FR3130758A1 - PROPULSION ASSEMBLY FOR AN AIRCRAFT - Google Patents

Abstract

The propulsion assembly (1) for an aircraft comprises a nacelle (3) surrounding a multi-flow turbomachine (2) which comprises a gas generator (4), a propeller (16) accelerating an air flow (FO) in the nacelle (3), an annular element (18) between the generator (4) and the nacelle (3) defining a first (20) and a second vein (21), comprising a nozzle (19) separating the flow (F0) into an air flow (F1) in the first stream (20) and into an air flow (F2) in the second stream (21), the assembly (1) comprising a stator vane (22) mounted between the slat (19) and the propeller (16), and a stator vane (24) between the generator (4) and the annular element (18), mounted between the slat (19) and a rotor vane of a compressor (8) of the generator (4), or between the annular element (18) and the nacelle (3). Figure for abstract: Figure 7

Description

PROPULSION ASSEMBLY FOR AN AIRCRAFT

Technical field of the invention

The present invention relates to the general field of aeronautics. It relates more particularly to a propulsion assembly for an aircraft comprising a turbomachine with several streams and a nacelle. The invention also relates to an aircraft comprising such a propulsion assembly.

Technical background

Conventionally, a propulsion assembly comprises a nacelle surrounding a turbomachine which makes it possible to generate the thrust necessary for the propulsion of an aircraft, whether it is an airliner or a fighter, etc. For this purpose, the turbomachine successively comprises at least one compressor which compresses a flow of air entering the nacelle, a combustion chamber in which the previously compressed air is mixed with fuel and then ignited in order to generate a flow of hot gas propellant, and at least one turbine which is rotated by this flow of hot gas, the turbine being connected by a shaft to the compressor. These elements form the engine also called gas generator. The flow of hot gas then escapes through a nozzle at the outlet of the turbomachine. A rotor blade also called a fan is generally mounted upstream of the gas generator in order to accelerate the flow of primary air.

There are also multi-flow turbomachines in which an annular separator is mounted between the nacelle and the gas generator so as to separate the flow entering the nacelle into a primary air flow flowing into the gas generator and a flow of cold secondary air which circulates in the vein formed by the space between the nacelle and the separator. These turbomachines have the main advantage of being less fuel-consuming and less noisy.

The use of multi-flow turbomachines is characterized by their dilution rate which corresponds to the ratio of the mass of the secondary/tertiary flow to the mass of the primary flow. This dilution rate can also vary according to the flight phases of the aircraft, in particular in variable cycle turbomachines. Nevertheless, variations in the bypass ratio can lead to losses of secondary/tertiary flow and therefore to reductions in the efficiency and operability of the turbomachine.

The purpose of the present invention is to overcome this drawback by proposing an architecture allowing both the straightening of the air flows entering the turbomachine and the minimization of the impact of the changes in dilution rate on the gas generator.

To this end, the invention relates to a propulsion assembly for an aircraft, this propulsion assembly comprising a turbomachine with several flows and a nacelle which surrounds the turbomachine, said turbomachine comprising:
- a gas generator comprising at least one compressor, one combustion chamber and one turbine, said gas generator being arranged along a longitudinal axis,
- at least a first propeller mounted inside the nacelle and around the longitudinal axis and configured to accelerate an air inlet flow entering the nacelle,
- at least one annular element arranged radially between the gas generator and the nacelle and defining a first internal annular vein supplying the gas generator, and a second external annular vein with the nacelle, said annular element comprising upstream a first spout of seperation annular which is configured to separate said air inlet flow into a first air flow flowing in said first vein and a second air flow flowing in said second external annular vein,

The propulsion assembly being characterized in that it further comprises:
- a first stator blade extending radially between a casing of the gas generator and the nacelle, upstream of the first annular separation nose and downstream of said first propeller, and
- a second stator blade extending radially between a casing of the gas generator and the annular element, downstream of said first separation nozzle and upstream of a first rotor blade of said at least one compressor of the gas generator, and/or between the annular element and the nacelle, downstream from said first annular separation spout,

and in that at least one of said first and second stator vanes is variable-pitch or includes a variable-pitch portion.

Thus, thanks to the invention, the straightening of the air flows entering the nacelle is carried out upstream of the separator so that the blades present in the veins only have the function of protecting the turbomachine from changes in the bypass ratio. Such an architecture makes it possible to simplify the construction and assembly of the various blades present in limited spaces such as veins. It also allows more freedom on the positioning of the variable stator blading (which is entirely variable-pitch, or which only includes a variable-pitch portion and therefore another fixed portion) depending on the space available. For example, the space available is often very limited at the level of the first separation nose (because the thickness available in this area is necessarily lower), so it may be more interesting to put it in the basket.

The propulsion assembly may also have one or more of the following characteristics, taken alone or in combination with each other:
- the first stator blading is variable-pitch,
- said second outer annular vein is devoid of stator blading from said first annular separation nozzle to a plane perpendicular to said longitudinal axis and passing substantially through a first stator blading of said at least one compressor of the gas generator,

-- In particular in the above configuration, when the first stator vane is variable-pitch, the second stator vane can be entirely fixed and not be variable-pitch; in this configuration, it is not necessarily necessary to have two consecutive variable-pitch stator blades,
- said second outer annular vein comprises a third stator blade downstream of said first annular separation nozzle,

-- In particular in the above configuration, when the first stator vane is variable-pitch and the second stator vane is entirely fixed, the third stator vane may comprise a variable-pitch portion and therefore a fixed portion; also in this configuration, it is not necessarily necessary to have two consecutive variable-pitch stator blades,
- the third stator blade is located downstream or to the right of the leading edges of the blades of said second stator blade, and upstream or to the right of the leading edges of the blades of a first stator blade of said at least one gas generator compressor,
- the second stator blade extends radially between the annular element and the nacelle, and said first internal annular vein has no stator blade upstream of a first rotor blade of said at least one compressor of the gas generator ,

-- in particular in the above configuration, the second stator blade is preferably variable-pitch,
- the second stator blade is located downstream or to the right of the leading edges of the blades of said first rotor blade of said at least one compressor of the gas generator, and upstream or to the right of the trailing edges of the blades of a first stator vane of said at least one compressor,
- the first stator blading and/or said second stator blading comprises blades, an upstream portion of which comprises a leading edge which can rotate around a substantially radial axis, and a downstream portion of which comprises a fixed trailing edge,

-- the first stator blading and/or the second stator blading comprises blades, a downstream portion of which comprises a trailing edge mobile in rotation about a substantially radial axis, and an upstream portion comprises a fixed trailing edge, and
- Said at least one first propeller and said first rotor blade are connected to the same shaft, preferably via a mechanical speed reducer.

The present invention also relates to an aircraft, in particular a transport aircraft, comprising a propulsion unit such as that mentioned above.

Brief description of figures

Other characteristics and advantages of the invention will appear during the reading of the detailed description which will follow for the understanding of which reference will be made to the appended drawings in which:

there shows a schematic longitudinal section of a propulsion assembly comprising a turbofan engine;

there shows a double longitudinal schematic section of a civil-type propulsion system comprising a bypass turbomachine;
there represents a double longitudinal schematic section of a propulsion assembly, comprising a turbofan engine;

[Fig. 4] FIG. 4 represents a schematic double longitudinal section of a propulsion assembly comprising a turbomachine with several streams;

there represents a double longitudinal schematic section of a propulsion assembly according to a first embodiment of the invention;

there shows a schematic double longitudinal section of a propulsion assembly according to a variant of this first embodiment of the invention;

there represents a double longitudinal schematic section of a propulsion assembly according to a second embodiment of the invention;

there represents a double longitudinal schematic section of a propulsion assembly according to a variant of the second embodiment of the invention; And

there shows a schematic radial section of a variable-pitch stator vane according to the second embodiment of the invention.

Claims (10)

Propulsion assembly for an aircraft, this propulsion assembly (1) comprising a turbomachine (2) with several flows and a nacelle (3) which surrounds the turbomachine (2), the said turbomachine (2) comprising:
- a gas generator (4) comprising at least a compressor (8), a combustion chamber and a turbine (7), said gas generator (4) being arranged along a longitudinal axis (X-X),
- at least a first propeller (16) mounted inside the nacelle (3) and around the longitudinal axis (X-X) and configured to accelerate an air intake flow (FO) entering the nacelle ( 3),
- at least one annular element (18) arranged radially between the gas generator (4) and the nacelle (3) and defining a first inner annular vein (20) for supplying the gas generator (4), and a second vein (21) external annular with the nacelle (3), said annular element (18) comprising upstream a first spout (19) separating annular which is configured to separate said air inlet flow (F0) into a first air flow (F1) flowing in said first vein (20) and into a second air flow (F2) s' flowing in said second external annular vein (21),
characterized in that it further comprises:
- a first stator blade (22) extending radially between a casing (5B) of the gas generator (4) and the nacelle (3), upstream of the first nozzle (19) for annular separation and downstream of said first propeller (16), and
- a second stator blading (24) extending radially between a casing (5B) of the gas generator (4) and the annular element (18), downstream of said first separating nose (19) and upstream of a first rotor blade (9) of said at least one compressor (8) of the gas generator (4), and/or between the annular element (18) and the nacelle (3), downstream of said first nozzle (19) annular separation,
and in that at least one of said first (22) and second (24) stator vanes is variable-pitch or includes a variable-pitch portion. A propulsion assembly (1) according to claim 1, wherein said second stator vane (24) extends radially between the gas generator (4) and the annular element (18), and said second outer annular vein (21) is devoid of stator blading from said first annular separating nose (19) to a plane perpendicular (P) to said longitudinal axis (X-X) and passing substantially through a first stator blading (22) of said at least one compressor ( 8) of the gas generator (4). A propulsion assembly (1) according to claim 1, wherein said second stator vane (24) extends radially between the gas generator (4) and the annular element (18), and said second outer annular vein (21) includes a third stator vane (26) downstream of said first annular separating nose (19). Propellant assembly (1) according to the preceding claim, in which said third stator vane (26) is located downstream or in line with the leading edges (25C) of the blades (25) of said second stator vane (24), and upstream or to the right of the leading edges (9A) of the blades of a first blade (9) of the stator of said at least one compressor (8) of the gas generator (4). A propulsion assembly (1) according to claim 1, wherein said second stator vane (24) extends radially between the annular element (18) and the nacelle (3), and said first internal annular vein (20) is devoid of stator vane upstream of a first rotor vane (9) of said at least one compressor (8) of the gas generator (4). Propellant assembly (1) according to the preceding claim, in which said second stator vane (24) is located downstream or at right angles to the leading edges (9A) of the blades of said first rotor vane (9) of said at least one compressor (8) of the gas generator (4), and upstream or in line with the trailing edges (10B) of the blades of a first stator vane (10) of said at least one compressor (8). Propulsion assembly (1) according to claim 5 or 6, in which said first stator vane (22) and/or said second stator vane (24) comprises blades (23, 25) of which an upstream portion (28) comprising a leading edge (23C, 25C) is rotatable around a substantially radial axis, and of which a downstream portion (29) comprising a trailing edge (23D, 25D) is fixed. Propellant assembly (1) according to any one of the preceding claims, in which the said at least one first propeller (16) and the said first rotor blade (9) are connected to the same shaft (S), preferably via a mechanical speed reducer. Propellant assembly (1) according to any one of the preceding claims, in which it comprises a second propeller (30) mounted upstream of the nacelle (3) and around the longitudinal axis (X-X) and configured to accelerate a main stream (FP), said nacelle (3) comprising upstream a second annular splitter nozzle (31) which is configured to separate said main air flow (FP) into said air inlet flow (F0) flowing in the nacelle (3), and in a third flow of air (F3) flowing around the nacelle (3). Aircraft comprising at least one propulsion assembly (1) according to any one of the preceding claims.