Turbine engine assembly comprising a device for varying the pitch of the propeller blades and an improved lubrication system

Abstract

The invention relates to an aircraft turbine engine comprising a propeller constituting an unducted fan, the turbine engine comprising a fan casing (55) defining a lubrication chamber, the turbine engine comprising a system comprising a wheel (60) that is able to rotate, relative to the fan casing (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) arranged on a first face (60A) of the wheel (60), and an outlet (62B) arranged 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 arranged at a first distance (R1) from the axis of rotation (X) and the outlet (62B) is arranged at a second distance (R2) from the axis of rotation (X), the second distance (R2) being strictly less than the first distance (R1).

Description

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 to be powered by at least one propulsion system incorporating a turbomachine, such as a turbofan engine. Each propulsion system is attached to the aircraft by a pylon, usually located under or on a wing, or at the level of the aircraft's fuselage. A turbofan engine primarily consists of a gas generator and a fan.

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

[0004] The gas generator includes, 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 enters a primary gas circulation channel passing through the turbojet's gas generator.

[0006] In the case of a shrouded fan, the secondary flow flows in a secondary stream surrounding the gas generator. The secondary stream is bounded, 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 turbofan engine. A portion of the secondary stream is further bounded radially outwards through a fan casing surrounding the fan, and through an intermediate casing located downstream of the fan casing. In the case of an uncased 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 unshrouded fan, it can be advantageous to incorporate a variable fan blade pitching system. In such a case, the fan blades are movable around a radial axis, and a mechanism allows the pitching of each blade to be varied, thus changing the blade angle relative to a radial plane. A cyclic blade pitching system can also be incorporated.

[0008] Timing control systems include components that require lubrication to function correctly. 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. Installing blower blades with a relatively large diameter can impede the return of oil by gravity.

[0009] Therefore, there is a need for an architecture that allows for a lubrication system that enables optimal oil circulation, and in particular ensures the return of oil located at the point where the blower blades are installed.

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

[0011] To this end, the invention relates to an aircraft turbomachine assembly, the turbomachine comprising a propeller constituting an unfaired fan generating an airflow called the secondary flow, the turbomachine comprising a gas generator through which a gas flow called the 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 circulation of lubricant within the lubrication chamber, the lubrication system comprising a rotating wheel relative to the blower 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 disposed at a first distance from the axis of rotation of the wheel and the outlet is disposed at a second distance from the axis of rotation of the wheel, 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 closer to the central axis, thereby bringing it closer to that axis. In an advantageous embodiment, the internal channels are arranged helically, which, thanks to the rotation of the wheel, promotes the circulation of the lubricant within the internal channels. In another 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, so as to bring the lubricant closer to that 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 characteristic, each internal channel extends along a helical curve whose axis coincides with the axis of rotation of the wheel.

[0015] According to one characteristic, 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 characteristic, the first face of the wheel is positioned 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 characteristic, 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 relative 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 characteristic, 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] Depending on one characteristic, the propeller has between 8 and 14 blades.

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

[0023] Figure 1 represents an aircraft equipped with a propulsion system comprising a turbomachine assembly according to the invention.

[0024] Figure 2 shows a schematic cross-sectional view of a turbomachine assembly according to the invention.

[0025] Figure 3 is a partial perspective view of the fan casing of the turbomachine assembly in Figure 2.

[0026] Figure 4 is a partial cross-sectional view of the blower housing of the turbomachine assembly in Figure 2.

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

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

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

[0030] Figure 1 depicts an aircraft 100, in this example an airplane, equipped with two propulsion systems comprising a turbomachine assembly 10, namely one propulsion system per wing 101, with only one propulsion system and one wing 101 shown in Figure 1. In one variant, the aircraft 100 can be equipped with more than one propulsion system per wing 101, each wing 101 having the same number of propulsion systems. The reference symbol "A" designates the axis of the fuselage 102 of the aircraft 100. The propulsion system can be configured to propel the aircraft 100 at a cruising speed between Mach 0.7 and Mach 0.9.

[0031] Figure 2 shows a schematic cross-sectional view of a turbomachine assembly 10, according to plane II of Figure 1. The turbomachine assembly 10 extends along a central axis X, and includes a propulsion module 20, a gas generator 30, and, in the example, a speed reduction device 40. When the turbomachine 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, forming 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 include an equal or smaller number of 24A blades, for example between 8 and 14 24A blades. The setting of the 24A blades of the rectifier 24 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 turbomachine assembly 10. This allows for an annular air inlet within the gas generator 30 coaxial with the central axis X, 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 turbomachine assembly 10 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 can include the low 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. In one variant, the low-pressure shaft 33A and high-pressure shaft 33B can be co-rotating, i.e., configured to rotate relative to each other in the same direction around the central axis X. In another variant, the low-pressure shaft 33A and high-pressure shaft 33B can be counter-rotating, i.e., configured to rotate relative to each other in opposite directions around the axis. central X. The rotational speed of the low-pressure body 33A may be lower than the rotational speed of the high-pressure body 33B.

[0035] According to an unshown variant, the turbomachine 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 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 can 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 can have at least 2 stages and at most 5 stages, for example 2 stages; the high-pressure compressor 32B can have between 8 and 11 stages (only two stages are shown for clarity in the figure); the high-pressure turbine 36B can have 2 stages; and the low-pressure turbine 36A can have between 3 and 8 stages (only two stages are shown for clarity in 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 to 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 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 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 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 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 is provided for the blades 22A of the propeller 22.

[0040] In operation, an airflow F (see figure 2) entering the turbomachine assembly 10 passes through the propeller 22 and is then divided into a primary flow Fl and a secondary flow F2, which flow upstream to downstream within the turbomachine 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 36A, and 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 rotation of the high-pressure and low-pressure turbines 36B and 36A, and therefore the shafts 33A and 33B.

[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 turbomachine 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 in Figure 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 includes a wheel 60 (visible in Figures 3 to 7), which rotates relative to the fan housing 55 about an axis of rotation coinciding with the central axis X of the turbomachine. The wheel 60 has 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 60A. The wheel 60 is configured so that, for each internal channel 62, the inlet 62A is located at a first distance RI from the axis of rotation X, and the outlet 62B is located at a second distance R2 from the axis of rotation X, the second distance R2 being strictly less than the first distance RI. Thus, for each internal channel 62, the output 62B is closer to the rotation axis 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 in a helix. 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 (shown by arrow F3 in Figure 5). This configuration promotes the capture of lubricant at the inlets 62A of the internal channels 62, and also promotes the circulation of lubricant within the internal channels 62, then the discharge of 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 in Figure 4, and in Figures 6 and 7. This path is shown in Figure 4 by the arrow F4. As shown in Figure 4 and Figure 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 discharged 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 discharged through 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 relative 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 21A of a hub 21 on which the blades 22A of the blower are fixed, the wall 21A of the hub 21 having openings 21B allowing the passage of lubricant to the inlets 62A 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. An aircraft turbomachine assembly (10), the turbomachine comprising a propeller (22) constituting an unfaired fan generating an airflow referred to as the secondary flow (F2), the turbomachine comprising a gas generator (30) through which a gas flow referred to as the 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 for generating lubricant circulation within 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 outlet (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 inlet (62A) is disposed at a first distance (RI) from the axis of rotation (X) of the wheel (60) and the outlet (62B) is disposed at a second distance (R2) from the axis of rotation (X) of the wheel (60), the second distance (R2) being strictly less than the first distance (RI), each internal channel (62) extending along a helical curve with its axis coinciding with the axis of rotation (X) of the wheel (60), the first face (60a) of the wheel (60) being located upstream of the second face (60B) with respect to the direction of gas flow in the turbomachine,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).

2. Turbomachine assembly (10) according to claim 1, 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).

3. Turbomachine assembly (10) 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 fan blades (22A) are attached, the wall (21A) of the hub (21) having openings (21B) allowing the passage of lubricant to the inlets (62A) of the internal channels (62) of the wheel (60).

4. Turbomachine assembly (10) 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).

5. Turbomachine assembly (10) according to the preceding claim, wherein the flow surface (60D) has a flared shape from upstream to downstream.

6. Turbomachine assembly (10) 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 includes, for example, at least 8 internal channels (62).

7. Turbomachine assembly (10) according to any one of the preceding claims, wherein the propeller (22) comprises between 8 and 14 blades (22A).

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

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