TURBOMACHINE SUBASSEMBLY FEATURING AN IMPROVED SWAN NECK CONFIGURATION AND TURBOMACHINE FEATURING SUCH A SUBASSEMBLY

Abstract

The invention relates to an aircraft turbomachine subassembly comprising an annular duct (36) extending concentrically around a longitudinal axis and in which an airflow flows from upstream to downstream, the annular duct comprising a duct portion (28) defining a gooseneck having an upstream inlet section and a downstream outlet section. A plurality of radial arms (B1, B2) extend circumferentially in the duct portion (28). One or more intermediate walls (P1, P2) extend between two circumferentially adjacent arms (B1, B2) which form a pair of circumferentially adjacent arms among the plurality of circumferentially adjacent arms, so as to locally separate the annular duct portion (28) between the two arms of the pair of arms into a radially internal portion (Ci) and a radially external portion (Ce) further from the longitudinal axis than the radially internal portion. Figure for the abridged version: Fig.3B.

Description

Title of the invention: TURBOMACHINE SUBASSEMBLY COMPRISING AN IMPROVED SWAN NECK CONFIGURATION AND TURBOMACHINE COMPRISING SUCH A SUBASSEMBLY technical field

[0001] The present exposition relates to a gooseneck aircraft turbomachine subassembly and a turbomachine comprising such a subassembly. Previous technique

[0002] Aircraft turbomachinery, such as aircraft turbojets, is known to have an internal architecture comprising, successively, from upstream to downstream in the direction of flow of a primary air stream originating from the turbojet's air intake, a low-pressure casing containing, in particular, a compressor and a low-pressure turbine, an intermediate casing, and a high-pressure casing containing, in particular, a compressor and a high-pressure turbine. These three casings are generally aligned along a longitudinal direction defined by a longitudinal axis of the turbojet. These casings are configured to define an internal annular duct extending concentrically around the longitudinal axis of the turbojet, from the low-pressure casing to the high-pressure casing. The primary air stream flows inside the annular duct from upstream to downstream of this duct.The portion of the annular duct defined by the intermediate housing has a gooseneck or S-shape (viewed in longitudinal section) which defines, for the primary air stream flowing in the annular duct, a transition zone between the low-pressure and high-pressure housings.

[0003] Such a gooseneck is characterized by several geometric parameters, including the gooseneck slope (radial deviation between the upstream inlet section of the gooseneck and its downstream outlet section), the gooseneck height (capacity to ensure the aforementioned radial deviation), and the gooseneck diffusion, namely the intrinsic slowing down of the airflow due to the gooseneck geometry. The criticality of the gooseneck is greater the steeper its slope and the greater its height. In such a case, guiding the airflow within the gooseneck is difficult, and the risks of aerodynamic separation of the air stream are high. Conversely, the criticality of the gooseneck is low when the slope and height of the gooseneck are small, since the airflow is then guided quite naturally by this gooseneck geometry. When the slope of the gooseneck When the swan neck has a steep slope and a shallow height, the airflow needs to be strongly guided, but the proximity of the walls defining the swan neck facilitates this guidance. When the swan neck has a shallow slope and a shallow height, the airflow does not need to be strongly guided and is naturally guided by this swan neck geometry.

[0004] Certain internal architectures of aircraft turbojet engines impose a steep gooseneck geometry (highly constrained gooseneck), which increases the criticality of the gooseneck from an aerodynamic point of view.

[0005] In order to reduce the criticality of the gooseneck, it is generally considered to modify one of its geometric characteristics. The simplest geometric modification consists of increasing the length (longitudinal dimension) of the gooseneck to decrease its slope in order to reduce its criticality.

[0006] However, such a geometric modification has a significant impact on the mass carried.

[0007] It would therefore be advantageous to design a new gooseneck configuration for aircraft turbomachine subassemblies that would limit the criticality of the gooseneck, particularly without significantly impacting the onboard mass. Description of the invention

[0008] The present invention thus relates to a turbomachine subassembly comprising an annular duct extending concentrically around a longitudinal axis XX' and in which an airflow flows from upstream to downstream, the annular duct comprising at least one portion of duct defining a gooseneck having an upstream inlet section and a downstream outlet section, the aircraft turbomachine subassembly comprising, in said at least one portion of duct, a plurality of arms extending circumferentially relative to the longitudinal axis XX' at intervals spaced apart from each other, each following a radial and axial extension,characterized in that the aircraft turbomachine subassembly comprises one or more intermediate walls extending between two circumferentially adjacent arms which form a pair of circumferentially adjacent arms among the plurality of circumferentially adjacent arms so as to locally separate between the two arms of said pair of arms at least a portion of annular duct into a radially internal portion and a radially external portion further from the longitudinal axis XX' than the radially internal portion.

[0009] The addition of one or more intermediate walls between at least two circumferentially adjacent arms of the gooseneck makes it possible to locally transform the airflow or air stream into two airflows or air streams, each of reduced height compared to the height of the airflow / stream. The airflow is therefore better guided in the gooseneck thus configured, and the criticality of the gooseneck is thereby limited. Such an arrangement makes it possible to significantly reduce the height of the airflow / stream and thus avoid the risk of flow separation. Furthermore, such an arrangement also allows for a gooseneck with a steep slope and a relatively short longitudinal or axial extent, i.e., relatively short axially (to limit the mass carried in the turbomachine subassembly), while considerably limiting the criticality of the gooseneck.

[0010] According to other possible characteristics: - the downstream outlet section of the gooseneck is closer to the longitudinal axis than the upstream inlet section; -the intermediate wall which extends between the two circumferentially adjacent arms is axially centered relative to the axial extension of the two arms; -the intermediate wall which extends between the two circumferentially adjacent arms is axially offset upstream from a position centered relative to the axial extension of the two arms; -The aircraft turbomachine subassembly includes means for modifying the inclination of the intermediate wall(s) relative to said at least a portion of annular duct -the intermediate wall which extends between the two circumferentially adjacent arms presents, according to a radial view, a general Y shape in which the common branch of the Y is arranged against one of the first of the two circumferentially adjacent arms and the two branches of the Y are arranged against the second of the two arms; -the intermediate wall which extends between the two circumferentially adjacent arms presents, according to a cross-sectional view relative to the longitudinal axis XX', a general Y shape in which the common branch of the Y is arranged against the first arm and the two end branches of the Y arranged against the second arm are each doubled so as to each form two end sub-branches conferring the Y shape according to this cross-section; -when the aircraft turbomachine subassembly includes several intermediate walls between the two circumferentially adjacent arms, the intermediate walls are distributed axially along the axial extension of the two arms; -the aircraft turbomachine subassembly comprises, on the one hand, an upstream intermediate wall disposed between the two arms at the level of a so-called leading edge of each of the two arms and, on the other hand, a downstream intermediate wall disposed between the two arms at the level of an opposite so-called trailing edge of each of the two arms; -the aircraft turbomachine subassembly comprises a series of intermediate walls which extend axially in a successive manner between a so-called leading edge of each of the two arms and a so-called trailing edge of each of the two arms;

[0011] -the series of intermediate walls is configured axially so that a first intermediate wall of the series of intermediate walls is disposed at the leading edge of each of the two arms and a last intermediate wall of the series of intermediate walls is disposed upstream of the trailing edge of each of the two arms or at this trailing edge; -each of the intermediate walls has a radial position which is different depending on the axial position it occupies relative to the axial extension of the two circumferentially adjacent arms along the longitudinal axis XX'; -the aircraft turbomachine subassembly comprises, for each of one or more other pairs of circumferentially adjacent arms of the plurality of circumferentially adjacent arms, one or more intermediate walls extending between the two circumferentially adjacent arms of the pair or other pairs of circumferentially adjacent arms so as to locally separate between the two arms of the pair or each of the other pairs of arms said at least a portion of annular duct into a radially internal portion and a radially external portion further away from the longitudinal axis XX' than the radially internal portion; -the circumferentially adjacent arms do not have a function of straightening the airflow flowing in the annular duct; -the aircraft turbomachine subassembly comprises at least two duct portions, each defining a gooseneck, and which are spaced apart from each other along the longitudinal axis XX'.

[0012] The invention also relates to an aircraft turbomachine comprising an aircraft turbomachine subassembly as briefly described above. Brief description of the drawings

[0013] Other features and advantages of the subject matter of this presentation will become apparent from the following description of embodiments, given by way of non-limiting examples, with reference to the attached figures.

[0014] [Fig-1] Fig. 1 is a schematic longitudinal section view of a mode of construction of a turbojet engine according to the invention;

[0015] [Fig.2] The [Fig.2] is a partial schematic view enlarged in longitudinal section of the portion of annular conduit defining the swan neck of the [Fig.1];

[0016] [Fig.3A] The [Fig.3A] is a schematic front perspective view of an intermediate casing of the reactor core of the [Fig.1] which incorporates a gooseneck;

[0017] [Fig.3B] The [Fig.3B] is a partial enlarged schematic view of an internal region of the housing of the [Fig.3A] incorporating an intermediate wall between two consecutive arms;

[0018] [Fig.3C] The [Fig.3C] is a very schematic view showing in longitudinal section the centered axial position of the intermediate wall of the [Fig.3B];

[0019] [Fig.4] The [Fig.4] is a schematic view of an alternative embodiment of the arrangement of the [Fig.3C];

[0020] [Fig.5] The [Fig.5] is a schematic view similar to that of figures 3C and 4 showing another embodiment with two intermediate walls, one upstream and one downstream;

[0021] [Fig.6] The [Fig.6] is a schematic view similar to that of the [Fig.5] and showing a variant with a series of intermediate walls;

[0022] [Fig.7A] The [Fig.7A] is a schematic view showing a possible variation in the downward direction of an inclination of an intermediate wall;

[0023] [Fig.7B] The [Fig.7B] is a schematic view showing a possible variation in the upward direction of an inclination of an intermediate wall;

[0024] [Fig.8A] Fig.8A is a schematic view following a radial view of walls intermediates in a general Y shape between two consecutive arms;

[0025] [Fig.8B] The [Fig.8B] is a schematic view following a cross-sectional view of intermediate walls in the general shape of a Y between two consecutive arms;

[0026] [Fig.9] The [Fig.9] is a schematic view in longitudinal half-section of another embodiment of a turbojet according to the invention. Detailed description

[0027] Fig. 1 represents a longitudinal section of a turbomachine such as an aircraft turbojet 10 according to an embodiment of the invention (this is a turbofan engine but the invention is not limited to this type of turbojet).

[0028] Such a turbojet has a general longitudinal shape centered around a longitudinal axis XX' and generally comprises, successively from upstream to downstream in the direction of airflow illustrated by the arrows F (i.e. from left to right on the [Fig.1]), an air inlet 12 and a fan 14 (known in Anglo-Saxon terminology as "fan") which supplies air to a primary duct 16 and a secondary duct 18 arranged coaxially with respect to the primary duct 16 and which are both configured in a generally annular shape.

[0029] As shown in [Fig. 1], the turbojet 10 comprises a subassembly such as a reactor core 20 which is internally configured to delimit The primary air stream 16 as described below. The reactor core 20 may comprise, more specifically, successively from upstream to downstream: -at the inlet, a plurality of guide stator blades 22 (IGV) arranged circumferentially relative to the longitudinal axis XX' and which generally have the function of guiding and / or straightening the incoming airflow 16, -a low-pressure compressor 24, -a transition zone 28 located downstream of the low-pressure compressor 24 and which defines a gooseneck (this zone will be represented more precisely in other figures described later), -a high-pressure compressor 26, -a combustion chamber 30, -a plurality of successive axial turbines generally referenced as 32 (high pressure, intermediate pressure and low pressure turbines), -an outlet 34 of the reactor core 20.

[0030] The set of components of the reactor core 20 defined above jointly define an annular conduit or passage 36 which extends concentrically around the longitudinal axis XX' and along the longitudinal or axial direction given by this axis over the entire length of the reactor core 20. The primary air stream 16 thus flows inside this annular conduit 36 ​​to be treated in a conventional manner: the air stream is first compressed successively by the low-pressure compressor 24 and the downstream high-pressure compressor 26, then the compressed air is introduced into the combustion chamber 30 where it is mixed in a known manner with a fuel to generate high-temperature combustion gases which then undergo successive expansions in the set of successive turbines 32 before being discharged through the outlet orifice 34 to generate thrust.

[0031] A nacelle 38 is shown in dotted lines on [Fig.1] and surrounds the turbojet 10 concentrically in a known manner, delimiting the upstream air inlet 12 and, downstream, an air outlet 40.

[0032] The secondary vein 18 circulates inside an annular conduit 42 arranged concentrically with respect to the annular conduit 36 ​​and generates a thrust which adds to the thrust produced by the expanded combustion gases exiting the orifice 34. Radial arms 43 form guide stator blades 22 (OGV).

[0033] The transition zone 28 corresponds to an intermediate portion of the swan-neck shaped annular conduit 36, which includes an upstream inlet section Se and a downstream outlet section Ss ([Fig. 1]). The inlet and outlet sections define the shape of the swan neck, which is more or less constrained, i.e., its slope is more or less pronounced. These sections are generally parallel between they insofar as the gooseneck shape connects between them a portion of annular conduit located upstream of the gooseneck (where the compressor 24 is located) and a portion of annular conduit located downstream of the gooseneck (where the compressor 26 is located) and which is parallel to the upstream portion of annular conduit.

[0034] In the configuration illustrated in [Fig.1] the downstream outlet section Ss is closer to the longitudinal axis XX' than the upstream inlet section Se, thus defining a swan neck with a downward slope.

[0035] Figure 2 schematically illustrates, in enlarged longitudinal section, the presence of an intermediate wall P in the gooseneck 28 which locally separates or divides (i.e., in the region where the intermediate wall extends axially, radially, and circumferentially) the portion of the annular duct into a radially internal portion Ci (internal channel) close to the longitudinal axis XX' and a radially external portion Ce (external channel) further from the longitudinal axis than the internal portion Ci and which is positioned above it in Figure 2. As shown in Figure 2, the gooseneck 28 is defined here as corresponding to the portion of the duct and the associated blades extending from the trailing edge of the upstream blade Aam (last blade of the upstream compressor 24) to the leading edge of the downstream blade Aav (first blade of the downstream compressor 26).

[0036] This arrangement allows the swan neck to be transformed into two swan necks of reduced height: a lower swan neck of height hl flowing in the radially internal portion Ci (internal channel) and an upper swan neck of height h2 flowing in the radially external portion Ce (external channel). The airflow in the swan neck 28 is thus better guided in each of the radially internal and external portions than in a swan neck without an intermediate wall P, and the criticality of the swan neck is thus limited.

[0037] Fig. 2 illustrates the height H, which is one of the two geometric parameters of the swan neck, namely its height H and its slope, which forms an angle with respect to the longitudinal direction of the axis XX'.

[0038] Figures 3A and 3B are detailed views that represent the application of the principle illustrated in [Fig. 2] to the intermediate casing 44 of the reactor core 20 of [Fig. 1]. It should be noted, however, that the principle illustrated in [Fig. 2] can also be applied to the upstream gooseneck 27, which here forms part of the inlet casing, or even to the downstream casing, in addition to or instead of the gooseneck 28.

[0039] As shown in Figures 3A and 3B, the housing 44 has an external portion forming an annular belt 46 which externally borders the internal gooseneck-shaped transition zone 28. The annular belt 46 and the gooseneck 28 are both arranged concentrically with respect to a portion of the hub internal hub 48 of reactor core 20, the annular conduit 36 ​​surrounding this internal hub. In Figures 3A and 3B only a portion of the internal hub 48 is shown.

[0040] A plurality of arms, here eight arms B1-B8 (this number may, however, differ depending on the turbojet configuration), are arranged inside the gooseneck 28, extending radially in a circumferential arrangement relative to the longitudinal axis XX', each with a distinct radial orientation as illustrated in [Fig. 3A]. The arms are circumferentially spaced from one another (for example, regularly as shown here). Each of the arms B1-B8 extends radially along the entire height of the gooseneck (radial dimension), i.e., from the inner hub 48 to the annular belt 46, and axially or longitudinally along all or part of the length of the gooseneck. Two consecutive arms (i.e., circumferentially adjacent) define, locally, a slice or volumetric segment of the portion of the annular duct 28 defining the gooseneck.This applies here to all arms of the plurality of arms distributed circumferentially. Such a slice or volumetric segment extends along an angular sector (defined by the angular spacing between the arms in a cross-sectional view) and axially or longitudinally and thus defines, between these arms, a local volume in which flows a fraction of the air stream that circulates in the portion of annular duct 28 defining the swan neck.

[0041] The arms extending inside the gooseneck are referred to here as structural arms, meaning that these arms have a mechanical function of connecting and transmitting forces between the annular belt 46 and the internal hub 48. In the present embodiment, these arms each have a straight / straight profile from upstream to downstream (along their axial extension) and without inclination (no aerodynamic profile), and therefore do not have a function of straightening / guiding the airflow. However, in other embodiments not shown, these structural arms can also be aerodynamically profiled so as to exert an aerodynamic function on the flow circulating in the portion of the annular duct 28 defining the gooseneck. Alternatively, the arms extending into the gooseneck may have only an aerodynamic function.In other possible configurations, only part of the arms may have a structural function, possibly with aerodynamic fairing.

[0042] In the present embodiment, an intermediate wall PI at P8 extends circumferentially between the two circumferentially adjacent (consecutive) arms of each pair of arms, namely the arm pairs B1 - B2, B2 - B3, B3 - B4, B4 - B5, B5 - B6, B6 - B7, B7 - B8 and B8 - B1. It should be noted that in other turbojet configurations not shown here, only certain arm pairs of the plurality of arms may include, between the consecutive arms pairs of arms concerned, one or more intermediate walls within the meaning of the present invention.

[0043] In the configuration shown in Figures 3A and 3B, a single intermediate wall is positioned between the two circumferentially adjacent arms of each pair of arms (and mechanically connected in a known manner to each of the arms) so as to subdivide the slice or segment of the annular duct portion delimited externally by these arms into a radially internal portion Ci (internal channel) and a radially external portion Ce (external channel) as illustrated in Figures 2 and 3B (volume space between arms B1 and B2). As shown in [Fig. 3B], the intermediate wall, such as P2 in [Fig. 3B], has a profile that follows the profile of the gooseneck, i.e., it conforms to the shape of the internal (internal hub wall 48) and external (internal wall of the annular belt 46) walls that define the gooseneck, so as not to introduce too much disturbance in the guidance of the air stream.

[0044] In the present embodiment, which is schematically represented in the enlarged longitudinal section of [Fig. 3C], the intermediate wall PI is shown axially centered with respect to the axial extension of arm B1 shown in the background. This wall can extend over the entire axial extension of arm B1 (and arm B2) or only a fraction of the axial extension of arm B1 (and arm B2), being equidistant from the leading edge ba and the trailing edge bf of the arm. The configuration where the intermediate wall extends over only a fraction of the axial extension of arm B1 (and arm B2) helps to limit aerodynamic losses. In the present embodiment, the wall PI thus has an axial extension or length that can vary between 50 and 100% of the axial extension or length of the arm.

[0045] When the wall extends only over a fraction of the axial extension of the arm, such an arrangement makes it possible to avoid limiting the ability of the flow to re-distribute itself radially upstream of the downstream compressor 26 of [Fig.1], in particular towards an area radially external to the internal channel Ci, in the case where the latter is obstructed or limited due to a possible abnormal behavior of this downstream compressor (for example a lack of flow draw at the foot of the compressor).

[0046] In an alternative embodiment schematically illustrated in the longitudinal section of [Fig. 4], the intermediate wall PI' extending between the two consecutive (circumferentially adjacent) arms of the same pair of arms is shorter than in [Fig. 3C] and is axially offset upstream (moving closer to the leading edge ba of arm B1 and away from the trailing edge bf of this arm) relative to the axially centered position of [Fig. 3C]. This arrangement provides the flow with a greater capacity, than in the mode of [Fig. 3C], to radially redistribute itself. upstream of the downstream compressor 26 of [Fig. 1] because the flow is less radially constrained in this region. According to this variant, the wall PI' thus has an axial extension or length less than 100% of the axial extension or length of the arm and which can be as low as 50% of the latter.

[0047] In the configurations of figures 3C and 4 the angle or inclination of the intermediate wall Pl / Pl' with respect to the longitudinal direction XX' can be identical to that of the swan neck (slope of the swan neck) or deviate from it by an angular deviation of + / - 20°.

[0048] Furthermore, in these configurations the implantation height of the intermediate wall, namely its radial position relative to the internal wall pi of the swan neck (here the lower wall in the figures), is located between 35 and 65% of the height H so that the heights of the two channels are not too disproportionate to each other.

[0049] Figure 5, which is a view analogous to that of Figures 3C and 4, illustrates another embodiment in which several intermediate walls extend between the two circumferentially adjacent arms of the same pair of arms, being distributed axially along the axial extension or length of the two arms. More particularly, an upstream intermediate wall PI” is disposed between the two arms at the leading edge ba of each arm (protruding slightly beyond this edge), and a downstream intermediate wall P2’’ is disposed between the two arms at the trailing edge bf of each arm.

[0050] In the present embodiment, each intermediate wall has an axial extension or length that can vary between 10% and 50% of the axial extension or length of the arm. The two walls may, for example, have radial positions offset from each other.

[0051] It should be noted that the installation height or radial position of each intermediate wall relative to the internal wall pi of the gooseneck (here the lower wall in the figures) can vary depending on the axial position (in the longitudinal direction of the axis XX') of the latter, in particular to optimize flow guidance and aerodynamic losses. The installation height of the intermediate wall PI” can thus be greater than or less than that of the intermediate wall P2” depending on the geometry. The installation height of the intermediate walls can vary between 25 and 75% of the height H.

[0052] In this embodiment, the angle or inclination of each intermediate wall can be identical to that of the swan neck (slope of the swan neck) or deviate from it by an angular deviation of + / - 20°.

[0053] Figure 6, which is a view analogous to that of Figure 5, illustrates an alternative embodiment in which a series of intermediate walls extend, on the one hand, circumferentially between the two circumferentially adjacent arms of the same pair of arms and, on the other hand, axially successively between the leading edge ba of each of the two arms and the trailing edge bf of each of them.

[0054] More particularly, the series of intermediate walls which here comprise five elements or fins (this number may however differ according to the configurations and for example be three, four, ...ten), is configured axially so that the first intermediate wall PI”' is disposed at the level of the leading edge ba of each of the two arms and so that the last intermediate wall P5”' is disposed at a distance and upstream of the trailing edge bf of each of the arms or at this trailing edge bf.

[0055] In this variant, each intermediate wall has an axial extension or length which can vary between 10 and 15% of the axial extension or length of the arm.

[0056] It should be noted that the installation height or radial position of each intermediate wall can vary depending on their axial position, in particular to optimize flow guidance and aerodynamic losses. Thus, the installation height of the first intermediate wall PI'" can be greater than that of the second intermediate wall immediately downstream, and so on up to the last intermediate wall P5'". The reverse arrangement is also possible. The installation height of the intermediate walls can vary between 25 and 75% of the height H.

[0057] In this variant the angle or inclination of each intermediate wall can be identical to that of the swan neck (slope of the swan neck) or deviate from it by an angular deviation of + / - 20°.

[0058] Figures 7A and 7B illustrate a possible configuration in which the aircraft turbojet includes means for modifying the inclination of an intermediate wall PI”” (identical to the wall PI' of [Fig.4]) relative to the annular duct portion 28. More particularly, the inclination of this intermediate wall relative to the internal wall pi (and therefore relative to the axis XX') can vary according to the inclination command communicated to it by conventional means such as a hydraulic or electric actuator in order to adopt an inclination angle a (downward slope) relative to a neutral inclination angle i corresponding to the inclination or slope of the gooseneck ([Fig.7A]) or an inclination angle [3 (upward slope) relative to a neutral inclination angle i corresponding to the inclination or slope of the gooseneck ([Fig.7B]). The value of each angle can vary within a range of 20°.

[0059] This variable tilt and position control can be used according to engine speed, for example to limit aerodynamic losses and optimize the efficiency of the intermediate wall. Thus, the configuration of [Fig. 7A] redirects the flow towards the hub (inner wall of the housing), thereby limiting the losses on the latter. The configuration of [Fig.7B] redirects the flow towards the outer wall of the housing, thus limiting losses on the latter.

[0060] Figures 8A and 8B illustrate another possible configuration of internal gooseneck arrangement according to the present invention, in which the intermediate wall extending between two circumferentially adjacent arms (and this applies to the entire plurality of radial arms arranged in the gooseneck; in other configurations not shown, the intermediate wall may only be present between certain circumferentially adjacent arms) has, according to a radial view illustrated in [Fig. 8A] for arms B1', B2' and B3' (i.e., a view taken along a radius of the gooseneck, from outside the latter), a general Y shape. The common branch Y1a, Y2a of each wall Y1, Y2 on [Fig.8 A] is arranged against one of the two circumferentially adjacent arms (here respectively Bl' and B2') and the two branches Ylb and Ylc, Y2b and Y2c of each wall Yl, Y2 are arranged against the second of the two arms (here respectively B2' and B3'). .

[0061] Moreover, as illustrated in the cross-sectional view (relative to the longitudinal axis XX') of [Fig.8B], each intermediate wall, such as the walls Y1 and Y2 defined above, which extends between the two circumferentially adjacent (e.g. structural) arms also has a general Y shape (following the radial extension): the common branch of the Y, here Yla and Y2a, is arranged against the first arm (here Bl' and B2') and the two end branches Ylb and Yle, Y2b and Y2c of each wall Yl, Y2 are in fact split to each adopt a Y shape. Only the branch Yle and Y2c of each wall is represented in a split form on [Fig.8B] in two sub-branches Ylcl and Ylc2, Y2cl and Y2c2 respectively.

[0062] Each intermediate wall thus has a double Y shape with a different axial position depending on whether the wall is to the left or right of the arm ([Fig. 8A]) and a different radial position depending on whether the wall is to the left or right of the arm ([Fig. 8B]). The double Y configuration of the intermediate walls has an aerodynamic function of straightening the flow. This arrangement can prove useful if the air entering the gooseneck has not been sufficiently straightened upstream. The asymmetry of the different branches allows adaptation to the asymmetry of the flow.

[0063] In a configuration not shown in the figures, the reverse arrangement can be envisaged, namely that the branches of the Y are oriented to the left.

[0064] In the described embodiments and their variants, the turbojet engine is of the twin-spool type. However, the turbojet engine may also be of the triple-spool type. The turbojet engine may be of the fan-driven type (known as a "turbofan" in Anglo-Saxon terminology) or a propeller-driven type. In the latter configuration, the difference in height between the inlet of the primary air stream and the casing or body of the high-pressure compressor is particularly significant, resulting in a gooseneck. in the transition zone between the low-pressure compressor and the high-pressure compressor, which is particularly constrained and therefore highly critical. The invention is thus particularly useful in this type of configuration. Furthermore, certain turbojet configurations, especially propeller-driven ones, include, in the annular duct through which the airflow circulates, at least two duct sections, each defining a downward-sloping gooseneck, spaced apart along the longitudinal axis XX' (at least two successive upstream and downstream goosenecks). The invention is particularly applicable to this type of configuration, as well as to any turbojet configuration featuring a downward-sloping gooseneck.

[0065] Fig. 9 illustrates a possible example of a turbofan engine architecture 60 with an unfaired fan 62, including here contra-rotating propellers, and in which a gooseneck 64 arranged upstream is likely to incorporate one of the embodiments or variants described above.

[0066] Although the present description refers to specific embodiments, modifications may be made to these examples without departing from the general scope of the invention as defined by the claims. Furthermore, individual features of the various embodiments illustrated or mentioned may be combined in additional embodiments. Therefore, the description and drawings should be considered in an illustrative rather than restrictive sense.

Claims

Claims

1. Aircraft turbomachine subassembly (10) comprising an annular duct (36) which extends concentrically around a longitudinal axis (XX') and in which an air flow flows from upstream to downstream, the annular duct comprising at least one duct portion (28) which defines a swan neck having an upstream inlet section (Se) and a downstream outlet section (Ss), the turbomachine subassembly comprising, in said at least one duct portion (28), a plurality of arms (B1-B8) which extend according to a circumferential distribution relative to the longitudinal axis (XX') in a manner spaced apart from each other, each following a radial and axial extension, characterized in that the aircraft turbomachine subassembly comprises several intermediate walls (P; P1-P8; PI'; PI”; PI'”; PI””;Y1-Y8) extending between two circumferentially adjacent arms which form a pair of circumferentially adjacent arms among the plurality of circumferentially adjacent arms, so as to locally separate between the two arms of the pair of arms said at least one portion of annular conduit (28) into a radially internal portion (Ci) and a radially external portion (Ce) further from the longitudinal axis (XX') than the radially internal portion, the intermediate walls being distributed axially along the axial extension of the two arms.;

2. Aircraft turbomachine subassembly according to claim 1, in which the downstream outlet section (Ss) of the swan neck is closer to the longitudinal axis (XX') than the upstream inlet section (Se).

3. Aircraft turbomachine subassembly according to claim 1 or 2, wherein the aircraft turbomachine subassembly comprises means for modifying an inclination of the intermediate walls (PI”) relative to said at least one portion of annular duct.

4. Aircraft turbomachine subassembly according to claim 1, in which the aircraft turbomachine subassembly comprises, on the one hand, an upstream intermediate wall (PI”)) arranged between the two arms at a so-called leading edge (ba) of each of the two arms and, on the other hand, a downstream intermediate wall (P2”) arranged between the two arms at the level of an opposite edge called the trailing edge (bf) of each of the two arms.

5. Aircraft turbomachine subassembly according to claim 1, wherein the aircraft turbomachine subassembly comprises a series of intermediate walls (Pl'”-P5”') which extend axially successively between a so-called leading edge (ba) of each of the two arms and a so-called trailing edge (bf) of each of the two arms.

6. Aircraft turbomachine subassembly according to claim 5, wherein the series of intermediate walls (Pl”'-P5'”) is configured axially so that a first intermediate wall (PI”') of the series of intermediate walls is arranged at the leading edge (ba) of each of the two arms and so that a last intermediate wall (P5'”) of the series of intermediate walls is arranged upstream of the trailing edge (bf) of each of the two arms or at this trailing edge.

7. Aircraft turbomachine subassembly according to one of claims 1, 4, 5 and 6, in which each of the intermediate walls has a radial position which is different depending on the axial position which it occupies relative to the axial extension of the two circumferentially adjacent arms along the longitudinal axis XX'.

8. Aircraft turbomachine subassembly according to one of claims 1 to 7, wherein the aircraft turbomachine subassembly comprises, for each of one or more other pairs of circumferentially adjacent arms of the plurality of circumferentially adjacent arms, one or more intermediate walls extending between the two circumferentially adjacent arms of the pair or other pairs of circumferentially adjacent arms, so as to locally separate between the two arms of the pair of arms or of each of the other pairs of arms said at least one annular duct portion into a radially inner portion and a radially outer portion further from the longitudinal axis (XX') than the radially inner portion.

9. Aircraft turbomachine subassembly according to one of claims 1 to 8, in which the circumferentially extending arms adjacent ones do not have a function of straightening the air flow flowing in the annular duct.

10. Aircraft turbomachine subassembly according to one of claims 1 to 9, in which the aircraft turbomachine subassembly comprises at least two duct portions each defining a swan neck and which are spaced from each other along the longitudinal axis (XX').

11. Aircraft turbomachine comprising an aircraft turbomachine subassembly according to one of claims 1 to 10.

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