FR3102148A1 - PROPULSION SYSTEM FOR AN AIRCRAFT - Google Patents

Abstract

The invention relates to a propulsion system (1, 1 ') for an aircraft, comprising a rotor (2) and a nacelle fairing (3) extending around said rotor with respect to an axis (X) and comprising a downstream section (20), a downstream end (21) of which forms an outlet section (BF) of the nacelle fairing (3); and characterized in that the downstream section (20) comprises radially inner (20a) and outer (20b) walls, at least part of one of said walls of which is made of a thermo-deformable shape memory material, this wall being provided with heating elements (23a, 23c) extending in different consecutive angular sectors around said axis (X), each heating element being operable independently and configured to deform this wall in a direction radial to said axis (X) and angularly centered with respect to its angular sector, under the effect of a predetermined tension control. Figure for abstract: Figure 3

Description

PROPULSION SYSTEM FOR AN AIRCRAFT

The present invention relates to the field of propulsion systems for aircraft. It relates in particular to a propulsion system capable of providing a lateral thrust component.

The present invention also relates to an aircraft comprising such a propulsion system.

A propulsion system for an aircraft comprises at least one rotor or one propeller comprising a plurality of blades mounted on a rotating shaft.

There are aircraft, and in particular Vertical Take-Off and Landing Aircraft (ADAV or VTOL acronym for " Vertical Take-Off and Landing " in English), having propulsion systems with simple rotors when they only have a single rotor or counter-rotating when they comprise rotors grouped in pairs rotating in opposite directions.

These propulsion systems are either with ducted rotors (the rotor is then surrounded by an annular nacelle fairing), or with free rotors, the propulsion systems and in particular the rotors (free or ducted) being able to be mounted on a pivot shaft allowing the orientation of the propulsion systems and therefore of the rotors between a vertical position and a horizontal position, for example the vertical orientation for a vertical take-off or landing and the horizontal orientation for forward flight or airplane mode .

Ducted rotors have several attractive advantages, such as:

• a significant decrease in the sound signature of the rotor in direct emission;
• protection of the rotor blades from surrounding obstacles;
• improved rotor performance, particularly when the aircraft is hovering or at low forward speed.

However, VTOL aircraft with ducted rotors, especially in hovering flight, do not have the same maneuverability as conventional helicopters.

For conventional helicopters, altitude and placement corrections can be made, in the reference trihedron in aeronautical standards X, Y, Z, by:

• corrections around the roll axis X (axis of rotation of the helicopter around its longitudinal axis) by an action to the left or to the right on the cyclic pitch (acting on the orientation of the thrust of the rotor, here in a plane perpendicular to the X axis);
• corrections around the pitch axis Y (axis of the helicopter around its transverse axis) by an action forward (to nose down) or backward (to nose up) on the cyclic pitch (acting on the orientation of the rotor thrust, here in a plane perpendicular to the Y axis);
• corrections around the yaw axis Z (axis of rotation in a horizontal plane of the helicopter around its vertical axis) by action on the rudder (acting on the thrust of the helicopter's anti-torque rotor);

• corrections in longitudinal translation along the X axis of the reference trihedron in aeronautical standards (longitudinal axis) by a control of the cyclic/collective pitch (acting on the orientation of the thrust of the rotor, here forwards or backwards back) ;
• corrections in lateral translation along the Y axis of the reference trihedron in aeronautical standards (transverse axis) by a command of the cyclic/collective pitch, (acting on the orientation of the thrust of the rotor, here to the right or to the left );
• corrections in vertical translation along the reference Z axis by a command of the collective pitch up, down (acting on the thrust of the rotor, pitch setting: command pulled up).

Of course, coordinated collective pitch and rudder pedal corrections, well known to pilots, must be made to counter the induced effects of loss of lift component and torque effect.

For VTOL aircraft with ducted rotors, positioning corrections, or corrections to counter gusts of wind in hovering flight, are generally complex and make the flight uncomfortable for the passengers of the aircraft.

For example, in the case of a quadrotor (i.e. an aircraft with four rotors), in order to make a correction in the direction of the reference X axis, it is necessary to tilt the aircraft forward around of the reference Y axis (according to the pitch axis), then to make a correction of the position reached (for example to counter a gust of wind) and to quickly return to a neutral position, called "flat" . Similarly, in order to make a correction in the direction of the reference Y axis, it is necessary to bank, i.e. to roll, the aircraft towards and around the X axis of reference (along the roll axis), then to correct the position reached and quickly return to a neutral position. With such a quadrotor, when transporting passengers, the latter may feel forward/backward and/or right/left swings.

A gas turbine engine has been proposed comprising a two-dimensional nozzle comprising a flexible panel capable of changing position, under the action of a jack, so as to regulate the exhaust from the engine.

Gas turbine engines have also been proposed comprising nozzles with variable geometries, so as to vary an outlet surface of the nozzle.

Thus, these proposed solutions propose nozzles in which the two-dimensional vector thrust technology is implemented.

However, none of these proposed solutions proposes the installation of three-dimensional vector thrust technology in the fairings of nacelles of propulsion systems, in particular in VTOL mode with rotors, guiding the secondary flow coming from said rotors.

There is therefore a need to provide a simple and effective solution to the problems mentioned above.

An object of the present invention is to propose a solution making it possible to improve the maneuverability of aircraft of the VTOL type, while reducing the impacts of mass and aerodynamic losses associated with the flight controls.

In particular, the invention proposes to improve the maneuverability of VTOL-type aircraft by providing a propulsion system capable of providing a thrust component laterally to the axis of the rotors.

To this end, the invention relates to a propulsion system for an aircraft, comprising at least one rotor and a nacelle fairing extending around said at least one rotor with respect to an axis of rotation of said rotor, this nacelle fairing comprising a downstream section of which a downstream end forms an exit section of the nacelle fairing; and characterized in that the downstream section comprises a radially internal wall and a radially external wall, at least a part of one of said walls being made of a thermo-deformable shape memory material, and in that said wall is provided of a plurality of heating elements, the heating elements extending in different consecutive angular sectors around said axis of rotation, each heating element being operable independently of the other heating elements and being configured to deform said wall in a direction which is radial to the axis of rotation and which is angularly centered with respect to the angular sector over which it extends, under the effect of a predetermined voltage command.

The propulsion system according to the invention can be with a single rotor or with counter-rotating rotors, installed in a fixed or pivoting nacelle, with a through or offset pivot axis.

According to the invention, the fairing is made at its air outlet of a thermo-deformable shape memory material and of a plurality of heating elements configured to automatically vary the shape of the nacelle and from there, the orientation of the air flow at the outlet thereof, in order to maneuver the aircraft on which said nacelle is installed. In particular, the heating elements allow the modification of the geometry of the radially outer wall or of the radially inner wall in a continuous manner as a function of the current passing through them (i.e. of the predetermined voltage command) and acting on the local retraction of the material at shape memory.

The profile of the fairing advantageously has a semi-rigid downstream part whose dimensions and shape of the exit section at the trailing edge can vary, so that the exit section at the trailing edge can be oriented laterally under the effect of a controlled device, to form a current tube producing a thrust with a lateral component, called vectored or oriented thrust.

The propulsion system according to the invention thus makes it possible to improve the maneuverability of the aircraft in which it is installed, in particular during maneuvers at low forward speed, such as take-offs and landings, while minimizing the noise pollution induced by the rotor of the propulsion system and by ensuring safety of this rotor by the presence of the nacelle fairing.

Indeed, the shape of the fairing is continuously adapted according to the deviation of the thrust commanded to obtain a precise placement of the aircraft. The fairing profile can therefore be oriented laterally according to the mechanical flight constraints sought in VTOL mode, so as to produce a thrust component perpendicular to the rotor axis.

The heating element system is simple, compact, and has low impacts in terms of mass compared to the solutions proposed in the prior art.

During flight phases in airplane mode, the heating element system, which is useful during phases at low forward speed, can be deactivated.

Each heating element may have the general shape of an annular portion. The heating elements can extend in different consecutive angular sectors around said axis of rotation so that the set of heating elements has a generally annular shape.

According to a first embodiment, at least part of the radially outer wall is made of a heat-deformable, and more precisely heat-shrinkable, shape-memory material.

According to a second embodiment, at least a part of the radially inner wall is made of a thermo-deformable, and more precisely thermo-extensible, shape-memory material.

According to a third embodiment, at least part of the radially outer wall is made of a heat-deformable, and more precisely heat-shrinkable, shape-memory material, and at least part of the radially inner wall is made of a thermo-deformable, and more precisely thermo-extensible, shape-memory material.

According to one embodiment, the plurality of heating elements include heating coatings.

According to another embodiment, the plurality of heating elements comprises heating resistors.

According to one embodiment, the wall made of a thermo-deformable shape memory material comprises at least one heating circuit configured to act on the heating elements.

According to another embodiment, the wall made of a thermo-deformable shape memory material comprises a plurality of heating circuits, each heating circuit being configured to act on at least one of the heating elements. This advantageously makes it possible to obtain a gradual heating of the heating elements, and thus a gradual deformation of the radially internal and radially external walls of the downstream section.

The heating circuits can operate independently of each other.

Each of the radially inner and radially outer walls can be provided with a plurality of heating elements, heated by heating circuits. The operation of the heating circuit(s) associated with the radially inner wall can be independent of the heating circuit(s) associated with the radially outer wall.

Advantageously, the propulsion system further comprises stiffening bridges connecting the radially inner and radially outer walls of the downstream section and making it possible to ensure a substantially constant gap between the radially inner and radially outer walls of the downstream section.

The nacelle fairing may comprise an upstream section forming an inlet section of the nacelle fairing and an intermediate section connecting the upstream and downstream sections.

Advantageously, the intermediate section is rigid and is connected by at least one mast to an engine of the propulsion system. This provides the nacelle fairing of the propulsion system with a rigid structure capable of performing a shielding function.

The present invention also relates to an aircraft characterized in that it comprises at least one propulsion system having at least any one of the aforementioned characteristics, the propulsion system being pivotally mounted on the aircraft by means of a pivot shaft offset or crossing with respect to the rotor.

The present invention will be better understood and other details, characteristics and advantages of the present invention will appear more clearly on reading the description of a non-limiting example which follows, with reference to the accompanying drawings in which:

FIG. 1A is a schematic perspective view of a first embodiment of a propulsion system shown with a nacelle mounted on an offset pivot axis, the propulsion system being in a horizontal position;

FIG. 1B is a view similar to FIG. 1A, illustrating the propulsion system in a vertical position;

FIG. 1C is a schematic perspective view of a second embodiment of a propulsion system shown with a nacelle mounted on a traversing pivot axis, the propulsion system being in a horizontal position;

FIG. 2 is a diagrammatic view in longitudinal section of the propulsion system according to the invention with its downstream section of the nacelle fairing in axial thrust mode;

FIGS. 3 and 4 are views similar to FIG. 2 of the propulsion system according to the invention with its downstream section of the nacelle fairing in asymmetrical thrust mode;

FIG. 5 is a schematic rear view of an embodiment of the heating elements according to the invention, in axial thrust mode;

FIGS. 6 and 7 are schematic rear views of an embodiment of the heating elements according to the invention, in asymmetrical thrust mode.

The elements having the same functions in the different implementations have the same references in the figures.

In this disclosure, the terms "axial", "internal" and "external" are used in reference to the axis of rotation of the propulsion system according to the invention.

A propulsion system generally consists of:

• a basket;
• an engine and its command and control system;
• and, in the case of propeller or rotor propulsion, its propeller or rotor(s).

The nacelle is the element that allows the engine to be integrated into the aircraft, it consists of:

• nacelle fairings (allowing the engine to be cowled, the rotors to be shrouded, the flowing air to be captured when the aircraft is operating, to create a thrust effect, to reverse the thrust on the propulsion systems, etc.);
• equipment to be mounted on the engine (such as the engine casing comprising the electrical, hydraulic and pneumatic networks known by the acronym EBU for “ Engine Build-Up ”); And
• aircraft attachment systems.

Figures 1A and 1B illustrate, in a simplified manner, a first embodiment of a propulsion system 1 for an aircraft according to the invention.

The propulsion system 1 here comprises at least one rotor 2 and a nacelle fairing 3 extending around said at least one rotor 2 with respect to an axis of rotation X of the rotor 2. The fairing 3 advantageously serves, by its shape and its materials, as an acoustic screen. The propulsion system 1 can be fixed mounted on the aircraft. The propulsion system 1 can still be mounted on a pivot shaft 4, offset with respect to the axis of rotation X of the rotor 2. The pivot shaft 4 is fixed by any means to the propulsion system 1, on the one hand, and to the aircraft, on the other hand, and allows the orientation of the propulsion system 1 on the aircraft, authorizing the tilting of the propulsion system 1 around the pivot shaft 4, according to the arrow F1, via known actuators, between a horizontal position as shown in Figure 1A, and a vertical position as shown in Figure 1B. This switchover makes it possible to switch the aircraft from a classic mode as for an airplane, to a VTOL or helicopter mode.

The rotor 2 of the propulsion system 1 is connected to the aircraft by a mast 5 supporting a motor 6, for example electric, driving the rotor 2 in rotation via a power shaft, in a manner known per se. According to the non-limiting example illustrated, each rotor 2 comprises two blades 7.

Figure 1C illustrates a second embodiment of a propulsion system 1' for an aircraft according to the invention in which the propulsion system 1' can be mounted on a pivot shaft 4', passing through the rotor 2 perpendicularly with respect to the axis of rotation X of the rotor 2. The rotor 2 of the propulsion system 1' is connected to the aircraft by a mast 5 supporting a motor 6, for example electric, driving the rotor 2 in rotation via a power shaft, in a manner known per se. According to the embodiment shown, the mast 5 of the rotor 2 coincides with the pivot shaft 4'.

With reference to Figures 2 to 4, the nacelle fairing 3 of the propulsion system 1, 1' according to the invention comprises:

• an upstream section 10;
• a downstream section 20; And
• an intermediate section 30 connecting said upstream 10 and downstream 20 sections.

The upstream section 10 forms an entry section (or in other words a leading edge) BA or air inlet of the fairing 3 of the nacelle.

The upstream section 10 is made of a material that can withstand temperatures making it suitable for providing an anti-icing function when it is supplied with hot air.

The intermediate section 30 is rigid. For example, it is made of aluminum alloy, titanium filled with 6% aluminum and 4% vanadium (TA6V), or carbon fiber composite with an organic matrix. The intermediate section 30 is advantageously connected to the engine 6 of the propulsion system 1, 1 'by at least one mast 5, and, preferably, by two masts 5 so as to mechanically secure the fairing 3 of the nacelle to the engine 6 of the propulsion system 1 , 1'. The intermediate section 30 thus confers, by its material and its configuration, a shielding function on the propulsion system 1.1'.

A downstream end 21 of the downstream section 20 forms an outlet section (or in other words a trailing edge) BF or an air outlet from the fairing 3 of the nacelle.

The downstream section 20 comprises a radially inner wall 20a and a radially outer wall 20b. The radially inner 20a and radially outer 20b walls of the downstream section 20 provide not only a structuring function of the downstream section 20 but also an aerodynamic function.

The radially inner 20a and radially outer 20b walls of the downstream section 20 are made of a semi-rigid deformable material with shape memory. In other words, the material constituting the radially inner 20a and radially outer 20b walls of the downstream section 20 is both rigid to give the downstream section 20 a structuring shape and flexible to give the downstream section 20 a possibility of deformation. Thus, the radially inner 20a and radially outer 20b walls of the downstream section 20 are made of a material capable of reacting under the effect of actuators as described below. When the radially inner 20a and radially outer 20b walls of the downstream section 20 are excited by one or more actuators, the walls are deformed and when the excitation stress of the actuator or actuators stops, the walls return to their initial shape. According to one embodiment, the radially inner wall 20a can be made of an alloy, a composite, or an organic material allowing the wall 20a to work in an elastic domain, while at least a part of the radially outer wall 20b can be made of heat-shrinkable shape-memory material. According to another embodiment, the radially outer wall 20b can be made of an alloy, a composite, or an organic material allowing the wall 20b to work in an elastic domain, while at least part of the radially inner wall 20a can be made of heat-extensible shape-memory material. According to another embodiment, at least part of the radially inner wall 20a can be made of thermo-extensible shape memory material and at least part of the radially outer wall 20b can be made of thermo-extensible shape memory material. -retractile. Preferably, the thermo-deformable shape memory material is arranged on the radially inner 20a and/or radially outer 20b walls only at the level of the actuators. As a variant, the radially inner 20a or radially outer 20b walls can be made entirely of heat-deformable shape-memory material. For example, the radially inner 20a and outer 20b walls can be made of a nickel and titanium alloy (also known as "Kiokalloy") such as NiTiNol or NiTiCu.

The shape memory material constituting the radially inner 20a and outer 20b walls is fail-safe, that is to say it is such that its rest position, in other words when no actuator acts on the shape memory material with a view to its deformation, corresponds to a natural geometry for the storage of said material or for a longer period of use. Thus, in the event of an actuator failure, the shape memory material will return to its natural shape at rest and the nacelle fairing 3 will return to a secure geometry providing axial thrust to ensure the proper functioning of the propulsion system 1, 1' of the aircraft.

The radially inner 20a and outer 20b walls can also have a variable thickness axially and also azimuthalally close to stiffening bridges 22 so as to locally modify the elasticity of the structure. It is also possible to locally optimize the mechanical characteristics of the shape memory material constituting the radially inner 20a and outer 20b walls according to the desired local properties along the downstream section 20. Thus, it can be envisaged that the downstream section 20 is made up of a plurality of sections of different materials.

The downstream section 20 being made of a deformable, semi-rigid material guaranteeing it a rigid structural shape so as to avoid its sagging both at rest and under the action of an air flow in operation of the propulsion system 1, 1' and thus allowing the nacelle fairing 3 to maintain a homogeneous aerodynamic profile of its BF exit section. Advantageously, the downstream end 21 of the downstream section 20 can be made of an orthotropic material having adequate elastic moduli.

In addition, in order to ensure a substantially constant gap between the radially inner 20a and radially outer 20b walls of the downstream section 20, stiffening bridges 22 are arranged angularly at regular intervals between these walls 20a, 20b.

An object of the present invention is to be able to benefit from a fairing 3 of the nacelle of the propulsion system 1, 1', of which it is possible to vary the outlet section BF, as well as the shape of the latter, so as to orient the propulsion system thrust.

Thus, the downstream section 20 comprises means making it possible to vary the shape of the outlet section BF.

To this end, at least one of the radially inner 20a and radially outer 20b walls of the downstream section 20 may be made of a material having thermo-deformable characteristics, that is to say capable of deforming under the effect of the heat.

According to one embodiment, at least part of the radially outer wall 20b of the downstream section 20 is made of a material having heat-shrinkable characteristics, that is to say which can shrink under the effect of heat.

More specifically, the radially outer wall 20b comprises a plurality of heating elements 23a, 23b, 23c, 23d. Each heating element 23a, 23b, 23c, 23d has the general shape of an annular portion. The heating elements 23a, 23b, 23c, 23d extend in different consecutive angular sectors around the X axis of rotation so that the set of heating elements 23a, 23b, 23c, 23d has a generally annular shape. The radially outer wall 20b of the downstream section 20 is therefore divided into annular portions, each annular portion being associated with a heating element in the general shape of an annular portion.

Each heating element 23a, 23b, 23c, 23d is configured to supply heat in order to deform, more precisely, to retract, the annular portion of the radially outer wall 20b of the downstream section 20 over which it extends. In particular, the radially outer wall 20b comprises one or a plurality of heating circuits (not shown) configured to act on the heating elements 23a, 23b, 23c, 23d. Advantageously, each heating element 23a, 23b, 23c, 23d of the radially outer wall 20b comprises a heating circuit of its own, the heating circuits then having operations independent of each other.

More specifically, according to a first embodiment, the radially outer wall 20b of the downstream section 20 is provided with a plurality of heating coatings, or according to a second embodiment, with a plurality of heating resistors. The heating circuit or circuits are configured to heat the heating coatings, or the heating resistors. Each heating coating or heating resistance is configured to supply the heat making it possible to deform, and more precisely making it possible to retract, the annular portion of the radially outer wall 20b on which it extends, which is then made of a heat-shrinkable material.

In particular, the radially outer wall 20b is made of a material having shape memory characteristics so as to allow it to regain its shape and dimensions when the heating element(s) 23a, 23b, 23c, 23d are no longer in use. heated.

The heating elements 23a, 23b, 23c, 23d can be heated by any means known per se, such as for example by resistive electric circuits circulating in the heating coatings.

Each heating element 23a, 23b, 23c, 23d is operable independently of the other heating elements and is configured to deform the radially outer wall 20b in a direction which is radial with respect to the axis X of rotation and which is angularly centered with respect to the angular sector over which it extends, under the effect of a predetermined voltage command. Thus, each heating element 23a, 23b, 23c, 23d is configured to deform the annular portion of the downstream section 20 with which it is associated.

Each heating element 23a, 23b, 23c, 23d connected to an automatic device (not shown) making it possible to send the heating current to the heating elements 23a, 23b, 23c, 23d by application of a voltage command adapted according to the configuration for the nacelle fairing 3 to achieve the desired lateral thrust component. In particular, the automatic device supplies current or not to the heating elements 23a, 23b, 23c, 23d independently of the other heating elements.

According to the embodiment illustrated in Figures 5 to 7, the propulsion system 1, 1' comprises four heating elements 23a, 23b, 23c, 23d in the shape of a quarter ring: a first heating element 23a extending over a first part (the upper part in the figures) of the radially outer wall 20b of the downstream section 20, a second heating element 23b extending over a second part (the part on the right in the figures) of the radially outer wall 20b of the downstream section 20, a third heating element 23c extending over a third part (the lower part in the figures) of the radially outer wall 20b of the downstream section 20 and a fourth heating element 23d extending over a fourth part (the left part on the figures) of the radially outer wall 20b of the downstream section 20. The heating elements 23a, 23b, 23c, 23d are arranged so as to form a ring. In the figures, the heating elements are shown in the upper part (upper), the lower part (lower), and the lateral parts (left and right) of the downstream section, but can of course be integrated in different angular sectors, these angular sectors depending on the orientation of the propulsion system when it is installed on the aircraft. Of course, the invention is in no way limited to this example embodiment, and the propulsion system 1, 1' can comprise two, three or more than four heating elements.

Figure 2 shows a propulsion system 1, 1' according to the invention whose nacelle fairing 3 is shown in neutral position, that is to say pure axial thrust. Figure 5 shows the heating elements 23a, 23b, 23c, 23d in this neutral position. None of the heating elements 23a, 23b, 23c, 23d is activated, and therefore receives no current. The local deformation of all the heating elements is therefore zero. All of the heating elements 23a, 23b, 23c, 23d form a ring.

FIG. 3 represents a propulsion system 1, 1′ according to the invention whose nacelle fairing 3 is shown in a position where the flow is deflected upwards. FIG. 6 represents the heating elements 23a, 23b, 23c, 23d in this asymmetrical downward thrust position. A command to create a lateral displacement of the nacelle downwards requires deforming the exit section of the nacelle upwards, so as to create a force directed from the top downwards. The heating element 23a is activated to retract it. A voltage setpoint is calculated according to the input of the flight command, then is sent to the heating element 23a, via the heating circuit. Under the effect of the heat produced by the heating element 23a, the radially outer wall 20b located at the level of this heating element retracts, which causes the local stretching of the radially inner wall 20a. Indeed, under the effect of the heat emitted by the heating element 23a subjected to the heating circuit, the length of the radially outer wall 20b of the downstream section 20 decreases locally compared to its length at rest, which causes bending of the radially inner wall 20a of the downstream section 20 according to the arrow F2 in FIGS. 3 and 6. The heating element 23a deforms the radially outer wall 20b in a radial direction (arrow F2 in FIG. 6) with respect to the axis X of rotation and angularly centered with respect to the angular sector α over which it extends, that is to say vertically and upwards in FIG. 6. The heating element 23a is therefore subjected to a control voltage , while the heating elements 23b, 23c, 23d are not powered. In other words, the heating elements 23b, 23c, 23d are in the rest position. The actuation of the heating element 23a results in a local deformation, that is to say a radial and local expansion E of the output section BF. The outlet section of the fairing is therefore asymmetrical, and its shape results from the mechanical characteristics of the structure subjected to the heat of the heating element 23a. When the heating element 23a is activated, there is a radial and local expansion E of the outlet section BF of the fairing 3, that is to say that a local deflection of the structure is observed on the side of the heating element 23a activated. The radial dimension (in the direction of the arrow F2) of all the heating elements 23a, 23b, 23c, 23d (the heating element 23a being activated and the heating elements 23b, 23c, 23d being at rest) is equal to the sum of the radial dimension D23 of all the heating elements at rest and the expansion E. All the heating elements 23a, 23b, 23c, 23d form a ring, part of which (here the upper part) is distorted.

FIG. 4 represents a propulsion system 1, 1′ according to the invention whose nacelle fairing 3 is shown in a position where the flow is deflected downwards. FIG. 7 represents the heating elements 23a, 23b, 23c, 23d in this asymmetric upward thrust position. A command to create a lateral displacement of the nacelle upwards requires deforming the output section of the nacelle downwards, so as to create a force directed from the bottom upwards. The heating element 23c is activated to retract it. A voltage setpoint is calculated according to the input of the flight command, then is sent to the heating element 23c, via the heating circuit. Under the effect of the heat produced by the heating element 23c, the radially outer wall 20b located at the level of this heating element retracts, which causes the local stretching of the radially inner wall 20a. Indeed, under the effect of the heat emitted by the heating element 23c subjected to the heating circuit, the length of the radially outer wall 20b of the downstream section 20 decreases locally compared to its length at rest, which causes bending of the radially inner wall 20a of the downstream section 20 according to the arrow F3 in FIGS. 4 and 7. The heating element 23c deforms the radially outer wall 20b in a radial direction (arrow F3 in FIG. 7) with respect to the axis X of rotation and angularly centered with respect to the angular sector β over which it extends, that is to say vertically and downwards in FIG. 7. The heating element 23c is therefore subjected to a control voltage , while the heating elements 23a, 23b, 23d are not powered. In other words, the heating elements 23a, 23b, 23d are in the rest position. The actuation of the heating element 23c results in a local deformation, that is to say a radial and local expansion E of the output section BF. The outlet section of the fairing is therefore asymmetrical, and its shape results from the mechanical characteristics of the structure subjected to the heat of the heating element 23c. When the heating element 23c is activated, there is a radial and local expansion E of the outlet section BF of the fairing 3, that is to say that a local deflection of the structure is observed on the side of the heating element 23c activated. The radial dimension (in the direction of the arrow F3) of all the heating elements 23a, 23b, 23c, 23d (the heating element 23c being activated and the heating elements 23a, 23b, 23d being at rest) is equal to the sum of the radial dimension D23 of all the heating elements at rest and the expansion E. All the heating elements 23a, 23b, 23c, 23d form a ring, part of which (here the lower part) is distorted.

A command to create a lateral displacement of the nacelle to the right requires deforming the exit section of the nacelle to the left, so as to create a force directed from the left to the right (and to deflect the flow to the left). In this case of asymmetrical thrust to the right (not shown), the heating element 23d is activated to retract it. A voltage setpoint is calculated according to the input of the flight command, then is sent to the heating element 23d, via the heating circuit. Under the effect of the heat produced by the heating element 23d, the radially outer wall 20b located at the level of this heating element retracts, which causes the local stretching of the radially inner wall 20a. Indeed, under the effect of the heat emitted by the heating element 23d subjected to the heating circuit, the length of the radially outer wall 20b of the downstream section 20 decreases locally compared to its length at rest, which causes bending local area of the radially inner wall 20a of the downstream section 20. The heating element 23d deforms the radially outer wall 20b in a radial direction with respect to the X axis of rotation and angularly centered with respect to the angular sector ϕ on which it is extends. The heating element 23d is therefore subjected to a control voltage, while the heating elements 23a, 23b, 23c are not powered. In other words, the heating elements 23a, 23b, 23c are in the rest position. The radial dimension of all the heating elements 23a, 23b, 23c, 23d (the heating element 23d being activated and the heating elements 23a, 23b, 23c being at rest) is equal to the sum of the radial dimension D23 of the all heating elements at rest and expansion E. All heating elements 23a, 23b, 23c, 23d form a ring, part of which (here the left side part) is deformed.

A command to create a lateral displacement of the nacelle to the left requires deforming the exit section of the nacelle to the right, so as to create a force directed from the right to the left (and to deflect the flow to the right). In this case of asymmetrical thrust to the left (not shown), the heating element 23b is activated to retract it. A voltage setpoint is calculated according to the input of the flight command, then is sent to the heating element 23b, via the heating circuit. Under the effect of the heat produced by the heating element 23b, the radially outer wall 20b located at the level of this heating element retracts, which causes the local stretching of the radially inner wall 20a. Indeed, under the effect of the heat emitted by the heating element 23b subjected to the heating circuit, the length of the radially outer wall 20b of the downstream section 20 decreases locally compared to its length at rest, which causes bending local area of the radially internal wall 20a of the downstream section 20. The heating element 23b deforms the radially external wall 20b in a radial direction with respect to the axis X of rotation and angularly centered with respect to the angular sector θ on which it is extends. The heating element 23b is therefore subjected to a control voltage, while the heating elements 23a, 23c, 23d are not powered. In other words, the heating elements 23a, 23c, 23d are in the rest position. The radial dimension of all the heating elements 23a, 23b, 23c, 23d (the heating element 23b being activated and the heating elements 23b, 23c, 23d being at rest) is equal to the sum of the radial dimension D23 of the all heating elements at rest and expansion E. All heating elements 23a, 23b, 23c, 23d form a ring, one part of which (here the right side part) is deformed.

According to another embodiment not shown, at least part of the radially inner wall 20a of the downstream section 20 is made of a material having thermo-extensible characteristics, which can expand under the effect of heat.

More specifically, the radially inner wall 20a comprises a plurality of heating elements. Each heating element has the general shape of an annular portion. The heating elements extend in different consecutive angular sectors around the X axis of rotation so that the assembly of heating elements has a generally annular shape. The radially internal wall 20a of the downstream section 20 is therefore divided into annular portions, each annular portion being associated with a heating element in the general shape of an annular portion.

Each heating element is configured to supply heat in order to deform, more precisely, to extend, the annular portion of the radially internal wall 20a of the downstream section 20 on which it extends. In particular, the radially internal wall 20a comprises one or a plurality of heating circuits (not shown) configured to act on the heating elements. Advantageously, each heating element of the radially internal wall 20a comprises a heating circuit which is specific to it, the heating circuits then having operations independent of each other.

More specifically, according to a first embodiment, the radially internal wall 20a of the downstream section 20 is provided with a plurality of heating coatings, or according to a second embodiment, with a plurality of heating resistors. The heating circuit or circuits are configured to heat the heating coatings, or the heating resistors. Each heating coating or heating resistance is configured to supply the heat making it possible to deform, and more precisely making it possible to extend, the annular portion of the radially internal wall 20a on which it extends, which is then made of a thermo-extensible material .

In particular, the radially internal wall 20a is made of a material having shape memory characteristics so as to allow it to regain its shape and dimensions when the heating element(s) are no longer heated.

The heating elements can be heated by any means known per se, such as for example by resistive electric circuits circulating in the heating coatings.

Each heating element is operable independently of the other heating elements and is configured to deform the radially internal wall 20a in a direction which is radial with respect to the axis X of rotation and which is angularly centered with respect to the angular sector on which it is extends, under the effect of a predetermined voltage command. Thus, each heating element is configured to deform the annular portion of the downstream section 20 with which it is associated.

Each heating element is connected to an automatic device (not shown) making it possible to send the heating current to the heating elements by applying a voltage command adapted according to the desired configuration for the fairing 3 of the nacelle to obtain the thrust component desired side. In particular, the automatic device supplies current or not to the heating elements independently of the other heating elements.

The plurality of heating elements has an operation similar to the operation described for the heating elements of the radially outer wall in relation to FIGS. 2 to 7.

According to another embodiment not shown, at least part of the radially inner wall 20a of the downstream section 20 is made of a material having heat-extensible characteristics and at least part of the radially outer wall 20b of the downstream section 20 is made of a material with heat-shrinkable characteristics.

The intermediate section 30 being rigid, the downstream section 20 is fixed via its radially internal 20a and external 20b walls integral with the intermediate section 30, and the outlet section BF being free, the retraction of the radially external wall 20b and the deflection of the radially inner wall 20a causes the displacement of the outlet section BF in a radially outer direction represented by the arrow F2 in Figure 6 and by the arrow F3 in Figure 7, thus causing an increase in one dimension in a direction radial to the X axis (up to a D23+E value) and a local change in the shape of the BF output section.

Since the radially inner 20a and radially outer 20b walls of the downstream section 20 are made of a material that also has shape memory characteristics, it is possible for them to recover their shape and their dimension at rest, so that the outlet section BF recovers also its minimum radial dimension D23 when the heating elements 23a, 23b, 23c, 23d are no longer active.

The gradual increase in the control voltage induced by the automatic device gradually varies the shape of the BF output section. In particular, there is a deformation, radially to the X axis and locally, of the radially inner 20a and radially outer 20b walls of deformable material with shape memory of the downstream section 20 associated with the heating element. This, therefore, varies the dimensions and shape of the output section BF which thus changes from a minimum radial dimension D23 and a circular shape in the configuration in axial thrust mode as illustrated in Figure 2 to a radial dimension D23+E greater than the minimum radial dimension D23, with deformation in at least one radial direction of the circular shape of the heater assembly in the asymmetric thrust mode configuration of the nacelle fairing 3 as illustrated in Figures 3 and 4.

Similarly, the progressive reduction of the control voltage induced by the automatic device gradually changes the fairing 3 of the nacelle from a configuration in asymmetrical thrust mode as illustrated in Figures 3 and 4 to a configuration in axial thrust mode such that 'shown in Figure 2.

The transition from the configuration of Figure 5 to the configuration of Figure 6 or Figure 7 of the downstream section 20 of the fairing 3 of the nacelle, and vice versa, is done continuously depending on the current supplying the heating circuit, the jumpers stiffening 22 ensuring a substantially constant gap between the radially inner 20a and radially outer 20b walls of the downstream section 20 during configuration changes of the BF outlet section (dimensions and shape) (in other words of the outlet section) of the fairing 3 of nacelle.

The local heating of the heating elements can be distributed according to the desired local sags and the local elasticity of the radially internal 20a and external 20b walls of the downstream section 20.

Advantageously, several heating circuits can be envisaged to heat the heating coatings 23a, 23b, 23c, 23d so as to obtain a gradual heating of the heating coating and thus a gradual deformation of the radially outer wall 20b of the downstream section 20.

The invention has mainly been described for a propulsion system comprising four heating elements, but the propulsion system can of course comprise fewer or more heating elements.

Claims (11)

Propulsion system (1, 1') for an aircraft, comprising at least one rotor (2) and a nacelle fairing (3) extending around said at least one rotor (2) with respect to an axis (X) of rotation said rotor (2), this nacelle fairing (3) comprising a downstream section (20) located downstream of the last stage of the rotor parts of a turbine of the propulsion system, and of which a downstream end (21) forms a section outlet (BF) of the nacelle fairing (3); and characterized in that the downstream section (20) comprises a radially inner wall (20a) and a radially outer wall (20b), at least a part of one of said walls being made of a thermo-deformable shape memory material , and in that said wall is provided with a plurality of heating elements, the heating elements (23a, 23b, 23c, 23d) extending in different consecutive angular sectors around said axis (X) of rotation, each element heater (23a, 23b, 23c, 23d) being operable independently of the other heating elements and being configured to deform said wall in a direction which is radial with respect to the axis (X) of rotation and which is angularly centered with respect to the sector angular over which it extends, under the effect of a predetermined voltage command. The propulsion system (1, 1') of claim 1, wherein each heating element (23a, 23b, 23c, 23d) has the general shape of an annular portion, the heating elements (23a, 23b, 23c, 23d) extending in different consecutive angular sectors around said axis (X) of rotation so that all of the heating elements (23a, 23b, 23c, 23d) have a generally annular shape. Propulsion system (1, 1') of claim 1 or 2, wherein said plurality of heating elements (23a, 23b, 23c, 23d) have heating coatings. Propulsion system (1, 1') of claim 1 or 2, wherein said plurality of heating elements (23a, 23b, 23c, 23d) comprise heating resistors. Propulsion system (1, 1') of one of the preceding claims, wherein said wall comprises at least one heating circuit configured to act on the heating elements (23a, 23b, 23c, 23d). Propulsion system (1, 1') according to one of the preceding claims, in which the said wall comprises a plurality of heating circuits, each heating circuit being configured to act on at least one of the heating elements (23a, 23b, 23c, 23d). Propulsion system (1, 1') according to the preceding claim, in which the heating circuits operate independently of each other. Propulsion system (1, 1') according to one of the preceding claims, in which each of the radially internal (20a) and radially external (20b) walls is provided with a plurality of heating elements, heated by heating, the operation of the heating circuit(s) associated with the radially inner wall (20a) being independent of the heating circuit(s) associated with the radially outer wall (20b). Propulsion system (1, 1') according to one of the preceding claims, further comprising stiffening bridges (22) connecting the radially inner (20a) and radially outer (20b) walls of the downstream section (20). Propulsion system (1, 1') according to one of the preceding claims, in which the nacelle fairing also comprises an upstream section (10) forming an inlet section of the nacelle fairing and an intermediate section (30) connecting the sections upstream (10) and downstream (20), and in which the intermediate section (30) is rigid and is connected by at least one mast (5) to an engine (6) of the propulsion system (1, 1'). Aircraft characterized in that it comprises at least one propulsion system (1, 1') according to any one of the preceding claims, the propulsion system (1, 1') being pivotally mounted on the aircraft via a pivot shaft (4, 4') offset or traversing with respect to the rotor (2).