System and method for supplying fuel to a combustion chamber of a turbine shaft engine of an aircraft

Abstract

A supply system (S) for supplying fuel (C) to a combustion chamber (5) of a turboshaft engine (6) of an aircraft comprises an injector (1) comprising a first pilot branch (10), a second pilot branch (16) and a main branch (13), the pilot branches (10, 16) being configured to deliver a constant flow of fuel (C), the main branch (13) being configured to deliver an intermittent flow of fuel (C); a distribution device (2) for distributing fuel (C) to the injector (1) comprises a main circuit (21), a secondary circuit (22) and a valve (20) configured to distribute fuel (C) between the main circuit (21) and the secondary circuit (22), the main circuit (21) being connected to the first pilot branch (10) and to the main branch (13), the secondary circuit (22) being connected to the second pilot branch (16); and a control (3) for controlling the valve (20).

Description

Technical Field

[0001] This invention relates to the field of supplying fuel to the combustion chamber of an aircraft turboshaft engine. Background Technology

[0002] Reference Figure 1 Knownly, the aircraft includes one or more turboshaft engines 600, each turboshaft engine including a combustion chamber 500, configured to produce a combustion reaction. Air A and fuel C enter the combustion chamber 500, thereby releasing the energy required for the thrust phase of the aircraft. Air A originates from outside the turboshaft engine 600 and is guided to the combustion chamber 500 by an airflow (not shown), and fuel C originates from a fuel circuit 400 leading to the combustion chamber 500.

[0003] Fuel system 400 includes a set of injectors 100 configured to inject fuel C in droplet form into combustion chamber 500, and a distribution device 200 for distributing fuel C to the injectors 100 controlled by control unit 300. Upstream, fuel circuit 400 typically includes the following elements (not shown) arranged sequentially from upstream to downstream in the direction of fuel C flow: a storage tank, a low-pressure pump, a filter, a high-pressure pump, and a device for metering the fuel C injected into combustion chamber 500, typically referred to as a fuel metering unit (“FMU”).

[0004] Still refer to Figure 1 A so-called "two-stage" injector 100 is known, comprising, on one hand, an ignition branch 110 configured to deliver a constant flow of fuel C optimized for low and medium speeds of the turboshaft engine 600, and on the other hand, a main branch 113 configured to deliver an intermittent flow of fuel C only for high speeds of the turboshaft engine 600. The ignition branch 110 is equipped with a sealing valve 111 and a nozzle 112. The sealing valve 111 is configured to circulate the fuel C when the pressure of the fuel C in the circuit reaches a set value, and the nozzle 112 is configured to generate a spray of fuel C at the center of the combustion chamber 500. The nozzle 112 is of aerodynamic injection type, i.e., it generates a fuel-air mixture that forms a spray by shearing the fuel C between two air vortices. Referring again... Figure 1 The main branch line 113 is equipped with a multi-point valve 114 and a multi-point nozzle 115. The multi-point nozzle 115 includes a set of circumferential injection points located around the nozzle 112 of the ignition branch line 110.

[0005] Still refer to Figure 1 Knownly, the distribution device 200 includes an ignition circuit 211 and a main circuit 212 that supply fuel C to the ignition branch line 110 and the main branch line 113 of each injector 100, respectively. The distribution device 200 also includes a distribution valve 210 that distributes fuel C between the ignition circuit 211 and the main circuit 212.

[0006] The two-stage injector 100 and dual-loop distributor 200 advantageously form a fuel supply system capable of optimizing the combustion response at any speed of the turboshaft engine 600, thereby reducing combustion gas exhaust while generating the desired thrust for the aircraft. However, in the event of a malfunction in the distributor valve 210 or the control unit 300, all fuel C may be introduced into a single loop 211, 212. If introduced into the ignition loop 211, it may result in an undesirable thrust loss from the turboshaft engine 600. If introduced into the main loop 212, it may impair the operation of the turboshaft engine 600 and could potentially cause engine shutdown.

[0007] A direct solution to avoid this drawback is to replicate and improve the safety of the distribution valve 210 and the control element 300 to prevent any risk of failure. However, improvements to the distribution valve 210 or the control element 300 are complex and would significantly increase the overall size and weight.

[0008] In addition to the aforementioned drawbacks, when the turboshaft engine 600 is at low speed, all fuel C is introduced into the ignition circuit 211, which increases the risk of coking (solidification through gum formation) of the remaining fuel C in the main circuit 212 and main branch line 113 of the injector 100. Furthermore, at high speeds, introducing high-velocity fuel C into the main branch line 113 of the injector may lead to unstable combustion in the combustion chamber 500. Prior art supply systems are known from documents US20150369489A1 and EP3211200A1. Summary of the Invention

[0009] The present invention provides a system and method for supplying fuel to a combustion chamber, thereby at least partially eliminating the aforementioned disadvantages.

[0010] This invention relates to a system for supplying fuel to the combustion chamber of an aircraft turboshaft engine, the system comprising:

[0011] - Multiple injectors configured to inject fuel into the combustion chamber, each injector being a three-stage injector comprising a first ignition branch, a second ignition branch, and a main branch. The first and second ignition branches are respectively equipped with a first ignition valve and a second ignition valve, which are configured to circulate fuel at a constant flow rate. The main branch is equipped with a main valve configured to circulate fuel at an intermittent flow rate, so as to adjust the supply to the combustion chamber according to the speed of the aircraft's turboshaft engine.

[0012] - A distribution device for distributing fuel in the injector and in the form of a dual-circuit distribution device including a main circuit and a secondary circuit, the main circuit being connected to a first ignition branch and the secondary circuit being connected to a second ignition branch, the distribution device including a distribution valve configured to distribute fuel between the main circuit and the secondary circuit, and

[0013] - A control unit for controlling the distribution valve according to the speed of the aircraft's turbine shaft engine.

[0014] Due to the three-stage structure of the injector and the dual-loop structure of the distribution device, a failure of the control components and / or the distribution valve could lead to the distribution valve being in an uncontrolled distribution position and thus shutting down. The power supply system of this invention advantageously increases safety in the event of a failure of the control components and / or the distribution valve. In fact, even if the closed position corresponds to either of the two end positions of the distribution valve (meaning only the main loop or the secondary loop is supplied), the ignition branch is still systematically supplied. Therefore, if all fuel is introduced into the main loop, it will only result in acceptable degraded operation without significantly reducing the performance of the aircraft turboshaft engine. However, since the fuel no longer circulates in the secondary loop, there is a risk of combustion instability and fuel coking in the secondary loop, but this risk will not lead to runaway engine shutdown and will only have an impact during injector maintenance. In the opposite possibility, where all fuel is introduced into the secondary loop, the aircraft turboshaft engine operation will only suffer a small thrust loss, especially when the secondary loop has relatively high permeability and allows a large jet flow rate. This operation will also not lead to runaway engine shutdown. Therefore, in the event of a malfunction in the distribution valve, the supply system of the present invention can reduce the impact of the malfunction and limit the performance loss to a moderate level.

[0015] Furthermore, this invention advantageously allows the risks of coking and combustion instability to be avoided during normal operation of the distribution valve. In fact, during operation, once the fuel flow rate entering the distribution valve exceeds a relatively low value, the distribution valve can be controlled in a suitable position to ensure continuous fuel circulation in the main and secondary circuits, thereby preventing fuel coking. Additionally, compared to the prior art, the three-stage structure of the injector allows for a reduction in fuel pressure in each branch, particularly in the main branch, because a portion of the fuel flow is distributed into the secondary circuit.

[0016] According to one aspect of the invention, the first ignition branch, the second ignition branch, and the main branch are respectively provided with a first ignition nozzle communicating with the output end of the first ignition valve, a second ignition nozzle communicating with the output end of the second ignition valve, and a main nozzle communicating with the output end of the main valve. The injection nozzles advantageously enable fuel to be injected into the combustion chamber.

[0017] Preferably, the permeability of the second ignition nozzle is greater than that of the first ignition nozzle to deliver a flow rate suitable for igniting the aircraft turboshaft engine. It is generally accepted that the nozzle permeability corresponds to the maximum fuel flow rate it can inject into the combustion chamber. Advantageously, the low permeability of the first ignition nozzle thus allows for a relatively low fuel flow rate suitable for igniting the combustion chamber when starting the aircraft turboshaft engine. Furthermore, since the second ignition nozzle can inject more fuel than the first ignition nozzle, more fuel flows to the secondary circuit rather than the primary circuit, thereby preventing accidental opening of the main valve. This also eliminates the need to increase the opening pressure of the main valve and the pressure in the main circuit.

[0018] Preferably, the permeability of the second ignition nozzle is lower than that of the main nozzle to ensure that high fuel flow passes through the high speed of the aircraft's turboshaft engine.

[0019] According to one aspect of the invention, the first ignition nozzle is a pneumatic injection type and is configured to inject a fuel spray into the combustion chamber. The first ignition nozzle is advantageously configured to form an air-fuel mixture by shearing the fuel between two air vortices. Preferably, the first ignition nozzle is located at the center of each injector to ensure a continuous supply of fuel to the combustion chamber.

[0020] According to one aspect of the invention, the second ignition nozzle is aerodynamically designed and configured to inject a fuel spray into the combustion chamber. Preferably, the second ignition nozzle is located around each injector, more preferably around the first ignition nozzle. This nozzle advantageously allows the fuel to be pressurized for injection as a spray. Preferably, the second ignition nozzle forms an annular shape around the first ignition nozzle, extending concentrically. Thus, the two ignition nozzles together ensure a continuous injection of fuel into the combustion chamber, which is sufficient at low and medium speeds.

[0021] According to one aspect of the invention, the main nozzle is in the form of a multi-point nozzle comprising a plurality of injection points, the injection points preferably located around each injector, and preferably around the first ignition nozzle. Advantageously, the main nozzle is configured to inject fuel as an auxiliary to the two ignition nozzles when the aircraft turboshaft engine is at high speed.

[0022] According to one aspect of the invention, the first ignition valve is a sealing valve, preferably without a metering function. This first ignition valve advantageously ensures that only the sealing of the first ignition branch can be guaranteed when the engine is stopped or before the combustion chamber is ignited.

[0023] According to a preferred aspect, the second ignition valve is a sealing valve and preferably does not have a metering function. This second ignition valve advantageously ensures that only the sealing of the second ignition branch is maintained when the engine is stopped or before the combustion chamber is ignited.

[0024] According to another preferred aspect, the main valve is a multi-point valve. This main valve advantageously ensures supply to each injection point of the main nozzle.

[0025] According to one aspect of the invention, the opening pressure of the main valve is greater than the opening pressure of the first ignition valve, so as to deliver a fuel flow rate only greater than a predetermined pressure in the main branch corresponding to a high fuel flow rate, i.e., at high speeds of the turboshaft engine or optionally at medium turboshaft engine speeds. Therefore, although the main valve and the first ignition valve are connected to the same main circuit, only the first ignition valve is configured to allow fuel passage when the main circuit pressure is lower than the predetermined opening pressure of the main valve. When the aircraft turboshaft engine is at low speeds, this restricts injection to both ignition valves.

[0026] It is generally believed that the valve's opening pressure corresponds to the minimum fuel pressure in the branch line, and the valve is configured to open at the minimum fuel pressure and allow fuel to pass through. For example, in the case of a spring valve, a spring is provided to press the ball against the sealing seat, and the spring's opening pressure is a function of the spring constant and the compression of the spring in the assembly.

[0027] According to one aspect of the invention, the opening pressure of the first ignition valve is substantially equal to the opening pressure of the second ignition valve, so as to ensure fuel sealing of the first and second ignition branches when the engine is stopped or during the start-up phase before the combustion chamber is ignited. The term "substantially" means that there is a deviation of up to 10% between the two opening pressure values.

[0028] According to another aspect of the invention, the second ignition valve is a metering valve, and the second ignition branch is provided with a sealing valve. Preferably, the second ignition valve is located downstream of the sealing valve of the second ignition branch. Also preferably, the opening pressure of the second ignition valve is greater than the opening pressure of the sealing valve of the second ignition branch. Preferably, the opening pressure of the sealing valve of the second ignition branch is substantially equal to the opening pressure of the first ignition valve. The combined use of the sealing valve and the metering valve can maintain the low pressure in the second ignition branch while controlling the fuel flow through the metering valve.

[0029] According to one aspect, the second ignition branch is provided with an anti-coking conduit, which opens on one side between the second ignition valve and the sealing valve of the second ignition branch, and on the other side upstream of the supply system, so as to allow fuel to circulate through the sealing valve when the second ignition valve is closed. The use of the anti-coking conduit allows fuel to circulate even at low pressure, where the sealing valve is open but the metering valve (i.e., the second ignition valve) is closed.

[0030] The present invention also relates to a fuel circuit for an aircraft turboshaft engine, the fuel circuit including the supply system as described above, the fuel circuit being configured to supply fuel to a distribution device.

[0031] The present invention also relates to the assembly of the fuel system and the combustion chamber of the aircraft turboshaft engine as described above, wherein the injectors of the supply system extend into the combustion chamber to supply fuel to the combustion chamber.

[0032] The present invention also relates to an assembly of a combustion chamber and fuel circuit for an aircraft turboshaft engine, the fuel circuit including a supply system as described above and configured to supply fuel to a distribution device, wherein an injector of the fuel system extends into the combustion chamber to supply fuel to the combustion chamber.

[0033] The present invention also relates to an aircraft turboshaft engine, the aircraft turboshaft engine including a combustion chamber and a supply system as described above, wherein the injectors of the supply system extend into the combustion chamber to supply fuel to the combustion chamber.

[0034] The present invention also relates to an aircraft turboshaft engine, the aircraft turboshaft engine including a combustion chamber and a fuel circuit, the fuel circuit including a supply system as described above, the fuel circuit being configured to supply fuel to a distribution device, and an injector of the fuel system extending into the combustion chamber to supply fuel to the combustion chamber.

[0035] The present invention also relates to an aircraft comprising at least one turboshaft engine as described above.

[0036] The present invention also relates to a method for supplying fuel to the combustion chamber of an aircraft turboshaft engine using the supply system as described above, the method comprising:

[0037] - The step of controlling the distribution valve via the control unit according to the speed of the aircraft's turbine shaft engine.

[0038] - The step of distributing the fuel between the main circuit and the secondary circuit via the distribution valve, and

[0039] - The step of injecting fuel into the combustion chamber via the injector.

[0040] - Wherein, regardless of the speed at which the aircraft turboshaft engine is, once the turboshaft engine is started, the distribution valve introduces fuel into the main circuit and the secondary circuit during the distribution step.

[0041] Advantageously, this supply method is effective at any operating speed of the aircraft's turboshaft engine through its three-stage injector architecture, and its dual-loop structure, each connected to an ignition branch line, enhances safety in the event of a failure. Because both loops are continuously supplied with fuel, this method is particularly effective in reducing risks such as fuel charring. Attached Figure Description

[0042] The invention will be better understood by reading the following description given as an example and by referring to the following drawings given as a non-limiting example, wherein the same reference numerals are used to denote similar objects.

[0043] Figure 1 This is a schematic diagram of the fuel supply system for the combustion chamber of a conventional aircraft turboshaft engine;

[0044] Figure 2 This is a schematic diagram of the fuel supply system of the combustion chamber of an aircraft turboshaft engine according to an embodiment of the present invention;

[0045] Figure 3A yes Figure 2 A schematic diagram of the injectors in the central supply system;

[0046] Figure 3B yes Figure 3A A schematic diagram of the injector nozzle;

[0047] Figure 4 This is a schematic diagram of a method for supplying fuel to the combustion chamber of an aircraft turboshaft engine according to an embodiment of the present invention;

[0048] Figure 5A When the aircraft's turboshaft engine is at low speed Figure 4 The diagram illustrates the steps of distributing and injecting fuel into the injector.

[0049] Figure 5B When the aircraft's turboshaft engine is at high speed Figure 4 The diagram illustrates the steps of distributing and injecting fuel into the injector.

[0050] Figure 6A and Figure 6B When the control and / or allocation steps fail Figure 4 Two schematic diagrams illustrating the steps of injecting fuel into the injector in the method shown;

[0051] Figure 7This is a schematic diagram of an injector according to an alternative embodiment of the present invention.

[0052] It should be noted that the accompanying drawings illustrate the invention in detail in order to implement the invention, and the drawings may, of course, be used to better define the invention if necessary. Detailed Implementation

[0053] Reference Figure 2 This invention relates to an aircraft comprising one or more turboshaft engines 6, each turboshaft engine 6 including a combustion chamber 5 into which air A and fuel C, which produce a combustion reaction, enter to release energy required for the thrust phase of the aircraft. Air A originates from outside the turboshaft engine 6 and is directed to the combustion chamber 5 by an airflow (not shown), while fuel C originates from a fuel circuit 4 leading to the combustion chamber 5. The fuel circuit 4 includes the following components (not shown) arranged from upstream to downstream in the direction of fuel C flow: a storage tank, a low-pressure pump, a filter, a high-pressure pump, and a fuel device for metering the fuel C injected into the combustion chamber 5, commonly referred to as a fuel metering unit (“FMU”).

[0054] Still refer to Figure 2 Furthermore, according to the present invention, the fuel circuit 4 also includes a fuel supply system S located in the downstream combustion chamber 5, the system comprising:

[0055] -Injector 1 extends into combustion chamber 5,

[0056] - Distribution device 2 for distributing fuel C in injector 1 and

[0057] - Control unit 3 for controlling the distribution device 2.

[0058] According to the present invention, such as Figure 2 As shown, each injector 1 includes three branches: a first ignition branch 10, a second ignition branch 16, and a main branch 13. The first ignition branch 10 and the second ignition branch 16 are respectively equipped with a first ignition valve 11 and a second ignition valve 17 configured to circulate fuel C at a constant flow rate. The main branch 13 is further equipped with a main valve 14 configured to circulate fuel C at an intermittent flow rate, so as to adjust the supply to the combustion chamber 5 according to the speed of the aircraft turboshaft engine 6. In practice, the first ignition valve 11 and the second ignition valve 17 are adapted to open when the turboshaft engine 6 is ignited, i.e., by the low pressure of fuel C in the fuel circuit 4. The main valve 14 is also adapted to open by the relatively high pressure of fuel C in the fuel circuit 4 corresponding to the high speed or, optionally, the moderate speed with acceleration of the aircraft turboshaft engine 6.

[0059] Reference Figure 2The first ignition branch line 10, the second ignition branch line 16, and the main branch line 13 are also respectively provided with a first ignition nozzle 12, a second ignition nozzle 18, and a main nozzle 15. Each nozzle 12, 15, 18 is installed downstream of an associated valve 11, 14, 17 relative to the flow direction of fuel C in the supply system S. Each nozzle 12, 15, 18 is configured to inject fuel C that has flowed through the corresponding valve 11, 14, 17 into the combustion chamber 5.

[0060] Still in accordance with the present invention and reference Figure 2 The distribution device 2 includes two circuits: a main circuit 21 connected to the first ignition branch line 10 and the main branch line 13, and a secondary circuit 22 connected to the second ignition branch line 16. The distribution device 2 also includes a distribution valve 20 configured to distribute fuel C between the main circuit 21 and the secondary circuit 22. A control unit 3 is connected to the distribution device 2 and configured to control the movement of the distribution valve 20 according to the speed of the aircraft turbine shaft engine 6, thereby adjusting the supply of fuel C to the combustion chamber 5.

[0061] Advantageously, the control element 3 and the dispensing device 2 are similar to those in the prior art, which makes it possible to achieve the advantages of the present invention by simply modifying the existing supply system.

[0062] exist Figure 2 In the example, for clarity and brevity, only three injectors 1 are shown, but it goes without saying that the number of injectors 1 can be any number, preferably greater than 10 and less than 30. The dispensing device 2 is also preferably unique, as is the control element 3.

[0063] Furthermore, within the scope of this invention, each injector 1 includes only three branches 10, 13, and 16, and the distribution device 2 includes only two loops 21 and 22. The main loop 21 and the secondary loop 22 are connected to each injector 1. Each loop 21 and 22 is divided into several delivery paths at a node, and each delivery path is connected to a given injector 1. Figure 2 In this example, the main circuit 21 and the secondary circuit 22 are thus each divided into three delivery paths, the number of which is equal to the number of injectors 1. The delivery paths are preferably identical to ensure an even distribution of fuel C among the different injectors 1. Therefore, in a manner similar to the prior art, each injector 1 is connected to a main circuit 21 and a secondary circuit 22.

[0064] Reference Figure 3AAs previously described, each injector 1 includes three valves 11, 14, and 17: two ignition valves 11 and 17 for continuous injection of fuel C, and a main valve 14 for intermittent injection of fuel C. More precisely, the main valve 14 is configured to open as a backup to the ignition valves 11 and 17 to ensure an adequate supply of fuel C when the aircraft turboshaft engine 6 is at high engine speeds. The ignition valves 11 and 17 together provide sufficient fuel C to the aircraft turboshaft engine 6 at low speeds. The use of two ignition valves 11 and 17, instead of a single ignition valve as in the prior art, ensures flame stability at high speeds.

[0065] Still refer to Figure 3A The first ignition valve 11 and the second ignition valve 17 respectively deliver fuel C, that is, fuel C from the main circuit 21 and the secondary circuit 22 respectively. In the event that a malfunction in the distribution valve 20 or the control element 3 introduces all fuel C into either circuit 21 or 22, the consequences can be minimized to a small thrust loss. In fact, regardless of which circuit 21 or 22 the entire fuel C is delivered to, it will be directed to one of the ignition valves 11 or 17, thus avoiding the risk of engine shutdown.

[0066] like Figure 3A As shown, the main circuit 21 supplies power to both the first ignition valve 11 and the main valve 14. At the internal node N of each injector 1, it splits into two paths: one corresponding to the first ignition branch 10, and the other to the main branch 13, thus limiting the number of circuits 21 and 22 in the distribution device 2 to two. Furthermore, since circuit 21 continuously supplies power to the ignition valve 11 and circuit 22 continuously supplies power to the ignition valve 17, the risk of coking of fuel C in circuits 21 and 22 is avoided.

[0067] Preferably, the first ignition valve 11 is a sealing valve configured to ensure the sealing of the first ignition branch 10 when the turbine engine 6 is stopped or before the combustion chamber 5 is ignited. According to a preferred aspect, the second ignition valve 17 is also a sealing valve for the same reason. Preferably, the main valve 14 is also a multi-point valve, configured to supply power to all injection points of the main nozzle 15, as described below.

[0068] Reference Figure 3AValves 11, 14, and 17 have opening pressures T11, T14, and T17, respectively, which correspond to the minimum pressure of fuel C in branches 10, 13, and 16. Valves 11, 14, and 17 are configured to inject fuel C into combustion chamber 5 at the minimum pressure. Preferably, the opening pressure T14 of the main valve 14 is greater than the opening pressure T11 of the first ignition valve 11, thereby allowing the fuel C supplied by the main valve 14 and the first ignition valve 11 to be managed independently, wherein both the main valve 14 and the first ignition valve 11 are supplied by the main circuit 21. In practice, below the predetermined pressure in the main circuit 21, i.e., when the aircraft turbine shaft engine is at a low speed, only the first ignition valve 11 causes fuel C to circulate. Above the predetermined pressure, i.e., when the aircraft turbine shaft engine is at a high speed, both the first ignition valve 11 and the main valve 14 cause fuel C to circulate. Preferably, the predetermined pressure is greater than 2 bar and less than 4 bar. In the two ignition valves 11 and 17 of the same type as in the previous example, the opening pressure T11 of ignition valve 11 is substantially equal to and lower than the opening pressure T17 of ignition valve 17, in order to continuously deliver fuel C to combustion chamber 5, especially when the pressures in the main circuit 21 and secondary circuit 22 are low. Alternatively, the second ignition valve 17 may be in a form other than a sealing valve, as shown later, for example, configured as a metering valve that meters fuel C according to the pressure in fuel circuit 4 and is associated with the sealing valve.

[0069] Preferably, the first ignition nozzle 12 is of aerodynamic injection type, that is, the first ignition nozzle 12 is configured to generate a spray of fuel C by shearing the fuel C between two air vortices. Also preferably, the second ignition nozzle 18 is of aerodynamic injection type, that is, the second ignition nozzle 18 is configured to generate a spray by pressurizing the fuel C. Therefore, the first ignition nozzle 12 and the second ignition nozzle 18 have different structures, thereby ensuring proper combustion throughout their entire operating range (from low to medium speeds). Figure 3B In this example, the first ignition nozzle 12 is located at the center of the injector 1. The second ignition nozzle 18 is located around the first ignition nozzle 12, and in this example, it is in a concentric ring with the first ignition nozzle 12. Therefore, the spray produced by the second ignition nozzle 18 completes the spray from the first ignition nozzle 12.

[0070] Still refer to Figure 3B The main nozzle 15 preferably includes a set of injection points 15-i for fuel C. Figure 3B In the example, the injection points 15-i of the main nozzle 15 are located on the periphery of each injector 1. In this example, the injection points 15-i are distributed in a ring and are concentric with the ignition nozzles 12 and 18, and the injection points 15-i extend outward along the ring.

[0071] Reference Figure 3A Each nozzle 12, 15, and 18 has a permeability P12, P15, and P18, respectively, which correspond to the maximum flow rate of fuel C that can be injected through nozzles 12, 15, and 18. Preferably, the permeability P12 of the first ignition nozzle 12 is less than the permeability P18 of the second ignition nozzle 18, meaning that the permeability of the first ignition branch 10 is less than the permeability of the second ignition branch 16, so that fuel C with a low flow rate can be injected when the aircraft turboshaft engine 6 is ignited. This also ensures that at low speeds, the injected fuel C mainly comes from the second ignition branch 16, thus reducing the pressure in the main circuit 21 to prevent the main valve 14 from being easily opened. Also preferably, the permeability P15 of the main nozzle 15 is the maximum, so that fuel C with a high flow rate can be delivered to the aircraft turboshaft engine 6 at high speeds.

[0072] In summary, the supply system S of the present invention enhances safety with its three-stage injector, comprising two ignition branches 10 and 16, a main branch 13, and a dual-loop distribution device 2 for distributing fuel C. In effect, the ignition branches 10 and 16 ensure the stability of the flame of the aircraft turboshaft engine 6 at high speeds, and the ignition branches 10 and 16 are supplied separately to ensure an acceptable degradation mode in the event of fuel C distribution failure. The first ignition branch 10 and the main branch 13 are supplied by the same main loop 21 to avoid any risk of charring. Existing fuel systems can therefore be converted into the system S of the present invention by simply modifying the injector 1.

[0073] The following describes a method for supplying the combustion chamber 5 using the aforementioned supply system S.

[0074] Reference Figure 4 According to the present invention, the method begins with control step E1, during which control unit 3 transmits a control signal S to distribution valve 20, the control signal S being a function of the speed of the aircraft turboshaft engine 6. In other words, control unit 3 controls the position of distribution valve 20 and adjusts the position of distribution valve 20 primarily based on the speed of the aircraft turboshaft engine 6.

[0075] like Figure 4 As shown, once the position of the distribution valve 20 is determined, the method includes step E2, which involves distributing fuel C between the main circuit 21 and the secondary circuit 22. In practice, regardless of the speed of the aircraft turboshaft engine 6, outside of the ignition period and under fault-free conditions, the distribution valve 20 introduces fuel C into the main circuit 21 and the secondary circuit 22, allowing the ignition valves 11 and 17 to continuously supply fuel to the combustion chamber 5. In other words, the position of the distribution valve 20 adjusts the relative proportion Q21 of fuel C delivered to the main circuit 21 and the relative proportion Q22 of fuel C delivered to the secondary circuit 22.

[0076] Still refer to Figure 4 Once fuel C is distributed between the main circuit 21 and the secondary circuit 22, the method includes an injection step E3, during which each injector 1 delivers fuel C into the combustion chamber 5. More precisely, in addition to the injection being performed by the main nozzle 15, the speed of the aircraft turbine shaft engine 6 is also achieved by the ignition nozzles 12 and 18 during injection step E3.

[0077] Figure 5A The invention's supply method is illustrated when the aircraft's turbine shaft engine 6 is at low speed. In step E2, the distribution valve 20 is positioned to distribute fuel C in the secondary circuit 22 at a larger proportion Q22 and in the main circuit 21 at a smaller proportion Q21. Due to their smaller opening pressures T11 and T17, in step E3, only the ignition valves 11 and 17 are opened to allow the ignition nozzles 12 and 18 to inject fuel C into the combustion chamber 5. In other words, the pressures in the main circuit 21 and secondary circuit 22 are lower than the predetermined opening pressure T14 of the main valve 14. Furthermore, the injection step E3 is primarily carried out by the second ignition nozzle 18, which has a larger permeability P18 than the first ignition nozzle 12.

[0078] Figure 5B The invention's fuel supply method is illustrated when the aircraft turbine shaft engine 6 is at high speed. In step E2, the distribution valve 20 is positioned to distribute fuel C in the main circuit 21 at a larger proportion Q21 and in the secondary circuit 22 at a smaller proportion Q22. Then, in step E3, all nozzles 12, 15, and 18 inject fuel C into the combustion chamber 5. In other words, the pressure in the main circuit 21 and the secondary circuit 22 is higher than the predetermined opening pressure T14 of the main valve 14, and therefore higher than the predetermined opening pressures of the ignition valves 11 and 17. Furthermore, the injection step E3 is primarily carried out by the main nozzle 15, which has a larger penetration rate P15.

[0079] Advantageously, by moving the distribution valve 20 and utilizing the differences in opening pressures T11, T14, T17 corresponding to valves 11, 14, 17 and the differences in permeability P12, P15, P18 corresponding to nozzles 12, 15, 18 associated with each injector 1, this supply method thus allows for optimization of the supply to the combustion chamber 5 based on the engine speed of the aircraft turboshaft engine 6. Furthermore, as... Figure 6A and Figure 6B As shown, in the event of a failure during control step E1 and / or allocation step E2, an acceptable degradation mode is guaranteed because each loop 21, 22 is connected to the ignition branch lines 10, 16. In fact, in Figure 6A In the example, the entire fuel flow Q21 is introduced into the main circuit 21, and fuel C is injected by the first ignition nozzle 12 and the main nozzle 15, thereby ensuring low thrust loss. Figure 6B In the opposite example, the entire fuel flow Q22 is introduced into the secondary loop 22, and fuel C is injected only by the second ignition nozzle 18, which has a greater permeability P18 than the first ignition nozzle 12, thereby reducing thrust loss.

[0080] Figure 7 An alternative embodiment of the invention is shown, which differs from the previous embodiment in that the second ignition valve 17' is a metering valve, and the ignition branch 16 further includes a sealing valve 19 and an anti-coking conduit CF. Figure 7 As shown, the second ignition valve 17' is located downstream of the sealing valve 19 to supply the second ignition nozzle 18. The second ignition valve 17', i.e., the metering valve, has an opening pressure 17' greater than the opening pressure T19 of the sealing valve 19. The anti-coking conduit CF is in the form of a leakage conduit, connecting on one side to the second ignition branch 16 between the second ignition valve 17 and the auxiliary valve 19, and on the other side to the upstream of the supply system S. Advantageously, the sealing valve 19 and the first ignition valve 11 can have substantially equal opening pressures T11, T19, i.e., low opening pressures, thereby allowing fuel C to circulate in the second ignition branch 16 even when the aircraft turbine shaft engine 6 is at low speed. When the pressure in the fuel circuit 14 is greater than the opening pressure T19 of the sealing valve 19 but less than the opening pressure 17' of the second ignition valve 17', the anti-coking conduit CF allows fuel C to leak controllably from the second ignition branch 16. In other words, when the sealing valve 19 is open and the second ignition valve 17' is closed, fuel C circulates via the anti-coking conduit CF. When the second ignition valve 17' is closed, the controlled circulation of fuel C within the sealing valve 19 enables cooling of the sealing valve 19 and thus prevents coking of the fuel C.

Claims

1. A supply system (S) for supplying fuel (C) to the combustion chamber (5) of an aircraft turboshaft engine (6), characterized in that, The supply system (S) includes: - Multiple injectors (1) configured to inject the fuel (C) into the combustion chamber (5), each injector (1) being a three-stage injector comprising a first ignition branch (10), a second ignition branch (16) and a main branch (13), the first ignition branch (10) and the second ignition branch (16) being respectively provided with a first ignition valve (11) and a second ignition valve (17, 17') configured to circulate fuel (C) with a constant flow rate, the main branch (13) being provided with a main valve (14) configured to circulate fuel (C) with an intermittent flow rate to adjust the supply to the combustion chamber (5) according to the speed of the aircraft turboshaft engine (6); - A distribution device (2) for distributing the fuel (C) in the injector (1) and in the form of a dual-loop distribution device including a main loop (21) and a secondary loop (22), the main loop (21) being connected to the first ignition branch (10) and the main branch (13), the secondary loop (22) being connected to the second ignition branch (16), the distribution device (2) including a distribution valve (20) configured to distribute the fuel (C) between the main loop (21) and the secondary loop (22); and - Control unit (3) for controlling the distribution valve (20) according to the speed of the aircraft turbine shaft engine (6).

2. The supply system (S) as described in claim 1, characterized in that, The first ignition branch (10), the second ignition branch (16) and the main branch (13) are respectively provided with a first ignition nozzle (12) connected to the output end of the first ignition valve (11), a second ignition nozzle (18) connected to the output end of the second ignition valve (17, 17') and a main nozzle (15) connected to the output end of the main valve (14).

3. The supply system (S) as described in claim 2, characterized in that, The first ignition nozzle (12) is a pneumatic injection type and is configured to spray the fuel (C) into the combustion chamber (5).

4. The supply system (S) as described in claim 2, characterized in that, The second ignition nozzle (18) is aerodynamically designed and configured to spray the fuel (C) into the combustion chamber (5).

5. The supply system (S) as described in claim 2, characterized in that, The main nozzle (15) is in the form of a multi-point nozzle that includes multiple injection points.

6. The supply system (S) as claimed in claim 1, characterized in that, The first ignition valve (11) is a sealing valve.

7. The supply system (S) as claimed in claim 1, characterized in that, The opening pressure (T14) of the main valve (14) is greater than the opening pressure (T11) of the first ignition valve (11) so that a set flow rate of fuel (C) is delivered only when the pressure is above a predetermined pressure in the main branch (13).

8. The supply system (S) as claimed in claim 1, characterized in that, The opening pressure (T11) of the first ignition valve (11) is substantially equal to the opening pressure (T17) of the second ignition valve (17).

9. The supply system (S) as claimed in claim 1, characterized in that, The second ignition valve (17') is a metering valve, and the second ignition branch (16) is equipped with a sealing valve (19).

10. The supply system (S) as claimed in claim 9, characterized in that, The second ignition branch (16) is provided with an anti-coking conduit (CF), which opens on one side between the second ignition valve (17') and the sealing valve (19), and on the other side upstream of the supply system (S), so that the fuel (C) can circulate through the sealing valve (19) when the second ignition valve (17') is closed.

11. An aircraft turboshaft engine (6), characterized in that, It includes a combustion chamber (5) and a supply system (S) as claimed in any one of claims 1-10, wherein the injector (1) of the supply system (S) extends into the combustion chamber (5) to supply fuel (C) to the combustion chamber (5).

12. A method for supplying fuel (C) to the combustion chamber (5) of an aircraft turboshaft engine (6) using the supply system (S) according to any one of claims 1-10, characterized in that, The method includes: - The step (E1) of controlling the distribution valve (20) via the control unit (3) according to the speed of the aircraft turbine shaft engine (6). - The step (E2) of distributing the fuel (C) between the main circuit (21) and the secondary circuit (22) via the distribution valve (20), and - Step (E3) of injecting the fuel (C) into the combustion chamber (5) through the injector (1), - Wherein, regardless of the speed at which the aircraft turboshaft engine (6) is, once the turboshaft engine (6) is started, the distribution valve (20) directs the fuel (C) into the main circuit (21) and the secondary circuit (22) during the distribution step (E2).

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