Helicopter turbomachine subassembly and helicopter turbomachine comprising such a subassembly

Abstract

A turbomachine subassembly (10) for a helicopter, comprising a primary nozzle (12) for the outlet of an exhaust gas stream (F1), which includes an outer casing (14) and an inner cone (16) arranged concentrically around a longitudinal axis (X) and connected together by radial arms (18), the casing (14) surrounding the cone (16) and defining between them an annular duct (C) for the flow of the stream (F1). The subassembly includes a distributor (20) encircling the casing (14) and configured to, on the one hand, receive at its inlet (20a) a cooling air stream (F2) having a temperature lower than the temperature of the stream (F1) and, on the other hand, distribute the stream (F2) at its outlet within radial arms (18). The internal cone has an external wall (16a; 16b') in which one or more openings (O) are provided to allow air to (F1). Figure for the abbreviation: Fig. 1.

Description

Title of the invention: Helicopter turbomachine subassembly and helicopter turbomachine comprising such a subassembly Technical field

[0001] The present invention relates to helicopter turbomachines. Prior art

[0002] Aircraft turbomachines, and more particularly helicopter turbomachines, comprise, in the case of a twin-engine helicopter, a propulsion unit which generally comprises a first engine configured to drive a main rotor and a second engine configured to drive a tail rotor. The two engines are generally connected to a main gearbox which has the role of distributing the power between the main rotor and the tail rotor. Each engine may be a heat engine, and more specifically a gas turbine engine comprising a compressor, a combustion chamber, a first turbine operatively connected by a rotating shaft to the compressor and a second turbine, or free turbine, coupled to a power take-off shaft for transmitting the movement to the rotor concerned.The assembly consisting of the compressor, the combustion chamber, the first turbine and the rotating shaft of the engine constitutes a gas generator.

[0003] The exhaust gases that come from the free turbine mentioned above are at very high temperatures. However, in certain flight situations, for example in the case of hovering flight, the exhaust gases can be redirected by the blast of the main rotor onto the tail boom of the helicopter. This can prove detrimental to the structure of the latter because, increasingly, composite materials are used to manufacture the structure of the helicopter and these materials are very often sensitive to thermal aggression. In order to withstand these high temperatures, it is therefore necessary to protect at least certain areas of the external fairing of the tail boom by thermal protection which, however, significantly increases the on-board mass and the economic cost of manufacturing the helicopter.

[0004] It would therefore be particularly useful to be able to remedy at least in part the drawbacks mentioned above. Statement of the invention

[0005] This disclosure is the result of technological research aimed at reducing the temperature of the exhaust gases in order to limit, or even eliminate, the need for thermal protection on the external fairing of the tail boom of the helicopter. For this purpose, it is notably planned to use air already available in the turbomachine, this air being at a temperature lower than the temperature of the exhaust gases, to cool the latter.

[0006] A first aspect of the present disclosure relates to a turbomachine subassembly intended for a helicopter, comprising a primary outlet nozzle for an exhaust gas flow, the primary nozzle comprising an outer casing and an inner cone which are arranged concentrically with respect to each other around a longitudinal axis of the turbomachine and connected to each other by a plurality of radial arms, the outer casing surrounding the inner cone and defining with the latter an annular duct for the flow of the exhaust gas flow, characterized in that the turbomachine subassembly comprises a distributor at least partially surrounding the outer casing, the distributor being configured to, on the one hand, receive, at the inlet of the distributor, a flow of cooling air existing in the turbomachine and having a temperature lower than the temperature of the exhaust gases and, on the other hand, distribute, at the outlet of the distributor,this flow of cooling air inside at least some radial arms of the plurality of radial arms (the at least some arms thus each ensure the distribution of a fraction of cooling air flow) to the inside of the internal cone, the internal cone (in which the fractions of cooling air flow from the radial arms are distributed) comprising an external wall in which one or more openings are arranged to allow the cooling air from the internal cone to exit said internal cone and to be injected at an internal periphery of the exhaust gas flow.

[0007] The use of an existing air flow in the turbomachine at the level of the internal cone of the primary nozzle (exhaust nozzle) makes it possible to cool the exhaust gas flow from the inside (on its internal periphery, i.e. at the center of the flow), i.e. in a part of this flow which is in depression, which proves to be particularly effective for the cooling operation. It follows that the external fairing of the tail boom of the helicopter integrating the turbomachine is less subject than previously to high temperatures which required the use of heat shields on the external fairing. It is therefore possible to limit the number of heat shields, or even to eliminate them (in certain circumstances), which makes it possible to reduce the mass on board the helicopter.Furthermore, the use of an air flow already existing in the turbomachine means that there is no need to take any samples from the engine(s), which therefore does not affect the performance of the latter.

[0008] According to other possible characteristics: -the outer wall of the inner cone has a downstream end face, following the direction of flow of the exhaust gas flow in the helicopter turbomachine, which is open to the outside so as to allow the cone cooling air internal to exit from the latter in a substantially axial direction and directed downstream; -the inner cone is configured to guide the cooling air exiting the inner cone; -the inner cone comprises a worm-like device having a helical groove which defines a helical path for guiding cooling air from the inner cone to the outside; -the inner cone comprises ducts which are each configured inside to convey a portion of the cooling air of the inner cone, the ducts being wound relative to each other so as to form a helix; -the inner cone comprises, upstream of the plurality of radial arms, a plurality of openings distributed circumferentially in the outer wall of the inner cone which is adjacent to the annular duct, thus allowing the fractions of cooling air flow distributed by said at least some radial arms of the plurality of radial arms to exit the inner cone by opening into the annular duct at an inner periphery of the exhaust gas flow; -the internal cone comprises at least one internal guiding device for the fractions of cooling air flow distributed by said at least some radial arms of the plurality of radial arms to the plurality of circumferential openings; -said at least one internal guide device further comprises a plurality of circumferential fins which are arranged on the path of the cooling air flow fractions in order to impart a gyrating movement to the air exiting through the circumferential openings; - the cooling airflow comes from a cooling system of a lubricant of the helicopter turbomachine; - the cooling airflow comes from a cooling system of a turbomachine starter-generator.

[0009] A second aspect of the present disclosure relates to a helicopter turbomachine comprising a turbomachine subassembly as briefly discussed above.

[0010] A third aspect of the present disclosure relates to a helicopter comprising a turbomachine as briefly described above. Brief description of the drawings

[0011] The invention will be better understood and its advantages will appear better, on reading the detailed description which follows, of embodiments represented by way of non-limiting examples. The description refers to the appended drawings which are schematic and aim above all to illustrate the principles of the disclosure.

[0012] In these drawings, from one figure to another, identical elements (or parts of elements) or equivalents are identified by the same reference signs. On these attached drawings:

[0013] [Fig. 1] [Fig. 1] schematically illustrates a first embodiment of the invention of a turbomachine subassembly.

[0014] [Fig.2] [Fig.2] schematically illustrates a second embodiment of the invention of a turbomachine subassembly.

[0015] [Fig.3] [Fig.3] schematically illustrates a possible embodiment of the first embodiment of the invention.

[0016] [Fig.4A] [Fig.4A] is an axial sectional view of a turbomachine subassembly according to the first embodiment of the invention and incorporating a worm screw device.

[0017] [Fig.4B] [Fig.4B] is a perspective view of the worm gear device of [Fig.4A].

[0018] [Fig.5] [Fig.5] is a perspective view of a turbomachine subassembly according to the first embodiment of the invention and integrating conduits wound in the manner of a propeller.

[0019] [Fig.6A] [Fig.6A] is a perspective view from upstream of a possible embodiment of a turbomachine subassembly according to the second embodiment of the invention.

[0020] [Fig.6B] [Fig.6B] is a partial axial sectional view of the turbomachine subassembly of [Fig.6A].

[0021] [Fig.6C] [Fig.6C] is a perspective view from upstream of the interior of the inner cone of the turbomachine subassembly of Figures 6A and 6B.

[0022] [Fig.7] [Fig.7] schematically illustrates an example of an aircraft with a set propulsion system comprising two engines.

[0023] [Fig.8] [Fig.8] more specifically illustrates an embodiment of the propulsion assembly of [Fig.7]. Description of the embodiments

[0024] In order to make the disclosure more concrete, embodiments are described in detail below, with reference to the accompanying drawings. It is recalled, however, that the invention is not limited to these embodiments.

[0025] Figures 1 and 2 schematically represent two main embodiments of the invention.

[0026] Generally, a subassembly 10 of an aircraft turbomachine such as a helicopter comprises a primary nozzle 12 partially shown and which comprises an external casing 14 centered on a longitudinal axis X of the turbomachine, along which a flow of exhaust gas (hot gas from a combustion process internal to the turbomachine, in particular coming from a gas turbine) flows from upstream to downstream, as indicated by the arrow Fl. The external casing 14 has an axial shape in an upstream part, then a divergent shape in a downstream part for the evacuation of the exhaust gas flow Fl out of the turbomachine.

[0027] The primary nozzle 12 also comprises an inner cone 16 which is arranged inside the outer casing 14, and arranged concentrically with respect to the latter around the longitudinal axis X of the turbomachine. The outer casing 14 surrounds the inner cone 16 so as to define between them an axial annular duct C for the flow of the exhaust gas flow FL. The outer casing 14 and the inner cone 16 are mechanically connected to each other by a plurality of arms 18 which extend radially relative to the longitudinal axis X inside the axial annular duct C. These radial arms 18 are distributed circumferentially inside the annular duct, and are for example arranged regularly with respect to each other.Generally, the plurality of these radial arms 18 provide a function of mechanical support and stiffening of the assembly (structural arms). At least part of the plurality of these radial arms 18 are hollow so as to be able to convey an air flow inside.

[0028] The turbomachine subassembly 10 also comprises a distributor 20 (device for distributing an air flow) which at least partially surrounds the external casing 14 of the nozzle 12. By way of example, in one possible embodiment, the distributor 20 takes the form of a hollow annular casing surrounding the external casing 14 and which is fixed to the latter. In another possible embodiment, the distributor may have a semi-annular shape (one or two half-rings) or take the form of several segments each forming ring portions.

[0029] The distributor 20 receives, at its inlet 20a, an existing air flow (air flow) coming from a source S internal to the turbomachine and which does not require taking additional power from the engine(s) of the turbomachine so as not to affect their performance.

[0030] In a possible embodiment, the existing air flow comes from a system for cooling a lubricant of the helicopter turbomachine. The lubricant is for example used for the lubrication of the engine or engines of the aircraft turbomachine and it may for example be oil. By way of example, the aforementioned cooling system may comprise an air-lubricant heat exchanger which is crossed by a circuit of hot lubricant coming from an engine of the turbomachine and by an air flow which may be outside air which heats up by heat exchange with the lubricant in order to cool the latter. The air flow leaving the exchanger is hotter than when it enters the exchanger but, in general, its temperature is nevertheless lower than the temperature of the exhaust gas flow F1, which will allow this air flow, hereinafter denoted F2, to be used for cooling the flow Fl. For example, the air flow is heated by oil to a temperature of 120°C while in the exhaust zone the exhaust flow can exceed 600°C. Thus, the source S of [Fig.l] generates an air flow F2, called the cooling air flow, which is conveyed via one or more conduits 22 (arranged here from upstream to downstream) from the source S to the inlet 20a of the distributor 20 to supply the latter with cooling air.

[0031] As a variant, it will be noted that the source of cooling air used to cool the flow F1 may come from a cooling system or circuit which is dedicated to cooling a generator-starter assembly and generally comprises a supply duct and an outlet for discharging the heated air. It is thus conceivable to use the heated air at the outlet of the generator-starter assembly to convey it to the distributor 20 described above. In known configurations of generator-starter assembly cooling systems, either a duct is provided at the outlet to convey the heated air to the outside of the aircraft, or the hot air is discharged directly into the engine compartment.

[0032] Other sources of cooling air may alternatively be used such as the airflow from the helicopter's main gearbox cooling system.

[0033] The distributor 20 is connected to the outer casing 14 and to the radial arms 18 and, more particularly, to one end of each of the hollow radial arms 18 which is open to the distributor, thus allowing fluid communication between the outlet of the distributor and the interior of the radial arms concerned. Thus, the flow of cooling air F2 which enters the distributor 20 is distributed or distributed inside each of the hollow radial arms 18 which thus each ensure the routing of a fraction of this flow of cooling air to the inner cone 16 to which each of the arms is connected. Each of the hollow radial arms 18 has an end opposite that in communication with the distributor and which is open to the inner cone 16 so as to establish fluid communication between these arms and the interior of the cone.It will also be noted that this opposite end of the arms can penetrate inside the cone and be shaped locally so as to direct the fraction of air flow conveyed by each arm in a substantially axial direction and directed downstream of the cone, as illustrated schematically in dotted lines in [Fig.l]. The fractions of cooling air flow thus distributed in the internal cone by the radial arms can be more or less mixed inside the cone 16 and flow axially downstream, as indicated by the arrows F3 in [Fig.l]. However, in a possible embodiment, the fractions of cooling air flow thus distributed in the internal cone 16 can be redirected in this . latter following an axial direction downstream without actually mixing before reaching, outside the cone, the exhaust gas flow by which they are sucked in. It will be noted that in the embodiment of [Fig.l] the internal cone 16 comprises an external wall which here forms an end face 16a, of transverse or radial extension relative to the longitudinal axis X, which is here open to the outside to allow the cooling air (flow F3) to exit the internal cone 16 and to be injected at an internal periphery of the exhaust gas flow F1 (this injection zone is a depression zone of the exhaust gas stream, which allows the cooling air exiting the cone to be sucked in).This air flow brought into contact with the exhaust gas flow allows the latter to be cooled and thus limits the undesirable effects of an excessively high temperature on the external fairing of the tail boom of the helicopter in which the turbomachine is installed.

[0034] In the embodiment of [Fig.l] the cone 16 can be configured internally to guide the cooling air F3 exiting the cone, which allows this air to have greater cooling efficiency on the exhaust gas flow. Guiding the flow in the cone makes it possible to reduce the pressure drop and therefore to distribute more cooling air flow. This also makes it possible to impart a gyration to the flow F3 in order to promote mixing with the hot flow F1. Different arrangements or devices, represented schematically in dotted lines by the reference 24, internal to the internal cone 16 are conceivable to ensure the guidance of the air in a substantially axial direction and oriented downstream of the subassembly. Possible examples of such internal arrangements will be described later.Such arrangements make it possible to reduce pressure losses and to impart a gyration to the air flow leaving the cone, in order to promote mixing and maximize the resulting reduction in temperatures. Generally, it is sought to favor high power operating points and, for this purpose, gyration angles between 0 and 40 degrees are envisaged. The gyration angle of the flow is generally defined as being equal to the arctangent of the ratio of the tangential velocity of the flow to the axial velocity of this flow. In an embodiment not shown, the interior of the internal cone may be left hollow and not incorporate such arrangements or devices.

[0035] It will be noted that only the turbomachine subassembly to which the invention applies is described and that the other constituent elements of the turbomachine which are not necessary for understanding the invention are not described here.

[0036] The description preceding [Fig.l] also applies to [Fig.2] with regard to the routing of a cooling air flow F2 to the distributor 20 and its distribution in the hollow radial arms 18, then the transport of the air flow fractions distributed in these arms to the internal cone. The references of the elements identical from one figure to another are preserved.

[0037] However, the description of the turbomachine subassembly 10' of [Fig.2] differs from that of [Fig.l] from the introduction of the cooling air flow fractions inside the internal cone 16' whose configuration differs from that of the internal cone 16.

[0038] The internal cone 16' here comprises one or more internal guiding devices for the cooling air, shown schematically in dotted lines by the reference 24'. This or these devices ensure a reorientation of the air fractions coming from the radial arms 18 so that the radial orientation of the air fractions during their transport by the arms 18 is modified by being redirected upstream, then again downstream so as to be able to be injected into the exhaust gas flow F1 in the same upstream-downstream flow direction as the latter. The change in orientation of the cooling air thus corresponds to a 360° turnaround. The injection of the cooling air into the exhaust gas flow takes place upstream of the radial arms.

[0039] Although not shown in [Fig. 2] (other figures described later will illustrate this aspect), the inner cone 16' comprises, upstream of the plurality of radial arms, a plurality of openings distributed circumferentially in the outer wall 16b' of the inner cone which is adjacent to the annular duct C. The openings make it possible to inject the cooling air at the inner periphery of the exhaust gas flow F1, opening into the annular duct C in the form of a plurality of cooling air fractions identified by the arrows F4 in [Fig. 2]. The cooling air fractions are directed into the exhaust gas flow F1 by the guide device(s) internal to the inner cone 16'.

[0040] The cooling air is thus guided (and redirected) inside the internal cone 16' by the internal guide device(s) to the plurality of circumferential openings which it passes through with the desired orientation in order to open into the annular duct at an angle of less than 90° relative to the axial direction of the flow FL

[0041] This air flow F4 brought into contact with the exhaust gas flow F1 makes it possible to cool the latter and thus to limit the undesirable effects of an excessively high temperature on the external fairing of the tail boom of the helicopter in which the turbomachine is installed.

[0042] [Fig. 3] illustrates a possible embodiment of a turbomachine subassembly according to the first embodiment of [Fig. 1], showing a conduit 22 for conveying the cooling air flow F2 (with a given flow rate) coming for example from a cooling system for a lubricant of the helicopter turbomachine, for example known by the acronym OCS (for “Oil Cooling System” in English terminology). The conduit connects the source S of the [Fig.l] (not shown here) at the inlet 20a of the distributor 20 which is flared so as to gently divide (without introducing excessive pressure losses) the inlet air flow F2 inside the distributor into two parts (following a large angular sector, for example, substantially 180°) which each use an annular half-ring to be distributed spatially throughout the entire annular ring of the flow distributor 20.

[0043] The internal cone 16 of the primary nozzle 16 is connected to the external casing 14 by means of the radial arms 18 which here are, for example, five in number (although a higher or lower number of arms is conceivable).

[0044] As shown in [Fig.4A], which is an axial sectional view of the subassembly 10 of [Fig.3], the internal cone 16 comprises a worm-type device 26 (shown in perspective in [Fig.4B]) which fluidly communicates with the cooling air fractions distributed by radial arms 18 so that each air fraction can be transported downstream by the device 26.

[0045] The downstream end face 16a (oriented transversely / radially relative to the longitudinal axis X) of the external wall of the internal cone 16 is open to the outside, so that the cooling air transported downstream by the worm-type device exits the internal cone 16 via the face 16a in a substantially axial direction and directed downstream and can thus reach the internal periphery of the exhaust gas flow FL

[0046] More particularly, the worm-type device 26 comprises a helical groove 26a which defines a helical path for the passage and guidance of the cooling air from the internal cone 16 to the outside thereof. The helical groove 26a extends, for example, from a base 26b which is arranged upstream of the radial arms 18 ([Fig.4A]), downstream in the direction of the open end face 16a.

[0047] In practice, the helical groove 26a comprises a number of turns which may be equal to the number of radial arms through which the cooling air fractions pass. Thus, each air fraction coming from an arm opens onto one of the turns of the propeller. Such a configuration makes it possible to reduce pressure losses by channeling the air flow leaving each of the arms to the outside of the cone, avoiding harmful turbulence linked to the mixing of the air fractions inside the cone. In this configuration, the air fractions coming from the arms remain independent of each other until they leave the endless screw device. Furthermore, the screw device finally also makes it possible to control the angle of gyration which is imparted to the air flow leaving the cone, by adapting the inclination of the turns appropriately.

[0048] As shown in [Fig.5], the internal cone 16 of the nozzle comprises another type of air flow guide which is here formed by ducts 28 internal to the cone which are each configured, within said ducts, to convey a portion of the cooling air of the internal cone. As illustrated in [Fig.5], the hollow ducts 28 are wound relative to each other so as to generally form a helix. Individually, each duct is shaped in the manner of a helix and all the corresponding helices are nested within each other as shown in [Fig.5].

[0049] In practice, in order to minimize pressure losses, the number of ducts 28 is equal to the number of radial arms 18 which each convey a fraction of cooling air. The ducts 28 can have various shapes and are each configured to allow the angle of gyration of the flow at the outlet of the cone to be controlled.

[0050] Each duct 28 is connected to a radial arm 18 and [Fig. 5] shows, in a given transverse plane (downstream of the connections between arms and ducts), the different sections 28a of the different ducts arranged circumferentially with respect to each other. If only certain radial arms are capable of conveying a flow of cooling air, only these arms are connected to a duct 28. It will be noted, however, that, in an embodiment not shown, the number of arms conveying a flow of cooling air is different from the number of ducts. For example, the same duct can receive air from several arms.

[0051] Figures 6A-6C illustrate a possible embodiment of a turbomachine subassembly 10' according to the second embodiment of [Fig.2].

[0052] In this embodiment, the internal cone 16' comprises, upstream of the plurality of radial arms 18 in the direction of the flow F1 (exhaust gas flow) upstream-downstream, a plurality of openings or perforations O distributed circumferentially in the external wall 16b' of the internal cone which is adjacent to the annular duct C (this wall 16b' here has a substantially cylindrical shape), as shown in [Fig. 6A]. In this exemplary embodiment, the number of openings is more or less in correspondence with the number of arms (to within one or two). Depending on the flow, it is however possible to adjust the dimensions and the number of openings.

[0053] It will be noted that these openings are configured to disturb as little as possible the main flow of flow F1 circulating in the aerodynamic vein.

[0054] These openings O are open on the annular duct C so as to allow the air to exit the internal cone 16' (following the arrow F4 in [Fig.6B]) by opening into the annular duct at the internal periphery of the exhaust gas flow FL. These openings are arranged on the internal meridian of the vein of the exhaust gas flow F1 in the primary nozzle, as close as possible to the outlet of the free turbine of the gas generator. This location corresponds to the zone of maximum depression in the primary nozzle, which thus makes it possible to maximize the suction of the cooling air flow by the exhaust gas flow and to optimize the efficiency of the cooling system.

[0055] As a variant, a device forming a “cap” may be added in order to promote the ingestion of the cooling air flow into the exhaust gas flow F1. Such an arrangement could, for example, take the form of a deflection element or deflector starting from an upstream edge of the opening and extending inside the channel C and obliquely downstream with an angle promoting the exit of the flow F4 and its encounter with the annular flow F1.

[0056] The internal cone 16' may also comprise, internally, one or more devices 24 for internally guiding the cooling air coming from the radial arms 18 (fractions of cooling air distributed by these arms) to the plurality of circumferential openings O. This or these internal guiding devices make it possible to reduce the pressure losses on the path of the flow coming from the radial arms and reaching the circumferential openings O. In [Fig.6B], the internal guide device(s) 24 take the form of partitions internal to the cone 16' and which channel each fraction of air flow coming from a radially internal end 18a of the arms 18 (relative to the opposite radially external end and in relation to the distributor 20) along a flow path having a generally triangular shape in axial section (this path is formed first by a vertical or radial portion leaving an arm, then by a horizontal or axial portion directed upstream, followed by an oblique portion directed downstream). The partitions internal to the cone 16' may comprise, on the path of each fraction of air flow, a first partition 24a starting from the end 18a of the arm (on a so-called upstream side 18al of this end) and extending obliquely in a radially internal direction and upstream in order to deflect the fraction of air flow in this direction.The internal partitions of the cone 16' may also comprise a partition 24b starting from the opposite side, called downstream, of the end 18a and which extends obliquely upstream, then axially, opposite the end 18a and the partition 24a so as to constrain obliquely, then axially the fraction of air flow while it is deflected obliquely by the partition 24a. This partition 24b having a concave shape which originates from the zone where the opposite partition 24a ends in order to strongly, but gently, deflect the flow so that it is deflected by more than 90° (turning), thus adopting an oblique orientation directed downstream (orientation according to the arrow F4). It will be noted that in this configuration, the downstream end face 16a' of [Fig.2] is closed so that the cooling air is oriented only upstream of the internal cone.

[0057] In a variant not shown, the interior of the internal cone 16' may be left hollow and not include an internal air flow guidance device.

[0058] Furthermore, as shown in [Fig.6C] which is an interior view of the cone internal 16' without the upstream end face, the internal guide device(s) 24 may comprise a plurality of circumferential fins A. These fins A are arranged on the path of the cooling air fractions guided by the internal guide device(s) 24 described above, immediately upstream of the circumferential openings O ([Fig.6B]), in order to impart a gyrating movement to the air exiting through the circumferential openings O. These fins A are for example mounted between the internal partitions 24a and 24b. The number, shape and orientation of the fins may be adjusted according to the results obtained downstream and according to the pressure of the flow F2.

[0059] The various embodiments and possible forms of embodiment of the invention described above each make it possible, in general, to reduce the thermal aggression of the exhaust gases on the structure of the helicopter, in particular by significantly reducing the temperature of the mixture formed by the flow of exhaust gases and the flow of cooling air at the center of the flow, at the outlet of the helicopter exhaust system.

[0060] Furthermore, the performance of the cooling system is thus improved thanks to the suction phenomenon created in the nozzle.

[0061] Furthermore, this also makes it possible to eliminate a mechanical interface between the engine manufacturer and the aircraft manufacturer. Indeed, in the absence of the invention, it is generally necessary to provide openings on the cowlings, an interface with the outlet of the OCS cooling system and sometimes other elements to evacuate the flow F2 to the outside.

[0062] Figures 7 and 8 illustrate an exemplary embodiment to which the present invention can be applied. [Fig.7] illustrates a rotary wing aircraft 100, more specifically a helicopter with a main rotor 102 and an anti-torque tail rotor 103 coupled to a propulsion unit 104 for their actuation. The propulsion unit 104 illustrated comprises a first engine 105a and a second engine 105b. Power take-off shafts 106a, 106b of these engines 105a, 105b are connected to a main gearbox 107 for actuating the main rotor 102 and the tail rotor 103.

[0063] The propulsion assembly 104 according to a first embodiment is illustrated in more detail in [Fig. 8]. As illustrated in this figure, each engine 105a, 105b may be a heat engine, and more specifically a gas turbine engine comprising a compressor 108a, 108b, a combustion chamber 109a, 109b, a first turbine 110a, 110b operatively connected by a rotating shaft 111a, 111b to the compressor 108a, 108b and a second turbine 112a, 112b, or free turbine, coupled to the power take-off shaft 106a, 106b. The assembly of the compressor 108a, 108b, combustion chamber 109a, 109b, first turbine 110a, 110b and rotating shaft 11a, 111b of each engine 105a, 105b can form a gas generator 120a, 120b. The rotating shaft 111a, 111b of each gas generator 120a, 120b can be operatively coupled to an actuator 113a, 113b and to a fan 114a, 114b.

[0064] Although in [Fig.8], the actuating device 113a, 113b associated with each gas generator 120a, 120b is illustrated arranged on the same rotating shaft as the respective fan 114a, 114b, it is also conceivable that the actuating device 113a, 113b and the fan 114a, 114b associated with each gas generator 120a, 120b are arranged on separate shafts, and operatively coupled to the gas generator 120a, 120b directly or through an intermediate transmission, which may in particular comprise one or more pinions.

[0065] The actuation device 113a, 113b may in particular be an electrical machine, more specifically an electric motor-generator connected to an electrical network of the aircraft 100. Thus, the actuation device 113a, 113b may be used both for starting the corresponding engine 105a, 105b and for generating electricity after this start. In the first case, the electrical machine of the actuation device 113a, 113b may be electrically powered by the electrical network of the aircraft to operate in motor mode. In the second case, the electrical machine of the actuation device 113a, 113b may operate in generator mode to power the electrical network of the aircraft.

[0066] Each fan 114a, 114b may be arranged in an air duct 115a, 115b (the air is for example fresh air coming from outside) passing through an air-oil heat exchanger 116a, 116b crossed by a lubricant circuit 117a, 117b of the corresponding engine 105a, 105b. Each lubricant circuit 117a, 117b may in particular also comprise a pump 118a, 118b and in particular be connected to a transmission casing 119a, 119b of the corresponding engine 105a, 105b. As illustrated, each fan 114a, 114b may be operatively coupled to the rotating shaft 111a, 111b of the corresponding gas generator 120a, 120b, so as to be driven thereby to provide airflow, through the air duct 115a, 115b, through the corresponding heat exchanger 116a, 116b.

[0067] In this exemplary embodiment, the air coming from the fan 114b (flow / debit F2 / F3 or F4 depending on the embodiment of [Fig. 1] or 2) is used to cool the exhaust gas flow F1 delivered by the free turbine 112b.

[0068] In this exemplary embodiment, in the propulsion assembly 104 illustrated in [Fig. 8], the second engine 105b may be in standby mode during the cruising speed of the aircraft 100, while the first engine 105a provides a majority, or even all, of the power used to drive the main rotor 102 and the tail rotor 103 through the main gearbox 107. In this exemplary embodiment, the propulsion assembly 104 may also include an air interconnecting duct 121, connecting the air duct 115a of the first engine 105a, downstream of its heat exchanger 116a, to the second engine 105b. More specifically, the air interconnection duct 121 may open into the air duct 115b of the second engine 105b, upstream of its heat exchanger 116b. The propulsion assembly 104 may further comprise a valve 122, which may be arranged on the air interconnection duct 121 and connected to an actuator 123. The actuator 123 may be controlled so as to open the valve 122 when the second engine 105b is in standby mode.

[0069] Although the present invention has been described with reference to specific exemplary embodiments, it is obvious that various modifications and changes 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 discussed may be combined in additional embodiments. Therefore, the description and drawings should be considered in an illustrative rather than restrictive sense.

Claims

Claims

1. A turbomachine subassembly (10; 10') for a helicopter, comprising a primary nozzle (12; 12') for discharging an exhaust gas flow (Fl), the primary nozzle comprising an outer casing (14) and an inner cone (16; 16') which are arranged concentrically with respect to each other around a longitudinal axis (X) of the turbomachine and connected to each other by a plurality of radial arms (18), the outer casing (14) surrounding the inner cone (16; 16') and defining with the latter an annular duct (C) for the flow of the exhaust gas flow (Fl), characterized in that the turbomachine subassembly (10;10') comprises a distributor (20) at least partially surrounding the outer casing (14), the distributor (20) being configured to, on the one hand, receive, at the inlet (20a) of the distributor, a flow of cooling air existing (F2) in the turbomachine and having a temperature lower than the temperature of the exhaust gases and, on the other hand, distribute, at the outlet of the distributor, this flow of cooling air inside at least some radial arms (18) of the plurality of radial arms to the inside of the inner cone (16; 16'), the inner cone comprising an outer wall (16a; 16b') in which one or more openings (O) are arranged to allow the cooling air (F3; F4) of the inner cone to exit said inner cone and to be injected at an inner periphery of the flow of exhaust gas.;

2. Turbomachine subassembly according to claim 1, in which the external wall of the internal cone comprises a downstream end face (16a), following the direction of flow of the exhaust gas flow (F1) in the helicopter turbomachine, which is open to the outside so as to allow the cooling air (F3) of the internal cone to exit the latter in a substantially axial direction and directed downstream.

3. A turbomachine subassembly according to claim 2, wherein the inner cone (16) is configured to guide the cooling air (F3) exiting the inner cone.

4. A turbomachine subassembly according to claim 3, wherein the inner cone (16) comprises a worm-like device (26) having a helical groove which defines a helical path for guiding cooling air from the inner cone to the outside.

5. A turbomachine subassembly according to claim 3, wherein the inner cone (16) comprises ducts (28) which are each configured internally to convey a portion of the cooling air of the inner cone, the ducts (28) being wound relative to each other so as to form a propeller.

6. A turbomachine subassembly according to claim 1, wherein the inner cone (16') comprises, upstream of the plurality of radial arms (18), a plurality of openings (0) distributed circumferentially in the outer wall (16b') of the inner cone which is adjacent to the annular duct (C), thus allowing the fractions of cooling air flow (F4) distributed by said at least some radial arms (18) of the plurality of radial arms to exit the inner cone by opening into the annular duct at an inner periphery of the exhaust gas flow (F1).

7. Turbomachine subassembly according to the preceding claim, in which the internal cone (16') comprises at least one internal guide device (24, 24a, 24b) for the fractions of cooling air flow distributed by said at least some radial arms of the plurality of radial arms to the plurality of circumferential openings (0).

8. Turbomachine subassembly according to the preceding claim, wherein said at least one internal guide device (24) further comprises a plurality of circumferential fins (A) which are arranged on the path of the cooling air flow fractions in order to impart a gyrating movement to the air (F4) exiting through the circumferential openings.

9. Turbomachine subassembly according to one of the preceding claims, in which the cooling air flow (F2) comes from a system for cooling a lubricant of the turbomachine.

10. Turbomachine subassembly according to one of claims 1 to 8, in which the cooling air flow (F2) comes from a cooling system of a generator-starter of a turbomachine.

11. Helicopter turbomachine (104) comprising a subassembly according to one of the preceding claims.

12. Helicopter (100) comprising a helicopter turbomachine according to the preceding claim.

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