Unducted air vehicle propulsion unit

Abstract

The invention relates to an aeronautical propulsion unit (100) comprising: - an outer casing (102); - a hub (104) mounted so as to pivot with respect to the outer casing (102) about a main axis (X) extending in an upstream-downstream direction of the aircraft; - a propulsion propeller (106) mounted on the hub (104) so as to be able to pivot with respect to the outer casing (102); and - a fixed flow straightener (112) mounted on the outer casing (102) downstream of the propulsion propeller (106) along the main axis (X), the fixed flow straightener (112) extending about the main axis (X), the fixed flow straightener (112) comprising stator blades (114). For at least one stator blade (114) of the fixed flow straightener (112), for all blade assembly heights (H'), the maximum thickness position is located between 0.1 and 0.5 and / or the maximum thickness position (Xepmax) is strictly decreasing over at least 60% of the blade assembly height H', preferably strictly decreasing over at least 75% of the blade assembly height (H').

Description

Technical Field

[0001] The present invention relates to a ductless aviation propulsion unit for an aircraft, and an aircraft comprising such an aviation propulsion unit. Background Technology

[0002] Efforts to minimize pollutant emissions from air transport focus particularly on all aspects of improving propulsion system efficiency, more specifically on improving propulsion efficiency, which characterizes the effectiveness of converting the energy transferred to the air flowing through the engine into useful thrust.

[0003] The factors that have the greatest impact on propulsion efficiency are those related to the low-pressure section of the propulsion system, which directly contributes to thrust generation. These include the low-pressure turbine, low-pressure transmission system, fan, and the secondary flow that guides the fan's airflow. A well-known guiding principle for improving propulsion efficiency is to reduce the fan's compression ratio, thereby reducing the flow velocity at the engine exit and the associated kinetic energy losses.

[0004] One of the main consequences of reducing the flow velocity at the engine outlet is that the low-pressure section (secondary flow) must handle a higher mass flow rate of air to ensure a given thrust level determined by the aircraft specifications: this therefore leads to an increase in the engine's dilution ratio. The bypass ratio, or BPR, is defined as the ratio of the mass flow through the secondary flow (cold flow) to the mass flow through the main flow (hot flow), specifically the mass flow supplied to the combustion chamber.

[0005] The direct consequence of this increase in secondary flow is the need to increase the fan diameter, and therefore the external dimensions of the housing surrounding the fan, as well as the external dimensions of the nacelle forming the aerodynamic envelope of that housing. To achieve a high dilution ratio, when the housing becomes too large and heavy (causing significant drag), the housing is removed to facilitate a ductless propeller configuration. Several designs are conceivable for ductless turbines, but the present invention relates to a ductless turbine comprising (at least) an upstream propeller with a variable pitch (referred to as an "open fan") and downstream outlet guide vanes with a fixed or variable pitch. In the case of such a turbine with a single propeller, such a turbine is called a ductless single-fan (USF).

[0006] Therefore, the prior art discloses a ductless aviation propulsion unit for an aircraft, which includes: - Outer shell; - The rotor hub is mounted to pivot relative to the outer shell about a main axis extending in the upstream-downstream direction of the aircraft; - A propeller, mounted on a hub, pivots relative to the outer casing; and - An exit guide vane, mounted on the housing and located downstream of the propeller along the main axis, extending about the main axis. The exit guide vane includes stator blades, each of which has: • The pressure side and suction side extending between the leading and trailing edges of the stator blades. • For each cross section of the stator blade perpendicular to the stacking axis, at the blade ring height along the stacking axis, the cross section has: ○ a curve located at the midpoint between the pressure side and the suction side, which is called the skeleton; ○ a distance located between the leading edge and the trailing edge, which is called the chord; and ○ a thickness perpendicular to the skeleton located between the pressure side and the suction side, which is normalized relative to the chord and has a maximum value normalized relative to the chord at a position on the chord.

[0007] This type of turbine can be configured as a "thruster" having a propeller and an outlet guide vane located downstream of the turbine (and mounted at the rear of the aircraft), or as a "puller" having a propeller and an outlet guide vane located upstream of the turbine (and mounted under the wings of the aircraft or at the rear of the fuselage).

[0008] From an aerodynamic perspective, the primary function of the stator guide vanes is to straighten all or part of the airflow from the propeller to generate thrust that helps propel the aircraft forward. This thrust is generated by the deflection of air on the stator blades, characterized by an angle of attack different from the launch angle. A secondary function of the stator guide vanes is to minimize the total losses of the propeller / stator guide vane duo by allowing the swirling flow downstream of the propeller to recover (return). In fact, without stator guide vanes, this swirling flow can cause significant aerodynamic losses, leading to a sharp decrease in thrust and consequently a sharp reduction in the efficiency of the propulsion system. In the case of a ductless propeller mounted on an aircraft, the stator guide vanes also serve to adapt the airflow to various components integrated into or near the engine, such as pylons, to ensure the operability and performance of the mounted propulsion system. Therefore, several families (or configurations) of stator guide vanes can exist, including those located on either side of the pylon. From an aero-acoustic perspective, the shape of the stator blades must minimize noise sources generated by the interaction of the propeller wake and propeller blade tip vortices with the surface of the stator guide vanes. It is well known that a flat plate or profile that is too thin at the leading edge can lead to separation and / or increased interaction noise. Finally, from a mechanical point of view, the shape of the blade and the materials used to manufacture it must enable the blade to withstand static loads, dynamic loads, and intake loads as specified by regulations.

[0009] The purpose of this invention is to provide a stator blade for a ductless exhaust guide vane that meets a range of technical standards (aerodynamic, mechanical, and acoustic standards) enabling the ductless engine to operate efficiently across its entire flight envelope. Furthermore, prior art includes the following documents: FR 3 125 797 A1, US 2023 / 249810 A1, US 2010 / 260609 A1, and US 9 340 277 B2. Summary of the Invention

[0010] Therefore, a ductless aero propulsion unit for aircraft is proposed, which includes: - Outer shell; - The rotor hub is mounted to pivot relative to the outer shell about a main axis extending in the upstream-downstream direction of the aircraft; - A propeller, mounted on a hub, pivots relative to the outer casing; and - An exit guide vane, mounted on the housing and located downstream of the propeller along the main axis, extending about the main axis. The exit guide vane includes stator blades, each of which has: • The pressure side and suction side extending between the leading and trailing edges of the stator blades. • For each cross-section of the stator blade perpendicular to the stacking axis, at the blade ring height along the stacking axis, the cross-section has: ○ a curve located at the midpoint between the pressure side and the suction side, called the skeleton; ○ a distance between the leading edge and the trailing edge, called the chord; and ○ a thickness perpendicular to the skeleton located between the pressure side and the suction side, normalized relative to the chord, and having a maximum value normalized relative to the chord at a position on the chord. The feature is that, for at least one stator blade of the outlet guide vane, the maximum thickness position is located between 0.1 and 0.5 for all blade ring heights, and / or the maximum thickness position is strictly reduced at at least 60% of the blade ring height H', preferably strictly reduced at at least 75% of the blade ring height.

[0011] From an aerodynamic perspective (meeting efficiency and operability standards), an acoustic perspective (meeting maximum noise emission limits), and a mechanical perspective (meeting service life and intake limits), the proposed thickness rule ensures the efficient operation of stator blades under various flight conditions.

[0012] The present invention may also include one or more of the following optional features in any technically feasible combination.

[0013] Optionally, the maximum thickness is located between 0% and 20% of the blade ring height at its maximum value.

[0014] Alternatively, the maximum thickness position may be greater than or equal to 0.15, preferably greater than or equal to 0.25, or more preferably greater than or equal to 0.3.

[0015] Alternatively, the maximum thickness position is reduced from the maximum value at a blade ring height of less than 0.1 to the maximum thickness position at a blade ring height R1 between 20% and 40%, with the maximum thickness position at the blade ring height R1 located between 0.2 and 0.4, preferably between 0.25 and 0.35.

[0016] Alternatively, at the upper portion of the blade ring, i.e., above 50% of the height, the maximum thickness position decreases strictly downward from 50% of the blade ring height with respect to a value between 0.15 and 0.35, preferably between 0.2 and 0.3, reaching this value at a blade ring height R2 between 55% and 100%.

[0017] Alternatively, the reduction in the maximum thickness at the location between the blade ring height R1 and the blade ring height R2 is more pronounced than the reduction at the location of maximum thickness between 0% and the blade ring height R1.

[0018] Alternatively, the maximum thickness position is reduced up to the blade ring height R3 between 65% and 100%, and from that point up to 100% of the blade ring height, the maximum thickness position is increased.

[0019] Alternatively, for all blade ring heights, the maximum thickness may be between 0.01 and 0.3, preferably between 0.02 and 0.3.

[0020] Alternatively, the maximum thickness is at its maximum value when the blade ring height is zero, and the maximum thickness is between 0.05 and 0.25, preferably between 0.08 and 0.18, or more preferably between 0.1 and 0.14.

[0021] Alternatively, the maximum thickness is also strictly reduced from a blade ring height of zero to a blade ring height R4 located between 35% and 85%.

[0022] Alternatively, the maximum thickness is at its minimum at the blade ring height R4, and at the blade ring height R4, the maximum thickness is between 35% and 85% of the maximum thickness at the blade ring height of zero, preferably between 50% and 80% of the maximum thickness at the blade ring height of zero.

[0023] Alternatively, the maximum thickness can be strictly increased between the blade ring height R4 and 100% of the blade ring height.

[0024] Alternatively, the maximum thickness at 100% of the blade ring height is between 125% and 400% of the maximum thickness at blade ring height R4.

[0025] An aircraft is also proposed, which includes an aviation propulsion unit (100) according to the invention. Attached Figure Description

[0026] The invention will be better understood through the following description, given only by way of example and with reference to the accompanying drawings, in which: - Figure 1 This is a side view of an aircraft propulsion unit according to the present invention. - Figure 2 It comes from Figure 1 The diagram shows a cross-sectional view of the outlet guide vanes of the propulsion unit. - Figure 3 The cross-section of the stator blade, taken perpendicular to the stacking axis, is shown. - Figure 4 It is similar to Figure 3 The view shows the skeleton of the stator blades. - Figure 5 It shows something similar to Figure 4 The view shows that the thickness curve is a function of the position along the chord of the stator blade. - Figure 6 Three curves are shown, illustrating the variation in the location of maximum thickness as a function of the blade ring height, and... - Figure 7 Three curves are shown, illustrating the variation in maximum thickness as a function of the blade ring height. Detailed Implementation

[0027] In the following description, if a characteristic is applied to at least one element, it can also be applied to all such elements. Similarly, if a characteristic is applicable to at least one value within a range, it can also be applied to all values ​​within that range.

[0028] refer to Figure 1 Now, an aviation propulsion unit 100 will be described, in which the present invention is implemented.

[0029] The aircraft propulsion unit 100 is a ductless single fan (USF), which is designed to help propel the aircraft forward. For example, the USF is designed to be supported by a pylon attached to one of the aircraft's wings.

[0030] The aircraft propulsion unit 100 first includes an outer shell 102 and a rotor hub 104, the rotor hub 104 being mounted to pivot relative to the outer shell 102 about a main axis X.

[0031] In the following text, the terms "upstream" and "downstream" will be used to specify the relative positions of the elements of the aero-propulsion unit 100 along the main axis X in the direction of airflow PHI when the aircraft is propelled by the aero-propulsion unit 100. For example, the aircraft can be propelled by the aero-propulsion unit 100 in cruise mode at a flight Mach number greater than 0.7.

[0032] For example, the hub 104 is located upstream of the housing 102.

[0033] The aircraft propulsion unit 100 also includes an engine 105 for driving the rotor hub 104. For example, the engine 105 extends into the housing 102.

[0034] For example, engine 105 includes at least one internal combustion engine, particularly a turbine, turbofan engine, turbojet engine or turbo blower, and / or at least one electric motor, and / or at least one hydrogen engine.

[0035] For example, the aircraft propulsion unit 100 also includes an air inlet 107 to supply the mains to the engine 105.

[0036] propeller The aircraft propulsion unit 100 also includes a ductless propulsion propeller 106, which is mounted on a hub 104 and pivots relative to the outer casing 102 about the main axis X. Therefore, the propeller 106 is driven by an engine 105 via the hub 104 to rotate about the main axis X.

[0037] For example, propeller 106 is located upstream of engine 105. This arrangement is called a "puller". Alternatively, engine 105 may be configured as a "thruster".

[0038] When propeller 106 rotates, propeller 106 is designed to drive a downstream airflow PHI to propel the aircraft. For this purpose, propeller 106 includes rotor blades 108 (e.g., between 3 and 25, preferably between 8 and 16), for example, the rotor blades 108 arranged in a single annular row around the main axis X. For example, all rotor blades 108 may be identical and spaced apart around the main axis X at a regular angular spacing.

[0039] For example, at least one rotor blade 108 has a variable pitch about a corresponding pitch axis Y. The pitch of each variable pitch rotor blade 108 is defined by the pitch angle C about the rotor pitch axis Y.

[0040] The pitch axis Y can pass through the main axis X or be slightly offset from the main axis X, for example, offset by 10cm, 5cm, 2cm, or 1cm.

[0041] Furthermore, the pitch axis Y can be perpendicular to the main axis X, as shown in the figure. Alternatively, due to manufacturing tolerances or by design, the pitch axis can form an angle slightly different from 90° relative to the main axis X. Therefore, the pitch axis Y can be perpendicular to the main axis X, for example, within 5°, or within 2°, or within 1°, or within 0.1°.

[0042] In a preferred embodiment, all rotor blades 108 have variable pitch.

[0043] Furthermore, each variable-pitch rotor blade 108 of the propeller 106 has an outer radius Re, which, by definition, is equal to the distance between the main axis X and the point on the rotor blade 108 furthest from the main axis X among all possible pitch angles C. Each variable-pitch rotor blade 108 also has an inner radius Ri, which, by definition, is equal to the distance between the main axis X and the point on the rotor blade 108, 114 closest to the main axis X among all possible pitch angles C.

[0044] Therefore, the diameter D of the propeller 106 is equal to twice the outer radius Re. This diameter D is also called the "drive diameter".

[0045] Furthermore, each rotor blade 108 is formed by a cross-section (also referred to as a "profile") stacked along an axis called the stacking axis. In other words, the cross-section in question is perpendicular to the stacking axis. When the rotor pitch axis Y is perpendicular to and intersects the main axis X, the stacking axis corresponds to the rotor pitch axis Y.

[0046] When the rotor pitch axis Y is offset from the main axis X and / or forms an angle other than 90° with the main axis X, the rotor pitch axis Y has a radial component as a stacking axis. This radial component corresponds to the projection of the rotor pitch axis Y onto a line perpendicular to the main axis X.

[0047] Export guide vanes The aircraft propulsion unit 100 also includes a fixed ductless outlet guide vane (OGV) 112, which is mounted on the housing 102 and located downstream of the propeller 106, for example, downstream of the air inlet 107, such that the air inlet is located between the propeller 106 and the outlet guide vane 112.

[0048] The exit guide vane 112 forms a stator, which is fixed to the housing 102 and extends about the main axis X, but cannot rotate about the main axis X. The exit guide vane 112 includes stator blades 114 (e.g., between 3 and 25, preferably between 8 and 16), for example, the stator blades 114 are arranged in a single annular row around the main axis X. Preferably, the number of stator blades 114 is different from the number of rotor blades 108 to reduce the noise of the aerospace propulsion unit 100. In particular, the number of rotor blades 108 is greater than the number of stator blades 114. In fact, if the number of rotor blades 108 and the number of stator blades 114 are equal, a wake may exist after the rotor blades 108, which can also interact with the stator blades 114, which can increase the noise level. For example, all stator blades 114 can be identical, or the stator blades 114 can be different and spaced apart around the main axis X with regular or irregular angular spacing, such that at least two stator blades 114 have different angular spacings around the main axis X.

[0049] The exit guide vane 112 is designed to straighten at least a portion of the airflow PHI passing through the propeller 106 to improve the performance of the aero propulsion unit 100. More specifically, the exit guide vane 112 aims to absorb the swirling of the flow caused by the propeller 106 to improve the performance of the ductless configuration. However, the presence of the exit guide vane 112 is a major source of noise due to its interaction with the wake of the propeller 106 (and blade tip vortices generated when the stator blades 114 are not sufficiently trunculated). Therefore, it is important to reduce the noise generated by the exit guide vane 112 and its interaction with the wake of the propeller 106 while maintaining good aerodynamic performance, as reducing noise emissions and fuel consumption are major challenges for ductless engine architectures.

[0050] For example, at least one stator blade 114 has a variable pitch around the corresponding stator pitch axis Y'.

[0051] The stator pitch axis Y' may pass through the main axis X or be slightly offset from the main axis X, for example, offset from the main axis X by no more than 10 cm, or offset from the main axis X by no more than 5 cm, or offset from the main axis X by no more than 2 cm, or offset from the main axis X by no more than 1 cm.

[0052] Furthermore, the stator pitch axis Y' may be perpendicular to the main axis X, as shown in the figure. Alternatively, due to manufacturing tolerances or by design, the stator pitch axis Y' may form an angle slightly different from 90° relative to the main axis X. Therefore, the stator pitch axis Y' may be perpendicular to the main axis X, for example, within 5°, or within 2°, or within 1°, or within 0.1°.

[0053] In a preferred embodiment, all stator blades 114 have variable pitch.

[0054] Furthermore, each variable-pitch stator blade 114 of the exit guide vane 112 has an outer radius Re', which, by definition, is equal to the distance between the main axis X and the point on the stator blade 114 furthest from the main axis X among all possible pitch angle settings C'. Each variable-pitch stator blade 114 also has an inner radius Ri', which, by definition, is equal to the distance between the main axis X and the point on the stator blade 114 closest to the main axis X among all possible pitch angles C'.

[0055] Furthermore, at a given radius between Ri' and Re', two consecutive stator blades 114 are spaced apart by a distance E, which corresponds to the radius multiplied by the radian of the circle connecting the two blades.

[0056] Furthermore, the pitch axes Y and Y' are separated by a distance S on the main axis X. This distance S is the distance between the point on the main axis X closest to the pitch axis Y and the point on the main axis X closest to the pitch axis Y'. When both the pitch axis Y and line Y' intersect the main axis X, this distance S is the distance between the two intersection points.

[0057] stator blades Now, one of the stator blades 114 will be described in more detail; all other stator blades are similar.

[0058] refer to Figure 2 First, the stator blade 114 includes a leading edge BA' and a trailing edge BF'. The airflow PHI from the propeller 106 reaches the leading edge BA' and moves away from the trailing edge BF'.

[0059] The leading edge BA' extends from the root BA'_P near the outer shell 102 to the tip BA'_T away from the outer shell 102. Similarly, the trailing edge BF' extends from the root BF'_P near the outer shell 102 to the tip BF'_T away from the outer shell 102.

[0060] As shown in the example, the stator blade 114 can be truncated, i.e., there is a truncated cross section 602 connecting the tips BA'_T and BF'_T, for example, the truncated cross section 602 is a straight cross section. Alternatively, the stator blade may not be truncated, in which case the tips BA'_T and BF'_T coincide.

[0061] Furthermore, the stator blades 114 are formed by stacking cross sections (also referred to as "cross sections" or "profiles") arranged along an axis called the stacking axis. In other words, the cross section in question is perpendicular to the stacking axis. When the stator pitch axis Y' is perpendicular to and intersects the main axis X, the stacking axis corresponds to the stator pitch axis Y'. This is the case shown in the figure.

[0062] When the stator pitch axis Y' is offset from the main axis X and / or forms an angle other than 90° with the main axis X, the stator pitch axis Y' has a radial component as a stacking axis. This radial component corresponds to the projection of the stator pitch axis Y' onto a line perpendicular to the main axis X.

[0063] In the following text, when used in the context of stator blade 114, the term “height” will refer to the distance between two points along the stack axis, that is, the distance between the orthogonal projections of these points onto the stack axis.

[0064] Therefore, for a blade ring positioned on the leading edge BA', the upstream blade ring height H' can be limited. 上游 Therefore, the upstream blade ring height H' 上游 It is the height h' from the root BA'_P BA The total height H' of the leading edge BA' between the root BA'_P and the tip BA'_T BA Ratio: H' 上游 =h' BA / H' BA Therefore, the upstream blade ring height H' 上游 This can be expressed as a percentage, varying between 0% (at the root BA'_P) and 100% (at the tip BA'_T). Similarly, for the one positioned on the trailing edge BF', the downstream blade ring height H' can be defined. 下游 Therefore, the downstream blade ring height H' 下游 It is the height h' from the root BF'_P BF The total height H' of the trailing edge BF' between the root BF'_P and the tip BF'_T BF Ratio: H' 下游 =h' BF / H' BF Therefore, the downstream blade ring height H' 下游It can be expressed as a percentage, and varies between 0% (located at the root BF'_P) and 100% (located at the tip BF'_T).

[0065] In the following text, when referring to the blade ring height represented by H' without specifying whether it is the downstream or upstream blade ring height, the blade ring height can refer to the upstream blade ring height H'. 上游 or downstream blade ring height H' 下游 .

[0066] Figure 3 The cross-section of the stator blade 114 taken at a certain height perpendicular to the stacking axis is shown.

[0067] As can be seen, the stator blade 114 has a pressure side 702 and a suction side 704, which are concave and convex, respectively, and are connected to each other by a leading edge BA' and a trailing edge BF'. Therefore, the leading edge BA' separates the pressure side 702 from the suction side 704 in the upstream portion of the stator blade 114, while the trailing edge BF' separates the pressure side 702 from the suction side 704 in the rear portion of the stator blade 114.

[0068] When a leading edge BA' and a trailing edge BF' exist in the considered cross-section (that is, for example, below the truncated cross-section 602 in the example shown), the leading edge BA' and the trailing edge BF' can be connected by a chord 706, the direction of which depends on the blade ring height H' (i.e., the considered upstream blade ring height H'). 上游 or downstream blade ring height H' 下游 The leading edge BA' and trailing edge BF' are separated along the chord line 706 by a distance called chord L', which varies depending on the blade ring height H'.

[0069] refer to Figure 4 For each cross-section of the stator blade 114, a frame 902 may be defined, for example, a curve located at the midpoint between the pressure side 702 and the suction side 704. For example, as... Figure 4 As shown, the skeleton 902 can be defined as the center of a set of circles inscribed in the cross-section (that is, placed flush with the pressure side 702 and the suction side 704).

[0070] refer to Figure 5For each cross-section of the stator blade 114, the thickness Ep of the stator blade 114 can then be defined along the chord 706 between the leading edge BA' and the trailing edge BF'. The thickness Ep is defined as a distance perpendicular to the frame 902 between the pressure side 702 and the suction side 704. More specifically, for each point identified by the x-coordinate along the chord L' between the leading edge BA' and the trailing edge BF', this distance is measured at a point on the frame 902 located on the perpendicular line to the chord 702 at the x-coordinate. The thickness Ep and the x-coordinate are normalized relative to the chord L': Ep = absolute thickness / L', x = absolute position / L'. Thus, in particular, the x-coordinate is 0 when the point is located on the leading edge BA', and 1 when the point is located on the trailing edge BF'. Therefore, the thickness Ep is a function of the blade ring height H' and the x-coordinate.

[0071] For a given blade ring height H', the thickness Ep along the chord 706 reaches its maximum value at a location denoted as Xepmax (i.e., the value of the x-coordinate), which is denoted as Epmax. Therefore, the location of the maximum thickness, Xepmax, is a function of the blade ring height H'.

[0072] Thickness regularity The present invention aims to optimize the radial variation of the maximum thickness Epmax; this variation is referred to as the "thickness law". To achieve this optimization, the maximum thickness Epmax of each stator blade in at least a portion of the stator blades 114, preferably the maximum thickness Epmax of all stator blades 114, has one or two of the following characteristics.

[0073] according to First characteristic To optimize the aerodynamic, acoustic, and mechanical behavior of the stator blades 114, the maximum thickness position Xepmax is located between 0.1 and 0.5 for all blade ring heights H'. This position provides the maximum thickness, ensuring optimal operation of the stator blades 114 from a cross-functional perspective.

[0074] according to Second characteristic The maximum thickness position Xepmax decreases monotonically at at least 60% of the blade ring height H', that is, within any length range falling within 60% of the 0%-100% range of the blade ring height H', for example, within the range of 15%-75%, preferably monotonically decreasing at at least 75% of the blade ring height H'. This criterion is associated with the reduction of tip mass (which, from a mechanical point of view, is required) and the need for de-icing at the leading edge of the stator, which imposes the maximum thickness near the leading edge to incorporate a de-icing solution (e.g., a heating pad). Regarding this second characteristic, the maximum thickness Epmax may have one or more of the following sub-characteristics.

[0075] according to First sub-characteristicThe maximum thickness location, Xepmax, is between 0% and 20% of the blade ring height H'. This prevents the formation of a sonic transition cross section near the hub or casing.

[0076] according to Second sub-characteristic The maximum thickness position Xepmax is located between 0% and 20% of the blade ring height H'. That is, for at least one value in the 0%-20% range, the maximum thickness position Xepmax is greater than or equal to 0.15, preferably greater than or equal to 0.25, or even more preferably greater than or equal to 0.3. This is particularly important for the outlet guide vanes on both sides of the hanger.

[0077] according to Third Sub-characteristics The maximum thickness position Xepmax is reduced from a maximum value at an annular height H' less than 0.1 to a maximum thickness position Xepmax at an annular height R1 between 20% and 40%, with the maximum thickness position Xepmax at an annular height R1 between 0.2 and 0.4, preferably between 0.25 and 0.35. This optimizes the aerodynamic performance of the profile within the height range between the blade root of the stator blade 114 and R1, while adhering to mechanical constraints primarily related to the thickness of the leading edge BA': for example, minimum manufacturable thickness and drag on intake.

[0078] according to Fourth sub-characteristic In the upper portion of the blade ring, i.e., above 50% of the height, the maximum thickness position Xepmax decreases strictly downwards from 50% of the blade ring height with respect to the blade ring height H' to a value between 0.15 and 0.35, preferably between 0.2 and 0.3, reaching this value at a blade ring height R2 between 55% and 100%. In practice, as the blade ring height H' increases, the profile thickness decreases towards the tip. A profile with decreasing thickness is more prone to separation, leading to aerodynamic losses and increased noise. An alternative is to position the maximum thickness position Xepmax closer to the leading edge BA', which makes the profile more robust under excessive angles of attack (such as those that may occur during landing and / or takeoff maneuvers).

[0079] according to Fifth Sub-characteristic The decrease in the maximum thickness position Xepmax between the blade ring heights R1 and R2 (i.e., in the upper part of the stator blade 114) is more significant than the decrease in the maximum thickness position Xepmax between 0% and the blade ring height R1 (i.e., in the lower part of the stator blade 114). In other words: Xepmax(R1) - Xepmax(R2) > Xepmax(0) - Xepmax(R1). This ensures that the location of the maximum thickness at the root cross section does not change too much, which is beneficial to the mechanical strength of the stator blade 114.

[0080] according to Sixth Sub-characteristic The maximum thickness position Xepmax decreases until it reaches 65% to 100% of the blade ring height R3. Beyond R3, Xepmax increases again until it reaches 100% of the blade ring height H'. This is because the reduction in chord length L' at the tip region of the stator blade 114 is greater than the reduction at the maximum thickness position Xepmax. In fact, the load on the outlet guide vane 112 is smaller in the upper portion of the stator blade 114 compared to the lower portion, and the vortices at the tip of the stator blade 114 are not a significant noise source. Therefore, it is worthwhile to reduce the chord length L' to reduce the mass of the stator blade 114 while maintaining good aerodynamics.

[0081] according to Seventh Sub-characteristic For all blade ring heights H', the maximum thickness Epmax is between 0.01 and 0.3, preferably between 0.02 and 0.3. This ensures good aerodynamic, acoustic, and mechanical performance.

[0082] according to Eighth sub-characteristic The maximum thickness Epmax is greatest at the root (H'=0) and lies between 0.05 and 0.25, preferably between 0.08 and 0.18, or more preferably between 0.1 and 0.14. This ensures that the thickness of the profile at the root (or root attachment) is sufficient to withstand and transmit the aerodynamic and mechanical forces acting on the blade.

[0083] according to Ninth Sub-characteristics The maximum thickness Epmax decreases strictly from the root (H'=0) to the blade ring height R4 between 35% and 85%.

[0084] according to Tenth Sub-characteristic The maximum thickness Epmax is at its minimum at the blade ring height R4. At the blade ring height R4, the maximum thickness Epmax is between 35% and 85% of the maximum thickness Epmax at the root (H'=0), preferably between 50% and 80% of the maximum thickness Epmax at the root (H'=0).

[0085] according to Eleventh Sub-characteristic Between the blade ring height R4 and the blade ring height H' at the tip (H'=100%), the maximum thickness Epmax increases strictly. This behavior is mainly caused by the reduction of the chord length L' in the upper portion of the stator blade 114. In fact, since the stator blade 114 experiences a smaller load in its upper portion, the chord length L' can be reduced (resulting in a reduction in mass) while maintaining good aerodynamic performance (without significant loss or separation).

[0086] according to Twelfth Sub-Characteristic The maximum thickness Epmax at the tip (H'=100%) is between 125% and 400% of the maximum thickness Epmax at the blade ring height R4.

[0087] Figure 6 Three curves, 602, 604, and 606, are shown. Curves 602, 604, and 606 show the maximum thickness position Xepmax as a function of the blade ring height H', which obeys all the properties and sub-properties related to the maximum thickness position Xepmax described above.

[0088] Figure 7 Three curves, 802, 804, and 806, are shown. Curves 802, 804, and 806 show the maximum thickness Epmax as a function of the blade ring height H', which obeys all the properties and sub-properties related to the maximum thickness Epmax described above.

[0089] Activity factors One parameter that provides an initial estimate of the chord distribution along the blade span of stator blade 114 is its activity factor (AF), which is constrained as follows: Where Ri' corresponds to the inner radius Ri' of the stator blade 114 at the leading edge BA'. BA Or, corresponding to the inner radius Ri' of the stator blade 114 at the trailing edge BF'. BF Re' corresponds to the outer radius Re' of stator blade 114 at its leading edge BA'. BA Or, corresponding to the outer radius Re' of stator blade 114 at trailing edge BF'. BF ; L' represents the radial distance relative to the principal axis X, excluding the radius Re'; ) represents the radial distance At point L', in a plane perpendicular to the radial component of the pitch axis Y', there is a chord L' between the leading edge BA' and the trailing edge BF' of the cross section (or aerodynamic profile) of the stator blade 114.

[0090] Preferably, the activity factor of the stator blade 114 is between 40 and 225, more preferably between 90 and 160.

[0091] This means that the L' chord of the stator blade 114 is relatively large in the lower portion of the stator blade 114, which, combined with the considerable thickness at the root of the stator blade, ensures the mechanical strength of the stator blade 114. Furthermore, the larger the chord length L' in the lower portion, the larger the vortex in the flow downstream of the propeller 106 (an aerodynamic advantage), which helps optimize noise reduction by removing the tips of the rotor blades 108 of the propeller 106 and the tips of the stator blades 114 of the outlet guide vanes 112 (where the flow velocity is higher).

[0092] Other possible characteristics of aircraft propulsion units Preferably, the number of rotor blades on propeller 106 differs from the number of stator blades 114 on exit guide vanes 112. This helps minimize noise from the aero-propulsion unit 100. In practice, with an equal number of rotor blades 108 and stator blades 114, the wake field of propeller 106 interacts with the stator blades 114 simultaneously, which increases the noise level. Preferably, there are more rotor blades 108 than stator blades 114; for example, two more rotor blades than stator blades 114.

[0093] Preferably, the strength is represented by П and is defined as the ratio of the chord L' to the distance E between two consecutive stator blades 114 at a given radius (the distance E corresponds to the radius value multiplied by the radian of the circle connecting the two blades with that radius), and the strength is less than 3 over the entire blade ring height H'. Furthermore, preferably, the strength П is less than 1 at the tip (blade ring height H' equals 100%).

[0094] Preferably, the ratio of the distance between the stacked axis of the propeller 106 and the stacked axis of the outlet guide vane 112 (in the example shown, this is distance S) to the engine diameter D is between 0.01 and 0.5, and preferably between 0.15 and 0.35, i.e., 0.01.

[0095] Preferably, the trailing edge of the rotor blade 108 of the propeller 106 is located at a more upstream axial position than the leading edge BA' of the stator blade 114 of the outlet guide vane 112, in order to prevent interference between the rotating propeller 106 and the outlet guide vane 112.

[0096] Fixed stator blades When one of the stator blades 114 is fixed (e.g., due to design constraints, such as lack of space under the hub to include the pitch control system, or to reduce weight), the stack axis is defined as an axis perpendicular to the main axis X and passing through the main axis X and also through the leading edge BA' at the blade root BA'_P. ​

[0097] Specific embodiments In a particular embodiment, at least two stator blades 114 have the same thickness pattern below the radial position R4. This allows a single geometric definition of the blade to be applied to two different maximum radius heights Re2 (or trimmed values) while satisfying conditions related to the thickness at the blade tip having one or more of the aforementioned characteristics.

[0098] In a particular embodiment, at least two stator blades have different thickness profiles, but each of these profiles satisfies at least the first and / or second characteristics described above. In the case of a ductless single-fan (USF) engine mounted on an aircraft, each stator blade 114 can effectively experience different aerodynamic loads (or forces exerted on the blade by the airflow). This behavior is caused by the mounting effect, resulting in a non-axisymmetric aerodynamic environment about the main axis X. Therefore, the thickness profile of each stator blade 114 can be optimized relative to the maximum aerodynamic load on each stator blade 114, where the maximum aerodynamic load can vary from one stator blade 114 to another. For example, this allows for a reduction in the thickness of some stator blades 114 relative to others, resulting in a reduction in the overall propulsion system mass and thus a reduction in specific fuel consumption (SFC).

[0099] Each stator blade has a leading edge thickness EpBA, defined as a (normalized) thickness Ep at a position x between 0 and 0.1. In a particular embodiment, the leading edge thickness EpBA is located between 0.005 and 0.12 for all blade ring heights H' to enable the blade to robustly resist changes in angle of attack. Therefore, even under operating conditions that result in a wide range of angles of attack experienced by the stator blades, the minimum value of the leading edge thickness EpBA still ensures proper aerodynamic operation of the stator blades 114 across the entire blade ring height H': there is no flow separation, which can lead to additional losses and thus reduce the overall efficiency of the engine.

[0100] In one particular embodiment, the leading edge thickness EpBA allows for the integration of a thermal system designed to de-ice the leading edge region during all or part of the aircraft's missions.

[0101] In a particular embodiment, at least two stator blades have different leading-edge thicknesses EpBA; however, each leading-edge thickness EpBA lies between 0.005 and 0.12. This design approach involves optimizing the leading edge of each stator blade mounted to the aircraft. In practice, the range of angles of attack experienced by each stator blade can vary due to mounting effects. Determining the leading-edge thickness (EpBA) of each stator blade ensures the robustness of each stator blade to variations in angle of attack, while avoiding excessively thick stator blades when subjected to smaller angles of attack. This optimization maximizes the aerodynamic performance of the exit guide vanes while minimizing their mass, resulting in reduced power consumption.

[0102] in conclusion In summary, it should be noted that the present invention is not limited to the embodiments described above. In fact, it will be apparent to those skilled in the art that various modifications can be made to the embodiments described above based on the teachings just disclosed.

[0103] In the foregoing detailed description of the invention, the terminology used should not be construed as limiting the invention to the embodiments disclosed herein, but rather as including all equivalents contemplated by those skilled in the art by applying their general knowledge to the implementation of the teachings just disclosed.

Claims

1. A ductless aero-propulsion unit (100) for an aircraft, comprising: - Outer shell (102); - A rotor hub (104) mounted to pivot relative to the outer shell (102) about a main axis (X) extending in the upstream-downstream direction of the aircraft; - A propulsion propeller (106) mounted on the hub (104) for pivoting relative to the outer casing (102); as well as - An outlet guide vane (112) is mounted on the housing (102) downstream of the propeller (106) along the main axis (X), the outlet guide vane (112) extending about the main axis (X), the outlet guide vane (112) including stator blades (114), each of the stator blades having: • The pressure side (702) and suction side (704) extending between the leading edge (BA') and trailing edge (BF') of the stator blade (114). • For each cross section of the stator blade (114) perpendicular to the stacking axis (Y'), at the blade ring height (H') along the stacking axis (Y'), the cross section has: ○ a curve located at the midpoint between the pressure side (702) and the suction side (704), the curve being called the skeleton (902); ○ a distance located between the leading edge (BA') and the trailing edge (BF'), the distance being called the chord (L'); and ○ a thickness (Ep) perpendicular to the skeleton (902) located between the pressure side (702) and the suction side, the thickness (Ep) being normalized relative to the chord (L'), and having a maximum value (Epmax) normalized relative to the chord (L') at a position (Xepmax) on the chord (L'). The feature is that, for at least one stator blade (114) of the outlet guide blade (112), for all the blade ring heights (H'), the maximum thickness position (Xepmax) is between 0.1 and 0.5, and / or the maximum thickness position (Xepmax) is strictly reduced at at least 60% of the blade ring height H', preferably strictly reduced at at least 75% of the blade ring height (H').

2. The aviation propulsion unit (100) according to claim 1, wherein, The maximum thickness location (Xepmax) is between 0% and 20% of the blade ring height (H') at its maximum value.

3. The aviation propulsion unit (100) according to claim 2, wherein, The maximum thickness position (Xepmax) is greater than or equal to 0.15, preferably greater than or equal to 0.25, or more preferably greater than or equal to 0.

3.

4. The aviation propulsion unit (100) according to any one of claims 1 to 3, wherein, The maximum thickness position (Xepmax) decreases from a maximum value at a blade ring height (H') less than 0.1 to a maximum thickness position (Xepmax) at a blade ring height R1 between 20% and 40%, and the maximum thickness position (Xepmax) at the blade ring height R1 is between 0.2 and 0.4, preferably between 0.25 and 0.

35.

5. The aviation propulsion unit (100) according to any one of claims 1 to 4, wherein, At the upper portion of the blade ring, that is, above 50% of the height, the maximum thickness position (Xepmax) decreases strictly downward from 50% of the blade ring height with respect to the blade ring height (H') to a value between 0.15 and 0.35, preferably between 0.2 and 0.3, reaching this value at a blade ring height R2 between 55% and 100%.

6. The aviation propulsion unit (100) according to claims 4 and 5, wherein, The reduction in the maximum thickness position (Xpemax) between the blade ring height R1 and the blade ring height R2 is more significant than the reduction in the maximum thickness position (Xpemax) between 0% and the blade ring height R1.

7. The aviation propulsion unit (100) according to any one of claims 1 to 6, wherein, The maximum thickness position (Xepmax) decreases until the blade ring height R3 is between 65% and 100%, and from that point up to 100% of the blade ring height (H'), the maximum thickness position (Xepmax) increases.

8. The aviation propulsion unit (100) according to any one of claims 1 to 7, wherein, For all the aforementioned blade ring heights (H'), the maximum thickness (Epmax) is between 0.01 and 0.3, preferably between 0.02 and 0.

3.

9. The aviation propulsion unit (100) according to any one of claims 1 to 8, wherein, The maximum thickness (Epmax) is at its maximum value when the blade ring height (H') is zero, and the maximum thickness (Epmax) is between 0.05 and 0.25, preferably between 0.08 and 0.18, or more preferably between 0.1 and 0.

14.

10. The aviation propulsion unit (100) according to any one of claims 1 to 9, wherein, The maximum thickness (Epmax) strictly decreases from the zero ring height (H') to the ring height R4 between 35% and 85%.

11. The aviation propulsion unit (100) according to claim 10, wherein, The maximum thickness (Epmax) is at its minimum at the blade ring height R4, and at the blade ring height R4, the maximum thickness (Epmax) is located between 35% and 85% of the maximum thickness (Epmax) at the blade ring height (H') being zero, preferably between 50% and 80% of the maximum thickness (Epmax) at the blade ring height (H') being zero.

12. The aircraft propulsion unit (100) according to claim 10 or 11, wherein, The maximum thickness (Epmax) increases strictly between the blade ring height R4 and 100% of the blade ring height (H').

13. The aircraft propulsion unit (100) according to any one of claims 10 to 12, wherein, The maximum thickness (Epmax) at the blade ring height of 100% is between 125% and 400% of the maximum thickness (Epmax) at the blade ring height R4.

14. An aircraft comprising an aviation propulsion unit (100) according to any one of claims 1 to 13.

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