Turbomachine assembly incorporating a variable propeller blade pitch device and an improved lubrication system

Abstract

The invention relates to an aircraft turbomachine comprising a propeller constituting an unfaired fan, the turbomachine comprising a fan housing (55) defining a lubrication chamber, the turbomachine comprising a system comprising a wheel (60) rotatable relative to the fan housing (55) about an axis of rotation (X), the wheel (60) comprising a set of internal channels (62), each internal channel (62) comprising an inlet (62A) disposed on a first face (60A) of the wheel (60), and an outlet (62B) disposed on a second face (60B), opposite the first face, the wheel (60) being configured such that, for each internal channel (62), the inlet (62A) is disposed at a first distance (R1) from the axis of rotation (X) and the outlet (62B) is disposed at a second distance (R2) from the axis of rotation (X), the second distance (R2) being strictly less than the first distance (R1). Figure 4.

Description

Title of the invention: Turbomachine assembly comprising a variable propeller blade pitching device and an improved lubrication system. Technical field

[0001] The present invention relates to the field of aeronautics, and more particularly to a propulsion assembly comprising a double-flow turbomachine and an unfaired fan. 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 pylon 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, a 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] In the case of an unshod fan, it may be advantageous to provide a variable fan blade pitching system. In such a case, the fan blades are movable around a radial axis, and a mechanism allows The adjustment of each blade's position allows for varying the blade's angle relative to a radial plane. Furthermore, a cyclic blade adjustment system can be implemented.

[0008] Timing control systems include components that require lubrication to ensure proper operation. The blower housing thus defines an oil chamber, within which oil circulates to the various components requiring lubrication. The entire oil chamber is filled with an oil mist generated by nozzles. The oil returns to the nozzles by gravity, and a minimum slope must be maintained for this purpose. The relatively large diameter of the blower blades can impede the return of oil by gravity.

[0009] There is therefore a need for an architecture that provides a lubrication system allowing optimal oil circulation, and in particular ensuring the return of the oil located at the level of the blower blade implantation.

[0010] The objective of the present invention is to propose a turbomachine that meets this need. Description of the invention

[0011] To this end, the invention relates to an aircraft turbomachine assembly 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 propeller comprising variable pitch blades, the turbomachine comprising a fan casing defining a lubrication chamber, the turbomachine comprising a lubrication system enabling the generation of lubricant circulation in the lubrication chamber, the lubrication system comprising a wheel rotating relative to the fan casing, around an axis of rotation coinciding with the central axis of the turbomachine, the wheel comprising a set of internal channels, each internal channel comprising an inlet disposed on a first face of the wheel, and comprising an outlet disposed on a second face of the wheel, opposite the first face,the wheel being configured so that, for each internal channel, the inlet is located at a first distance from the axis of rotation of the wheel and the outlet is located at a second distance from the axis of rotation of the road, the second distance being strictly less than the first distance.

[0012] Thus, by providing a rotating wheel with internal channels, each channel configured so that its inlet and outlet are located at different distances from the central axis of the turbomachine, the invention creates a path that brings the lubricant to a smaller distance from the central axis, thereby bringing the lubricant closer to this axis. In one embodiment Advantageously, the internal channels are arranged in a helix, which, thanks to the rotation of the wheel, promotes the circulation of lubricant within the internal channels. In an advantageous embodiment, the internal channels are arranged in a conical helix. In other words, the diameter of the helix in which the internal channels are arranged decreases from upstream to downstream along the direction of the central axis of the turbomachine engine, so as to bring the lubricant closer to this axis.

[0013] The turbomachine assembly according to the invention may include one or more of the following optional features, considered alone or in all possible combinations.

[0014] According to one feature, each internal channel extends along a helical curve whose axis coincides with the axis of rotation of the wheel.

[0015] According to one feature, the first face of the wheel is located upstream of the second face with respect to the direction of gas flow in the turbomachine, and the direction of the helical curve along which each internal channel extends from the first face of the wheel to the second face is opposite to the direction of rotation of the wheel.

[0016] According to one characteristic, each internal channel extends along a conical helical curve whose diameter decreases from upstream to downstream along the axis of rotation of the wheel.

[0017] According to one feature, the first face of the wheel is arranged near a wall of a hub to which the blower blades are attached, the hub wall having openings allowing the passage of lubricant to the inlets of the internal channels of the wheel.

[0018] According to one feature, the wheel has an annular wall extending from the second face and forming a flow surface for the lubricant expelled from the outlets of the internal channels, the flow surface being inclined with respect to the axis of rotation of the wheel so as to promote the return of the lubricant downstream of the blower housing.

[0019] According to one characteristic, the flow surface has a flared shape from upstream to downstream.

[0020] According to one feature, the wheel has a number of internal channels greater than or equal to the number of blades of the propeller, and has for example at least 8 internal channels.

[0021] According to one characteristic, the propeller has between 8 and 14 blades.

[0022] The invention also relates to an aircraft comprising at least one propulsion unit including a turbomachine assembly conforming to that defined above. Brief description of the drawings

[0023] Fig. 1 represents an aircraft equipped with a propulsion system according to the invention.

[0024] [Fig.2] Fig.2 represents a schematic cross-sectional view of a propulsion assembly according to the invention.

[0025] [Fig.3] The [Fig.3] is a partial perspective view of the blower housing of the propulsion assembly of the [Fig.1].

[0026] [Fig.4] The [Fig.4] is a partial cross-sectional view of the blower housing of the propulsion assembly of the [Fig.1]

[0027] [Fig.5] The [Fig.5] is a detailed perspective view showing the wheel.

[0028] [Fig.6] The [Fig.6] is a detailed perspective view showing the wheel.

[0029] [Fig.7] The [Fig.7] is a detailed perspective view showing the wheel. Detailed description

[0030] 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. The reference symbol "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.

[0031] 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 a central axis X and comprises a propulsion module 20, a gas generator 30, and, in the example, a speed reduction device 40. When the propulsion assembly 10 is mounted on the aircraft 100, the central axis X is not necessarily parallel to the axis A.

[0032] 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 central axis X. 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.

[0033] The gas generator 30 has a drive shaft 33A. The drive shaft can extend along the central axis X. The propeller shaft 26 can be coaxial with the drive shaft 33A, and their respective axes of rotation can coincide with the central axis X of the propulsion unit 10. This allows for an annular air inlet within the gas generator 30 coaxial with the central axis X, resulting in a relatively simple shape and a degree of rotational symmetry in the outer casing of the gas generator, which tends to reduce potential airflow disturbances. In this example, the gas generator 30 comprises, from upstream to downstream, the gases flowing within the propulsion unit 100, and from upstream to downstream, a compressor 32 (or compressor section 32), a combustion chamber 34, and a turbine 36 (or turbine section 36).

[0034] 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 may comprise the low-pressure and high-pressure turbines 36A and 36B. The low-pressure and high-pressure shafts 33A and 33B may be coaxial. The high-pressure shaft 33B may 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 central axis X. 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 central axis X. The rotational speed of the low-pressure shaft 33A may be lower than the rotational speed of the high-pressure shaft 33B.

[0035] According to an alternative (not shown), the propulsion unit 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 may be located between the low-pressure shaft 33A and the high-pressure shaft 33B. The intermediate shaft and the low-pressure shaft 33B may rotate co- or counter-rotating with respect to each other.

[0036] 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 central axis X (or rotor) and a blade wheel, respectively 32AB, 32BB, 36AB, 36BB, fixed about the central axis X (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 around the central axis X, can be arranged downstream of the combustion chamber 34 and upstream of the high-pressure turbine 36B.

[0037] 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 body 30A) to an inlet of the speed reduction device 40, while the propeller shaft 26 connects an output 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 body 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 arranged, considered along the central axis X, between an upstream end of the drive shaft 33A and a downstream end of the propeller shaft 36.

[0038] 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.

[0039] In the example, a variable timing device 50 of the blades 22A of the propeller 22 is provided.

[0040] In operation, an airflow F (see [Fig.2]) entering the propulsion assembly 10 passes through the propeller 22 and is then divided into a primary flow Fl and a secondary flow F2, which circulate from upstream to downstream within the propulsion assembly 10.

[0041] The primary airflow Fl flows in a channel called the "primary channel" inside the gas generator 30, sometimes also called the primary body, passing successively through the low-pressure compressor 32A, the high-pressure compressor 32B, the combustion chamber 34, the high-pressure turbine 36B, the low-pressure turbine pressure 36A, then through the outlet nozzle. The expansion of the combustion gases downstream of the combustion chamber 34 within the turbine 36 provides the energy to drive the high and low pressure turbines 36B, 36A, and therefore the shafts 33A and 33B, into rotation.

[0042] The secondary airflow F2 flows through the rectifier 24, then along the gas generator 30, outside the gas generator 30. This secondary airflow F2 provides, by reaction, the majority of the thrust generated by the propulsion assembly 10. The secondary airflow F2 can also be used to cool the gas generator 30 from the outside.

[0043] The speed reduction device 40 and the variable timing device 50 described above are integrated into a blower housing 55, which forms a lubricated enclosure.

[0044] The blower housing 55 is partially visible on [Fig.3] which is a partial perspective view of the propulsion module 20.

[0045] According to the invention, the turbomachine includes a lubrication system for generating a circulation of lubricant in the lubrication chamber, in order to lubricate the various elements located in the blower housing 55. The lubrication system is in particular configured to transfer lubricant from a point located at a first distance from the central axis X to a point located at a second distance from the central axis X, the second distance being less than the first distance.

[0046] The lubrication system comprises a wheel 60 that rotates relative to the blower housing 55 about an axis of rotation coinciding with the central axis X of the turbomachine. The wheel 60 comprises a set of internal channels 62. Each internal channel 62 has an inlet 62A located on a first face 60A of the wheel 60, and an outlet 62B located on a second face 60B of the wheel, opposite the first face 60B. The wheel 60 is configured so that, for each internal channel 62, the input 62A is disposed at a first distance RI from the axis of rotation X, and so that the output 62B is disposed at a second distance R2 from the axis of rotation X, the second distance being strictly less than the first distance RL. Thus, for each internal channel 62, the output 62B is closer to the axis of rotation X of the wheel 60 (i.e. the central axis X of the turbomachine) than the input 62A.

[0047] Advantageously, each internal channel 62 is arranged helically. In other words, each internal channel 62 extends along a helical curve whose axis coincides with the axis of rotation X of the wheel 60, i.e., the central axis of the turbomachine. Preferably, the direction of the helix formed by each internal channel 62 from the first face 60A of the wheel 60 to the second face 60B is opposite to the direction of rotation of the wheel (indicated by arrow F3 in [Fig. 5]). This configuration allows to promote the capture of the lubricant at the inlets 62A of the internal channels 62, and also to promote the circulation of the lubricant within the internal channels 62, then the discharge of the lubricant at the outlets 62B of the internal channels 62. Advantageously, each internal channel 62 extends along a conical helical curve, the diameter of which decreases from upstream to downstream along the axis of rotation X of the wheel 60.

[0048] The path of the lubricant within the internal channels 62 is visible in particular on [Fig.4] and on figures 6 and 7. This path is materialized on [Fig.4] by the arrow F4. As shown in [Fig. 4] and [Fig. 7], the lubricant is drawn in at the first face 60A of the wheel 60 at a first height, corresponding to the distance RI between the axis of rotation X of the wheel 60 and the inlet 62A of each internal channel 62. The lubricant is then forced out towards the outlet 62B of the internal channel 62, located at a second height, corresponding to the distance R2 between the axis of rotation X of the wheel 60 and the outlet 62B of each internal channel 62. The distance R2 is less than the distance RI, so the lubricant is forced out of each internal channel 62 at a distance from the axis of rotation X that is less than the distance at which the lubricant is drawn in by each internal channel 62.

[0049] In the example, the impeller 60 has an annular wall 60C extending from the second face 60B and forming a flow surface 60D for the lubricant discharged from the outlets 62B of the internal channels 62. The flow surface 60D is inclined with respect to the axis of rotation X of the impeller 60 so as to promote the return of the lubricant downstream of the blower housing 55, particularly under the action of centrifugal force. In other words, the flow surface 60D has a flared shape (from upstream to downstream of the turbomachine), allowing, during the operation of the turbomachine, the lubricant from the internal channels 62 to be moved downstream under the action of centrifugal force.

[0050] Advantageously, the first face 60A of the wheel 60 is arranged near a wall of a hub 21 on which the blades 22A of the blower are fixed, the wall of the hub 21 having openings 21A allowing the passage of lubricant towards the inlets of the internal channels 62 of the wheel 60. In the example in the figures, the wheel 60 is fixed to the hub 21.

[0051] Advantageously, the wheel 60 has a number of internal channels 62 which is greater than or equal to the number of blades 22A in the propeller 22. The wheel 60 has, for example, between 8 and 20 channels.

Claims

Demands

1. Aircraft turbomachine assembly comprising a propeller (22) constituting an unfaired fan generating an airflow called secondary flow (F2), the turbomachine comprising a gas generator (30) through which a gas flow called primary flow (F1) passes, the propeller (22) comprising variable pitch blades (22A), the turbomachine comprising a fan casing (55) defining a lubrication chamber, the turbomachine comprising a lubrication system enabling the generation of lubricant circulation in the lubrication chamber, the lubrication system comprising a wheel (60) rotatable relative to the fan casing (55) about an axis of rotation (X) coinciding with the central axis of the turbomachine, the wheel (60) comprising a set of internal channels (62), each internal channel (62) comprising an inlet (62A) disposed on a first face (60A) of the wheel (60),and comprising an output (62B) disposed on a second face (60B) of the wheel (60), opposite the first face, the wheel (60) being configured such that, for each internal channel (62), the input (62A) is disposed at a first distance (RI) from the axis of rotation (X) of the wheel (60) and the output (62B) is disposed at a second distance (R2) from the axis of rotation (X) of the road (60), the second distance (R2) being strictly less than the first distance (RI).

2. Turbomachine assembly according to the preceding claim, wherein each internal channel (62) extends along a helical curve with axis coinciding with the axis of rotation (X) of the wheel (60).

3. Turbomachine assembly according to the preceding claim, wherein the first face (60a) of the wheel (60) is located upstream of the second face (60B) with respect to the direction of gas flow in the turbomachine, and wherein the direction of the helical curve along which each internal channel (62) extends from the first face (60A) of the wheel (60) to the second face (60B) is opposite to the direction of rotation of the wheel (60).

4. Turbomachine assembly according to any one of claims 2 and 3, wherein each internal channel (62) extends along a conical helical curve whose diameter decreases from upstream to downstream along the axis of rotation (X) of the wheel (60).

5. Turbomachine assembly according to any one of the preceding claims, wherein the first face (60A) of the wheel (60) is disposed near a wall (21A) of a hub (21) to which the blades (22A) of the blower are attached, the wall of the hub having openings (21B) allowing the passage of lubricant to the inlets (62A) of the internal channels (62) of the wheel (60).

6. Turbomachine assembly according to any one of the preceding claims, wherein the wheel (60) has an annular wall (60C) extending from the second face (60B) and forming a flow surface (60D) for the lubricant discharged from the outlets (62B) of the internal channels (60), the flow surface (60D) being inclined with respect to the axis of rotation (X) of the wheel (60) so as to promote the return of the lubricant downstream of the blower housing (55).

7. Turbomachine assembly according to the preceding claim, wherein the flow surface (60D) has a flared shape from upstream to downstream.

8. Turbomachine assembly according to any one of the preceding claims, wherein the wheel (60) has a number of internal channels (62) greater than or equal to the number of blades (22A) of the propeller (22), and has, for example, at least 8 internal channels (62).

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

10. Aircraft (100) comprising at least one propulsion unit (10) comprising a turbomachine assembly conforming to one of the preceding claims.

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