Aircraft turbomachine incorporating an improved input rectifier

Abstract

The invention relates to a turbofan aircraft turbomachine comprising a propeller constituting an unfaired fan (22) generating an open secondary airflow (F2), the turbomachine comprising a gas generator (30) through which a primary gasflow (F1) passes, the gas generator (30) comprising an inlet straightener (38) comprising straightener blades (38A), each straightener blade comprising a blade 4 disposed in a primary flow channel, or primary channel 4, the blade 4 comprising a leading edge extending between a radially inner end and a radially outer end, the line joining the leading edge ends of each blade extending in a direction forming an angle of less than or equal to 60° with a central axis (X) of the turbomachine. Figure 2.

Description

Title of the invention: Aircraft turbomachine comprising an improved input rectifier. Technical field

[0001] The present invention relates to the field of aeronautics, and more particularly to a double-flow turbomachine and to a propulsion assembly comprising a double-flow turbomachine. State of the art

[0002] Aircraft are known that are propelled by at least one propulsion unit comprising a turbomachine, such as a turbofan engine. Each propulsion unit is attached to the aircraft by a mast located generally under or on a wing, or at the level of the aircraft fuselage. A turbofan engine primarily comprises a gas generator and a fan.

[0003] The fan may be enclosed, in which case the turbojet is housed in a nacelle. The fan may also be unenclosed, as is the case with turbojets known as "open fan" or "unducted single fan".

[0004] The gas generator includes, in particular, from upstream to downstream with respect to the direction of gas flow, an inlet rectifier, a low pressure compressor and a high pressure compressor.

[0005] During operation, an airflow is accelerated by the fan, then splits into a primary flow and a secondary flow. The primary flow flows in a primary gas circulation channel through the turbojet's gas generator.

[0006] In the case of a shrouded fan, the secondary flow runs in a secondary channel surrounding the gas generator. The secondary channel is delimited, radially inward, partly by an internal structure of the nacelle that encloses the gas generator, and radially outward, partly by an external structure of the nacelle that surrounds the turbojet. A portion of the secondary channel is further delimited radially outward by a fan casing surrounding the fan, and by an intermediate casing located downstream of the fan casing. In the case of an unshrouded fan, the secondary flow is also generated by the fan, but is open and flows around the gas generator.

[0007] The inlet rectifier of the gas generator includes blades which are intended to operate in a severe environment, and which are in particular subjected to mechanical shocks, linked for example to the ingestion of foreign bodies by the turbomachine (for example in the case of ingestion of a bird or hailstones).

[0008] In the case of a shrouded fan, the presence of a fan cone of generally conical shape having a relatively large apex angle, as well as the implantation of the fan blades on a hub having a generally conical shape, promotes a dispersion of foreign bodies towards the secondary flow stream rather than in the primary flow, towards the gas generator.

[0009] In the case of an unshrouded fan, the presence of a fan cone with a generally cylindrical shape, as well as the mounting of the fan blades on a hub with a generally cylindrical shape, does not promote the dispersion of foreign bodies outside the primary flow. Furthermore, the number of fan blades and the fan's rotational speed can be reduced compared to a shrouded fan. There is therefore a greater risk that a foreign body heading towards the engine will not be at least partially intercepted and cut by the fan blades before coming into contact with components of the gas generator, particularly the rectifier blades.

[0010] There is therefore a need to prevent excessive damage to the rectifier blades in the event of a collision with a foreign body in the case of an architecture with an unshod blower.

[0011] The objective of the present invention is to propose a turbomachine that meets this need. Description of the invention

[0012] For this purpose, the invention relates to a twin-flow aircraft turbomachine comprising a propeller constituting an unfaired fan generating an airflow called secondary flow, the turbomachine comprising a gas generator through which a gas flow called primary flow passes, the gas generator comprising an inlet straightener comprising straightener blades, each straightener blade comprising a blade disposed in a circulation channel of the primary flow, or primary channel, the blade comprising a leading edge extending between a radially internal end and a radially external end, the line joining the ends of the leading edge of each blade extending in a direction forming with a central axis of the turbomachine an angle less than or equal to 60°.

[0013] Thus, by providing stator blades with a leading edge that is only slightly inclined relative to the central axis of the turbomachine, i.e., inclined at most 60° to this axis, the deflection of foreign bodies coming into contact with these blades is facilitated, away from the primary flow. Furthermore, the mechanical resistance of the blades to impacts is improved.

[0014] In one embodiment, the angle is between 20° and 60°, or between 30° and 50°.

[0015] In one embodiment, the leading edge of each blade is substantially straight between its radially inner end and its radially outer end.

[0016] In one embodiment, the turbomachine includes a structure separating the primary and secondary flows, or separation nozzle, the separation nozzle having an end portion which is aligned with the direction formed by the ends of the leading edge, or which is slightly inclined with respect to this direction, for example inclined by less than 15°.

[0017] In one embodiment, the radially external end of the leading edge of each stator blade coincides with a free end of the separation nozzle, where it is offset, along the central axis, by a distance less than or equal to 50 millimeters from this free end.

[0018] In one embodiment, the blower comprises between 8 and 14 blower blades.

[0019] The invention also relates to an aircraft comprising at least one propulsion assembly comprising a turbomachine as defined above. Brief description of the drawings

[0020] [Fig-1] Fig. 1 represents an aircraft equipped with a propulsion system conforming to the invention.

[0021] [Fig.2] Fig.2 represents a schematic cross-sectional view of a propulsion assembly according to the invention.

[0022] [Fig.3] The [Fig.3] represents a detailed view of the [Fig.2], schematically illustrating the geometry of an inlet rectifier blade.

[0023] [Fig.4] The [Fig.4] is a perspective view showing the arrangement of an inlet rectifier of a propulsion assembly according to the invention.

[0024] [Fig.5] The [Fig.5] is a cross-sectional view in a diametrical plane of the input rectifier of the [Fig.4]. Detailed description

[0025] Figure 1 represents an aircraft 100, in this example an airplane, equipped with two propulsion units 10, namely one propulsion unit 10 per wing 101, with only one propulsion unit 10 and one wing 101 being shown in Figure 1. According to one embodiment, the aircraft 100 can be equipped with more than one propulsion unit 10 per wing 101, each wing 101 having the same number of propulsion units 10. Axis A designates the axis of the fuselage 102 of the aircraft 100. The propulsion unit 10 can be configured to propel the aircraft 10 at a cruising speed between Mach 0.7 and Mach 0.9.

[0026] Figure 2 shows a schematic cross-sectional view of the propulsion assembly 10, according to plane II of Figure 1. The propulsion assembly 10 extends along an axis X. The propulsion assembly 10 comprises a turbomachine which includes a module propulsion unit 20, a gas generator 30, and, in the example, a speed reduction device 40. When the propulsion unit 10 is mounted on the aircraft 100, the X axis is not necessarily parallel to the A axis.

[0027] The propulsion module 20 has a propeller 22, constituting an unfaired fan. The propeller is provided with a plurality of blades 22A. The propulsion module also includes a stator 24 provided with a plurality of blades 24A, and a propeller shaft 26 configured to drive the propeller 22 in rotation. The propeller shaft 26 can extend along the X-axis. The blades 22A of the propeller 22 can be made entirely or partially of composite material. The blades 24A of the stator 24 can be made entirely or partially of composite material. The propeller 22 can comprise between 8 and 14 blades 22A and the stator 24 can comprise the same or fewer number of blades 24A, for example, between 8 and 14 blades 24A. The setting of the 24A blades of the 24 rectifier can be fixed or variable.

[0028] The gas generator 30 has a drive shaft 33A. The drive shaft can extend along the X axis. The propeller shaft 26 can be coaxial with the drive shaft 33A, and their respective axes of rotation can coincide with the X axis of the propulsion assembly 10. This allows for an annular air inlet within the gas generator 30 coaxial with the X axis, thanks to which the outer casing of the gas generator has a relatively simple shape and exhibits a certain rotational symmetry, which tends to reduce possible airflow disturbances. In this example, the gas generator 30 comprises, from upstream to downstream, the gases flowing within the propulsion assembly 100, from upstream to downstream, a compressor 32 (or compressor section 32), a combustion chamber 34, and a turbine 36 (or turbine section 36).

[0029] The gas generator 30 may be of the twin-spool type and comprise a low-pressure spool 30A and a high-pressure spool 30B. The low-pressure spool 30A may comprise a low-pressure compressor 32A rotationally coupled to a low-pressure turbine 36A via a low-pressure shaft 33A, which may form the drive shaft of the gas generator 30. The high-pressure spool 30B may comprise a high-pressure compressor 32B located downstream of the low-pressure compressor 32A and upstream of the combustion chamber 34, and a high-pressure turbine 36B located downstream of the combustion chamber 34 and upstream of the low-pressure turbine 36A, and rotationally coupled to the high-pressure compressor 32B via a high-pressure shaft 33B. The compressor 32 of the gas generator 30 may comprise the low-pressure and high-pressure compressors 32A and 32B.The turbine 36 of the gas generator 30 can comprise the low-pressure and high-pressure turbines 36A and 36B. The low-pressure and high-pressure shafts 33A and 33B can be coaxial. The high-pressure shaft 33B can receive a portion of the low-pressure shaft 33A. According to one embodiment, the low-pressure shaft 33A and high-pressure shaft 33B may be co-rotating, i.e., configured to rotate relative to each other in the same direction around the X-axis. According to another embodiment, the low-pressure shaft 33A and high-pressure shaft 33B may be counter-rotating, i.e., configured to rotate relative to each other in opposite directions around the X-axis. The rotational speed of the low-pressure shaft 33A may be lower than the rotational speed of the high-pressure shaft 33B.

[0030] According to an embodiment not shown, the propulsion assembly may be of the three-shaft type. The turbine 36 may include an intermediate turbine arranged axially between the high-pressure turbine 36B and the low-pressure turbine 36A and configured to drive an intermediate compressor arranged axially between the low-pressure compressor 32A and the high-pressure compressor 32B via an intermediate shaft.The intermediate shaft can be housed between the low-pressure shaft 33A and the high-pressure shaft 33B. The intermediate shaft and the low-pressure shaft 33B can be co-rotating or counter-rotating with respect to each other.

[0031] Each compressor 32A, 32B and turbine 36A, 36B may comprise a plurality of stages, each stage comprising a blade wheel, respectively 32AA, 32BA, 36AA, 36BA, movable in rotation about the X-axis (or rotor) and a blade wheel, respectively 32AB, 32BB, 36AB, 36BB, fixed about the X-axis (or stator). In this example, the low-pressure compressor 32A may have at least 2 stages and at most 5 stages, for example 2 stages, the high-pressure compressor 32B may have between 8 and 11 stages (only two stages being shown for clarity of the figure), the high-pressure turbine 36B may have 2 stages, and the low-pressure turbine 36A may have between 3 and 8 stages (only two stages being shown for clarity of the figure). A rectifier 37, or fixed paddle wheel rotating about the X axis, can be arranged downstream of the combustion chamber 34 and upstream of the high-pressure turbine 36B.

[0032] A speed reduction device 40 can indirectly couple the drive shaft 33A in rotation with the propeller shaft 26. The speed reduction device 40 can be configured to drive the propeller shaft 26 at a rotational speed lower than the rotational speed of the drive shaft 33A. The drive shaft 33A connects the low-pressure turbine 36A (or the low-pressure housing 30A) to an inlet of the speed reduction device 40, while the propeller shaft 26 connects an outlet of the speed reduction device 40 to the propeller 22. The propeller 22 is therefore driven by the low-pressure turbine 36A (or the low-pressure housing 30A) via the drive shaft 33A (or low-pressure shaft), the speed reduction device 40, and the propeller shaft 26. In this example, the speed reduction device 40 can be positioned, considered along the X-axis, between an upstream end of the drive shaft 33A and a downstream end of the propeller shaft 36.

[0033] For example, the speed reduction device 40 may be an epicyclic gear train reduction device, for example of the "epicyclic" or "planetary" type, according to the terminology sometimes used by those skilled in the art. Such a mechanism may comprise one stage, two stages, or more than two stages.

[0034] The gas generator includes, upstream of the elements described above, an input rectifier 38 according to the invention, comprising a plurality of rectifier blades 38a, and which is described in more detail below.

[0035] Figures 3 to 5 are views showing more precisely the geometry of the inlet rectifier blades.

[0036] Fig. 3 is a partial schematic view of Fig. 2, showing the geometry of the input rectifier 38.

[0037] The input rectifier 38 includes input guide vanes (commonly referred to as IGV, an acronym corresponding to the English term "inlet guide vanes"), or rectifier vanes 38A. The rectifier vanes 38A are fixed vanes, and may or may not be variable stator vanes (commonly referred to as VSV, an acronym corresponding to the English term "variable stator vanes").

[0038] Each stator blade 38A comprises a blade 380 which is disposed in the primary flow channel, or primary stream 300. Each blade has a first radially internal end 382, ​​or blade root 382, ​​and a second radially external end 384, or blade tip 384. The blade 380 is extended, at its first end 382, ​​successively by an internal mounting plate (not shown), and, at its second end 384, by an external mounting plate (not shown). The terms "external" and "internal" are used in this application to designate an element radially external, and radially internal, respectively, with respect to the principal axis X of the turbomachine and with reference to the normal operating position of the blade in the turbomachine.

[0039] Each blade 380 has a leading edge 386 and a trailing edge 388. The leading edge 386 and the trailing edge 388 each extend within the primary vein 300, between the blade root 382 and the blade tip 384.

[0040] According to the invention, the leading edge 386 extends in a general direction that is slightly inclined with respect to the central axis X of the turbomachine. More precisely, the straight line BB joining the inner end 386A and the outer end 386B of the leading edge forms an angle α with the central axis X of the turbomachine that is less than or equal to 60°. Preferably, this angle is between 20° and 60°, or between 30° and 50°.

[0041] In the event of ingestion of a foreign body, such as a bird or a block of ice, the slight inclination of the inlet stator blades 38 relative to the central axis of the turbomachine allows a foreign body coming into contact with the leading edge 386 of these blades to be deflected, thus preventing the foreign body from entering the primary flow stream 300. Furthermore, this orientation of the leading edge 386 improves the mechanical strength of the stator blades 38A in the event of a collision with a foreign body, since the mechanical stress experienced by the blade during such a collision will be better distributed over time than with a conventional configuration (in which the stator blades are oriented radially or very slightly inclined relative to the radial direction). Indeed, the pressure caused by the impact on the blade will be progressively distributed over the entire height of the leading edge (or at least over a major portion thereof).This increases the probability that, in the event of impacts with a foreign object, the 38A rectifier blades will be able to withstand the impact without being damaged.

[0042] Advantageously, the structure 39 separating the primary flow from the secondary flow at the level of the inlet straightener, called the separation nozzle 39, includes an end portion 390 at which the primary and secondary flows separate, this end portion 390 being located in the extension of the leading edge of the blades, and aligned with the line joining the inner 386A and outer 386B ends of the leading edge of the blade 380. In other words, the outer end 386B of the leading edge of the straightener blades is not set back (with respect to the direction of circulation of the primary flow) with respect to the free end 392 of the separation nozzle 39, and vice versa. The free end 392 of the separating beak can however be slightly offset (upstream) relative to the radially external end 386B of the leading edge 386, for example offset (along the central axis X) by a distance d less than or equal to 50 millimeters (cf. [Fig.2]).This configuration facilitates the deflection of a foreign body coming into contact with one or more blades of the inlet stator towards the secondary flow. In particular, a zero or relatively small offset of the free end 392 of the separation nozzle 39 relative to the radially external end 386B of the leading edge 386 of each blade 380 prevents excessive protrusion of the separation nozzle relative to the leading edge 386, which would limit or, in some cases, negate the deflection effect of foreign bodies by the blades 380.

Claims

Demands

1. A turbofan aircraft turbomachine comprising a propeller constituting an unfaired fan (22) generating an open secondary airflow (F2), the turbomachine comprising a gas generator (30) through which a primary gasflow (F1) passes, the gas generator (30) comprising an inlet straightener (38) comprising straightener blades (38A), each straightener blade comprising a blade (380) disposed in a primary flow channel, or primary channel (300), the blade (380) comprising a leading edge (386) extending between a radially inner end (386A) and a radially outer end (386B), the line (BB) joining the ends (386A, 386B) of the leading edge of each blade (380) extending in a direction forming an angle (a) with a central axis (X) of the turbomachine less than or equal to 60°.

2. Turbomachine according to the preceding claim, wherein the angle (a) is between 20° and 60°, or between 30° and 50°.

3. Turbomachine according to any one of the preceding claims, wherein the leading edge (386) of each blade (380) is substantially straight between its radially inner end (386A) and its radially outer end (386B).

4. Turbomachine according to any one of the preceding claims, comprising a structure (39) separating the primary flow (F1) and the secondary flow (F2), or separation nozzle (39), the separation nozzle (39) comprising an end portion (390) which is aligned with the direction formed by the ends of the leading edge (386), or which is slightly inclined with respect to this direction, for example inclined by less than 15°.

5. Turbomachine according to the preceding claim, wherein the radially external end (386B) of the leading edge (386) of each stator blade (380) coincides with a free end (392) of the separation nozzle, where is offset, along the central axis (X), by a distance (d) less than or equal to 50 millimeters from this free end (392).

6. Turbomachine according to any one of the preceding claims, wherein the blower (22) comprises between 8 and 14 blower blades (22A).

7. Aircraft comprising at least one propulsion unit (10) comprising a turbomachine conforming to one of the preceding claims.

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