Assembly for an aircraft propulsion unit, comprising an item of equipment intended to be passed through by a controlled flow of fluid

Abstract

The invention relates to an assembly (15) for an aircraft propulsion unit, comprising - a radial delimiting structure (31, 34, 36) for the circulation of a gas flow (12b, 10b); - an item of equipment (45) of the propulsion unit through which a fluid (P) passes; - an exhaust channel (54) for the fluid, communicating with an outlet (55) of the item of equipment and comprising an orifice (56) for ejecting the fluid (P) into the gas flow (12b, 10b). According to the invention, the ejection orifice (56) is at least partially delimited by a flow control member (58) for controlling the flow through the item of equipment, the flow control member (58) being mounted so as to be translatable relative to the radial delimiting structure (31, 34, 36) so as to vary an outlet cross section of the fluid ejection orifice (56).

Description

[0001] Aircraft propulsion system assembly, comprising equipment intended to be traversed by a controlled fluid flow

[0002] TECHNICAL FIELD

[0003] The present invention relates to the field of aircraft propulsion systems, comprising equipment configured to be traversed by a flow of fluid, such as air from a secondary flow of the propulsion system, or from an external air flow of this propulsion system.

[0004] The invention applies to propulsion assemblies comprising turbomachinery of any type, such as turbojets and turboprops, with single or double fan or propeller, preferably ducted or unducted, and preferably of dual-flow design.

[0005] STATE OF PRIOR ART

[0006] In aircraft propulsion systems, equipment is used that allows fluids, such as air, to pass through it to ensure proper operation. This includes heat exchangers, such as air-to-air, air-to-oil, and air-to-fuel heat exchangers. For example, ACOC (Air-Cooled Oil-cooler) heat exchangers are designed to be supplied with air from a secondary flow within the propulsion system or from an external airflow. This type of exchanger is described, for instance, in document EP 2336525 A1.

[0007] For such equipment, the fluid flow requirement may vary depending on the point of flight, and / or depending on the engine operating point.

[0008] This usually leads to designing the propulsion system so that the fluid flow rate through the equipment meets the identified maximum requirement. However, for flight points and / or engine operating points that can operate with a lower fluid flow rate through the equipment, excessive fluid withdrawal negatively impacts the overall performance of the propulsion system, thus increasing fuel consumption. To address this issue, control valves can be installed at the fluid inlets in the equipment. However, the placement of these valves can be problematic, particularly in terms of size, weight, and control. Furthermore, locating the heat exchanger in a portion of the secondary flow path also affects the performance of the propulsion system.

[0009] DESCRIPTION OF THE INVENTION

[0010] To address at least partially the aforementioned drawbacks relating to prior art achievements, the invention first relates to an assembly for an aircraft propulsion system, comprising the characteristics of claim 1. The invention thus proposes a solution for varying the fluid flow through the equipment in a simple, reliable, easy-to-implement manner, based on varying the outlet section of the fluid ejection orifice provided at the end of the fluid exhaust channel.

[0011] The fluid sampling, preferably of air, can therefore be easily adapted to the specific needs associated with different flight points and / or engine operating points. This adaptation consequently reduces the amount of fluid drawn, and in this respect, the invention is the result of technological research aimed at significantly improving aircraft performance and, in this sense, contributes to reducing the environmental impact of these aircraft (decarbonization).

[0012] Furthermore, the proposed solution allows for a wide variability in the outlet cross-section of the fluid ejection orifice, thus enabling control of the fluid flow over a broad operating range, while simultaneously optimizing the fluid flow through the exhaust channel. In particular, when the flow control element is retracted, the induced drag on the fluid is low, and this drag remains limited at every operating point of the flow control element, comparable, for example, to a sliding flap.

[0013] The nature of the flow control device's movement, in translation, allows the fluid to be directed optimally within the gas flow at every operating point of this flow control device. The ejection of the fluid exiting the equipment into the gas flow can then occur with a satisfactory thrust coefficient.

[0014] The design specific to the invention allows for easy installation within the propulsion assembly, even for a flow control device extending over a high angular sector, for example greater than 10°, 20° or 30°, but which can go up to 180°.

[0015] Finally, the placement of the equipment outside the propulsion system's flow paths, particularly outside the primary and secondary flow paths, reduces disturbances in the flow, and thus advantageously increases the system's performance.

[0016] The invention also preferably provides for at least one of the following optional features, taken individually or in combination.

[0017] Preferably, the assembly is configured so that a downstream movement of the flow control device generates an increase in the outlet area of ​​the ejection orifice, or vice versa.

[0018] Preferably, the ejection orifice is also partly delimited by the radial delimiting structure.

[0019] Preferably, the flow control element is mounted to move in translation relative to the radial boundary structure, along a direction of movement parallel or substantially parallel to an axial direction of the assembly. This helps to improve the thrust coefficient of the fluid exiting the ejection orifice, in all positions of the control element. However, other directions of translation may be adopted without departing from the scope of the invention.

[0020] Preferably, the flow control element is mounted to move in translation on the radial boundary structure, between a retracted position within the radial boundary structure and a deployed position relative to this radial boundary structure. The retracted position may correspond to the position resulting in the smallest outlet area for the fluid ejection orifice, and the deployed position to the position resulting in the largest outlet area for this ejection orifice, or vice versa.

[0021] Preferably, the assembly includes a supply channel for said fluid, communicating with an inlet of the equipment, and comprising a fluid inlet orifice delimited by the radial delimiting structure, the fluid corresponding to a sample taken from said gas flow. However, the supply channel may be supplied by a fluid other than said gas flow as defined in the invention, and for example taken from an outside air flow or a secondary flow.

[0022] Preferably, the equipment is a heat exchanger, and the fluid is cooling air. Other equipment could nevertheless be considered, without departing from the scope of the invention.

[0023] Preferably, the assembly includes an electronic control system, configured to drive the translational control device of the flow control element, according to one or more data representative of a fluid flow requirement through the equipment.

[0024] Preferably, the translational control device includes at least one actuator for the flow control element, the actuator being circumferentially spaced from this flow control element. This facilitates the installation of the actuator in a very dense environment with limited space available in the radial direction. The invention also relates to an aircraft propulsion system comprising at least one such assembly, as well as a turbomachine. Preferably, the turbomachine is a turbofan engine, and even more preferably a turbofan or even a turbojet engine.

[0025] More preferably, the turbomachine includes an unshod fan, said radial boundary structure includes the radially external nacelle surface, and said gas flow corresponds to the external air flow to the propulsion assembly, conforming to the radially external nacelle surface.

[0026] Other advantages and features of the invention will appear in the detailed, non-limiting description below.

[0027] BRIEF DESCRIPTION OF THE DRAWINGS

[0028] This description will be made with reference to the attached drawings, among which are;

[0029] - [Fig. 1] represents a schematic longitudinal cross-sectional view of an aircraft propulsion assembly, according to the invention; - [Fig. 2] represents a more detailed half longitudinal cross-sectional view of an assembly for the propulsion assembly shown in the previous figure, the assembly being in the form of a preferred embodiment of the invention and comprising a flow control element through equipment, the flow control element being shown in a retracted position;

[0030] - [Fig. 3] represents a half longitudinal sectional view similar to the previous one, with the flow control device shown in a deployed position;

[0031] - [Fig. 4] is a cross-sectional view taken along line IV-IV of figure 2;

[0032] - [Fig. 5] is a cross-sectional view taken along line VV of figure 2;

[0033] - [Fig. 6] is a cross-sectional view taken along line VI-VI of figure 3.

[0034] DETAILED EXPLANATION OF PREFERRED METHODS OF IMPLEMENTATION

[0035] With reference first to Figure 1, a propulsion unit 100 for an aircraft is shown, comprising a turbomachine 1 and a nacelle 9 surrounding the turbomachine. In the preferred embodiment of the invention, which will be described below, the turbomachine preferably corresponds to a twin-spool turbofan engine, comprising a single ducted fan. However, it could be a turbomachine of another type, for example, a turboprop. More generally, it could be any type of turbojet or turboprop, with a single or twin fan or propeller, ducted or unducted.

[0036] Subsequently, the terms "upstream" and "downstream" are defined relative to a general direction 5 of gas flow through the propulsion unit 1 when it generates direct thrust, this direction being parallel or substantially parallel to axis 2. These terms "upstream" and "downstream" could respectively be replaced by the terms "front" and "rear," with the same meaning. Furthermore, the propulsion unit 1 is represented in a coordinate system formed by three orthogonal directions, namely the longitudinal direction L parallel to axis 2, which will be defined below, the circumferential direction C, and the radial direction R.

[0037] The turbojet 1 has a central longitudinal axis 2 around which its various components extend. It comprises, from upstream to downstream along the main direction 5 of gas flow through this turbomachine, a fan 3, a low-pressure compressor 4, a high-pressure compressor 6, a combustion chamber 11, a high-pressure turbine 7 and a low-pressure turbine 8. The fan 3 can be driven directly by a low-pressure unit comprising the compressor 4 and the turbine 8, or indirectly by a reduction gear (not shown).

[0038] Conventionally, an airflow F arriving at an air inlet of the blower splits into an internal airflow 10a which enters the nacelle, and an external airflow 10b which follows the radially external surface of the nacelle 36a. This surface 36a is defined by an outer skin of the nacelle 36.

[0039] After passing through the blower 3, the internal airflow 10a splits into a central primary flow 12a and a secondary airflow 12b, or secondary flow, the latter surrounding the primary flow. The primary flow 12a flows into a main gas circulation channel 14a passing through the compressors 4, 6, the combustion chamber 11, and the turbines 7, 8. The secondary flow 12b flows into a secondary air channel 14b, radially delimited outwards by a casing surrounded by the nacelle 9. More specifically, the casing includes a fan casing 20 surrounding the fan blades, this casing 20 being extended downstream by an outer ring 22 of an intermediate casing 24. This intermediate casing 24 has a hub 26 centered on the axis 2, and may include a flow separation nozzle 28.The intermediate casing is completed by radial arms 30, which form outlet guide vanes, allowing the secondary airflow 12b to be straightened and are conventionally called OGVs (from the English "Outlet Guide vanes"). The arms 30 thus connect the outer ferrule 22 of the intermediate casing to its hub 26, at the level of the low-pressure compressor 4.

[0040] The turbojet engine exhibits a bypass ratio, or BPR (Bypass Ratio), of approximately three to forty, and more specifically, approximately three to eighteen in the case of a shrouded fan corresponding to the described mode. In the case of one or more unshrouded fans / propellers, this bypass ratio is more likely to be in the range of eighteen to forty. Whether the fan(s) / propeller(s) are shrouded or unshrouded, their diameter is preferably in the range of 0.5 to 5 m, while the compression ratio is preferably in the range of 1 to 1.8.

[0041] In the context of a turbomachine equipped with a reduction gear, with the propeller(s) / fan(s) shrouded or unshrouded, the reduction ratio can be between 1.1 and 20. The outer shell 22 forms an external radial delimitation structure of the secondary air stream 14b. It is extended downstream by another annular delimitation structure 31 of this type, integrated into a movable thrust reverser hood 32. This external radial boundary structure of the propulsion assembly 31 is also called OFS (Outer Fixed Structure). It is located radially opposite an internal radial boundary annular structure 34 of the vein 14b, also called IFS (Inner Fixed Structure), and initiated from upstream by the nozzle 28. The internal radial boundary structure 34 includes the hub 26 of the intermediate housing.

[0042] Structure 34 has an internal radial boundary surface 34a of the secondary vein, just as structure 31 has an external radial boundary surface 31a of the secondary vein.

[0043] The rear parts of the skin 36 and the internal radial delimiting annular structure 34 of the vein 14b belong to the hood 32, which is preferably a sliding hood of a thrust reverser integrated into the propulsion assembly.

[0044] In a preferred embodiment of the invention shown in Figures 1 to 6, the propulsion assembly comprises an assembly 40 including the cowling 32, and therefore the outer nacelle skin 36, forming or belonging to an external radial boundary structure for the circulation of a gas flow, corresponding here to the external airflow 10b, which is not radially confined outwards. Hereafter, the outer nacelle skin, or the external radial boundary structure comprising this skin, will be referred to as the external radial boundary structure 36.

[0045] The assembly 40 could alternatively be mounted upstream of the cowling 32 in the nacelle, with its equipment 45 arranged axially at or near the blades 30, as shown in Figure 1. Combining several of these assemblies 40 remains possible, with identical or similar designs. However, in the remainder of this description, only the assembly 40 integrated into the cowling 32 will be described.

[0046] Within the thin radial thickness of the cowling 32, between the two elements 31 and 36 of this cowling, the assembly 40 includes a component 45 configured to allow the flow of a fluid, here corresponding to a sample from the outside airflow 10b. Preferably, the component 45 is a heat exchanger designed to allow the sample from the outside airflow 10b to pass through it, in order to create a cooling flow. For example, it is an ACOC type heat exchanger (Air-Cooled Oil-cooler). Another fluid is also intended to pass through the exchanger 45, such as oil for the lubrication and / or cooling of turbojet components, such as bearing housings or the fan drive gearbox.

[0047] The assembly includes a fluid supply channel 48, communicating with an inlet 50 of the equipment 45, and comprising a fluid inlet orifice 52 delimited by the structure 36, and thus opening at surface 36a into the outside airflow 10b. After passing through the inlet orifice 52, the sample P from the outside airflow 10b flows through the supply channel 48, located radially between the two elements 31 and 36. This sample P then passes inside the heat exchanger 45 for heat exchange with the oil. It is then extracted from the heat exchanger via an exhaust channel 54, communicating with an outlet 55 of the equipment. The exhaust channel 54 includes an orifice 56 for ejecting the sample P into the outside airflow 10b, which it therefore joins after following this bifurcation path within the hood 32. Alternatively, the fluid sample P could come from another source, for example from the secondary flow.

[0048] In this embodiment, the exchanger 45 is therefore arranged radially between the surface 36a and the secondary vein 14b of the assembly, being preferentially separated and at a distance from this vein radially outwards.

[0049] One of the distinctive features of the invention lies in the fact that the assembly 40 is designed so that the flow rate of the fluid sample P, which passes through the heat exchanger 45, can be controlled in a simple, robust, and efficient manner. To this end, the invention provides for the implementation of a flow control element 58, which at least partially delimits the ejection orifice 56. This control element is preferably a sliding cover or flap, which is mounted to move in translation relative to the structure 36, from upstream to downstream and vice versa. More precisely, the flow control element 58 is mounted to move in translation on the skin of the structure 36, or on any other element of the radial delimiting structure to which this skin belongs. The mounting preferably ensures a direction of movement of the element 58 parallel or substantially parallel to an axial direction of the assembly, corresponding to axis 2.

[0050] Thus, the flow control element 58 is mounted to move in translation between a retracted position within the radial boundary structure 36 and a deployed position relative to this same structure. The retracted position is shown in Figures 2 and 5, while the deployed position of the flow control element 58 is shown in Figures 3 and 6. A translational control device 60 for the flow control element 58 is provided. An electronic control system 62, included in or communicating with the turbofan engine's FADEC (Full Authority Digital Engine Control), is also provided. The system 62 is configured to control the control device 60 based on one or more data points representing a fluid flow requirement P through the heat exchanger 45.For example, this could be temperature data of fluid P, typically at outlet 55 of the exchanger, and / or temperature data of the other fluid passing through this exchanger.

[0051] Based on this data, the aim is to induce a precise displacement of the flow control element 58, in order to vary the outlet cross-section of the fluid P ejection orifice 56. In subsonic operation, it is this outlet cross-section of the ejection orifice 56 that determines the flow rate of the fluid P drawn through the heat exchanger 45. The ejection orifice 56 is delimited not only by the flow control element 58, but also by the radial delimiting structure 36. In the retracted position of the element 58, its rear end is located axially at or near a portion 64 of maximum diameter of the structure 36. Between these two elements, the outlet cross-section S of the fluid P ejection orifice 56 is minimal.After being displaced downstream in its deployed position, the rear end of the component 58 is axially aligned with a portion 66 of reduced diameter in the structure 36, which has a slightly conical shape and converges rearward from the portion 64. Between these two elements 58 and 36, the outlet area S of the fluid P ejection orifice 56 is at its maximum. Thus, the downstream displacement of the ejection orifice 56 leads to an increase in the outlet area S of this orifice. Conversely, a downstream displacement of the component 58 could be implemented, with a reduction in the outlet area.

[0052] In this regard, it is noted that the rear end of the flow control element 58 preferably has a slight inclination with respect to the axial direction, this inclination being identical or similar to the angle of the cone formed by the structure 36 with respect to the axial direction, between its two portions 64, 66. This allows the ejection of the flow P to be directed under good conditions into the external airflow 10b, at any operating point of the flow control element 58, in particular with a satisfactory thrust coefficient. Indeed, at the level of the ejection orifice 56, the external airflow 10b and the fluid P ejected by this orifice 56 locally have parallel or slightly inclined directions with respect to each other.Preferably, the angle between these two flow directions, locally at the ejection orifice 56, is preferably less than 20°, or even less than 10°, and more preferably on the order of 5°. These values ​​apply to all positions of the flow control element 58, throughout its entire operating range.

[0053] In this regard, it is noted that the radially external surface of the nacelle 36a, formed by the structure 36, may exhibit a radial step at the ejection orifice 56 defined in the retracted position of the component 58, this step being downwards as it moves downstream. This same step, also called a level break in the radial direction, is most clearly visible in Figure 2.

[0054] The flow control element 58 extends over an angular sector, for example, greater than 30° relative to axis 2, or even up to 180°, to define a single continuous ejection orifice 56 along direction C. However, several elements 58 could be arranged adjacently circumferentially to define a single continuous ejection orifice 56. Similarly, several elements 58 could be spaced circumferentially from one another around axis 2, so as to define several distinct ejection orifices 56, also spaced circumferentially from one another. The translational control device 60 for the flow control element 58 includes actuators 72, such as cylinders, for example, hydraulic, electric, or pneumatic.Here, for example, two actuators 72 are arranged circumferentially on either side of the control member 58, each thus being circumferentially spaced from this member 58, as schematically shown in Figure 4. This particular arrangement addresses the space constraint resulting from the limited space in the radial direction R. The two actuators 72, preferably synchronized to drive the same member 58 in translation, are connected to the latter by mechanical connecting means 74, or alternatively connected to a reinforcing member 76 fixed to this control member 58. To facilitate sliding, rails 78 are provided to guide the opposite longitudinal edges of the control member 58. The rails 78 are thus fixed to the external radial boundary structure 36, namely to the outer skin of the nacelle, and / or to another element of this structure. 36.

[0055] Of course, various modifications can be made by a person skilled in the art to the invention just described, solely by way of non-limiting examples. For instance, if the turbomachine described in the preferred embodiment above takes the form of a twin-flow turbojet with a shrouded fan, any other type of turbomachine remains conceivable.

Claims

DEMANDS 1. Assembly (40) for an aircraft propulsion system (100), comprising: - a radial delimitation structure (36) for the circulation of a gas flow (10b); - equipment (45) of the propulsion assembly configured to be traversed by a fluid (P); - an exhaust channel (54) for said fluid, communicating with an outlet (55) of the equipment and comprising an orifice (56) for ejecting the fluid (P) into said gas stream (10b), characterized in that the ejection orifice (56) is at least partially delimited by a flow control member (58) through the equipment, the flow control member (58) being mounted movable in translation relative to the radial delimiting structure (36) so as to vary an outlet section (S) of the ejection orifice (56) of the fluid, the assembly also comprising a translational control device (60) of the flow control member (58), the radial delimiting structure having a radially external nacelle surface (36a), said gas flow corresponding to an air flow (10b) external to the propulsion assembly, conforming to the radially external nacelle surface, the equipment (45) and the exhaust channel (54) being located radially between the radially external nacelle surface (36a), and a secondary vein (14b) of the propulsion assembly.

2. Assembly according to claim 1, characterized in that it is configured so that a downstream displacement of the flow control member (58) generates an increase in the outlet section (S) of the ejection orifice (56).

3. Assembly according to claim 1 or 2, characterized in that the ejection orifice (56) is also partly delimited by the radial delimiting structure (36).

4. Assembly according to any one of the preceding claims, characterized in that the flow control member (58) is mounted to move in translation relative to the radial delimiting structure (36), along a direction of movement parallel or substantially parallel to an axial direction of the assembly.

5. Assembly according to any one of the preceding claims, characterized in that the flow control member (58) is mounted to move in translation on the radial delimiting structure (36), between a position retracted into the radial delimiting structure, and a position deployed relative to this radial delimiting structure (36).

6. Assembly according to any one of the preceding claims, characterized in that it comprises a supply channel (48) for said fluid (P), communicating with an inlet (50) of the equipment (45), and comprising an inlet orifice (52) for the fluid delimited by the radial delimiting structure (36), the fluid corresponding to a sample from said gas flow (10b).

7. Assembly according to any one of the preceding claims, characterized in that the equipment (45) is a heat exchanger, and in that said fluid (P) is cooling air.

8. Aircraft propulsion assembly (100) comprising at least one assembly (15) according to any one of the preceding claims, as well as a turbomachine (1).

9. Assembly according to the preceding claim, characterized in that the turbomachine (1) comprises an unshod fan (3), in that said radial delimiting structure (36) comprises the radially external nacelle surface (36a), and in that said gas flow corresponds to the external air flow to the propulsion assembly, conforming to the radially external nacelle surface (36a).

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