Helicopter turbine subassembly and helicopter turbine comprising such subassembly

Abstract

Turbine engine subassembly (10) for helicopters, comprising a main outlet nozzle (12) for discharging an exhaust gas flow (F1), the main outlet nozzle (12) comprising an outer casing (14) and an inner cone (16) coaxially arranged about a longitudinal axis (X) and connected to each other by radial arms (18), the casing (14) surrounding the cone (16) and defining therebetween an annular passage (C) for the flow of the gas flow (F1). The subassembly comprises a distributor (20) surrounding the casing (14), the distributor (20) being configured so as to receive at its inlet (20a) an airflow (F2) of cooling air at a temperature lower than the temperature of the airflow (F1) and to distribute at its outlet the airflow (F2) inside the radial arm (18). The inner cone has an outer wall (16a, 16b ') in which one or more openings (O) are arranged to allow air injection into the airflow (F1).

Description

Technical Field

[0001] This invention relates to helicopter turbines. Background Technology

[0002] In the case of a twin-engine helicopter, the aircraft turbine, and more specifically, the helicopter turbine, includes a propulsion assembly that typically comprises a first engine configured to actuate the main rotor and a second engine configured to actuate the tail rotor. The two engines are typically connected to a main gearbox, which distributes power between the main rotor and the tail rotor. Each engine may be a heat engine, more specifically a gas turbine, including a compressor, a combustion chamber, a first turbine operatively connected to the compressor via a rotating shaft, and a second turbine or free turbine coupled to a power output shaft for transmitting motion to the respective rotor. The assembly formed by the compressor, combustion chamber, first turbine, and the rotating shaft of the engine constitutes a gas generator.

[0003] The exhaust gas from the aforementioned free turbine operates at extremely high temperatures. However, under certain flight conditions, such as during hovering, the exhaust gas may be redirected to the helicopter's tail boom by the main rotor's downwash. This can be detrimental to the tail boom's structure, as composite materials are increasingly used in helicopter construction, and these materials are generally sensitive to thermal stress. To withstand these high temperatures, it is therefore necessary to protect at least some areas of the tail boom's external fairing with thermal shielding; however, this significantly increases the helicopter's onboard mass and manufacturing costs.

[0004] Therefore, being able to overcome at least some of the above-mentioned shortcomings would be particularly useful. Summary of the Invention

[0005] This disclosure is the result of technical research aimed at reducing exhaust temperature to limit or even eliminate the need for thermal protection on the external fairing of a helicopter tail boom. For this purpose, it is specifically planned to use air already present in the turbine, at a temperature lower than the exhaust temperature, to cool the exhaust.

[0006] The first aspect of this disclosure relates to a sub-assembly of an aircraft turbine, particularly a helicopter turbine, comprising a main outlet nozzle for exhaust gas flow. The main nozzle includes an outer casing and an inner cone, which are coaxially arranged relative to each other about the longitudinal axis of the turbine and connected to each other by a plurality of radial arms. The outer casing surrounds and defines an annular channel for exhaust gas flow around the inner cone. The turbine sub-assembly is characterized in that it includes a distributor at least partially surrounding the outer casing, the distributor being configured to receive, at its inlet, a cooling gas flow present in the turbine at a temperature lower than that of the exhaust gas, and at its outlet, a plurality of radial arms... Cooling airflow from the interior of at least some radial arms (at least some arms thus each ensures a small portion of cooling airflow is distributed) is distributed to the interior of an inner cone. The inner cone (in which a small portion of cooling airflow from the radial arms is distributed) includes an outer wall with one or more openings arranged in the outer wall to allow cooling air from the inner cone to exit the inner cone and be injected at the inner circumferential surface of the exhaust airflow. The outer wall of the inner cone includes a downstream end face along the direction of exhaust airflow flow in the helicopter turbine, the downstream end face opening to the outside to allow cooling air from the inner cone to exit the inner cone in a generally axially downstream direction. The inner cone is internally configured to guide the cooling air exiting the inner cone and impart a rotation angle to it.

[0007] Utilizing the existing airflow from the turbine at the level of the inner cone of the main nozzle (exhaust nozzle) allows for cooling of the exhaust airflow from the inside (on its inner circumferential surface, i.e., at the flow center), specifically in the low-pressure portion of the airflow. This proves particularly effective for cooling operations. Consequently, the external fairing of the tail boom of a helicopter with an integrated turbine experiences less of the high temperatures that previously required thermal protection on the external fairing. Therefore, the amount of thermal protection can be limited, or even (in some cases) eliminated, allowing for a reduction in the helicopter's onboard mass. Furthermore, using the existing airflow from the turbine eliminates the need to draw power from (one or more) engines, thus not affecting performance. The internal structure of the inner cone allows for imparting a rotational component (rotational effect) to the cooling airflow, which is particularly effective for cooling the exhaust airflow from the inside.

[0008] Based on other possible characteristics: - The rotation angle is greater than 0° and less than 40°. The rotation angle of the airflow is defined as the arctangent of the ratio of the tangential velocity of the airflow to the axial velocity of the airflow. - The inner cone includes a worm gear-type device that includes a helical spline that defines a helical path for guiding cooling air from the inner cone to the outside; - The inner cone includes channels, each internally configured to deliver a portion of the cooling air from the inner cone, the channels being intertwined relative to each other to form a spiral; - The inner cone includes multiple openings upstream of the multiple radial arms, the multiple openings being circumferentially distributed in the outer wall of the adjacent annular channel of the inner cone, thereby allowing a small portion of the cooling airflow dispersed by at least some of the radial arms to exit the inner cone by passing through the annular channel at the inner circumferential surface of the exhaust airflow (this feature may be combined with at least one of the aforementioned features). - The inner cone includes at least one means for internally guiding a small portion of the cooling airflow, which is dispersed by at least some of the radial arms, to a plurality of circumferential openings (this feature depends on the preceding feature). - The at least one internal guide device further includes a plurality of circumferential fins arranged in the path of a small portion of the cooling airflow to impart rotational motion to the air discharged through the circumferential opening (this feature depends on the former feature). - Airflow from a system used to cool lubricant in helicopter turbines can be used as cooling airflow in aircraft turbine subassemblies as briefly described above; The airflow from the cooling system of the generator-starter of the aircraft turbine can be used as the cooling airflow in the aircraft turbine sub-assemblies as briefly described above.

[0009] A second aspect of this disclosure relates to a helicopter turbine, including the aircraft turbine subassembly as briefly described above.

[0010] A third aspect of this disclosure relates to a helicopter, including a helicopter turbine as briefly described above. Attached Figure Description

[0011] The invention will be better understood and its advantages will become more apparent from the following detailed description of embodiments given as non-limiting examples. This description refers to illustrative drawings, which are primarily used to illustrate the principles of this disclosure.

[0012] In these accompanying figures, identical or equivalent elements (or parts thereof) are identified using the same reference numerals from one figure to the next. In these accompanying figures: [ Figure 1 ] Figure 1 A first embodiment of the present invention, comprising a turbine subassembly, is schematically illustrated.

[0013] [ Figure 2 ] Figure 2 A second embodiment of the present invention, of a turbine subassembly, is schematically illustrated.

[0014] [ Figure 3 ] Figure 3 The illustration schematically shows one possible form of the first embodiment of the present invention.

[0015] [ Figure 4A ] Figure 4A This is an axial sectional view of a turbine subassembly integrating a worm gear device according to the first embodiment of the present invention.

[0016] [ Figure 4B ] Figure 4B yes Figure 4A A three-dimensional diagram of a worm gear mechanism.

[0017] [ Figure 5 ] Figure 5 This is a perspective view of a turbine sub-assembly integrating a helical winding channel according to the first embodiment of the present invention.

[0018] [ Figure 6A ] Figure 6A This is a perspective view from upstream of one possible form of a turbine subassembly according to a second embodiment of the present invention.

[0019] [ Figure 6B ] Figure 6B yes Figure 6A A partial axial sectional view of the turbine subassembly.

[0020] [ Figure 6C ] Figure 6C yes Figure 6A and Figure 6B A three-dimensional view of the inner cone of the turbine subassembly from upstream.

[0021] [ Figure 7 ] Figure 7 An example of an aircraft with a propulsion system comprising two engines is shown schematically.

[0022] [ Figure 8 ] Figure 8 More specifically shown Figure 7 An embodiment of the propulsion component. Detailed Implementation

[0023] To make this disclosure more specific, embodiments will now be described in detail with reference to the accompanying drawings. However, it should be noted that the present invention is not limited to these embodiments.

[0024] Figure 1 and Figure 2 Two main embodiments of the present invention are illustrated schematically.

[0025] Typically, a sub-assembly 10 of an aircraft turbine, such as a helicopter, includes a main nozzle 12, partially shown, which includes an outer casing 14 centered on the turbine's longitudinal axis X. Along this longitudinal axis, an exhaust gas flow (originating from hot gases from internal combustion processes within the turbine, particularly from a gas turbine) flows from upstream to downstream as indicated by arrow F1. The outer casing 14 has an axial shape in its upstream portion and then an expanding shape in its downstream portion for discharging the exhaust gas flow F1 from the turbine.

[0026] The main nozzle 12 also includes an inner cone 16 disposed within the outer casing 14 and coaxially arranged relative to the outer casing 14 about the longitudinal axis X of the turbine. The outer casing 14 surrounds the inner cone 16, thereby defining an axial annular passage C between them for the flow of exhaust gas F1. The outer casing 14 and the inner cone 16 are mechanically connected to each other by a plurality of arms 18 extending radially relative to the longitudinal axis X within the axial annular passage C. These radial arms 18 are circumferentially distributed within the annular passage and are arranged, for example, regularly relative to each other. Typically, the plurality of these radial arms 18 ensure the mechanical strength and reinforcing function of the assembly (structural arms). At least a portion of the plurality of these radial arms 18 is hollow to allow airflow to be delivered internally.

[0027] The turbine subassembly 10 also includes a distributor 20 (a means for distributing airflow) that at least partially surrounds the outer casing 14 of the nozzle 12. For example, in one possible embodiment, the distributor 20 takes the form of a hollow annular casing that surrounds and is fixed to the outer casing 14. 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 an annular portion.

[0028] Distributor 20 receives existing airflow (air volume) from the turbine's internal source S at its inlet 20a, which does not require drawing additional power from the turbine(s) and thus does not affect its performance.

[0029] In one possible exemplary embodiment, the existing airflow originates from a system used to cool lubricant in a helicopter turbine. The lubricant, for example, is used for lubrication of one or more engines in an aircraft turbine and can be, for example, oil. For instance, the aforementioned cooling system may include an air-lubricant heat exchanger, a hot lubricant circuit from the turbine engine, and an airflow, which may be external air, passing through the air-lubricant heat exchanger. This external air is heated by heat exchange with the lubricant to cool it. The airflow leaving the exchanger is hotter than when it enters; however, its temperature is generally still lower than that of the exhaust airflow F1, which would allow the airflow (hereinafter referred to as F2) to be used to cool airflow F1. For example, the airflow is heated to a temperature of 120°C by oil, while in the exhaust region, the exhaust airflow can exceed 600°C. Therefore, Figure 1 Source S generates an airflow F2 called cooling airflow, which is guided from source S to inlet 20a of distributor 20 via one or more channels 22 (arranged from upstream to downstream) to supply cooling air to distributor 20.

[0030] As a variation, it will be noted that the source of cooling air for the cooling airflow F1 can be a cooling system or circuit dedicated to cooling the generator-starter assembly, and typically includes a supply channel and outlet for discharging heated air. Therefore, it is conceivable to use the heated air at the generator-starter assembly outlet, directing it to the aforementioned distributor 20. In known configurations of the generator-starter assembly cooling system, either a duct is provided at the outlet to direct the heated air to the outside of the aircraft, or the hot air is directly discharged into the engine nacelle.

[0031] Alternatively, other sources of cooling air can be used, such as airflow from the helicopter's main gearbox cooling system.

[0032] Distributor 20 is connected to the outer casing 14 and the radial arms 18, and more specifically, to one end of each hollow radial arm 18 leading to the distributor, thereby allowing fluid communication between the distributor outlet and the interior of the associated radial arm. Thus, the cooling airflow F2 entering distributor 20 is distributed or dispersed within each hollow radial arm 18, so that each hollow radial arm 18 ensures that a small portion of this cooling airflow is directed to the inner cone 16 connected to each arm. Each hollow radial arm 18 includes an end opposite to the end communicating with the distributor, which leads to the inner cone 16, in order to establish fluid communication between these arms and the interior of the cone. It will also be noted that this opposite end of the arm can extend into the interior of the cone and be locally shaped to guide the small portion of the airflow delivered by each arm in a generally axially downstream direction along the cone, such as... Figure 1 The dashed lines in the diagram illustrate this. Therefore, a small portion of the cooling airflow dispersed within the inner cone by the radial arms can mix more or less inside the cone 16 and flow axially downstream, as shown... Figure 1As indicated by arrow F3. However, in one possible implementation, a small portion of the cooling airflow, thus dispersed within the inner cone 16, can be redirected axially downstream within the inner cone 16, and thus not actually mixed before reaching the exhaust airflow that draws it in outside the cone. It will be noted that in Figure 1 In one embodiment, the inner cone 16 includes an outer wall forming an end face 16a that extends laterally or radially relative to the longitudinal axis X. This end face 16a opens outward to allow cooling air (airflow F3) to exit the inner cone 16 and be injected into the inner circumferential surface of the exhaust airflow F1 (this injection area is a low-pressure region in the exhaust flow path, allowing the intake of cooling air exiting the cone). This airflow in contact with the exhaust airflow allows for cooling of the exhaust airflow, thereby limiting the adverse effects of excessively high temperatures on the external fairing of the helicopter tail boom on which the turbine is mounted.

[0033] exist Figure 1 In one embodiment, cone 16 may be internally configured to guide cooling air F3 exiting the cone, allowing for higher cooling efficiency of the airflow to the exhaust gas stream. Guiding the flow within the cone allows for reduced pressure drop, thereby distributing more cooling airflow. This also allows for rotational attenuation of the airflow F3 within the cone to promote mixing with the hot airflow F1. Different arrangements or devices (schematically indicated by dashed lines in reference numeral 24) within the inner cone 16 are conceivable to ensure air guidance in a generally axial direction and oriented downstream of the subassembly. Possible examples of such internal arrangements will be described later. This arrangement allows for reduced pressure drop and rotational at the cone outlet to promote mixing and maximize the resulting temperature reduction. Typically, a high-power operating point is sought, and for this purpose, rotation angles are conceivable to be between 0 and 40 degrees. The rotation angle of the airflow is typically defined as the arctangent of the ratio of the airflow's tangential velocity to its axial velocity. In an embodiment not shown, the interior of the inner cone may remain hollow and without such an arrangement or device.

[0034] It will be noted that only turbine sub-assemblies to which the present invention is applicable are described herein, and other turbine components not required for understanding the present invention are not described.

[0035] Figure 1 The foregoing description also applies to Figure 2 Regarding the guidance of the cooling airflow F2 to the distributor 20 and its distribution in the hollow radial arms 18, the small portions of airflow distributed in these arms are then delivered to the inner cone. Reference numerals for the same elements are retained from one figure to another.

[0036] However, Figure 2 The description of turbine sub-assembly 10' and Figure 1The difference in the description is that a small portion of the cooling airflow is introduced into the interior of the inner cone 16', and the construction of the inner cone 16' is different from that of the inner cone 16.

[0037] The inner cone 16' here includes one or more means for internally guiding cooling air, schematically indicated by dashed lines by reference numeral 24'. This or these means ensure the redirection of a small portion of air originating from the radial arm 18, such that the radial orientation of this small portion of air during its transport through arm 18 is altered by redirection upstream and then downstream, thereby enabling it to be injected into the exhaust gas flow F1 in the same upstream-downstream flow direction as the exhaust gas flow F1. The change in the orientation of the cooling air thus corresponds to a 360° reversal. The injection of cooling air into the exhaust gas flow occurs upstream of the radial arm.

[0038] Although this is in Figure 2 Not shown in the figure (other figures described later will illustrate this aspect), but the inner cone 16' includes a plurality of openings upstream of the plurality of radial arms, these openings being circumferentially distributed in the outer wall 16b' of the adjacent annular channel C of the inner cone. These openings allow passage through... Figure 2 Arrow F4 indicates that multiple small portions of cooling air are introduced into the annular channel C, thereby injecting cooling air into the inner circumferential surface of the exhaust airflow F1. These small portions of cooling air are directed into the exhaust airflow F1 through one or more guide devices inside the inner cone 16'.

[0039] Therefore, the cooling air is guided (and redirected) inside the inner cone 16' by one or more internal guides to a plurality of circumferential openings and exits through these circumferential openings in a desired orientation so as to enter the annular channel at an angle of less than 90° relative to the axial direction of the airflow F1.

[0040] The airflow F4 that comes into contact with the exhaust airflow F1 allows for cooling of the exhaust airflow F1, thereby limiting the adverse effects of excessively high temperatures on the external fairing of the helicopter tail boom where the turbine is mounted.

[0041] Figure 3 It shows that according to Figure 1 One possible implementation of the turbine subassembly in the first embodiment shows a channel 22 for guiding a cooling airflow F2 (with a given flow rate), which, for example, originates from a system for cooling lubricant in a helicopter turbine, as known for example by the abbreviation OCS (Lubricant Cooling System). This channel will... Figure 1The source S (not shown here) is connected to the inlet 20a of the distributor 20, which is flared so as to smoothly (without introducing excessive pressure drop) divide the inlet airflow F2 inside the distributor into two parts (along a larger angular sector, for example, approximately 180°), each of which is spatially distributed over the entire annular crown of the flow distributor 20 using an annular semi-crown.

[0042] The inner cone 16 of the main nozzle 16 is connected to the outer casing 14 by means of radial arms 18, of which there are, for example, five (although a higher or lower number of arms may be conceived).

[0043] like Figure 4A As shown, this figure is Figure 3 An axial sectional view of sub-assembly 10 shows that the inner cone 16 includes a worm gear type device 26 (in... Figure 4B (As shown in a perspective view), the worm gear type device 26 is in fluid communication with a small portion of cooling air distributed by the radial arm 18, so that each small portion of air can be conveyed downstream through the device 26.

[0044] The downstream end face 16a of the outer wall of the inner cone 16 (oriented laterally / radially relative to the longitudinal axis X) opens to the outside, so that the cooling air delivered downstream by the worm gear device leaves the inner cone 16 in a generally axial and downstream direction via face 16a, thereby reaching the inner circumferential surface of the exhaust airflow F1.

[0045] More specifically, the worm gear type device 26 includes a helical spline 26a that defines a helical path for guiding cooling air from the inner cone 16 to its exterior. The helical spline 26a extends, for example, from the base 26b located upstream of the radial arm 18. Figure 4A It extends downstream in the direction of the open end face 16a.

[0046] In practice, the helical spline 26a comprises a number of spiral coils, which can be equal to the number of radial arms through which a small portion of the cooling air passes. Thus, each small portion of air originating from an arm is directed onto one of the spiral coils. This configuration allows for a reduction in pressure drop by guiding the airflow leaving each arm to the outside of the cone, thereby avoiding harmful turbulence associated with mixing with the small portion of air inside the cone. In this configuration, the small portions of air originating from the arms remain independent of each other before leaving the worm gear assembly. Furthermore, by appropriately adapting the inclination of the spiral coils, the worm gear assembly also allows for control of the rotation angle imparted to the airflow at the cone outlet. The rotation angle values ​​given above also apply here.

[0047] like Figure 5 As shown, the inner cone 16 of the nozzle includes another type of internal airflow guidance, which is formed by channels 28 inside the cone, each channel 28 being configured internally to deliver a portion of the cooling air to the inner cone. Figure 5 As shown, the hollow channels 28 are intertwined relative to each other to form a general spiral. Individually, each channel is shaped in a spiral manner, and all the corresponding spirals are as follows: Figure 5 The structures shown are nested within each other.

[0048] In practice, to minimize pressure drop, the number of channels 28 is equal to the number of radial arms 18, each delivering a small portion of the cooling air. The channels 28 can have various shapes and are each configured to allow control of the rotation angle of the flow at the cone outlet. The rotation angle values ​​given above also apply here.

[0049] Each channel 28 is connected to the radial arm 18. Figure 5 Different cross sections 28a of the different channels 28 are shown in a given transverse plane (downstream of the connection between the arm and the channel), these sections being arranged circumferentially relative to each other. If only some radial arms are capable of delivering cooling airflow, then only these arms are connected to the channel 28. However, it will be noted that in one embodiment not shown, the number of arms delivering cooling airflow differs from the number of channels. For example, the same channel may receive air from multiple arms.

[0050] Figures 6A-6C It shows that according to Figure 2 One possible implementation of the turbine sub-assembly 10' of the second embodiment.

[0051] In this embodiment, the inner cone 16' includes multiple openings or holes O upstream of the multiple radial arms 18 in the upstream and downstream direction of the airflow F1 (exhaust airflow). These openings or holes O are circumferentially distributed in the outer wall 16b' of the adjacent annular channel C of the inner cone (which has a generally cylindrical shape here), such as... Figure 6A As shown. In this exemplary embodiment, the number of openings corresponds more or less to the number of arms (difference of one or two). However, depending on the airflow, it is conceivable that the size and number of openings can be adjusted.

[0052] It should be noted that these openings are designed to minimize disturbance to the main airflow F1 circulating in the aerodynamic flow path.

[0053] These openings O lead to the annular channel C, allowing air to pass through the annular channel at the inner circumferential surface of the exhaust airflow F1 (along...). Figure 6B Arrow F4 in the diagram exits the inner cone 16'. These openings are positioned on the inner meridian of the exhaust gas flow path F1 in the main nozzle, as close as possible to the outlet of the free turbine of the gas generator. This location corresponds to the region of maximum low pressure in the main nozzle, thereby allowing the exhaust gas flow to maximize the intake of cooling airflow and optimize the efficiency of the cooling system.

[0054] As a variation, a device forming a "cap" can be added to facilitate the intake of cooling airflow into the exhaust airflow F1. This arrangement can take the form, for example, a deflector or baffle that begins at the upstream edge of the opening, extends downstream at an angle that facilitates the exhaust of airflow F4 and its encounter with the annular airflow F1.

[0055] The inner cone 16' may also internally include one or more means 24 for internally guiding cooling air (a small portion of the cooling air dispersed by these arms) originating from the radial arms 18 to the plurality of circumferential openings O. This or these internal guiding means allow for reduction of pressure drop along the flow path originating from the radial arms and reaching the circumferential openings O. Figure 6B In this configuration, one or more internal guide devices 24 take the form of baffles inside the cone 16' that guide each small portion of airflow originating from the radially inner end 18a of the arm 18 (relative to the opposite radially outer end and associated with the distributor 20) along a flow path that is generally triangular in axial section (this path is first formed by a vertical or radial portion exiting the arm, then by an upstream-guided horizontal or axial portion, and subsequently by a downstream-guided inclined portion). The internal baffles of the cone 16' may include a first baffle 24a in the path of each small portion of airflow, which begins at the end 18a of the arm (on a side referred to as the upstream side 18a1 at that end) and extends inclined in a radially inward and upstream direction to deflect the small portion of airflow in that direction. The baffle inside cone 16' may also include baffle 24b, which begins from the opposite side of end 18a, referred to as the downstream side, and faces end 18a and baffle 24a, extending obliquely upstream and then axially to obliquely constrain, and then axially constrain, a small portion of the airflow that is obliquely deflected by baffle 24a. The baffle 24b has a concave shape originating from the region where the facing baffle 24a ends, so as to forcefully but gently deflect the flow, causing it to deflect more than 90° (rotation), thereby adopting a downstream-guided oblique orientation (along the orientation of arrow F4). It will be noted that in this configuration, Figure 2 The downstream end face 16a' is closed, so that the cooling air is directed only upstream of the inner cone.

[0056] In a variant not shown, the interior of the inner cone 16' may remain hollow and does not include means for guiding airflow internally.

[0057] In addition, such as Figure 6C As shown, Figure 6CThis is an internal view of the inner cone 16' without an upstream end face. One or more internal guide devices 24 may include a plurality of circumferential fins A. These fins A are positioned on the path of a small portion of the cooling air guided by the aforementioned internal guide devices 24, immediately upstream of the circumferential opening O. Figure 6B This imparts rotational motion to the air exhausted through the circumferential opening O. These fins A are, for example, installed between internal partitions 24a and 24b. The number, shape, and orientation of the fins can be adjusted based on results obtained downstream and according to the pressure of the airflow F2.

[0058] The various possible embodiments and implementations of the invention described above generally allow for a reduction in the thermal stress of exhaust gas on the helicopter structure, particularly by significantly reducing the temperature of the mixture formed by the exhaust gas flow and the cooling gas flow at the flow center at the helicopter exhaust system outlet.

[0059] Furthermore, the performance of the cooling system is improved due to the suction phenomenon generated in the nozzle.

[0060] Furthermore, this also allows for the elimination of mechanical interfaces between aircraft manufacturers and aircraft manufacturers. In fact, without this invention, it would typically be necessary to provide openings on the fairing, interfaces with the OCS cooling system outlet, and sometimes other components for discharging the F2 airflow to the outside.

[0061] It will be noted that Figure 1 and Figure 2 The patterns (and, for example, more detailed illustrations of each of these patterns) can be combined to design an aircraft turbine subassembly in which cooling airflow from the radial arm and entering the inner cone is distributed / distributed in a downstream-guided airflow section (such as...). Figure 1 And for example Figure 4A -B and Figure 5 As shown, and according to the corresponding description above) and the upstream guided airflow section (such as... Figure 2 And for example Figure 6A -C as shown, and according to the corresponding description above) between.

[0062] Figure 7 and Figure 8 An exemplary embodiment to which the present invention can be applied is shown. Figure 7 A rotorcraft 100 is shown, more specifically a helicopter having a main rotor 102 and a counter-torque tail rotor 103, which are coupled to a propulsion assembly 104 for its actuation. The illustrated propulsion assembly 104 includes a first engine 105a and a second engine 105b. The power output shafts 106a and 106b of these engines 105a and 105b are connected to a main gearbox 107 to actuate the main rotor 102 and the tail rotor 103.

[0063] According to the first embodiment, the propulsion component 104 is in Figure 8 The diagram is shown in more detail below. As shown in the figure, each engine 105a, 105b can be a heat engine, more specifically a gas turbine engine, including compressors 108a, 108b, combustion chambers 109a, 109b, a first turbine 110a, 110b operatively connected to the compressors 108a, 108b via rotating shafts 111a, 111b, and a second turbine 112a, 112b, or a free turbine, coupled to power output shafts 106a, 106b. The assembly of the compressors 108a, 108b, combustion chambers 109a, 109b, first turbines 110a, 110b, and rotating shafts 111a, 111b of each engine 105a, 105b can form gas generators 120a, 120b. The rotating shafts 111a and 111b of each gas generator 120a and 120b are operatively connected to actuators 113a and 113b and fans 114a and 114b.

[0064] Despite Figure 8 In the diagram, the actuators 113a, 113b associated with each gas generator 120a, 120b are shown as being mounted on the same rotating shaft as the corresponding fans 114a, 114b. However, it is also conceivable that the actuators 113a, 113b associated with each gas generator 120a, 120b and the fans 114a, 114b are mounted on separate shafts and operatively coupled to the gas generators 120a, 120b directly or via an intermediate drive, which may specifically include one or more pinions.

[0065] Actuators 113a and 113b can be, in particular, electric motors, and more specifically, electric motor generators connected to the electrical grid of aircraft 100. Therefore, actuators 113a and 113b can be used both to start the corresponding engines 105a and 105b, and after starting, to generate electricity. In the first case, the motors of actuators 113a and 113b can be powered by the aircraft's electrical grid and operate in motor mode. In the second case, the motors of actuators 113a and 113b can operate in generator mode to supply power to the aircraft's electrical grid.

[0066] Each fan 114a, 114b can be disposed in an air passage 115a, 115b (the air being, for example, fresh air from the outside), which passes through an air-oil heat exchanger 116a, 116b through which the lubricant circuits 117a, 117b of the respective engines 105a, 105b pass. Each lubricant circuit 117a, 117b may also specifically include a pump 118a, 118b, and is specifically connected to the transmission housing 119a, 119b of the respective engines 105a, 105b. As shown, each fan 114a, 114b can be operatively coupled to the rotating shaft 111a, 111b of the respective gas generators 120a, 120b so as to be driven by the rotating shaft 111a, 111b, thereby ensuring airflow through the air passage 115a, 115b through the respective heat exchangers 116a, 116b.

[0067] In this exemplary embodiment, the air (airflow / flow rate F2 / F3 or F4, depending on the fan 114b) is supplied by the fan. Figure 1 or Figure 2 (An embodiment) is used to cool the exhaust gas flow F1 delivered by the free turbine 112b.

[0068] In this exemplary embodiment, in Figure 8 In the propulsion assembly 104 shown, during the cruise speed of the aircraft 100, the second engine 105b can be in standby mode, while the first engine 105a provides most or all of the power used to drive the main rotor 102 and tail rotor 103 via the main gearbox 107. In this exemplary embodiment, the propulsion assembly 104 may also include an air interconnection channel 121 that connects an air passage 115a downstream of the heat exchanger 116a of the first engine 105a to the second engine 105b. More specifically, the air interconnection channel 121 may open into an air passage 115b upstream of the heat exchanger 116b of the second engine 105b. The propulsion assembly 104 may also include a valve 122 that may be disposed on the air interconnection channel 121 and connected to an actuator 123. When the second engine 105b is in standby mode, the actuator 123 can be controlled to open the valve 122.

[0069] Although the invention has been described with reference to specific exemplary embodiments, it will be apparent that various modifications and changes can be made to these examples without departing from the overall scope of the invention as defined by the claims. Furthermore, various features of the discussed embodiments may be combined in additional embodiments. Therefore, the specification and drawings should be considered in an illustrative rather than restrictive sense.

Claims

1. A sub-assembly (10, 10') of an aircraft turbine, particularly a helicopter turbine, comprising a main outlet nozzle (12, 12') for an exhaust gas flow (F1), the main nozzle comprising an outer casing (14) and an inner cone (16, 16'), the outer casing (14) and the inner cone (16, 16') being coaxially arranged relative to each other about a longitudinal axis (X) of the turbine 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 inner cone (16, 16') an annular passage (C) for the flow of the exhaust gas flow (F1), characterized in that, The turbine subassemblies (10, 10') include a distributor (20) that at least partially surrounds the outer casing (14). The distributor (20) is configured to, on the one hand, receive a cooling airflow (F2) present in the turbine at a temperature lower than that of the exhaust gas at the inlet (20a) of the distributor, and on the other hand, distribute the cooling airflow inside at least some of the radial arms (18) of the plurality of radial arms into the interior of the inner cone (16, 16') at the outlet of the distributor. The inner cone includes an outer wall (16a, 16b') in which the cooling airflow is distributed. One or more openings (O) are arranged in the inner cone to allow the cooling air (F3, F4) of the inner cone to leave the inner cone and be injected at the inner circumferential surface of the exhaust airflow. The outer wall of the inner cone includes a downstream end face (16a) along the flow direction of the exhaust airflow (F1) in the helicopter turbine. The downstream end face (16a) opens to the outside to allow the cooling air (F3) of the inner cone to leave the inner cone in a generally axially downstream direction. The inner cone (16) is internally configured to guide the cooling air (F3) leaving the inner cone and give it a rotation angle.

2. The aircraft turbine sub-assembly according to claim 1, characterized in that, The rotation angle is greater than 0° and less than 40°. The rotation angle of the airflow is defined as the arctangent of the ratio of the tangential velocity of the airflow to the axial velocity of the airflow.

3. The aircraft turbine subassembly according to claim 1 or 2, characterized in that, The inner cone (16) includes a worm gear-type device (26) that includes a helical spline that defines a helical path for guiding the cooling air from the inner cone to the outside.

4. The aircraft turbine subassembly according to claim 1 or 2, characterized in that, The inner cone (16) includes channels (28), each internally configured to deliver a portion of the cooling air from the inner cone, the channels (28) being wound around each other to form a spiral.

5. The aircraft turbine subassembly according to any one of the preceding claims, characterized in that, The inner cone (16') includes a plurality of openings (O) upstream of the plurality of radial arms (18), the plurality of openings (O) being circumferentially distributed in the outer wall (16b') of the inner cone adjacent to the annular channel (C), thereby allowing a small portion of the cooling airflow (F4) dispersed by at least some of the radial arms (18) to exit the inner cone by entering the annular channel at the inner circumferential surface of the exhaust airflow (F1).

6. The aircraft turbine subassembly according to the preceding claims, characterized in that, The inner cone (16') includes at least one device (24, 24a, 24b) for internally guiding a small portion of the cooling airflow dispersed by at least some of the radial arms to the plurality of circumferential openings (O).

7. The aircraft turbine subassembly according to the preceding claims, characterized in that, The at least one internal guide device (24) further includes a plurality of circumferential fins (A) arranged in the path of the small portion of the cooling airflow to impart rotational motion to the air (F4) discharged through the circumferential opening.

8. Use of airflow from a system for cooling lubricant in an aircraft turbine subassembly as a cooling airflow (F2) according to any of the preceding claims.

9. Use of airflow from the cooling system of the generator-starter of the aircraft turbine as cooling airflow (F2) in an aircraft turbine subassembly according to any one of claims 1 to 7.

10. A helicopter turbine (104) comprising an aircraft turbine subassembly according to any one of claims 1 to 7.

11. A helicopter (100) comprising a helicopter turbine according to the preceding claims.

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