DESCRIPTION
TECHNICAL FIELD
The present invention relates to an annular combustion
chamber for turbine engine, especially for an aircraft propelling assembly.
The invention particularly but not exclusively applies to
combustion chambers fitted with an annular row of fuel injectors each
comprising a head provided with a central nose for injecting fuel and a peripheral
fuel injection device, for example of the multipoint type. Injectors of this type are
used in combustion chambers referred to as having “a staged lean combustion”.
The invention also relates to a combustion chamber module and
to a turbine engine comprising such a combustion chamber module.
STATE OF PRIOR ART
Figure 1 illustrates a typical example of a turbine engine 1 of a
known type, for example an aircraft twin spool turbofan engine.
The turbine engine 1 successively comprises, according to the
thrust direction represented by the arrow 2 also corresponding to the general
direction of gas flow in the turbine engine, a low pressure compressor 4, a high
pressure compressor 6, an annular combustion chamber 8, a high pressure
turbine 10 and a low pressure turbine 11.
In the following description, the upstream and downstream
directions are defined relative to the general direction of gas flow within the
combustion chamber and more generally of the turbine engine.
In a well-known manner, the combustion chamber 8 is mounted
downstream of the high pressure compressor 6 for supplying pressurized air to
this chamber, and upstream of the high pressure turbine 10 for rotating the high
pressure compressor 6 under the effect of gas thrust coming from the
combustion chamber.
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Figure 2 illustrates on a larger scale the combustion chamber 8
and its close environment.
The combustion chamber 8 comprises two respectively radially
inner 12, and radially outer 13 coaxial annular walls, which extend around the
longitudinal axis 14 of the combustion chamber.
These two annular walls 12 and 13 are fixed downstream to
inner 15 and outer 16 casings of the chamber, and are connected to each other at
their upstream end by an annular end wall 18 of the combustion chamber.
The annular end wall 18 includes an annular row of ports evenly
distributed around the axis 14 of the combustion chamber, and in which injection
systems 20 are mounted, in which respective heads 21 of fuel injectors 22 are
respectively mounted fitted. These fuel injectors 22 each have a fuel emission
axis which merges with an axis 24 of the corresponding injection system 20. The
injection systems 20 are mounted in the end wall 18 so as to be able to move
slightly along a direction orthogonal to the axis 24 and thus support the
differential expansions affecting the combustion chamber 8, the injection
systems 20 and the casings 15 and 16, in operation.
The assembly formed by the combustion chamber 8 and by the
fuel injectors 22 is referred to as a “combustion chamber module” in the present
description.
Each injection system 20 includes an upstream end forming a
bushing 26, a downstream end taking the shape of a flared bowl 28 opening into
the combustion chamber 8, and an annular air inlet 30 arranged between the
bushing 26 and the bowl 28 and for letting in part 31 of the airflow 32 coming
from a diffuser 34 mounted at the outlet of the high pressure compressor of the
turbine engine, so as to pre-mix the admitted air with the fuel coming from the
fuel injector 22 mounted in the bushing 26, within the injection system.
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In the illustrated example, the annular air inlet 30 is crossed by
fins 36 for imparting a rotary movement to the airflow which crosses them. The
air inlet is thus of the type commonly referred to as a “swirler”.
Furthermore, the annular walls 12 and 13 of the combustion
chamber are connected at their upstream end to an annular shroud 40 including
ports 42 arranged facing the injection systems 20 for passing fuel injectors 22 and
air 31 supplying the injection systems 20. The main functions of this shroud 40
are to protect the end wall 18 of the combustion chamber and to guide parts 44
and 46 of the airflow 32 which travel downstream respectively along the inner 12
and outer 13 annular walls of the combustion chamber, within two respectively
inner 48 and outer 50 bypass spaces. Hereinafter, these parts 44 and 46 of the
airflow 32 are respectively referred to as “inner by pass airflow” and “outer by
pass airflow”. The inner 48 and outer 50 by pass spaces form, together with an
upstream space 52 which connects them to each other, an enclosure in which the
combustion chamber 8 extends. Of course, each port 42 is located upstream of
the annular air inlet 30 relative to the axis 24 of the corresponding injection
system.
However, the air supply of the annular air inlet 30 of the
injection systems 20 has an inhomogeneous nature around the axis 24 of each
injection system, likely to induce a reduction in the performances of the
combustion chamber especially in terms of limiting emissions of pollutants and in
terms of controlling the thermal profile of the exhaust gases at the outlet of the
combustion chamber.
This problem is all the more significant in the case of
combustion chambers implementing a combustion mode referred to as “staged
lean combustion”, such as the combustion chamber of Figure 2. In this type of
combustion chamber, the head 21 of each injector 22 includes a central nose 54
for injecting fuel, an axial air intake device 56 arranged around the central
nose 54, and a peripheral fuel injection device 58 arranged around the axial air
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intake device 56. Fuel coming from the central nose 54 mixes, within each
injection system, with the air admitted through the axial air intake device 56,
whereas the fuel coming from the peripheral fuel injection device 58 mixes,
within each injection system, with the air admitted through the annular air
inlet 30 of the injection system. But, the injection systems of this type have
relatively large transverse overall dimensions, likely to increase the
inhomogeneities of the airflow 31. Moreover, these injection systems require a
relatively considerable airflow in order to operate, which tends to emphasize
heterogeneities in the combustion chamber 8.
The above-described problems are particularly significant in the
case of a combustion chamber arranged at the outlet of a centrifugal type
diffuser 34’, as illustrated in Figure 3.
In this case indeed, a radially outer part of the annular air
inlet 30 of each injection system 20 receives a direct airflow 31a whereas a
radially inner part of the annular inlet 30 only receives an indirect airflow 31b.
DISCLOSURE OF THE INVENTION
The object of the invention is especially to bring a simple
economical, and efficient solution to these problems, enabling the
abovementioned drawbacks to be at least in part avoided.
To do so, the invention provides an annular combustion
chamber for a turbine engine, comprising:
an annular end wall provided with a plurality of injection systems each
centred on a respective axis and each having an upstream end forming a bushing
for receiving a head of a fuel injector, a downstream end opening into said
combustion chamber, and an annular air inlet arranged between said upstream
and downstream ends so that the air admitted through said annular air inlet
mixes, within the injection system, with the fuel coming from the fuel injector,
and
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an annular shroud covering an upstream side of said end wall and
comprising a plurality of injector ports respectively arranged facing said injection
systems, said annular shroud and said end wall delimiting together an annular
space into which the annular air inlet of each injection system opens.
According to the invention, said annular shroud includes a
plurality of air intake ports separate from said injector ports.
Furthermore, said bushing of each of said injection systems
crosses the corresponding injector port of said annular shroud and comprises at
its upstream end an annular collar having a free end remote from said axis of the
injection system by a first distance greater than or equal to a second distance
separating a rim of said corresponding injector port from said axis of the injection
system.
The annular collar of the bushing of each injection system
enables the inlet of the corresponding injector port of the annular shroud to be
concealed and thus the airflow supplying the annular air inlet of the injection
system via said injector port to be reduced to substantially nothing. The supply of
the annular air inlet is thus nearly exclusively indirectly provided by air passing
through the air intake ports of the annular shroud.
This results in a better homogeneity of the air supply of the
annular air inlet of each injection system, as will appear more clearly hereinafter.
Moreover, this configuration makes it possible to maintain the
mobile nature of each injection system relative to the annular shroud and to the
end wall of the combustion chamber.
Preferably, said air intake ports and said injector ports are
distributed so that at least one air intake port is circumferentially arranged
between each pair of consecutive injector ports along the circumference of said
annular shroud.
Such a distribution enables the homogeneity of the air supply of
the annular air inlet of each injection system to be optimized.
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In this case, said air intake ports are preferably alternately
distributed with said injector ports along the circumference of said annular
shroud.
The invention also relates to a combustion chamber module for
a turbine engine, comprising:
an annular combustion chamber of the above-described type, and
an annular row of fuel injectors comprising respective injector heads
mounted fitted respectively in said bushings of the injection systems of said
combustion chamber.
The invention advantageously applies to such a combustion
chamber module, wherein each injector head includes a central nose for injecting
fuel, an axial air intake device arranged around said central nose, and a peripheral
fuel injection device arranged around said axial air intake device.
Furthermore, said injector ports of said annular shroud
advantageously have respective isobarycentres inscribed on a first circle centred
on an axis of said combustion chamber and having a first diameter.
In a first preferred embodiment of the invention, said air intake
ports of said annular shroud have respective isobarycentres inscribed on a second
circle centred on the axis of said combustion chamber and having a second
diameter strictly greater than said first diameter of said first circle.
In a second preferred embodiment of the invention, said air
intake ports of said annular shroud have respective isobarycentres inscribed on
said first circle.
The invention finally relates to a turbine engine for an aircraft,
comprising a combustion chamber module of the above-described type.
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BRIEF DESCRIPTION OF THE DRAWINGS
The invention will be better understood, and further details,
advantages and features of the invention will appear upon reading the following
description made by way of non-limiting example and with reference to the
accompanying drawings in which:
Figure 1, already described, is an axial cross-section partial schematic view
of a known type turbine engine;
Figure 2, already described, is an axial cross-section partial schematic view
of a combustion chamber module of the turbine engine of Figure 1, comprising an
axial diffuser;
Figure 3, already described, is an axial cross-section partial schematic view
of a combustion chamber module of a known type turbine engine, comprising a
centrifugal diffuser;
Figure 4 is an axial cross-section partial schematic view of a combustion
chamber module of a turbine engine according to a first preferred embodiment
of the invention;
Figure 5 is an axial cross-section partial schematic view of a combustion
chamber belonging to the combustion chamber module of Figure 4, shown
isolated;
Figure 6 is a partial schematic view of the combustion chamber module of
Figure 4, from upstream;
Figure 7 is a view similar to Figure 6, illustrating an alternative
embodiment of the combustion chamber module of Figure 4;
Figure 8 is a view similar to Figure 6, illustrating a combustion chamber
module of a turbine engine according to a second preferred embodiment of the
invention.
Throughout these figures, identical references can refer to
identical or analogous elements.
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DETAILED DISCLOSURE OF PREFERRED EMBODIMENTS
Figures 4 to 6 illustrate part of a combustion chamber
module 59 according to a first preferred embodiment of the invention. This
combustion chamber module is part of a turbine engine the other parts of which
can be of a conventional type, such as illustrated in above-described Figure 1.
Figures 4 to 6 more particularly show a rear part of the
combustion chamber 8 as well as the injectors 22 of the combustion chamber
module, whereas Figure 5 illustrates only the rear part of the combustion
chamber 8.
As appears in Figure 6, the annular shroud 40’ which covers the
upstream side of the combustion chamber 8 includes a plurality of air intake
ports 60 separate from the injector ports 42. In the illustrated example, the air
intake ports 60 are alternately distributed with the injector ports 42 along the
circumference of the annular shroud 40’. In a known manner per se, each injector
port 42 is located upstream of the annular air inlet 30 relative to the axis 24 of
the corresponding injection system.
As shown in Figures 4 and 5, the bushing 26’ of each injection
system 20 crosses the corresponding injector port 42 of the annular shroud 40’.
The bushing 26’ comprises at its upstream end an annular
collar 62. This annular collar 62 has a free end 64 remote from the axis 24 of the
injection system 20 by a first distance d1 (Figure 5) greater than or equal to a
second distance d2 separating a rim of said injector port 42 from the axis 24 of
the injection system.
In the illustrated example, the annular collar 62 does not have a
rotational symmetry. Indeed, the first distance d1 slightly varies around the
axis 24 of the injection system.
More precisely, a radially outer part of the annular collar 62 is
more extended than a radially inner part of the latter. Thus, in the axial section
plane of Figure 5, the radially outer side 66 of the free end 64 is more distant
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from the axis 24 of the injection system than the radially inner side 66’ of the free
end 64.
Similarly, the injector port 42 does not have a rotational
symmetry, so that the second distance d2 slightly varies around the axis 24 of the
injection system.
The above disparity between the first distance d1 and the
second distance d2 stands within each axial section plane of the combustion
chamber module.
Besides, in the illustrated example, the combustion chamber is
of the staged lean combustion type. Thus, each injector head includes a central
nose 54 for injecting fuel, an axial air intake device 56 arranged around said
central nose 54, and a peripheral fuel injection device 58 arranged around said
axial air intake device. This peripheral device 58 is for example of the “multipoint”
type, that is including an annular row of fuel ejection ports.
As shown in Figure 6, the injector ports 42 of the annular
shroud 40’ have respective isobarycentres 68 inscribed on a first circle 70 centred
on the axis 14 of the combustion chamber 8 and having a first diameter D1.
In the first embodiment of the invention, the air intake ports 60
of the annular shroud 40’ have respective isobarycentres 72 inscribed on a
second circle 74 centred on the axis 14 of the combustion chamber 8 and having
a second diameter D2 strictly greater than the first diameter D1 of said first
circle 70.
Thus, the air intake ports 60 are radially offset outwardly of the
annular shroud 40’. This configuration is particularly advantageous when the
diffuser supplying air to the combustion chamber is of the centrifugal type, as in
the prior art example illustrated in Figure 3.
In the example of Figure 6, the air intake ports 60 have an
oblong shape along the circumferential direction. Furthermore, each air intake
port 60 is remote from the first abovementionned circle 70.
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Alternatively, each air intake port 60 can extend up to the first
circle 70, as illustrated in Figure 7. In this case, the rim of each air intake port 60
advantageously has curved side regions so as to substantially follow the curve of
the injector port 42 located nearby.
In a second preferred embodiment of the invention illustrated
in Figure 8, the air intake ports 60 of the annular shroud 40’ have respective
isobarycentres 72 inscribed on the first circle 70.
This configuration is particularly advantageous when the
diffuser supplying air to the combustion chamber is of the axial type, as in the
prior art example illustrated in Figure 2.
In every case, in operation, the air supplying the annular air
inlet 30 of each injection system 20 exclusively or nearly exclusively passes
through the air intake ports 60 of the annular shroud 40’. Indeed, the annular
collar 62 of the bushing of each injection system 20 substantially prevents the
passage of air around each injection system through the corresponding injector
port 42. The annular collar 62 and the rim of the injector port 42 indeed form an
annular baffle for the airflow coming from the diffuser supplying the combustion
chamber with pressurized air.
For this reason, the air supplying the annular air inlet 30 of each
injection system 20 first travels by swirling within a space 78 (Figure 4) defined
between the annular end wall 18 and the annular shroud 40’ of the combustion
chamber.
This results in an improved homogeneity of the air supply of the
annular air inlet 30 around its respective axis.
In the above-described preferred embodiments, the injector
and air intake ports are distributed so as to alternate.
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As an alternative, other configurations are possible without
departing from the scope of the invention.
Generally speaking, to provide an optimum homogeneity of the
air supply of the annular air inlet 30 of each injection system, the air intake
ports 60 and the injector ports 42 are preferably distributed so that at least one
air intake port 60 is circumferentially arranged between each pair of consecutive
injector ports 42 along the circumference of the annular shroud 40’.
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We claim:
1. An annular combustion chamber (8) for a turbine engine,
comprising :
an annular end wall (18) provided with a plurality of injection systems (20)
each centred on a respective axis (24) and each having an upstream end forming
a bushing (26’) for receiving a head (21) of a fuel injector (22), a downstream
end (28) opening into said combustion chamber, and an annular air inlet (30)
arranged between said upstream and downstream ends so that the air admitted
through said annular air inlet (30) mixes, within the injection system, with the
fuel coming from the fuel injector (22), and
an annular shroud (40’) covering an upstream side of said end wall (18) and
comprising a plurality of injector ports (42) respectively arranged facing said
injection systems (20), said annular shroud (40’) and said end wall (18) delimiting
together an annular space (78) into which the annular air inlet (30) of each
injection system (20) opens,
said combustion chamber being characterized in that said annular shroud (40’)
includes a plurality of air intake ports (60) separate from said injector ports (42),
and in that said bushing (26’) of each of said injection systems crosses the
corresponding injector port (42) of said annular shroud and comprises, at its
upstream end, an annular collar (62) having a free end (64) remote from said
axis (24) of the injection system by a first distance (d1) greater than or equal to a
second distance (d2) separating a rim of said corresponding injector port (42)
from said axis (24) of the injection system.
2. The annular combustion chamber according to claim 1,
wherein said air intake ports (60) and said injector ports (42) are distributed so
that at least one air intake port (60) is circumferentially arranged between each
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pair of consecutive injector ports (42) along the circumference of said annular
shroud (40’).
3. The annular combustion chamber according to claim 2,
wherein said air intake ports (60) are alternately distributed with said injector
ports (42) along the circumference of said annular shroud (40’).
4. A combustion chamber module (59) for a turbine engine,
comprising:
an annular combustion chamber (8) according to any of claims 1 to 3, and
an annular row of fuel injectors (22) comprising respective injector
heads (21) mounted respectively fitted in said bushings (26’) of the injection
systems (20) of said combustion chamber.
5. The combustion chamber module according to claim 4,
wherein each injector head (21) includes a central nose (54) for injecting fuel, an
axial air intake device (56) arranged around said central nose, and a peripheral
fuel injection device (58) arranged around said axial air intake device.
6. The combustion chamber module according to claim 4
or 5, wherein said injector ports (42) of said annular shroud (40’) have respective
isobarycentres (68) inscribed on a first circle (70) centred on an axis (14) of said
combustion chamber (8) and having a first diameter (D1).
7. The combustion chamber module according to claim 6,
wherein said air intake ports (60) of said annular shroud (40’) have respective
isobarycentres (72) inscribed on a second circle (74) centred on the axis (14) of
said combustion chamber (8) and having a second diameter (D2) strictly greater
than said first diameter (D1) of said first circle (70).
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8. The combustion chamber module according to claim 6,
wherein said air intake ports (60) of said annular shroud (40’) have respective
isobarycentres (72) inscribed on said first circle (70).
9. A turbine engine for an aircraft, characterized in that it
comprises a combustion chamber module (59) according to any of claims 4 to 8.