A COMBUSTION CHAMBER FOR A TURBINE ENGINE
The present invention relates to a combustion
chamber for a turbine engine such as an airplane
turboprop or turbojet.
Such a combustion chamber has coaxial walls forming
bodies of revolution that extend one inside the other and
that are connected together at their upstream ends by an
annular chamber end wall including air feed openings and
means for delivering fuel, in particular constituted by
injectors.
The inner and outer walls of the chamber have inlet
orifices for primary air and for dilution air, and zones
presenting multiple perforations for passing cooling air.
In order to withstand extreme temperatures better,
it is known to fit thermal barriers on the walls of the
combustion chamber, such barriers being in the form of
additional thicknesses of material applied against the
walls in question.
Document JP 61/67245 describes a combustion chamber
in which the inner wall presents thickness that is
constant and that is covered in a thermal barrier of
varying thickness.
The use of a thermal barrier increases the ability
of the chamber to withstand high temperatures, but also
increases its weight.
In order to satisfy market requirements, it is
necessary to reduce combustion chamber weight.
Nevertheless, the lifetime of the combustion chamber must
not be shortened. In particular, the walls must be
dimensioned in such a manner as to withstand being
damaged by creep. The term "creep" is used to designate
the irreversible deformation to which a material is
subjected by a constant stress that is applied for a
sufficient duration. This deformation is made worse by
the high temperatures to which the walls of a combustion
chamber are subjected.
A particular object of the invention is to provide a
solution to this problem that is simple, effective, and
inexpensive.
To this end, the invention provides a combustion
5 chamber for a turbine engine such as an airplane
turboprop or turbojet, the combustion chamber comprising
inner and outer annular walls forming bodies of
revolution that are connected together by an annular
chamber end wall, the combustion chamber being
10 characterized in that its inner wall is constituted by a
single thickness of material that presents thickness
and/or nature varying along the longitudinal axis and in
the circumferential direction of said wall, while its
outer annular wall presents thickness that is
15 substantially constant.
The invention makes it possible to increase the
ability of the combustion chamber to withstand extreme
temperatures without using thermal barriers and without
increasing weight, by locally modifying the thickness
20 and/or the nature of the material constituting the walls
of the chamber.
The outer annular wall is generally not as hot as
the inner annular wall, and therefore does not require
any particular adaptation of its structure.
25 In an embodiment of the invention, the inner wall of
the combustion chamber that is constituted by a single
thickness of material has at least one "hot" zone with a
large temperature gradient that is of greater thickness,
and at least one "cool" zone of smaller temperature
30 gradient that is of smaller thickness.
Since the "hot" zones are the zones that are
subjected to the greatest temperature gradients, it is
advantageous to increase their thickness.
According to another characteristic of the
35 invention, the inner wall of the combustion chamber that
is constituted as a single thickness of material presents
at least two adjacent zones made of different materials.
As mentioned above, in the zones having the highest
temperatures or the zones that are subjected to the
greatest temperature gradients, it is possible, locally,
to use a material that withstands those conditions
5 better, and in the zones that are of lower temperature or
that are subjected to smaller temperature gradients, it
is possible, locally, to use a material that withstands
those conditions less well, but that is lighter in
weight.
10 Preferably, the inner wall of the combustion chamber
of varying thickness is made by machining.
Machining makes it possible to obtain dimensional
tolerances that are smaller than those that can be
obtained by the sheet metal forming that is
15 conventionally used for making combustion chambers.
Furthermore, machining enables the thickness of the
inner wall to be varied both along the longitudinal axis
and in a circumferential direction.
In a variant, the inner wall of the combustion
20 chamber of varying thickness is made by stretching and
forming sheet metal.
This method is simpler and less expensive than
machining.
The zones of varying thickness and/or of varying
25 nature of the inner wall of the combustion chamber
comprise one zone forming part of the group comprising
the zones situated between the injectors, the zones
including primary air and dilution air holes, the zones
including annular fastening flanges, and the zones
30 including multiple perforations.
The invention also provides a turbine engine such as
an airplane .turboprop or turbojet, the engine including a
combustion chamber of the above-described type.
The inventi,on can be better understood and other
35 details, characteristics, and advantages of the invention
appear on reading the following description made by way
of non-limiting example with reference to the
accompanying drawings, in which:
Figure 1 is a diagrammatic half-view in axial
section of an annular combustion chamber of a turbine
5 engine;
Figure 2 is a perspective view of a sector of the
Figure 1 combustion chamber; and
Figure 3 is a detail view of the invention, in the
form of a section of the inner annular wall of the
10 Figure 1 combustion chamber.
As shown in Figures 1 and 2, an annular combustion
chamber 1 of a turbine engine is arranged at the outlet
from a diffuser 2, itself situated at the outlet from a
compressor (not shown), and comprises inner and outer
15 annular walls 3 and 4 constituting bodies of revolution
that are connected at the upstream end by an annular
chamber end wall 5 and that are fastened at their
downstream ends by inner and outer annular flanges 6 and
7 respectively to an inner frustoconical shroud 8 of the
20 diffuser 2 and to one end of an outer casing 9 of the
chamber 1, the upstream end of the casing 9 being
connected by an outer frustoconical shroud 10 to the
diffuser 2.
The annular chamber end wall 5 has openings 11
25 (Figure 2) through which air from the diffuser 2 passes
and that serve for mounting fuel injectors 12 that are
fastened to the outer casing 9 and that are regularly
distributed on a circumference around the longitudinal
axis of the chamber. Each injector 12 comprises a fuel
30 injector head 13 centered in an opening 11 in the annular
wall 5 and extending along the axis of the opening 11.
Some of the air flow delivered by the compressor and
leaving the diffuser 2 passes through the openings 11 and
feeds the combustion chamber, while the remainder of the
35 air flow is fed to inner and outer annular channels 14
and 15 that extend past the combustion chamber.
The inner channel 14 is formed between the inner
shroud 8 of the diffuser 2 and the inner wall 3 of the
chamber, and the air that passes along this channel is
shared between a flow that penetrates into the chamber
5 via orifices 16 and 17 for primary air and for dilution
air (Figure 2) in the inner wall 3, and a flow that
passes through holes in the inner flange 6 of the chamber
1 for the purpose of cooling components (not shown) that
are situated downstream from the chamber 1.
10 The outer channel 15 is formed between the outer
casing 9 and the outer wall 4 of the chamber 1, and the
air that passes along this channel is shared between a
flow that penetrates into the chamber via orifices 18 and
19 for primary air and for dilution air (Figure 2) in the
15 outer wall 4, and a flow that passes through holes in the
outer flange 7 in order to cool components situated
downstream.
The primary air inlet orifices 16 and 18 are
\ regularly distributed around the circumferences of the
20 inner and outer walls 3 and 4 respectively, being
centered on the axis of the chamber 1, and the dilution
air inlet orifices 17 and 19 are regularly distributed on
circumferences of the inner and outer walls 3 and 4
respectively, being centered on the axis of the chamber
25 1, downstream from the orifices 16 and 18.
The inner and outer annular walls 3 and 4 also
include microperforations (not shown) for passing cooling
air.
In operation, the outer and inner annular walls 4
30 and 3 present zones having different temperatures, this
temperature non-uniformity being represented
diagrammatically in Figure 2 in the form of zones 20, 21,
22, and 23 that are shaded differently from one another.
This phenomenon relates in particular to the inner
35 annular wall 3. These temperature zones are given
numbers that increase with increasing temperature. Thus,
the zones 20 are the zones that are relatively "cool"
being subjected to the smallest temperature gradients,
whereas the zone 23 is the zone that is the "hottest",
being subjected to the greatest temperature gradient.
This distribution of the zones is given purely by way of
5 example and results specifically from the particular
structure of the combustion chamber 1.
The presence and the locations of the various zones
20 to 23 can be revealed by simulation by computation or
by applying a paint that reacts to temperature so that
10 its color, after the combustion chamber has been in
operation, varies locally as a function of temperature.
According to the invention, the inner wall 3 is
constituted by a single thickness of material that
presents thickness and/or nature varying along the
15 longitudinal axis and/or along the circumferential
direction of said wall.
In the embodiment shown in the figures, the
thickness of the inner wall is caused to vary locally,
which wall has the zones 20 to 23 of different
20 temperatures.
Thus, as shown in Figure 3, the inner annular wall 3
is made as a single thickness of material and it has
zones of greater thickness el (see Figure 3), e.g. the
zones 22 and 23, and zones of smaller thickness e2, e.g.
25 the zones 20 and 21.
The thicker zones are those that are subjected in
operation to the highest temperature, e.g. temperatures
of the order of 1000"~. These zones present thickness el
lying in the range 1 millimeter (mm) to 2 mm, and
30 preferably being about 1.5 mm. Conversely, the thinner
zones are zones that, in operation, are subjected to
temperatures that are lower. These zones have a
thickness e2 lying in the range 0.5 mm to 1 mm, and they
are preferably about 1 mm thick.
35 The outer annular wall 4 presents thickness that is
substantially constant, lying in the range 1 mm to
1.5 mm, and preferably being about 1.2 rnm.
By way of example, it is thus possible, starting
from a prior art combustion chamber in which the walls
forming bodies of revolution present a constant thickness
of 1.5 mm, to produce a combustion chamber that is of
lighter weight, having an outer wall with a thickness of
1.2 mm and-an inner wall with a thickness of 1.5 mm in
hot zones and of 1 mm in cooler zones, the weight of this
chamber being equal to the weight of a chamber in which
all of the walls have a constant thickness equal to
1.2 mm.
The combustion chamber of the invention, and in
particular its inner wall 3 of varying thickness, is made
by machining.
Alternatively, the inner wall 3 of varying thickness
is made by stretching and forming sheet metal.
In an embodiment that is not shown in the drawings,
the zones of varying thickness may be replaced by or may
include zones of different natures, so as to comprise
zones made of material that withstand high temperatures
in the hottest zones, and zones made of a material that
does not withstand such high temperatures but that is
lighter in weight in the cooler zones.
Similarly, the zones of different natures may serve
to avoid cracks forming, with the material being changed
locally so that zones that are initially stressed in
traction, i.e. zones in which cracks may start, are
actually stressed in compression as a result of the
behavior of contiguous zones.
Each of these embodiments enable the weight of the
combustion chamber to be reduced while improving its
ability to withstanding high temperatures, and thus its
lifetime.
The zones of varying thickness and/or nature in the
inner wall 3 are, in particular, zones that are situated
between the injectors 12, the zones including the primary
air and dilution air holes 16 and 17, the zones including
the annular fastening flanges 6, and the zones including
multiple perforations.

CLAIMS
1. A combustion chamber (1) for a turbine engine such as
an airplane turboprop or turbojet, the combustion chamber
comprising inner and outer annular walls (3, 4) forming
5 bodies of revolution that are connected together by an
annular chamber end wall (5), the combustion chamber
being characterized in that its inner wall (3) is
constituted by a single thickness of material that
presents thickness (el, e2) and/or nature varying along
10 the longitudinal axis and in the circumferential
direction of said wall ( 3 ) , while its outer annular wall
(4) presents thickness that is substantially constant.
2. A combustion chamber (1) according to claim 1,
15 characterized in that its inner wall that is constituted
by a single thickness of material has at least one "hot"
zone (23, 22) with a large temperature gradient that is
of greater thickness (el), and at least one "cool" zone
(21, 20) of smaller temperature gradient that is of
20 smaller thickness (e2).
3. A combustion chamber (1) according to claim 1 or claim
2, characterized in that its inner wall that is
constituted as a single thickness of material presents at
25 least two adjacent zones made of different materials.
4. A combustion chamber (1) according to any one of
claims 1 to 3, characterized in that its inner wall of
varying thickness is made by machining.
30
5. A combustion chamber (1) according to any one of
claims 1 to 3, characterized in that its inner wall of
varying thickness is made by stretching and forming sheet
metal.
35
6. A combustion chamber (1) according to any one of
claims 1 to 5, characterized in that the zones of varying
thickness (el, e2) and/or of varying nature of its inner
walls comprise one zone forming part of the group
comprising the zones situated between the injectors (12),
the zones including primary air and dilution air holes
5 (16, 18; 17, 19), the zones including annular fastening
flanges (6, 7), and the zones including multiple
perforations.
7. A turbine engine such as an airplane turboprop or
10 turbojet, the engine including a combustion chamber (1)
according to any one of claims 1 to 6.
Dated this 07/06/2012
ATTORNEY