and such modifications are intended to fall within the appended
claims.
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