FLOW SEPARATION NOZZLE FOR AN AIRCRAFT TURBOMACHINE

Abstract

The invention relates to a flow separation nozzle (307) for an aircraft turbomachine (301). The nozzle (307) is intended to be positioned around a rotor (307) extending longitudinally along a drive axis (X), downstream of a moving blade wheel (305) and upstream of a fixed blade wheel according to the direction of flow in the turbomachine (301). The nozzle (307) is configured to separate a main flow into a primary flow and a secondary flow and is characterized in that a straight line (D) passing through an upstream end and a downstream end of a leading edge of the nozzle (307) forms a non-zero angle (β) with a plane (P0) orthogonal to the driving axis (X) and / or in that the cross-section of the nozzle (307), at the leading edge, has an elliptical shape whose first radius (Re) varies as a function of an azimuth (θ). Figure for the abbreviation: Fig. 3

Description

Title of the invention: FLOW SEPARATION NOZZLE FOR AN AIRCRAFT TURBOMACHINE technical field

[0001] The invention relates to the field of flow management in a dual-flow turbomachine. It relates in particular to a flow separation nozzle for an aircraft turbomachine and a turbomachine equipped with such a nozzle. Previous technique

[0002] Reducing polluting emissions related to the use of aircraft involves, in particular, improving the propulsive efficiency of turbomachines, that is to say, improving the efficiency with which the energy that is communicated to the air passing through a turbomachine is converted into useful thrust.

[0003] In a turbofan engine, the primary flow passes through the high-pressure modules (compressor + turbine) and the low-pressure modules, passing through a chamber where it is heated. The secondary flow passes through a secondary duct (when defined) and is generated by a fan (in the case of a shrouded turbofan engine, for example). It contributes primarily (on the order of 80%) to the engine's thrust.

[0004] A known approach to improve this efficiency is to decrease the compression ratio of the blower (also called fan), which leads to a decrease in the flow velocity at the outlet of the turbomachine and the associated kinetic energy losses.

[0005] One consequence of this approach is the need to treat a higher mass flow rate of air in the low-pressure section in order to ensure a given thrust level. This therefore leads to an increase in the engine's bypass ratio, or BPR (Bypass Ratio), which defines the ratio between the mass flow rate of air passing through the secondary flow (also called the cold flow) and the mass flow rate passing through the primary flow (also called the hot flow).

[0006] This increase in the mass flow rate of the secondary flow necessitates an increase in the diameter of the fan, and consequently, an increase in the external dimensions of the containment housing (or nacelle) that surrounds it. Therefore, to obtain high dilution ratios without the size and weight of the housing becoming a hindrance, so-called unshrouded turbomachinery (without the containment housing) is being considered.

[0007] However, the use of uncased turbomachinery involves overcoming several difficulties. Indeed, the absence of a casing notably results in the absence of The blade retention devices, the lack of acoustic absorption capacity with acoustic treatments located on the casing, and the exposure of the blades to the external environment all contribute to the overall vulnerability of the turbine. Consequently, the absence of a casing makes the assembly highly susceptible to external stresses such as wind and temperature (especially cold).

[0008] Furthermore, in a known manner, an unfaired turbomachine comprises a first rotor called a "propeller wheel" (i.e. a wheel of moving blades) which is followed by a non-rotating grid called a "propeller stator" (i.e. a wheel of fixed blades) and between these two elements is an air inlet supplying, via the primary flow, the power generator of the turbomachine.

[0009] The separation of the so-called main flow (upstream of this air inlet) into a primary flow mainly supplying the power generation section and a secondary flow used for propulsion is achieved by means of a flow separation nozzle. In the prior art, this nozzle is axisymmetric.

[0010] However, given the absence of fairing (retention casing), the main flow is subject to the constraints mentioned above, and, in particular, to possible disturbances related to the incidence of the flow, the possible presence of a crosswind or even a distortion of the flow.

[0011] These disturbances can have consequences which lead to a reduction in the performance or stability of the primary flow or even prevent the correct supply of the primary flow.

[0012] Fig. 1 schematically illustrates the effects of an incidence of a flow relative to an axis X, called the driving axis, which is the longitudinal extent axis of a rotating hub 101 of a turbomachine 103. The rotating hub 101 is configured to rotate around the driving axis X and drive in rotation a wheel of movable blades 105.

[0013] The turbomachine 103 further includes a flow separation nozzle 107 positioned around the rotating hub 101. The rotating hub 101 on the one hand and the nozzle 107 on the other hand define an air inlet for a primary flow which subsequently flows into a vein of a high-pressure body (not shown).

[0014] In the example shown, the upstream flow has an effect on the turbomachine 103, such as during takeoff, with a relative wind upstream of the turbomachine 103 having a velocity Vo that is not collinear with the engine axis X. Since the turbomachine 103 is not ducted, this flow is not fully straightened along the engine axis before reaching the flow separation nozzle 107. Thus, depending on the operating conditions of the turbomachine, the supply of air to the primary flow inlet can be critical depending on the azimuth.

[0015] In what follows, the terms "upstream" and "downstream" are defined with respect to the direction of flow in the operating turbomachine. For example, on the [Fig. 1], the upstream side is on the left and the downstream side on the right. The term "azimuth" defines an angle in the (Y,Z) plane with respect to the Z-axis. The XYZ frame of reference is an orthonormal frame of reference and the Z-axis is vertical with respect to the aircraft's frame of reference.

[0016] Thus, [Fig. 1] illustrates how, depending on the angle of attack α (defined in the (X,Z) plane of the flow incidence relative to the motor axis X), the air intake can be favorable or unfavorable depending on the azimuth. In this case, the air intake in the area designated 12 o'clock (visible in the upper part of the figure) is unfavorable, while the air intake in the area designated 6 o'clock (visible in the lower part of the figure) is favorable. The 6 o'clock and 12 o'clock positions define an azimuth by analogy with a clock face.

[0017] Fig. 2 shows an aircraft turbomachine 201 according to an embodiment of the prior art and according to a profile view on the left and a front view in section (at the level of the AA axis) on the right.

[0018] The turbomachine 201 comprises a rotor 203, a moving blade wheel 205, a fixed blade wheel 207 and a flow separation nozzle 209 located between the two wheels along the longitudinal extent of the turbomachine 201 along the drive axis X. In addition, as in the case of [Fig.1], this turbomachine is an unshrouded turbomachine in which the blades are not surrounded and therefore protected externally by a casing and the flow upstream of the turbomachine is therefore not straightened by a shroud.

[0019] As can be seen in the figure, the air inlet 211 is an annular surface whose inlet section is perpendicular to the motor axis X (i.e. in a plane orthogonal to said "motor axis X"), with internal radius Ri and external radius Re. More precisely, the flow capture surface (i.e. air inlet) is therefore defined as being equal to ir(Re2- Ri2) in the case of an axisymmetric air inlet.

[0020] In the case where the flow has a so-called positive incidence (i.e. with a>0) the vertical velocity component VOjZ which is equal to Vo sin(a) is non-zero, and, as already described with reference to [Fig.1], the annular zone at 6h (according to azimuth 0) presents a favorable situation for supplying air flow to the primary flow and conversely the zone at 12h presents an unfavorable situation.

[0021] In summary, when the flow upstream of the flow separation nozzle has an angle of incidence with respect to the turbomachine and cannot be straightened by said turbomachine (i.e., redirected around the motor axis X), the air intake conditions for the primary flow can be critical depending on the azimuth 0. Summary of the invention

[0022] The present invention proposes a solution to these drawbacks.

[0023] Thus, one objective of the invention is to enable the optimization of the primary flow air intake of a dual-flow turbomachine according to the actual operating conditions of said turbomachine.

[0024] To this end, the invention according to a first aspect relates to a flow separation nozzle for an aircraft turbomachine, said nozzle being intended to be positioned around a rotor extending longitudinally along a drive axis, downstream of a rotating blade wheel and upstream of a fixed blade wheel according to the direction of flow in said turbomachine,

[0025] said nozzle being further configured to separate a main flow, originating from said moving blade wheel, into a primary flow flowing inside said nozzle, and a secondary flow flowing outside said nozzle,

[0026] and being characterized in that a straight line passing through an upstream end and a downstream end of a leading edge of said nozzle forms a non-zero angle with a plane orthogonal to the motor axis and / or in that the section of said nozzle, at the level of said leading edge, has the shape of an ellipse whose first radius varies as a function of an azimuth.

[0027] The spout according to the invention may comprise one or more of the following features, taken individually or in combination with each other:

[0028] - the line forms an angle between +2° and +10° with the plane.

[0029] - the axial position of the leading edge of the beak varies according to a law, depending on the azimuth, included among: a linear law, a parabolic law, a polynomial law and an exponential law.

[0030] - the beak section, at the leading edge, has an elliptical shape whose origin is not located on the drive axis and is for example located at a radial distance from the drive axis between 0 and 0.2D, where D is the diameter of the moving blade wheel of the turbomachine, preferably between 0.01D and 0.1D.

[0031] - the beak section, at the leading edge, has an elliptical shape of which The eccentricity is between 0 and 0.75, preferably between 0.01 and 0.55.

[0032] The invention according to a second aspect further relates to an aircraft turbomachine comprising a rotor extending along a drive axis, a movable blade wheel driven in rotation around the drive axis by said rotor and a fixed blade wheel, located downstream of said movable blade wheel in the direction of flow of a flow in said turbomachine, said turbomachine being characterized in that it further comprises a flow separation nozzle according to the first aspect, positioned around the rotor, downstream of the movable blade wheel and upstream of the fixed blade wheel.

[0033] The turbomachine according to the second aspect may comprise one or more of the following features, taken individually or in combination with each other:

[0034] - the turbomachine is of the unfaired type or of the short nacelle type.

[0035] - the beak section, at the leading edge, has an elliptical shape and the difference The distance between the first ray and the second ray varies monotonically as a function of the azimuth.

[0036] - the turbomachine further comprises means for actuation of the nozzle, configured to drive the rotation of the nozzle around the motor axis. Brief description of the drawings

[0037] The invention will be better understood with the aid of the following description, given solely by way of example and made with reference to the accompanying drawings in which:

[0038] [Fig.1] is a schematic representation of the effects of an incidence of a flow at the level of a flow separation nozzle of an aircraft turbomachine;

[0039] [Fig.2] is a schematic representation of an aircraft turbomachine according to an embodiment of the prior art;

[0040] [Fig.3] is a schematic representation of an aircraft turbomachine according to an embodiment of the invention;

[0041] [Fig. 4] is a profile view of a flow separation nozzle according to different embodiments of the invention

[0042] [Fig. 5] is a schematic representation of an air inlet section of a primary flow of an aircraft turbomachine according to different embodiments of the invention; and,

[0043] [Fig.6] is a schematic representation of an aircraft turbomachine according to an embodiment of the invention. Description of the implementation methods

[0044] With reference to [Fig.3], we will now describe an aircraft turbomachine according to one embodiment of the invention.

[0045] The turbomachine 301 comprises a rotor 303 which extends longitudinally along an axis X, called the drive axis. A movable blade wheel 305 is driven in rotation about the drive axis X by the rotor 303. In other words, the drive axis X is the axis of revolution of the rotor 303 on which the movable blade wheel 305 is mounted.

[0046] The turbomachine 301 also includes a fixed blade wheel which is not shown in this figure but which conforms to a fixed blade wheel of a prior art turbomachine such as that shown in [Fig.2].

[0047] In this example, the turbomachine 301 is an unshrouded turbomachine; however, the invention also applies to other types of turbomachines in which an azimuthal distortion in the inlet flow of the primary stream is present, such as, for example, a so-called short-nacelle turbomachine, that is to say, a turbomachine in which the dimensions of the nacelle (of the casing) are such that the upstream flow The turbomachine cannot necessarily be straightened if it is not collinear with the engine axis. For example, it may be a turbomachine for which L / D < 0.5, preferably L / D < 0.35, where L is the distance between the leading edge of the nacelle and the leading edge of the moving blade wheel at the radially outer end of the blades and D is the diameter of the moving blade wheel.

[0048] The fixed blade wheel is therefore located downstream of the moving blade wheel 305 according to the direction of flow of the flow in the turbomachine (i.e. the direction of the arrow representing the motor axis X in the figure).

[0049] The turbomachine 301 further includes a flow separation nozzle 307 positioned around the rotor 303 (i.e. radially outside the rotor 303), downstream of the moving blade wheel 305 and upstream of the fixed blade wheel.

[0050] The nozzle 307 is configured to separate a main flow, originating from the moving blade wheel 305, into a primary flow flowing inside the nozzle 307, and a secondary flow flowing outside the nozzle 307. The terms "inside" and "outside", and "internal" and "external", are defined radially with respect to the drive axis X.

[0051] The main flow is the flow which comes from the upstream of the turbomachine 301 and which passes through the moving blade wheel 305 before being separated by the nozzle 307 into two distinct flows: the primary flow and the secondary flow.

[0052] The primary flow then flows into a channel formed between the outer periphery of the rotor 303 and the inner periphery of the nozzle 307 while the secondary flow flows along the outer periphery of the nozzle 307. In other words, the space between the nozzle 307 and the rotor 303, at the level of the upstream edge of the nozzle, called the leading edge, defines the dimensions of the air inlet of the primary flow.

[0053] In the example shown, the nozzle 307 is shown in two distinct positions. A position 307a which corresponds to a prior art nozzle and a position 307b which illustrates an embodiment of a flow separation nozzle according to the invention.

[0054] Thus, the nozzle in position 307b is defined by the fact that a straight line D which passes through an upstream end of its leading edge (visible in the upper part of the figure) and through a downstream end of its leading edge (visible in the lower part of the figure) forms a non-zero angle [3 with a plane PO which is orthogonal to the motor axis X.

[0055] In other words, the axial position of the leading edge of the nozzle 307, that is to say its position along the motor axis X, varies as a function of the azimuth 0 (as defined above with reference to [Fig.2]) so that the leading edge has an inclination with respect to the plane PO perpendicular to the motor axis X. This inclination makes it possible to modify the air inlet section of the primary flow.

[0056] By way of example, in a particular embodiment, the leading edge of the beak may be further forward (i.e. positioned further upstream) at 12 o'clock than at 6 o'clock, for example to adapt to a positive incidence of the flow (i.e. with a > 0) relative to the turbomachine.

[0057] Furthermore, [Fig. 3] shows a turbomachine in profile and illustrates an example in which the leading edge inclination defined by the angle [3] is achieved relative to the vertical (i.e., to the azimuth at 12 o'clock). However, those skilled in the art will appreciate that in different embodiments, this angle can correspond to an inclination (relative to the PO plane) along any azimuth.

[0058] Indeed, the turbomachine may have to withstand distortion (of the flow) due, for example, to a crosswind during takeoff and landing. This distortion potentially generates angles of attack in any direction. Thus, the maximum and minimum angles of attack can be located at 10 o'clock and 4 o'clock, or at 2 o'clock and 8 o'clock, respectively. This therefore necessitates robustness at an angle of attack that varies with the azimuth.

[0059] Thus, advantageously, it is possible to optimize the inclination according to critical conditions associated with an identified incidence of the main flow.

[0060] By way of non-limiting example, a value [3] defining an angle from the plane PO to a line D through which the extremum points (i.e. the most upstream and the most downstream) of the leading edge pass can be between +2° and +10° (according to the trigonometric direction or according to the anti-trigonometric direction).

[0061] By way of example, [Fig.4] shows the embodiment described so far in which, from left to right, the angle [3] has respectively the value of 5°, 10° and 15°.

[0062] In addition, the axial position of the leading edge of the nozzle can vary according to a law including: a linear law, a parabolic law, a polynomial law and an exponential law as a function of the azimuth 0. Such a law governs the way in which the points of the leading edge are connected (i.e. the shape of the curve which connects them).

[0063] Figure 5 shows other embodiments of the invention in which the section of the nozzle 307, at the leading edge, has an elliptical shape (section varying according to the azimuth 0). The section being understood here as an annular line corresponding to the leading edge of the nozzle in a plane orthogonal to the motor axis.

[0064] In particular, [Fig. 5] schematically shows the shape of the outer periphery of the rotor 303 (Ri, in this case a circle) and the shape of the leading edge of the flow-separating nozzle 307 (Re, in this case an ellipse) for three distinct embodiments. The space between the two radii Re and Ri thus corresponds to the primary flow inlet section. Furthermore, the center of the circle and the ellipse may be the same or different.

[0065] In this configuration, the shape of the leading edge is no longer axisymmetric and characterized by a constant radius Re but by an oval-shaped line where Re(0) is a function of the azimuth angle 0.

[0066] Furthermore, in the non-limiting examples shown, Re(0) defines an ellipse that is offset with respect to the X-axis. In this case, Re(0) is defined as follows with respect to the origin O of the ellipse:

[0067] [Math.l]

[0068] where b corresponds to the semi-minor axis and e is the eccentricity such that:

[0069] [Math.2]

[0070] Where a corresponds to the semi-major axis of the ellipse.

[0071] Besides the displacement of the origin O of the ellipse relative to the motor axis X, as illustrated by the three embodiments shown in [Fig.5], according to the position of the semi-major axis of the ellipse a (direction A), the ellipse can be preferentially oriented along a direction where the supply of the primary flow is considered critical.

[0072] Thus, in the three examples shown, the area where the supply is judged to be critical is symbolized by the oval 409 and corresponds respectively to a first configuration (on the left) in which the primary flow air inlet section is increased to 3h (for example in response to the effects of a crosswind coming from the left), to a second configuration in which the primary flow air inlet section is increased to 12h (for example in response to an incidence of the flow relative to the turbomachine as shown with reference to [Fig.1]) and a third configuration in which the primary flow air inlet section is increased to Ih.

[0073] Generally speaking, a person skilled in the art will be able to adapt the shape of the ellipse, its orientation and the position of its origin to predetermined critical conditions.

[0074] By way of non-limiting example, in a particular embodiment, the origin O of the ellipse that defines the shape of the section of the nozzle 307 is not located on the drive axis X, but is located at a radial distance from the drive axis X of between 0 and 0.2D of said drive axis X, where D is the diameter of the moving blade wheel of the turbomachine. More precisely, this distance is between 0.01D and 0.1D.

[0075] Advantageously, this embodiment allows for more precise targeting of a specific azimuth at which the air inlet section must be increased.

[0076] In another particular embodiment, the section of the beak 307, at the leading edge, has an elliptical shape whose eccentricity e is between 0 and 0.75 and preferably between 0.01 and 0.55.

[0077] Figure 6 also illustrates this latter approach in which the shape of the leading edge section of the flow separation nozzle 307 (in position 307b) illustrating the invention) is modified according to the azimuth in order to locally increase the air inlet section of the primary flow.

[0078] Furthermore, this figure illustrates the fact that such a modification can be made while maintaining an iso axial position, that is to say so that the leading edge of the nozzle 307 extends in the plane PO orthogonal to the motor axis X contrary to the embodiment presented with reference to [Fig.3].

[0079] However, those skilled in the art will appreciate that, in different embodiments, the two approaches presented so far, which both contribute to a modification of the primary flow air inlet section but in different ways, can be combined according to the desired effects related to the known constraints.

[0080] Furthermore, in a particular case, the difference between the radius Re and the radius Ri can vary monotonically as a function of the azimuth 0. In other words, the difference Re(0) - Ri(0) can follow a strictly monotonic law between min{S(0)} and max{S(0)}, that is to say that the difference between the internal radius Ri and external radius Re increases in a strictly monotonic way between the azimuthal positions having the minimum section and the maximum section at the level of the inlet of the primary flow.

[0081] In the embodiments described so far, the nozzle is fixed in the sense that the air inlet cross-section cannot be modified in real time. The shape and / or inclination of the nozzle may have been adapted (optimized) beforehand according to known operating conditions.

[0082] However, in a particular embodiment, the turbomachine 301 may also include means for actuation of the nozzle 307, configured to drive the rotation of the nozzle 307 around the motor axis X, such as, for example, an electric motor.

[0083] Advantageously, the primary flow air inlet section can thus be adapted in real time to known critical incidence conditions, observed or measured.

[0084] In conclusion, the invention makes it possible to make the primary flow air supply more robust (and therefore the control of the dilution rate) at the flow separation nozzle, in particular in flight phases likely to present significant distortions at the level of the main flow.

Claims

Demands

1. A flow separation nozzle (307) for an aircraft turbomachine (301), said nozzle (307) being intended to be positioned around a rotor (307) extending longitudinally along a drive axis (X), downstream of a moving blade wheel (305) and upstream of a fixed blade wheel in the direction of flow in said turbomachine (301), said nozzle (307) further being configured to separate a main flow, originating from said moving blade wheel (305), into a primary flow flowing inside said nozzle (307), and a secondary flow flowing outside said nozzle (307), and being characterized in that a straight line (D) passing through an upstream end and a downstream end of a leading edge of said nozzle (307) forms a non-zero angle (|3) with a plane (PO) orthogonal to the motor axis (X) such that the leading edge is inclined at said angle (|3), relative to the plane (PO) orthogonal to the motor axis,according to a determined azimuth of said beak (307) and / or in that the section of said beak (307), at the level of said leading edge, has an elliptical shape whose first radius (Re) varies according to an azimuth (0).

2. Flow separation nozzle (307) for an aircraft turbomachine (301) according to claim 1, wherein the line (D) forms an angle (|3) between +2° and +10° with the plane (PO).

3. Flow separation nozzle (307) for an aircraft turbomachine (301) according to claim 1 or claim 2, wherein the axial position of the leading edge of said nozzle (307) varies according to a law, as a function of the azimuth (0), comprising: a linear law, a parabolic law, a polynomial law and an exponential law.

4. Flow separation nozzle (307) for an aircraft turbomachine (301) according to any one of the preceding claims, wherein the section of said nozzle (307), at the leading edge, has an elliptical shape whose origin (O) is not located on the drive axis (X), and is for example located at a radial distance from the drive axis (X) of between 0 and 0.2D, where D is the diameter of the moving blade wheel of the turbomachine, preferably between 0.01D and 0.1D.

5. A flow separation nozzle (307) for an aircraft turbomachine (301) according to any one of the preceding claims, in which the section of said beak (307), at the level of the leading edge, has an elliptical shape whose eccentricity (e) is between 0 and 0.75, preferably between 0.01 and 0.

55.

6. Aircraft turbomachine (301) comprising a rotor (303) extending along a drive axis (X), a movable blade wheel (305) driven in rotation about the drive axis (X) by said rotor (305) and a fixed blade wheel (207), located downstream of said movable blade wheel (305) in the direction of flow of a stream in said turbomachine (301), said turbomachine (301) being characterized in that it further comprises a nozzle (307) for separating a stream according to any one of the preceding claims, positioned around the rotor (303), downstream of the movable blade wheel (305) and upstream of the fixed blade wheel (207).

7. Aircraft turbomachine (301) according to claim 6, said turbomachine (301) being of the unfaired type or of the short nacelle type.

8. Aircraft turbomachine (301) according to any one of claims 6 or 7, wherein the section of the nozzle (307), at the leading edge, has an elliptical shape and the difference between the first radius (Re) and the second radius (Ri) varies monotonically as a function of the azimuth (9).

9. Aircraft turbomachine (301) according to any one of claims 6 to 8, said turbomachine further comprising nozzle (307) actuating means configured to drive the rotation of said nozzle (307) about the motor axis (X).

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