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. Prior art

[0002] Reducing polluting emissions linked 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 which is communicated to the air passing through a turbomachine is converted into useful thrust force.

[0003] In a dual-flow turbomachine, the primary flow passes through the high-pressure modules (compressor + turbine) and the low-pressure modules, passing through a chamber to be heated there. The secondary flow passes through a secondary vein (when it is defined) and is generated by a fan (in the case of a shrouded turbomachine, for example). It contributes mainly (around 80%) to the engine thrust.

[0004] A known approach to improving this efficiency consists of reducing the compression ratio of the blower (also called fan), which results in a reduction in the flow velocity at the outlet of the turbomachine and the kinetic energy losses linked to it.

[0005] A consequence of this approach is the need to have the low pressure part process a greater mass flow of air in order to ensure a given thrust level. This therefore leads to increasing the engine's dilution ratio or BPR (ByPass Ratio) which defines the ratio between the mass flow of air passing through the secondary flow (also called cold flow), and the mass flow passing through the primary flow (also called hot flow).

[0006] This increase in the mass flow rate of the secondary flow has the consequence of requiring an increase in the diameter of the fan, and consequently, an increase in the external dimensions of the retention casing (or nacelle) which surrounds it. Therefore, to obtain high dilution rates without the dimensions and weight of the casing becoming a handicap, so-called unducted turbomachines (without the retention casing) are envisaged.

[0007] However, the use of unshrouded turbomachines involves overcoming several difficulties. Indeed, the absence of a casing results in particular in the absence of dis positive blade retention properties, the lack of sound absorption capacity with acoustic treatments located on the casing and the exposure of the blades to the external environment. Consequently, the absence of a casing makes the assembly very sensitive to external constraints such as wind and temperature (especially cold) for example.

[0008] Furthermore, in a known manner, an unducted turbomachine comprises a first rotor called a “propulsive propeller” (i.e. a wheel of moving blades) which is followed by a non-rotating grille called a “propeller stator” (i.e. a wheel of fixed blades) and between these two elements there 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 part and a secondary flow used for propulsion is achieved by means of a flow separation nozzle. In the state of the 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 linked to the incidence of the flow, to the possible presence of a crosswind or even to 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 which prevent the correct supply of the primary flow.

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

[0013] The turbomachine 103 further comprises 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 delimit 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 incidence relative to the turbomachine 103, such as for example at takeoff, with a relative wind upstream of the turbomachine 103 having a speed Vo, not collinear with the engine axis X. The turbomachine 103 being unducted, this flow is not entirely straightened in the engine axis before reaching the flow separation nozzle 107. Thus, depending on the operating points of the turbomachine, the supply of the air inlet of the primary flow can be critical depending on the azimuth.

[0015] In the following, the terms "upstream" and "downstream" are defined in relation to the direction of flow circulation in the turbomachine in operation. For example, on the [Fig.l], upstream is on the left and downstream is on the right. The term "azimuth" defines an angle in the (Y,Z) plane relative to the Z axis. The XYZ reference frame is an orthonormal reference frame and the Z axis is vertical relative to the aircraft reference frame.

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

[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 axis AA) on the right.

[0018] The turbomachine 201 comprises a rotor 203, a wheel of moving blades 205, a wheel of fixed blades 207 and a nozzle 209 for separating the flow located between the two wheels along the engine axis X of longitudinal extent of the turbomachine 201. Furthermore, as in the case of [Fig.l], this turbomachine is an unducted turbomachine for 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 fairing.

[0019] As can be seen in the figure, the air inlet 211 is an annular surface whose inlet section is perpendicular to the engine axis X (i.e. in a plane orthogonal to said “engine axis X”), with an internal radius Ri and an 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 has already been described with reference to [Fig.l], the annular zone at 6 o'clock (along the azimuth 0) presents a favorable situation for the supply of air flow to the primary flow and conversely the zone at 12 o'clock presents an unfavorable situation.

[0021] In summary, when the flow upstream of the flow separation nozzle has an incidence relative to the turbomachine and cannot be straightened by said turbomachine (i.e. reoriented along the engine 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 provides a solution to these drawbacks.

[0023] Thus, an objective of the invention is to make it possible to optimize the air intake of the primary flow of a dual-flow turbomachine according to real conditions. of use 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 an engine axis, downstream of a wheel of moving blades and upstream of a wheel of fixed blades according to the direction of flow of the flow in said turbomachine,

[0025] said nozzle being further configured to separate a main flow, coming from said mobile 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 slat forms a non-zero angle with a plane orthogonal to the engine axis and / or in that the section of said slat, at the level of said leading edge, has the shape of an ellipse of which a first radius varies as a function of an azimuth.

[0027] The spout according to the invention may comprise one or more of the following characteristics, taken in isolation from one another or in combination with one another:

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

[0031] - the section of the beak, at the level of 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 an engine axis, a wheel of moving blades driven in rotation about the engine axis by said rotor and a wheel of fixed blades, located downstream of said wheel of moving blades according to the direction of flow of a flow in said turbomachine, said turbomachine being characterized in that it further comprises a nozzle for separating a flow according to the first aspect, positioned around the rotor, downstream of the wheel of moving blades and upstream of the wheel of fixed blades.

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

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

[0035] - the section of the beak, at the level of the leading edge, has an elliptical shape and the difference between the first ray and the second ray varies monotonically with the azimuth.

[0036] - the turbomachine further comprises means for actuating 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 with reference to the appended drawings in which:

[0038] [Fig.l] is a schematic representation of the effects of an incidence of a flow at the level of a separation nozzle of a flow 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 one embodiment of the invention;

[0041] [Fig.4] is a side 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 one embodiment of the invention. Description of the embodiments

[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 engine axis. A movable blade wheel 305 is driven in rotation about the engine axis X by the rotor 303. In other words, the engine 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 comprises a fixed blade wheel which is not shown in this figure but which conforms to a fixed blade wheel of a turbomachine of the state of the art such as that shown in [Fig.2].

[0047] In this example, the turbomachine 301 is an unducted turbomachine, however the invention also applies to other types of turbomachine for which an azimuthal distortion in the inlet flow of the primary flow is present, such as, for example, a so-called short-nacelle turbomachine, i.e. a turbomachine for which the dimensions of the nacelle (of the casing) are such that the flow upstream of 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 impeller at the radially outer end of the blades and D is the diameter of the impeller.

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

[0049] The turbomachine 301 further comprises 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, coming from the mobile 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 relative to the engine axis X.

[0051] The main flow is the flow which comes from upstream of the turbomachine 301 and which passes through the mobile 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 in a vein 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 located between the nozzle 307 and the rotor 303, at the most 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 spout 307 is shown in two distinct positions. A position 307a which corresponds to a spout of the state of the art and a position 307b which illustrates an embodiment of a flow separation spout according to the invention.

[0054] Thus, the beak 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 engine 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 engine 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 relative to the plane PO perpendicular to the engine 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 nozzle may be more advanced (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 inclination of the leading edge defined by the angle [3 is carried out relative to the vertical (ie 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 resist distortion (of the flow) due for example to a crosswind during the takeoff and landing phases. This distortion potentially causes incidences in any direction. Thus, the maximum and minimum incidences may respectively be positioned at 10 o'clock and 4 o'clock or at 2 o'clock and 8 o'clock. This therefore results in a need for robustness to an incidence that varies depending on the azimuth.

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

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

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

[0062] Furthermore, the axial position of the leading edge of the beak may vary according to a law included among: 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] [Fig. 5] shows other embodiments of the invention in which the section of the slat 307, at the leading edge, has an elliptical shape (section variable as a function of the azimuth 0). The section being understood here as an annular line corresponding to the leading edge of the slat in a plane orthogonal to the engine axis.

[0064] In particular, [Fig. 5] shows schematically the shape of the external periphery of the rotor 303 (Ri in this case a circle) and the shape of the leading edge of the flow separation nozzle 307 (Re in this case an ellipse) for three distinct embodiments. The intermediate space between the two radii Re and Ri therefore corresponds to the inlet section of the primary flow. Furthermore, the center of the circle and the ellipse may be identical 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 angle 0 of the azimuth.

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

[0067] [Math.l] Re2 (0) b2 l-t'-cos2 ( 8 )

[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] In addition to 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], depending on the position of the semi-major axis of the ellipse a (direction A), the ellipse can be oriented preferentially in a direction where the supply of the primary flow is judged to be critical.

[0072] Thus, in the three examples shown, the zone 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 air inlet section of the primary flow 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 air inlet section of the primary flow is increased to 12h (for example in response to an incidence of the flow relative to the turbomachine as shown with reference to [Fig.l]) and a third configuration in which the air inlet section of the primary flow is increased to 1h.

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

[0074] As a non-limiting example, in a particular embodiment, the origin O of the ellipse which defines the shape of the section of the nozzle 307 is not located on the engine axis X, but it is located at a radial distance from the engine axis X of between 0 and 0.2D of said engine 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 makes it possible to further target a determined azimuth at which the air inlet section must be increased.

[0076] In another particular embodiment, the section of the beak 307, at the level of 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] [Fig.6] also illustrates this latter approach in which the shape of the section of the leading edge of the flow separation nozzle 307 (in position 307b illustrating the invention) is modified as a function of 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 axial iso position, that is to say so that the leading edge of the beak 307 extends in the plane PO orthogonal to the engine axis X unlike 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 air inlet section of the primary flow but in a different manner, can be combined depending on the desired effects linked 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 strictly monotonically between the azimuthal positions having the minimum section and the maximum section at the level of the primary flow inlet.

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

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

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

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

Claims

Claims

1. A nozzle (307) for separating a flow for an aircraft turbomachine (301), said nozzle (307) being intended to be positioned around a rotor (307) extending longitudinally along an engine axis (X), downstream of a wheel of moving blades (305) and upstream of a wheel of fixed blades according to the direction of flow of the flow in said turbomachine (301), said nozzle (307) being further configured to separate a main flow, coming from said wheel of moving blades (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 engine axis (X) and / or in that the section of said beak (307), at the level of said leading edge, has the shape of an ellipse of which a first radius (Re) varies as a function of an azimuth (0).

2. Flow separation nozzle (307) for an aircraft turbomachine (301) according to claim 1, in which the straight line (D) forms an angle (|3) of 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), included among: 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 the shape of an ellipse whose origin (O) is not located on the engine axis (X), and is for example located at a radial distance from the engine axis (X) of between 0 and 0.2D, where D is the diameter of the moving blade wheel of the turbomachine, preferably of between 0.01D and 0.1D.

5. Flow separation nozzle (307) for an aircraft turbomachine (301) according to any one of the preceding claims, in which the section of said nozzle (307), at 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 an engine axis (X), a movable blade wheel (305) driven in rotation about the engine 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 flow in said turbomachine (301), said turbomachine (301) being characterized in that it further comprises a nozzle (307) for separating a flow 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 non-ducted type or of the short nacelle type.

8. Aircraft turbomachine (301) according to one of claims 6 or 7, in which the section of the beak (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 (0).

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

Images