Stator of an aircraft turbine engine
TECHNICAL FIELD
The present invention relates to a stator of an aircraft turbine engine, and in particular
to a stator comprising at least one annular row of fixed vanes and one annular row of arms.
PRIOR ART
The prior art in particular includes WO-A 1-2005/119028 and EP-A2-0 942 150.
In general, an aircraft turbine engine comprises, from upstream to downstream, in the
· direction of flow of the gases, a fan, at least one compressor, an annular combustion
chamber, at least one turbine and a pipe for ejecting combustion gases.
For a bypass turbojet engine, the air flow that passes through the fan is divided into a
primary flow which supplies the engine and a secondary flow which flows around the engine.
Conventionally, the engine comprises at least one stator and at least one rotor. For a
multi-spool turbojet engine, the engine may for example comprise two rotors or spools, one
being low pressure and the other high pressure. The turbojet engine may thus comprise a
low-pressure spool having a first shaft connecting a low-pressure compressor to a lowpressure
turbine, and a high-pressure spool having a second shaft connecting a high-
20 pressure compressor to a high-pressure turbine.
The air entering the engine is compressed successively in the low-pressure
compressor and the high-pressure compressor before being mixed with fuel which is burned
in the combustion chamber. The combustion gases then expand in the high-pressure turbine
and then the low-pressure turbine in order to rotate the low-pressure shaft, which in turn
25 drives the fan shaft.
The stator of the turbine engine comprises structural annular casings, i.e. casings
which are stiff enough to transmit forces. A turbine engine in particular comprises an inlet
casing that extends downstream of the fan and an intermediate casing that extends between
the low-pressure and high-pressure compressors.
In general, each of these stator casings comprises generally structural arms that
extend substantially radially between two annular walls that are inner and outer, respectively,
these arms being tubular in order for auxiliary systems to pass therethrough from the inside
of the inner wall as far as the outside of the outer wall.
In order to reduce the consumption of the engines, some turbine engines comprise a
35 reduction gear. The stators of these turbine engines generally comprise an annular row of
fixed vanes upstream of the structural arms of the casing and an annular row of variablepitch
vanes downstream of the structural arms. The intermediate casing of this type of
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turbine engine, which is positioned downstream of the low-pressure compressor, thus
comprises an annular row of structural arms, also referred to as primary arms, inserted
between an annular row of fixed vanes (also referred to as inlet guide vanes (IGV)) and an
annular row of variable-pitch vanes (also referred to as variable stator vanes (VSV)). The
5 inlet casing, which is positioned upstream of the low-pressure compressor, also comprises
an annular row of structural arms inserted between an annular row of fixed vanes and an
annular row of variable-pitch vanes. ·
This type of stator having three successive rows of vanes and arms (two rows of
vanes and one row of arms) is very disadvantageous since it has a negative impact on
10 weight and drops in pressure due to its significant axial size.
The present invention provides a simple, effective and economical solution to this
problem.
SUMMARY OF THE INVENTION
The invention proposes a stator of an aircraft turbine engine, comprising an annular
15 row of fixed vanes and an annular row of arms, for example structural arms, said· stator
having an axis of revolution, characterised in that the trailing edges of the fixed vanes are
positioned substantially in a first plane that is transverse to said axis of revolution and is
positioned downstream of a second plane that is transverse to said axis of revolution and is
positioned downstream of a second plane that is transverse to said axis and passes
20 substantially through the leading edges of the arms, and the leading edges of the fixed
vanes are positioned substantially in a third plane that is transverse to said axis of
revolution and is positioned upstream of the second transverse plane, and in that it further
comprises an annular row of variable-pitch vanes that is positioned directly downstream of
the annular row of arms, the annular row of variable-pitch vanes comprising first variable-
25 pitch vanes that are positioned substantially in the extension of the arms, and second
variable-pitch vanes that are positioned between the first variable-pitch vanes, the first
variable-pitch vanes having different aerodynamic profiles to those of the second variablepitch
vanes.
The invention makes it possible to reduce the axial size of a stator comprising an
30 annular row of fixed vanes and an annular row of arms by axially interlinking these rows at
least in part. Therefore, by contrast with the prior art, in which the fixed vanes are
positioned upstream of the arms, in this case the fixed vanes are positioned between the
arms at least in part. The axial size of the stator can thus be considerably reduced
compared with those from the prior art, and this has a positive effect on the weight and the
35 drops in pressure through the stator.
Although the rows of fixed vanes and arms are arranged in a particular manner
according to the invention, the parameters and the aerodynamic properties of the profiles
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of the vanes and the arms are advantageously retained. The arms therefore preferably
retain their thicknesses, shapes and secondary functions for auxiliary systems to pass
therethrough. The fixed vanes may retain their function of guiding the air flow (which is
discharged from the fan of the turbine engine, for example). This makes it possible to retain
5 the speetl triangles upstream and downstream of the stator.
The invention makes it possible to improve the aerodynamic performance of the
stator in the different flight phases. Indeed, offsetting the leading edges of the fixed vanes
in the upstream direction relative to those of the arms means that it is easier for the flow to
pass into the fixed vanes. This also makes it possible to lessen the deflection of the flow at
10 the outlet of the fan by preventing said flow from being directly deflected by the arms.
Therefore, there is less of a drop in pressure, and the performance of the turbine engine is
improved thereby.
In the present application, "transverse plane" means a plane that is substantially
perpendicular to the longitudinal axis or the axis of revolution of the stator, which is
15 generally the longitudinal axis of the turbine engine.
Advantageously, the central longitudinal planes of the arms are inclined relative to
the longitudinal axis of the stator. The central longitudinal planes of at least some of the
arms and of at least some of the fixed vanes are preferably substantially parallel or slightly
inclined. This makes it easier to interlink the row of vanes with the row of arms.
20 The inclination or pitch of the arms makes it possible to further reduce the axial size
of the stator without reducing the actual chord of the arms, and therefore without modifying
the thickness thereof (the thickness/chord ratio as well as the number and distribution of
the arms can thus be retained in comparison with the prior art).
The stator further comprises an annular row of variable-pitch vanes that is positioned
25 directly downstream of the annular row of arms. As in the above-mentioned case, the stator
according to the invention thus comprises two rows of vanes and one row of arms.
The annular row of variable-pitch vanes comprises first variable-pitch vanes that are
positioned substantially in the extension of the arms, and second variable-pitch vanes that
are positioned between the first variable-pitch vanes. The first variable-pitch vanes have
30 different aerodynamic profiles to those of the second variable-pitch vanes. Advantageously,
the first variable-pitch vanes have aerodynamic profiles or curvatures that are more
aceentuated than those of the second variable-pitch vanes.
Interlinking the fixed vanes between the arms has an impact on the aerodynamic
performance of the stator because the speed triangle is not constant in azimuth
35 downstream of the arms. In order to overcome this drawback, the variable-pitch vanes
positioned downstream of the arms have profiles that are not all identical. Two different
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profiles are used depending on whether the vane is directly downstream of an arm or
downstream of a fixed vane.
The leading edges of the first variable-pitch vanes are preferably positioned as
closely as possible to the trailing edges of the arms so as to further reduce the axial size of
5 the stator, but also so that these first vanes act as "flaps" that ensure that the low-pressure
and high-pressure compressors operate under all flight conditions.
The stator may comprise two annular walls that are inner and outer, respectively,
between which the rows of fixed vanes and arms extend.
Each variable-pitch vane may comprise, at its outer radial end, a cylindrical pivot that
10 is mounted in a duct on the outer wall.
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The outer wall may comprise, upstream of the arms, an annular row of discharge air
through-slots and means, which are preferably controllable, for closing said slots. The slots
and the closing means form a discharge valve that is also referred to as a variable bleed
valve (VBV).
In a variant, the outer wall comprises, between the arms, at least one discharge air
through-opening arid at least one door, which is preferably controllable, for closing said
opening.
In a particular embodiment of the invention, the invention can be estimated to reduce
the weight of the module by 5 to 10 % using a casing of this type. This is mainly explained
20 by the reduction in axial length (and therefore the reduction in the inner and outer casing
walls in the primary and secondary flow), linked to the integration of the fixed vanes
between the arms, the reduction in axial clearances (in particular that between the rows of
arms and of variable-pitch vanes), and the pitch of the arms (which may be between 20
and 30°, which makes it possible to gain approximately 10% in chord, as opposed to oo in
25 the prior art).
An additional, unquantifiable benefit is the reduction in the weight of the pipelines and
harnesses that extend around the fan casing, by reducing the distances travelled.
The present invention also relates to an aircraft turbine engine, such as a turbojet
engine or turboprop engine, characterised in that it comprises at least one stator as
30 described above. If the turbine engine comprises a fan and low-pressure and high-pressure
compressors, the stator can be mounted between the fan and the low-pressure
compressor and/or between the low-pressure and high-pressure compressors.
Alternatively, when the turboprop engine comprises for example two contra-rotating
external propellers, positioned in the vicinity of the downstream end of the turboprop
35 engine in relation to the flow of the gases therein, the stator can be mounted at the inlet of
the low-pressure compressor.
DESCRIPTION OF THE FIGURES
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The invention will be better understood, and other details, features and advantages of
the invention will become apparent upon reading the following description, given by way of
non-limiting example and with reference to the accompanying drawings, in which:
- Fig. 1 is a schematic, axial sectional half view of an aircraft turbine engine according to the
5 prior art, viewed from the side,
-Fig. 2 is a highly schematic plan view of a part of the turbine engine from Fig. 1,
- Fig. 3 is a schematic, axial sectional half view of an aircraft turbine engine according to the
invention, viewed from the side,
- Fig. 4 is a highly schematic plan view of a part of the turbine engine from Fig. 3,
10 - Fig. 5 is a view corresponding to Fig. 3 that shows an embodiment of an inlet casing
according to the invention, and
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- Fig. 6 and 7 are views corresponding to Fig. 3 that show variants of an intermediate casing
according to the invention.
DETAILED DESCRIPTION
Reference is first made to Fig. 1, which shows a turbine engine 10 according to the
prior art, said turbine engine 10 in this case being a bypass turbojet engine. The invention is
described with reference to this example, but it is clear that the invention is applicable to
other turbine engine architectures.
The turbine engine 10 comprises, from upstream to downstream, in the direction of
20 flow of the gases, a fan 12 which generates a flow which divides into two coaxial flows, the
primary flow powering the engine which comprises a low-pressure compressor 14, a highpressure
compressor (not shown), a combustion chamber (not shown), high-pressure and
low-pressure turbines (not shown) and a pipe (not shown) for ejecting combustion gases.
These modules of the engine (fan, compressors, combustion chamber, turbines) are
25 surrounded by structural annular stator casings. The turbine engine 10 thus comprises a
plurality of successive annular casings, including an inlet casing 18 upstream of the lowpressure
compressor 14, and an intermediate casing 20 between the low-pressure and highpressure
compressors.
The inlet casing 18 comprises an annular row of structural arms 24 (or primary arms)
30 which is inserted between an annular row of fixed vanes 22 (or inlet guide vanes (IGV)) and
an annular row of variable-pitch vanes 26 (or variable stator vanes (VSV)).
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As can also be seen in Fig. 2, the row of fixed vanes 22 is positioned between the fan
12 and the row of arms 24, and the row of variable-pitch vanes 26 is positioned between the
row of arms 24 and a movable rotor wheel 28 of the low-pressure compressor 14.
In the same way, the intermediate casing 20 comprises an annular row of structural
arms (or primary arms) which is inserted between an annular row of fixed vanes (or inlet
guide vanes (IGV)) and an annular row of variable-pitch vanes (or variable stator vanes
6
0fSV)). The row of fixed vanes is positioned between a rotor wheel of the low-pressure
compressor and the row of arms, and the row of variable-pitch vanes is positioned between
the row of arms and a rotor wheel of the high-pressure compressor.
Reference numeral 34 in Fig. 2 represents the speed triangles of the air flow entering
5 the fan 12, and reference numerals 36 and 38 represent the speed triangles of the primary
air flow upstream of the fixed vanes 22 and downstream of the variable-pitch vanes 26.
As explained above, the inlet casing 18 and the intermediate casing 20 each form,
axially together with the rows of vanes 22, 26, a bulky stator.
The invention makes it possible to overcome this problem by axially interlinking the
I 0 row of fixed vanes with the row of arms of the stator.
Fig. 3 and 4 show an embodiment of a stator or inlet casing according to the
invention, this embodiment of course also being applicable to an intermediate casing.
P1 denotes a transverse plane (perpendicular to the longitudinal axis A or axis of
revolution of the stator 40 and of the turbine engine) passing substantially through the trailing
15 edges 42 of the fixed vanes 44, P2 denotes a transverse plane passing substantially through
the leading edges 46 of the arms 48, P3 denotes a transverse plane passing substantially
through the leading edges 50 of the fixed vanes 44, P4 denotes a transverse plane passing
substantially through the trailing edges 52 of the arms 48, and P5 denotes a transverse
plane passing substantially through the leading edges 54 of the variable-pitch vanes 56, 58.
20 In the example shown, P1 is downstream of P2 which is downstream of P3.
Moreover, P5 is downstream of P4 and at a short axial distance therefrom. The variablepitch
vanes 56, 58 are therefore positioned directly downstream of the arms 48.
In addition, H1 denotes a central longitudinal plane for each arm 48 and H2 denotes
a central longitudinal plane for each fixed vane 44.
25 In this case, the planes H1 and H2 are inclined relative to the axis A and are
substantially mutually parallel. In comparison with the prior art, the arms 48 therefore do not
have an axial orientation, but are by contrast "provided with a pitch" (at an angle of 20-30•,
for example) around an axis that is substantially radial relative to the axis A.
Each arm 48 is symmetrical to its plane H1. Each arm 48 is tubular in order to allow
30 auxiliary systems to pass therethrough, and in addition it may be structural or non-structural.
Each arm 48 preferably retains its dimensions, such as its chord and its thickness, in
comparison with the prior art shown in Fig. 1 and 2.
The fixed vanes 44 may be similar to those from the prior art.
The annular row