TURBOMACHINE MODULE COMPRISING VARIABLE-PITCH ROTOR BLADES AND BLADES MOUNTED ON A ROTOR, AND CORRESPONDING TURBOMACHINE

Abstract

The invention relates to a turbomachine module comprising: - a rotor (4) with a longitudinal axis (X), - a plurality of variable-pitch rotor blades (3) mounted on the rotor and extending radially, the rotor blades pivoting between a first and a second position, - a separation nozzle (9) separating an air flow passing through the rotor blades into a primary flow (F1) and a secondary flow (F2), the separation nozzle comprising an upstream edge (14) delimiting an inlet (17) of a flow vein (10) where the primary flow circulates, and, - stator blades (18) arranged at the inlet (17). According to the invention, the module comprises fins (21) mounted integral in rotation with the rotor and extending radially, the fins being arranged downstream of the rotor blades and upstream of the stator blades, the fins being configured to deflect the air flow leaving the rotor blades towards the stator blades. Figure for abstract: Fig.1

Description

Title of the invention: TURBOMACHINE MODULE COMPRISING PITCHED ROTOR BLADES VARIABLE AND BLADES MOUNTED ON A ROTOR, AND CORRESPONDING TURBOMACHINE Technical field of the invention

[0001] The present invention relates to the field of aircraft turbomachines. The invention relates more specifically to means for preventing the ingestion of foreign objects by components of the turbomachine. Technological background

[0002] Aircraft turbomachines are often confronted with the ingestion of foreign bodies called FOD which is the acronym for the English expression "Foreign Object Damage" and which can present a risk of damage to the internal organs of the turbomachine, and in particular to the combustion chamber. This risk of ingestion is increased with the reduction in the number of blades which are arranged at the inlet of the turbomachine and their rotation speeds. Turbomachines of the "open rotor" or "open fan" type which are equipped with a propeller with a number for example of fourteen unducted rotor blades are the most exposed to this risk of ingestion.

[0003] Foreign objects such as birds which enter the core of the turbomachine have significant impacts on the sizing of the various components of the turbomachine, such as the compressor, the combustion chamber, and / or the turbine and on the axial clearances between the various stages in the compressor and turbine assembly.

[0004] However, within the framework of the new aeronautical certifications which were imposed after numerous incidents resulting from collisions with birds, the most well-known of which resulted in the landing of an aircraft, foreign objects must not cause significant damage to the structure of the turbomachine components, nor cause the combustion chamber to go out if they reach the latter with a dramatic risk.

[0005] Furthermore, the supply of the combustion chamber is always problematic when the rotor blades which are located upstream of the flow vein of the primary flow occupy a thrust reversal position which closes and limits the flow rate of circulation of the air flow towards it.

[0006] There is a need to resolve all or part of the aforementioned drawbacks. Summary of the invention

[0007] The objective of the present invention is to provide a solution making it possible to avoid the ingestion of foreign bodies at the core of the turbomachine while improving the distribution of the air flow entering the core of the turbomachine.

[0008] We achieve this objective in accordance with the invention by means of a turbomachine module comprising: - a rotor centered on a longitudinal axis, - a plurality of rotor blades mounted on the rotor and extending radially relative to the longitudinal axis, the rotor blades being variable in pitch about a pitch axis transverse to the longitudinal axis between at least a first position corresponding to cruising flight and a second position corresponding to takeoff, - a separation nozzle arranged downstream of the rotor blades and separating an air flow passing through the rotor blades into a primary flow and a secondary flow, the separation nozzle comprising an upstream edge delimiting an inlet of a primary flow vein in which the primary flow circulates, and, - stator blades arranged at the inlet of the primary flow vein, the module comprising fins mounted integral in rotation with the rotor and extending radially relative to the longitudinal axis, the fins being arranged downstream of the rotor blades and upstream of the stator blades, the fins being configured so as to deflect the air flow leaving the rotor blades towards the stator blades between the first position and the second position.

[0009] Thus, this solution makes it possible to achieve the aforementioned objective. The installation of fins on the rotor of the module which rotates around the longitudinal axis, or engine axis, gives a centrifugal effect to the fins which makes it possible, on the one hand, to cut or chop each foreign body such as ice, volatile and therefore to reduce the volume of the foreign body which could penetrate into the core of the turbomachine downstream of the fins and, on the other hand, to drive the foreign body radially outwards, and away from the inlet of the primary vein, by centrifugal effect. The configuration of the fins is such that it makes it possible to act on the air flow at the outlet of the rotor blades and to improve the performance of the turbomachine. In particular, the profile of the fins is associated with the timing of the rotor blades to achieve the equivalent of a variable cycle.

[0010] The turbomachine module also includes one or more of the following features, taken alone or in combination:

[0011] - the number of blades is greater than the number of rotor blades.

[0012] - the number of blades is equal to twice the number of rotor blades.

[0013] - the fins are immobile in rotation relative to the rotor.

[0014] - each rotor blade comprises a blade having a leading edge and a leading edge trailing edges which are opposed and each fin has a leading edge and a trailing edge which are opposed, each blade having a blade chord, connecting the leading edge to the trailing edge, which is substantially parallel to a fin chord connecting the leading edge to the trailing edge in the first position.

[0015] - in the first position, an angle of incidence formed at the leading edge of the fins is equal to + / -10°. The angle of incidence corresponding to the difference between a skeleton angle at the leading edge of the fin, formed between a skeleton line connecting the leading edge to the trailing edge of the fin and the longitudinal axis, and an airflow angle at the trailing edge of the propeller, formed between a skeleton line connecting the leading edge to the trailing edge of the propeller and the longitudinal axis.

[0016] - the angle of incidence of each fin varies from one end of the foot to another end fin head following a stacking line.

[0017] - each fin comprises a free head end arranged radially at inside a straight line inclined relative to the longitudinal axis and passing through a minimum point of the separation beak, forming an angle of inclination of between 15° and 30°.

[0018] - each fin has an external radius which is a function of the axial position of each fin relative to the separating beak, a relationship between the external radius and the axial position being expressed by the following formula:

[0019] Rexlaiiette (x) > Rbec + D2 *tan (ah) - p

[0020] with, Rext being the outer radius at the tip of each fin between a leading edge and a trailing edge of the fin, measured from the longitudinal axis, Rbec being the radius of the separation beak measured at the upstream edge of the beak and from the longitudinal axis, D2 being the predetermined axial distance between two points and defined between the separation beak and a tip point of the fin located between its leading edge and its trailing edge, p having a value corresponding substantially to a radius of a body of a foreign object of substantially circular shape and having a weight of approximately 1 kg, ah being the angle of inclination of the inclined line.

[0021] The invention relates to a turbomachine comprising at least one turbomachine module as described above.

[0022] The invention also relates to an aircraft equipped with such a turbomachine. Brief description of the figures

[0023] The invention will be better understood, and other aims, details, characteristics and advantages thereof will appear more clearly on reading the detailed explanatory description which follows, of embodiments of the invention given as purely illustrative and non-limiting examples, with reference to the appended schematic drawings in which:

[0024] - [Fig.l] illustrates an axial and partial section of an example of a tur- machine equipped with a propeller or blower, and downstream fins arranged between the propeller or blower and the separation nozzle according to the invention;

[0025] - [Fig.2a] shows in more detail a cross-section of a fin according to the invention;

[0026] - [Fig.2b] shows in more detail a cross-section of a blade. rotor and a vane in the second position;

[0027] - [Fig.3] schematically represents a cylindrical section of the ro- torics and fins in a first cruising position according to the invention;

[0028] - [Fig.4] schematically represents the difference between the flow angles in upstream of the stator blades of a primary flow with and without the invention in takeoff and cruise conditions;

[0029] - [Fig.5] schematically represents a cylindrical section of the ro- torics and fins in a second take-off position according to the invention;

[0030] - [Fig.6] schematically represents the pressure of the flow upstream of the blades of stator as a function of the angles of incidence of the fins according to the invention;

[0031] - [Fig.7] is a partial axial sectional view of an example of a tur- machine equipped with a propeller or blower and downstream fins arranged between the propeller or blower and the separation nozzle according to the invention;

[0032] - [Fig.8] schematically represents the action of a vortex on the blades sta torics downstream of the fins according to the invention;

[0033] - [Fig.9] schematically represents a cylindrical section of the ro- torics and fins in a third thrust reversal position according to the invention;

[0034] - [Fig. 10] is an axial sectional view of a turbomachine module in which The distribution of the air flow around the rotor blades of a propeller or fan in a thrust reversal position according to the invention is shown. Detailed description of the invention

[0035] [Fig.l] represents a multi-flow turbomachine 1, intended to be mounted on an aircraft such as an airplane. The turbomachine 1 represented comprises an unducted propeller 2. Such a turbomachine is a turboprop and is known under the expression English for "open rotor" or "unducted fan" or "open fan". In this category of turbomachine, there are those which have two unducted and counter-rotating propellers (known by the English acronym UDF for "Unducted Dual Fan") or those having a single unducted propeller and a rectifier which is also unducted and which includes several stator blades (known by the English acronym USF for "Unducted Single Fan"). Of course, the invention applies to other types of turbomachines such as turbojets, and in particular double flow and double spool. The invention applies generally to a turbomachine which includes a fan or a propeller which is either ducted or unducted.

[0036] In the present invention, and generally, the terms "upstream", "downstream", "axial" and "axially" are defined with respect to the circulation of gases in the turbomachine and with respect to the longitudinal axis X, or engine axis X, of the turbomachine. Similarly, the terms "radial", "radially", "internal" and "external" are defined with respect to a radial axis Z perpendicular to the longitudinal axis X and with respect to the distance from the longitudinal axis X.

[0037] A turbomachine generally comprises, from upstream to downstream, a low pressure compressor or ("booster" in English) (not shown), a high pressure compressor (not shown), a combustion chamber (not shown), a high pressure turbine (not shown) and a low pressure turbine (not shown). The rotors of the low pressure compressor and the low pressure turbine are mechanically connected by a low pressure shaft so as to form a low pressure body. The rotors of the high pressure compressor and the high pressure turbine are mechanically connected by a high pressure shaft so as to form a high pressure body. The low pressure shaft extends inside the high pressure shaft and are coaxial with the longitudinal axis X.

[0038] Each compressor and each turbine comprises stages of rotor blades and stator blades which are alternated along the longitudinal axis. In the present invention, we understand by the term "stator blade" or "fixed blade", a blade which is not driven in rotation around the longitudinal axis X of the turbomachine.

[0039] In [Fig. 1], the propeller 2 is mounted upstream of the low-pressure compressor. The propeller 2 comprises a plurality of rotor blades 3 or moving blades arranged around the longitudinal axis X and extending radially from a rotor 4 forming the hub of the propeller 2. The rotor 4 is advantageously centered on the longitudinal axis X.

[0040] Each rotor blade 3 comprises a root 5 and a blade 6 which extends radially outwards from the root 5. Each blade 6 has an upstream leading edge 6a and a downstream trailing edge 6b, opposite each other. The blades 6 each have a transversely opposite intrados surface 6i and an extrados surface 6e, which connect the leading and trailing edges 6a, 6b.

[0041] A rectifier 7 is arranged downstream of the propeller 2. The rectifier 7 comprises a plurality of stator vanes 8 (or fixed vanes) known by the English acronym “OGV” for Outlet Guide Vane. The stator vanes 8 are distributed around the longitudinal axis X and are arranged downstream of the rotor vanes 3 of the propeller 2 so as to straighten the air flow generated by them.

[0042] The air flow F which passes through the propeller 2 is split into a primary flow F1 and a secondary flow F2 by a separation nozzle 9. The latter is advantageously arranged downstream of the rotor blades 3. The primary air flow F1 circulates in a primary flow vein 10 while the secondary flow F2 circulates radially outside the primary flow vein 10. The primary flow F1 passes through all of the compressors, the combustion chamber and all of the turbines, and is ejected outside by a primary nozzle (not shown). The secondary flow F2 circulates radially outside the casings and sweeps the unducted rectifier 7. In the case of a dual-flow or three-annular-flow turbojet comprising rotor blades of a ducted fan, the secondary flow F2 circulates in a secondary flow vein.

[0043] The separation nozzle 9 has a shape of revolution. More precisely, the separation nozzle 9 has a first annular wall 11 which makes it possible to guide the secondary air flow. The first annular wall 11 extends along the longitudinal axis X and is connected to a radially internal wall 12 which extends the first annular wall 11 downstream. The secondary flow F2 is also guided along the radially internal wall 12. Advantageously, but not limitingly, the rectifier 7 is mounted on the radially internal wall 12.

[0044] The separation nozzle 9 advantageously comprises a second annular wall 13 which makes it possible to guide the primary flow. The second annular wall 13 extends radially inside the first annular wall 11 and is connected to the latter at an upstream edge 14. The second annular wall 13 is connected to a radially external wall 15 which makes it possible to guide the primary flow. The radially external wall 15 extends, for example, the second annular wall 13 axially towards the downstream of the turbomachine. In the present example, the radially external wall 15 forms, with a radially internal wall 16, the primary flow vein 10.

[0045] Advantageously, but not limitatively, the upstream edge 14 of the separation nozzle 9 delimits an inlet 17 of the primary flow vein 10.

[0046] As can be seen in [Fig.l], stator vanes 18 known as IGV for “Inlet Guide Vane” in English are arranged at the inlet 17 of the primary flow vein 10. The stator vanes 18 extend between the radially outer wall 15 and the radially inner wall 16, and are distributed regularly around the longitudinal axis X. These stator vanes 18 are arranged in upstream of the low pressure compressor.

[0047] Advantageously, but not limitingly, each stator blade 18 comprises a leading edge 18a, upstream and a trailing edge 18b, downstream, opposite substantially along the longitudinal axis X. The leading and trailing edges 18a, 18b are connected by an intrados surface and an extrados surface, opposite transversely. Each stator blade 18 also comprises a root end fixed to the radially internal wall 16 and a head end fixed to the radially external wall 15.

[0048] According to one embodiment, each rotor blade 3 has variable pitch around a pitch axis A. The pitch axis A is perpendicular to the longitudinal axis. However, the pitch axis A may have an inclination of, for example, between 5° and 10° relative to the radial axis Z. The turbomachine 1 is equipped for this purpose with a pitch change system 20 which makes it possible to vary the pitch or the pitch of the rotor blades 3 around their pitch axis A so that they can have different angular positions depending on the operating conditions of the turbomachine and the flight phases.

[0049] Advantageously, but not limitingly, the blades of the rotor blades 3 pivot between at least a first position and a second position. In the first position, the rotor blades 3 allow cruising flight of the aircraft while in the second position the rotor blades 3 allow takeoff of the aircraft. In particular, in the first cruising flight position, the rotor blades 3 have a pitch angle of the chord relative to the longitudinal axis X which is between 35° and 50°. The chord of each rotor blade 3 is the straight segment which connects a leading edge 6a and a trailing edge 6b for each cross section of a blade 6 of rotor blade 3. In the second takeoff position, the rotor blades 3 have a pitch angle which is between 50° and 90° relative to the longitudinal axis X.

[0050] Advantageously, but not limitatively, the blades of the rotor blades 3 are also capable of equipping a third position in which they operate a thrust reversal which is known in English as "reverse". In the thrust reversal position, the blades 6 participate in the braking of the aircraft, in the manner of conventional thrust reversers. In this last position, the pitch angle of the blades is of the order of 90°.

[0051] The pitch change system 20 may comprise a control means (not shown) and a connecting mechanism (not shown) which connects the root of each blade 3 to the control means. The control means may be an actuator, for example linear, with a fixed body and a movable body. The connecting mechanism may comprise connecting rods or links, and / or a chain link, etc.

[0052] The pitch change system 20 is optionally installed in the rotor 4.

[0053] The turbomachine comprises fins 21 which are configured so as to prevent a foreign object from entering the primary flow stream 10 and reaching the combustion chamber for example. The fins 21 are mounted on the rotor 4 and extend radially outwards. These are advantageously arranged downstream of the rotor blades 3. They are more precisely arranged upstream of the separation nozzle 9. In other words, the fins 21 are arranged upstream of the stator blades 18. The fins 21 are advantageously arranged around the longitudinal axis X and preferably in a regular manner. Installing the vanes 21 on the rotor 4 allows them to be driven in rotation also at the same time as the rotor blades 3 of the fan or the propeller, which creates a centrifugal effect helping to avoid the ingestion of foreign bodies, such as birds or ice.

[0054] According to an exemplary embodiment, the number of fins 21 is greater than the number of rotor blades 3. Advantageously, but not limitingly, the number of fins 21 is equal to at least twice the number of rotor blades 3. This configuration makes it possible to obtain an inter-fin pitch 21 (along the circumferential direction around the longitudinal axis X) which is sufficiently small so that, for example, a bird having a weight of the order of 1 kg is cut by the fins 21 and partly driven outwards by the centrifugal effect.

[0055] Each fin 21 comprises a leading edge 21a and a trailing edge 21b which are opposite, here along the longitudinal axis. Advantageously, but not limitingly, each fin 21 has a radius of curvature which is minimal at the leading edge 21a and the trailing edge 21b. Each fin 21 comprises a lower surface face 21i and an upper surface face 21e which connects the leading edge 21a and the trailing edge 21b. Each fin 21 also extends between a root end 21c and a head end 21d which is free. The root end 21c is implanted in the rotor 4. There is no fairing around the fins 15.

[0056] Advantageously, the fins 21 are immobile in rotation relative to the rotor 4. That is to say that they do not pivot around a setting axis and rotate only with the rotor 4.

[0057] Advantageously, but not limitingly, each fin 21 comprises a plurality of cross sections or profiles which are stacked along a stacking line L between the root end 21c and the head end 21d. The stacking line L passes through the center of gravity of each transverse fin section 21.

[0058] [Fig.2a] represents an example of cross-section Ni or profile of a fin 21. The cross-section Ni is taken at a given radius measured from the axis Ion- gitudinal X. Each fin 21 has a chord Ci which is a straight segment connecting the leading edge 21a and the trailing edge 21b, for each cross-section. Furthermore, a skeleton line Si also connects each leading edge 21a to each trailing edge 21b and is equidistant from the intrados surface 21i and the extrados surface 21e. In other words, the skeleton line Si is a centerline between the intrados surface 21i and the extrados surface 21e of the fin 21.

[0059] [Fig.2b] illustrates a section of a rotor blade 6 of a rotor blade 3 and a fin 21 shown in the second position. The fin 21 has a skeleton angle ai at the leading edge 21a and for each cross-section. The skeleton angle ai is measured between the skeleton line Si and the longitudinal axis X at the leading edge 21a. The fin 21 also comprises a trailing angle [3i at the trailing edge 21b and for each cross-section. The trailing angle [3i is measured between the skeleton line Si and the longitudinal axis X at the trailing edge 21b. The skeleton angle ai may be identical from the tip end 21d to the root end 21c or may be different. Similarly, the vanishing angle [3i] may be the same at the root end 21c and the tip end 21d or different along the height of the fin 21.

[0060] [Fig.2b] also illustrates for example the air flow F arriving at the leading edge 21a and forming an angle aF of the air flow F measured from the longitudinal axis X. The incidence of the air flow F at the leading edge 21a of the fin 21 is represented by the angle of incidence a, corresponding to the difference between the angle aF of the air flow F and the skeleton angle ai.

[0061] Advantageously, but not limitatively, the blades 6 of the rotor blades 3 of the propeller 2 have the same parameters as those of the fins 21, i.e., a chord C6, a skeleton line S6, a skeleton angle a6, a leakage angle [36 for each cross section. Of course, the value of these parameters may be different from those of the fins 21.

[0062] Advantageously, the fins 21 are configured so as to deflect the air flow leaving the rotor blades 3 towards the stator blades 18 between the first position and the second position. Each fin 21 has a configuration and / or a geometry making it possible to achieve the deflection but also to improve the distribution of the pressure of the air flow along the leading edge 18a of the stator blades 18.

[0063] Advantageously, but not limitingly, the fins 21 are configured so as to reduce the difference in angle between the trailing edge of the rotor blades 3 and the leading edge 18a of the stator blades 18.

[0064] [Fig. 3] represents a cylindrical section at a given radius of the propeller 2 and the blades 21. In this [Fig. 3], the rotor blades 3 occupy the first cruising position. The blades 21 are configured so that their blade chords are substantially parallel to the blade chords of the blades 6 of the propeller 2 in at least one of the first, second and third positions, and preferably in the first cruising position as shown.

[0065] According to an exemplary embodiment, the angle of incidence α of the fin 21 is equal to + / -10° in the first cruising position. Preferably, each angle of incidence α of the fin 21 is equal to + / -5° in the first cruising position. Advantageously, these values ​​apply to the cross section at the level of the roots of the propeller 2 and the root end of the fins 21.

[0066] [Fig.4] represents on the abscissa axis, the angle aF of the air flow F at the leading edge 18a of the stator blades 18 and on the ordinate axis, the height hf of the air flow F along the radial axis Z. There is a first curve C3 which represents the angle aF of the air flow F at the leading edge 18a of the stator blades 18 when the propeller 2 is in the first position, a second curve C21 which represents the angle aF of the air flow F at the leading edge 18a of the stator blades 18 when the propeller 2 is in the second position without the presence of the fins 21 and a third curve C18 which represents the angle aF of the air flow F at the leading edge 18a of the stator blades 18 in the second position with the presence of the fins 21.We can see from this diagram that the third curve Cl8 is closer to the first curve C3 than the second curve C21, which implies that the presence of fins 21 allows a reduction in the angle difference aF of the air flow, and therefore the angle of incidence difference, at the level of the leading edge 18a of the stator blades 18 between take-off (second position) and cruise (first position).

[0067] By setting the fins 21 on the angle of escape [3i of a cross-section at the level of the root end of the rotor blade 3 of the propeller 2 on an operating point corresponding to a cruise flight, this makes it possible to be transparent on this operating point whereas for operating points such as takeoff, the fins 21 will provide more work. Here we mean by the term "transparent" that the fins 21 will not work or will do little work on this operating point because they will be in the axis of the flow of the air flow F. Similarly, in the present description we mean by the term "work", a deviation of the air flow carried out by an aerodynamic body such as a fin body 21 or a rotor blade body 3. The deviation of the air flow causes an increase in pressure.The work provided by the vanes 21 in the second position (take-off position) makes it possible to reduce the work of the compressor stages downstream of the vanes 21 and to obtain a gain in mass and performance. Indeed, a rotating rotor blade 3 has the purpose of generating work to the air flow. Having vanes 21 on the rotor 4 makes it possible to add additional work to the air flow.

[0068] [Fig.5] illustrates the rotor blades 3 in the second position, i.e. the position of takeoff. The fins 21 remain fixed in position. In this case, the airflow F arriving at the fins 21 is in positive incidence, and the fins 21 increase the deflection of the airflow at the rotor 4 (the rotor reference mark 4) which has the effect of providing more work to the airflow F when the rotor blades 3 are in the second position corresponding to takeoff than in the first position which corresponds to the cruising flight of the aircraft. The fins 21 make it possible to reduce the airflow angle upstream of the stator blades 18 and to find an airflow angle closer to the airflow angle observed in the first position, which makes it possible to reduce the airflow incidence range under which the stator blades 18 must be effective.

[0069] The angle of incidence a of each fin 21 varies from the root end 21c to the tip end 21d along the stacking line L. By varying the angle of incidence a, the work of the air flow increases at the root end 21c of the fins 21. The pressure of the air flow and the air flow angle can also be homogenized and allow for a less aggressive gradient along the leading edge of the stator blades 18.

[0070] [Fig.6] represents on the abscissa axis the total pressure Pt at the leading edge 18a of the stator blades 18 and on the ordinate axis the height hf of the air flow F along the radial axis Z. A first curve C4 represents the pressure at the leading edge 18a of the stator blades 18 in the case where the vanes 21 deflect the flow in a similar manner between the root end 18c and the tip end 18d. A second curve C5 represents the pressure at the leading edge 18a of the stator blades 18 with a greater deflection at the root end 18c relative to the tip end 18d in the first cruising position. The second curve C5 is substantially parallel to a vertical axis in the plane of [Fig.6] relative to the first curve C4 which has an inflection. This allows for better performance from the stator blade 18 and the compressor (booster).

[0071] According to an advantageous, but non-limiting, characteristic, each fin 21 has an external radius Rext which is radial to the longitudinal axis X. In the remainder of the description, the term “external radius” will be used for the fin 21. The external radius Rext is measured between the longitudinal axis X and the free tip end 21d, and preferably at a midpoint between the leading edge 21a and the trailing edge 21b. Advantageously, the external radius Rext is less than the height h2 of the blades 6 of the propeller 2. The height h2 of the rotor blades 3 is measured between the longitudinal axis X and a tip end 6d of the blade 6. The external radius Rext can be a function of several parameters, at least one of which is chosen from the group comprising the predetermined axial position of each fin 21 and a critical ingestion angle.

[0072] In the present application, we understand by the expression “critical angle ingestion angle”, an angle of inclination ah which is measured between the longitudinal axis X and an inclined line DI which defines the trajectory of the foreign object, here a bird. The critical ingestion angle depends on the missions of the aircraft and the certification of the turbomachine. The missions of the aircraft define its flight envelope (set of points flight speed, incidence, etc.) and mainly its rate of incidence in climb (the angle of the aircraft in its takeoff phase) or descent in its flight plan. The inclined line DI passes through a minimum point of the separation slat 9. The minimum point (measured radially from the longitudinal axis X) is located at the upstream edge 14.

[0073] In the present example, each fin 21 is inscribed in this angle of inclination ah formed between the longitudinal axis X and the inclined line D1, so that the foreign object O traveling along the inclined line D1 comes into contact with the fin 21 over its entire height. In particular, the free head end 21d of each fin 21 is radially inside the inclined line DL

[0074] Advantageously, but not limitatively, the angle of inclination ah is between 15° and 30°.

[0075] Advantageously, but not limitatively, each fin 21 is arranged at a predetermined axial position (or predetermined axial distance D2) relative to the separation beak 9. The axial position is determined relative to the upstream edge 14 of the separation beak 9. The predetermined axial position allows that whatever the angle of arrival of the foreign object O, the latter will come into contact with at least one fin 15.

[0076] Advantageously, but not limitingly, the predetermined axial distance D2 measured from the upstream edge 14 and a point between the trailing edge 21b and the leading edge 21a of each fin 21 in a radial plane is between 30 and 130 mm. Advantageously, but not limitingly, the greater the predetermined axial distance D2, the greater the external radius Rext of the fins 21 must be. The relationship between the external radius Rext and the axial position (predetermined axial distance D2 between the two points) of each fin 21 is expressed by the following mathematical formula: Rextailette (x) > Rbec+ D2 -tan (ah)- p

[0077] With:

[0078] - Rext being the external radius at the head end 21d of each fin 21, in mm, at level of a point between the leading edge 21a and the trailing edge 21b, measured from the longitudinal axis X,

[0079] - Rbec being the radius of the separation beak measured at the upstream edge 14 of the beak and from the longitudinal axis X (the upstream edge 14 being the most upstream point relative to the longitudinal axis X), in mm,

[0080] - D2 being the predetermined axial distance, in (mm), and is defined between the nozzle of se preparation and a leading point 21d of the fin 21 located between its leading edge 21a and its trailing edge 21b.

[0081] - p having a value, preferably 40mm, corresponding substantially to a radius of a body of a foreign object O, in particular a bird, of a generally substantially circular shape and having a weight of approximately 1 kg,

[0082] - ah being the angle of inclination of an inclined line expressed in radians.

[0083] The above mathematical formula gives a minimum radius (Rext) that the fin 21 must reach at its leading end 21d, at a point, between the leading edge 21a and the trailing edge 21b. The above mathematical formula therefore defines the radius Rext of the fin 21 as a function of its position along the longitudinal axis X relative to the upstream edge 14 of the separating lip 9.

[0084] With reference to [Fig.7], as each fin 21 stops at a small radius (the external radius Rext is less than the height h2 of the rotor blades 3), these will generate a vortex V at the level of the free head ends 21d of the fins 21 which will have the effect of homogenizing the pressure along the leading edge of the stator blades 18 and the angle of the air flow at the level of the leading edge of the stator blades 18.

[0085] [Fig.8] represents the action of the vortex on curves C4 and C5. The vortex reduces the load at the tip end of the stator blades 18 and increases the load at the root end of the stator blades 18. Curves C4 and C5 intersect at mid-height.

[0086] When the rotor blades 3 occupy the third thrust reversal position as shown in FIGS. 9 and 10, the setting of the rotor blades 3 of the propeller 2 at the root end greatly limits the passage of the air flow which comes upstream of the propeller 2. In fact, the spacing between two root ends of adjacent rotor blades is smaller than at their tip ends. By installing the fins 21 downstream of the rotor blades, the speed of the air flow arriving at the root ends of the rotor blades 3 is higher compared to the speed of the air flow circulating at the tip ends of the rotor blades 3. The presence of the fins 21 also makes it possible to maintain dynamics (the fins make it possible to advance the flow downstream, i.e. towards the stator blades 8) and to suck in the air flow circulating from above the separation nozzle 9.

Claims

Claims

1. A turbomachine module comprising: - a rotor (4) centered on a longitudinal axis (X), - a plurality of rotor blades (3) mounted on the rotor (4) and extending radially relative to the longitudinal axis (X), the rotor blades (3) being variable pitch about a pitch axis (A) transverse to the longitudinal axis (X) between at least a first position corresponding to cruising flight and a second position corresponding to takeoff, - a separation nozzle (9) arranged downstream of the rotor blades (3) and separating an air flow passing through the rotor blades (3) into a primary flow (F1) and a secondary flow (F2), the separation nozzle (9) comprising an upstream edge (14) delimiting an inlet (17) of a primary flow vein (10) in which the primary flow (F1) circulates, and stator blades (18) arranged at the level of the inlet (17) of the primary flow vein (10),characterized in that the module comprises fins (21) mounted integral in rotation with the rotor (4) and extending radially relative to the longitudinal axis (X), the fins (21) being arranged downstream of the rotor blades (3) and upstream of the stator blades (18), the fins (21) being configured so as to deflect the air flow leaving the rotor blades (3) towards the stator blades (18) between the first position and the second position.,

2. Turbomachine module according to the preceding claim, characterized in that the number of fins (21) is greater than the number of rotor blades (3).

3. Turbomachine module according to claim 1 or 2, characterized in that the number of fins (21) is equal to twice the number of rotor blades (3).

4. Turbomachine module according to one of the preceding claims, characterized in that the fins (21) are immobile in rotation relative to the rotor (4).

5. Turbomachine module according to any one of the preceding claims, characterized in that each rotor blade (3) comprises a blade (6) having a leading edge (6a) and a trailing edge (6b) which are opposite and each fin (21) has a leading edge (21a) and a trailing edge (21b) which are opposed, each blade (6) having a blade chord (C6), connecting the leading edge (6a) to the trailing edge (6b), which is substantially parallel to a winglet chord (Ci) connecting the leading edge (21a) to the trailing edge (21b) in the first position.

6. Turbomachine module according to the preceding claim, characterized in that, in the first position, an angle of incidence (a) formed at the leading edge (21a) of the fins (21) is equal to + / -10°, wherein the angle of incidence (a) corresponding to the difference between a skeleton angle (ai) at the leading edge (21a) of the fin (21), formed between a skeleton line (Si) connecting the leading edge (21a) to the trailing edge (21b) of the fin (21) and the longitudinal axis (X), and an airflow angle (aF) at the trailing edge (6b) of each blade of the rotor blades (3), formed between a skeleton line (S6) connecting the leading edge (6a) to the trailing edge (6b) of each blade of the rotor blades (3) and the longitudinal axis (X).

7. Turbomachine module according to the preceding claim, characterized in that the angle of incidence (a) of each fin (21) varies from a root end (21c) to a head end (21d) of the fin (21) along a stacking line (L).

8. Turbomachine module according to the preceding claim, characterized in that each fin (21) comprises a free head end (21d) which is arranged radially inside a straight line (Dl) inclined relative to the longitudinal axis (X) and passing through a minimum point of the separation nozzle (9) forming an angle of inclination (ah) of between 15° and 30°.

9. Turbomachine module according to the preceding claim, characterized in that each fin (21) has an external radius (Rext) which is a function of the axial position of each fin (21) relative to the separation nozzle (9), a relationship between the external radius and the axial position being expressed by the following formula: ReXtai]ette ( x ) Rfec + *tan (^) ~ P With, - Rext being the external radius at the tip end (21d) of each fin (21) between a leading edge (21a) and a trailing edge (21b) of the fin (21), - Rbec being the radius of the separation nozzle (9) measured at the upstream edge (14), - D2 being the predetermined axial distance between two points and being defined between the separating beak (9) and the set of points defining the head of each fin (21), - p having a value corresponding substantially to a radius of a body of a foreign object (O) of substantially circular shape and having a weight of approximately 1 kg, and, - ah being the angle of inclination of the inclined line (Dl).

10. A turbomachine (1) comprising a turbomachine module according to any one of the preceding claims.

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