Control device for aircraft engine, control method, and aircraft engine

Abstract

To ensure stable ignition of fuel supplied to the combustor during aircraft engine startup. [Solution] The aircraft engine control device comprises a fuel supply system that supplies fuel from a fuel pump to a combustor via a fuel supply passage, and an igniter that ignites the fuel in the combustor, and includes a processing circuit for controlling the fuel supply system. The processing circuit commands the igniter to perform an ignition operation and controls the fuel supply system such that the flow rate of fuel flowing through the fuel supply passage increases over time from the time the ignition operation is commanded to the igniter until the completion of fuel ignition in the combustor is detected.

Description

Technical Field

[0001] The present disclosure relates to a control device, a control method, and a control program for an aircraft engine.

Background Art

[0002] Patent Document 1 discloses a fuel control method at the start of a gas turbine engine.

Prior Art Documents

Patent Documents

[0008] An aircraft engine control program according to one aspect of the present disclosure causes at least one processor to execute the control method. The control program may be stored in a computer-readable, non-temporary, and tangible storage medium. [Effects of the Invention]

[0009] According to one aspect of this disclosure, fuel supplied to the combustor can be stably ignited when starting an aircraft engine. [Brief explanation of the drawing]

[0010] [Figure 1] Figure 1 is a cross-sectional view of an aircraft engine according to an embodiment. [Figure 2] Figure 2 is a cross-sectional view of the combustor shown in Figure 1, viewed from the axial direction. [Figure 3] Figure 3 is a cross-sectional view taken along the line III-III in Figure 2, showing the arrangement of the fuel nozzles and igniters. [Figure 4] Figure 4 is a block diagram of the control device shown in Figure 1. [Figure 5] Figure 5 is a flowchart illustrating the startup control of the control device shown in Figure 4. [Figure 6] Figure 6 is a graph showing the change in fuel flow rate over time in the starting control shown in Figure 5. [Figure 7] Figure 7 is a cross-sectional view, equivalent to Figure 2, showing a modified fuel chamber. [Modes for carrying out the invention]

[0011] The embodiments will be described below with reference to the drawings. In the following description, the axial direction X means the direction in which the axis L of the engine rotation shaft 2 of the aircraft engine 1 extends, the radial direction Y means the direction perpendicular to the axis L, and the circumferential direction Z means the direction extending circumferentially around the axis L. Furthermore, the axial direction X is the front-rear direction of the aircraft engine 1. "Front" means the side of the aircraft engine 1 in which air is introduced from the outside in the axial direction X, and "rear" means the side of the aircraft engine 1 in which exhaust gas is discharged in the axial direction X.

[0012] Figure 1 is a cross-sectional view of an aircraft engine 1 according to an embodiment. The aircraft engine 1 is used as an engine to generate thrust for an aircraft such as an unmanned aerial vehicle. As shown in Figure 1, the aircraft engine 1 is a gas turbine engine. The aircraft engine 1 is, for example, a turbofan engine. The aircraft engine 1 comprises a rotating shaft 2, a fan 3, a compressor 4, a combustor 5, a turbine 6, and a casing 7.

[0013] Fan 3 is connected to the front of the rotating shaft 2 and rotates with the rotating shaft 2. The compressor 4, combustor 5, and turbine 6 are arranged in this order from front to rear along the rotating shaft 2. The casing 7 includes a roughly cylindrical inner shell 17 and outer shell 18 arranged concentrically with respect to each other. The inner shell 17 houses the compressor 4, combustor 5, and turbine 6. The front of the outer shell 18 houses fan 3. A cylindrical bypass passage B is formed between the inner shell 17 and the outer shell 18.

[0014] The aircraft engine 1 is, for example, a twin-shaft gas turbine engine. The rotating shaft 2 includes a low-pressure shaft 12 and a high-pressure shaft 11 which is arranged on the same axis as the low-pressure shaft 12 and is rotatable relative to the low-pressure shaft 12. The high-pressure shaft 11 is a tubular hollow shaft. The low-pressure shaft 12 is inserted through the hollow space of the high-pressure shaft 11. The low-pressure shaft 12 is longer than the high-pressure shaft 11 in the front-rear direction, and the front and rear ends of the low-pressure shaft 12 are exposed to the outside of the high-pressure shaft 11. The low-pressure shaft 12 is mechanically connected to a fan 3.

[0015] The compressor 4 includes a low-pressure compressor 13 and a high-pressure compressor 14 disposed behind the low-pressure compressor 13. The low-pressure compressor 13 is an axial-flow compressor, and the high-pressure compressor 14 is a centrifugal compressor. An outer periphery of the high-pressure compressor 14 is provided with a diffuser 8 that sends out the air flowing out from the high-pressure compressor 14 rearward. A combustor 5 is disposed behind the diffuser 8.

[0016] The combustor 5 is a reverse-flow combustor. The combustor 5 defines a combustion chamber 20 having an annular shape around an axis L. The combustor 5 is provided with a plurality of fuel nozzles 30 that inject fuel into the combustion chamber 20 facing the combustion chamber 20. The combustor 5 is provided with a plurality of igniters 40 that ignite the fuel in the combustion chamber 20 facing the combustion chamber 20. The fuel nozzles 30 and the igniters 40 are arranged side by side on a virtual plane VP orthogonal to the axis L.

[0017] The turbine 6 includes a high-pressure turbine 15 and a low-pressure turbine 16 disposed behind the high-pressure turbine 15. The low-pressure shaft 12 mechanically connects the low-pressure compressor 13 to the low-pressure turbine 16. The high-pressure shaft 11 mechanically connects the high-pressure compressor 14 to the high-pressure turbine 15.

[0018] A part of the air sucked in by the fan 3 driven by the low-pressure shaft 12 flows through the bypass passage B and is discharged rearward. The air sucked in by the fan 3 and introduced into the interior of the inner casing 17 flows into the low-pressure compressor 13. The air that has passed through the low-pressure compressor 13 and the high-pressure compressor 14 flows into the combustor 5 via the diffuser 8. The combustion gas flowing out from the outlet of the combustor 5 passes through the nozzle unit 9, passes through the high-pressure turbine 15 and the low-pressure turbine 16, and is discharged rearward.

[0019] The aircraft engine 1 includes a fuel supply system 21 that supplies fuel to the combustor 5. The fuel supply system 21 includes a fuel tank 22, a fuel pump 23, and a fuel supply line 24. The fuel supply system 21 supplies fuel from the fuel pump 23 to the combustor 5 via the fuel supply line 24. The fuel tank 22 stores fuel. The fuel tank 22 may be arranged outside the outer casing 18, or may be arranged in the fuselage of the aircraft on which the aircraft engine 1 is mounted. The fuel pump 23 is an electric pump having a pump and an electric motor for driving the pump. The fuel pump 23 sucks in the fuel stored in the fuel tank 22 and discharges it into the fuel supply line 24.

[0020] The fuel supply line 24 fluidly connects the discharge port of the fuel pump 23 to the fuel nozzle 30. A flow sensor 26 for detecting the flow rate of the fuel flowing through the fuel supply line 24 is arranged in the fuel supply line 24. Here, the flow rate means the flow rate per unit time. The aircraft engine 1 includes a rotation speed sensor 27 that is arranged opposite to the low-pressure shaft 12 and detects the rotation speed of the low-pressure shaft 12. Here, the rotation speed means the rotation speed per unit time. The rotation speed sensor 27 may detect the rotation speed of the high-pressure shaft 11 instead of detecting the rotation speed of the low-pressure shaft 12.

[0021] The aircraft engine 1 includes a control device 25 that includes a processing circuit 50. The control device 25 is connected to the flow sensor 26, the rotation speed sensor 27, the fuel pump 23, and the igniter 40 via communication lines respectively. The processing circuit 50 controls the fuel pump 23 and the igniter 40 by referring to the detection signals of the flow sensor 26 and the rotation speed sensor 27.

[0022] Figure 2 is a cross-sectional view of the combustor 5 in FIG. 1 as viewed from the axial direction X. As shown in FIG. 2, the combustor 5 has an annular combustion chamber 20 when viewed from the axial direction X. The plurality of fuel nozzles 30 are arranged at intervals in the circumferential direction Z around the axis L of the combustion chamber 20. For example, the fuel nozzles 30 include a first fuel nozzle 31, a second fuel nozzle 32, a third fuel nozzle 33, a fourth fuel nozzle 34, a fifth fuel nozzle 35, and a sixth fuel nozzle 36.

[0023] In Figure 2, the clockwise side of the circumferential direction Z is considered one side, and the counterclockwise side of the circumferential direction Z is considered the other side. The second fuel nozzle 32 is located on the other side of the circumferential direction Z of the first fuel nozzle 31. The third fuel nozzle 33 is located on the other side of the circumferential direction Z of the second fuel nozzle 32. The fourth fuel nozzle 34 is located on the other side of the circumferential direction Z of the third fuel nozzle 33. The fifth fuel nozzle 35 is located on the other side of the circumferential direction Z of the fourth fuel nozzle 34. The sixth fuel nozzle 36 is located on the other side of the circumferential direction Z of the fifth fuel nozzle 35.

[0024] Multiple igniters 40 are arranged at intervals from each other in the circumferential direction Z. The number of igniters 40 is less than the number of fuel nozzles 30. For example, the igniters 40 include a first igniter 41, a second igniter 42, and a third igniter 43. The first igniter 41 is located between the first fuel nozzle 31 and the second fuel nozzle 32 in the circumferential direction Z. The second igniter 42 is located between the third fuel nozzle 33 and the fourth fuel nozzle 34 in the circumferential direction Z. The third igniter 43 is located between the fifth fuel nozzle 35 and the sixth fuel nozzle 36 in the circumferential direction Z. No igniters are located in the region between the second fuel nozzle 32 and the third fuel nozzle 33 in the circumferential direction Z, the region between the fourth fuel nozzle 34 and the fifth fuel nozzle 35 in the circumferential direction Z, and the region between the sixth fuel nozzle 36 and the first fuel nozzle 31 in the circumferential direction Z.

[0025] Figure 3 is a cross-sectional view taken along the line III-III in Figure 2, showing the arrangement of the fuel nozzle 30 and the igniter 40. As shown in Figure 3, the fuel nozzle 30 has a discharge port 30a that discharges fuel F toward the combustion chamber 20. The igniter 40 is an ignition cartridge capable of ignition only once. The igniter 40 includes a powder containment section 40a containing a powder containment chamber S containing gunpowder P, a first outlet 40b communicating with the powder containment chamber S, a second outlet 40c communicating with the powder containment chamber S, and a heat source 40d.

[0026] The heating source 40d heats and burns the gunpowder P in the gunpowder chamber S. The heating source 40d may be a coil capable of generating sparks, or a heating wire that generates heat when energized. The heating source 40d heats and burns the gunpowder P in the gunpowder chamber S in response to a command from the processing circuit 50.

[0027] The first outlet 40b and the second outlet 40c are connected to a common explosive chamber S. In the ignition cartridge, the igniter 40, sparks generated by the burning of the explosive P are injected into the combustion chamber 20 from the first outlet 40b and the second outlet 40c for several seconds. By using explosive P, the igniter 40 can be made small and simple in configuration. The igniter 40 only needs to produce ignition heat into the combustion chamber 20, and may produce a flame or high-temperature gas instead of a spark.

[0028] The first outlet 40b is oriented to one side in the circumferential direction Z so as to emit ignition heat in a direction inclined with respect to the axial direction X toward the fuel nozzle 30 adjacent to one side in the circumferential direction Z of the igniter 40. The second outlet 40c is oriented to the other side in the circumferential direction Z so as to emit ignition heat in a direction inclined with respect to the axial direction X toward the other side in the circumferential direction Z of the igniter 40.

[0029] Specifically, the first outlet 40b of the first igniter 41 is oriented in a direction inclined with respect to the axial direction X toward the first fuel nozzle 31, which is located on one side of the first igniter 41 in the circumferential direction Z. The first outlet 40b of the first igniter 41 is directed toward the discharge region R1 of the first fuel nozzle 31 in the combustion chamber 20. The first outlet 40b of the first igniter 41 injects a spark toward the region in the combustion chamber 20 that is closer to the first fuel nozzle 31 than to the second fuel nozzle 32.

[0030] The second outlet 40c of the first igniter 41 is oriented in a direction inclined with respect to the axial direction X toward the second fuel nozzle 32, which is located on the other side of the first igniter 41 in the circumferential direction Z. The second outlet 40c of the first igniter 41 is directed toward the discharge region R2 of the second fuel nozzle 32 in the combustion chamber 20. The second outlet 40c of the first igniter 41 injects a spark toward the region in the combustion chamber 20 that is closer to the second fuel nozzle 32 than to the first fuel nozzle 31.

[0031] The first outlet 40b of the second igniter 42 is oriented in a direction inclined with respect to the axial direction X toward the third fuel nozzle 33, which is located on one side of the second igniter 42 in the circumferential direction Z. The first outlet 40b of the second igniter 42 is directed toward the discharge region R3 of the third fuel nozzle 33 in the combustion chamber 20. The first outlet 40b of the second igniter 42 injects a spark toward the region in the combustion chamber 20 that is closer to the third fuel nozzle 33 than to the fourth fuel nozzle 34.

[0032] The second outlet 40c of the second igniter 42 is oriented in a direction inclined with respect to the axial direction X toward the fourth fuel nozzle 34, which is located on the other side of the second igniter 42 in the circumferential direction Z. The second outlet 40c of the second igniter 42 is directed toward the discharge region R4 of the fourth fuel nozzle 34 in the combustion chamber 20. The second outlet 40c of the second igniter 42 injects a spark toward the region in the combustion chamber 20 that is closer to the fourth fuel nozzle 34 than to the third fuel nozzle 33.

[0033] Although not shown in Figure 3, the first outlet 40b and second outlet 40c of the third igniter 43 are the same as those of the first igniter 41 and second igniter 42 described above.

[0034] With the igniter 40 described above, a single igniter 40 can stably ignite the fuel F discharged from multiple fuel nozzles 30. Furthermore, since the igniter 40 emits a spark in a direction inclined toward the fuel nozzle 30 with respect to the axial direction X of the combustion chamber 20, it can accurately direct the spark to the fuel F discharged from the fuel nozzle 30. Therefore, the ignition of the fuel F in the combustion chamber 20 can be completed within the spark injection period of the igniter 40, which is the ignition cartridge. Thus, the aircraft engine 1 can be started quickly.

[0035] Furthermore, even without arranging an igniter between the second fuel nozzle 32 and the third fuel nozzle 33 in the circumferential direction Z, the fuel F discharged from the second fuel nozzle 32 and the third fuel nozzle 33 can be stably ignited. Therefore, the aircraft engine 1 can be started quickly with a small number of igniters 40.

[0036] Figure 4 is a block diagram of the control device 25 of Figure 1. As shown in Figure 4, the control device 25 controls the fuel pump 23 and the igniter 40 in response to an input of a start command from an external source. For example, the control device 25 includes a processor 51, system memory 52, storage memory 53, and interface 54. The processor 51 may include a CPU (Central Processing Unit). The system memory 52 may include volatile memory. The storage memory 53 may include non-volatile memory. The storage memory 53 may include a hard disk, flash memory, or a combination thereof. The storage memory 53 stores a control program 53a. An example of a processing circuit 50 is a configuration in which the processor 51 executes the control program 53a read from the storage memory 53 to the system memory 52.

[0037] Interface 54 includes an input interface connected to the flow sensor 26 and the rotational speed sensor 27, and an output interface connected to the fuel pump 23 and the igniter 40. Interface 54 receives an external input signal indicating a start command for the aircraft engine 1.

[0038] The processing circuit 50 receives a start command from an external source to instruct the aircraft engine 1 to start. The processing circuit 50 also receives a detection signal from the flow sensor 26. The processing circuit 50 controls the fuel pump 23 and the igniter 40. That is, when the processing circuit 50 receives a start command from an external source, it controls the fuel pump 23 and the igniter 40 while referring to the detection signals from the flow sensor 26 and the rotational speed sensor 27.

[0039] Figure 5 is a flowchart illustrating the start-up control of the control device 25 in Figure 4. Figure 6 is a graph showing the change in fuel flow rate over time in the start-up control of Figure 5. The vertical axis of the graph in Figure 6 represents the fuel flow rate detected by the flow sensor 26. The start-up control will be explained below following the flow shown in Figure 5, with appropriate reference to the configurations in Figures 1 to 4 and the graph in Figure 6.

[0040] When the control device 25 receives a start command to start the aircraft engine 1 in step S1, it performs pre-ignition control in step S2. Pre-ignition control is performed before commanding the igniter 40 to perform ignition. Pre-ignition control is performed to push any air present in the fuel supply passage 24 into the combustion chamber 20 of the combustor 5 with the fuel discharged by the fuel pump 23, thereby filling the fuel supply passage 24 with fuel. Pre-ignition control prevents air present in the fuel supply passage 24 from being supplied to the combustor 5 along with the fuel when the igniter 40 subsequently performs ignition, thereby improving ignition stability.

[0041] As shown in Figure 6, when the control device 25 receives a start command at time T1 as pre-ignition control, it drives the fuel pump 23 without activating the igniter 40. In pre-ignition control, the fuel pump 23 is controlled so that the fuel flow rate A4 discharged by the fuel pump 23 in pre-ignition control is greater than a predetermined ignition start flow rate A2, which is the fuel flow rate discharged by the fuel pump 23 at the start of ignition at time T3, which will be described later.

[0042] For example, in pre-ignition control, the fuel pump 23 is controlled so that the fuel flow rate A4 discharged by the fuel pump 23 in pre-ignition control is greater than the idling flow rate A3, which is the fuel flow rate discharged by the fuel pump 23 during idling, as described later. In pre-ignition control, the fuel pump 23 may also be controlled so that the fuel flow rate A4 discharged by the fuel pump 23 in pre-ignition control is greater than the maximum fuel flow rate discharged by the fuel pump 23 during aircraft flight. By increasing the fuel flow rate A4 discharged by the fuel pump 23 during pre-ignition control, the air present in the fuel supply passage 24 can be effectively pushed out by the vigorous fuel flow.

[0043] The control device 25 terminates pre-ignition control when it determines that a parameter that increases with increasing cumulative fuel flow rate from the fuel pump 23 exceeds a threshold value during pre-ignition control. The threshold value of the parameter is the value at which the amount of fuel supplied to the combustor 5 from the start time T1 of pre-ignition control exceeds the volume of the fuel supply passage 24.

[0044] For example, the control device 25 accumulates the flow rate detected by the flow sensor 26 from the start time T1 of pre-ignition control, and terminates the pre-ignition control when it determines that the accumulated value exceeds a predetermined threshold at time T2. Alternatively, the control device 25 may terminate the pre-ignition control when it determines that the elapsed time from the start time T1 of pre-ignition control exceeds a predetermined threshold. Between time T1 and time T4, air is supplied to the combustor 5 by rotating the high-pressure shaft 11 or the low-pressure shaft 12 by some means such as an auxiliary power unit (APU) or a starter motor.

[0045] Once pre-ignition control is complete, in step S3, the control device 25 controls the fuel pump 23 so that the flow rate of fuel discharged by the fuel pump 23 decreases to the ignition start flow rate A2. The ignition start flow rate A2 is a value smaller than the idling flow rate A3.

[0046] At time T3, when the fuel flow rate discharged by the fuel pump 23 reaches the ignition start flow rate A2, in step S4, the control device 25 commands the igniter 40 to start ignition. In response to the ignition command, the igniter 40 causes the heat source 40d to heat the gunpowder P in the gunpowder chamber S. As a result, the gunpowder P in the gunpowder chamber S burns, and sparks are ejected from the first outlet 40b and the second outlet 40c of the igniter 40 for several seconds.

[0047] In step S5, the control device 25 controls the fuel pump 23 so that the flow rate of fuel flowing through the fuel supply passage 24 gradually increases over time, from the ignition start time T3, when the ignition operation is commanded to the igniter 40, to the ignition completion time T4, when the completion of fuel ignition in the combustor 5 is detected. In the range from the ignition start time T3 to the ignition completion time T4, the control device 25 gradually increases the flow rate of fuel flowing through the fuel supply passage 24 while keeping the amount of air supplied to the combustor 5 constant, thereby gradually increasing the fuel concentration in the combustion chamber 20. In this way, the fuel in the combustion chamber 20 can be ignited at a low concentration within the ignitionable concentration range, and the temperature of the combustor 5 is prevented from becoming too high.

[0048] The control device 25 controls the fuel pump 23 so that the rate of increase in the discharge volume of the fuel pump 23 is smaller than the maximum capacity value of the rate of increase in the discharge volume of the fuel pump 23 during the period from the start of ignition T3 to the completion of ignition T4. For example, the rate of increase in the flow rate of fuel discharged by the fuel pump 23 during the period from the start of ignition T3 to the completion of ignition T4 is smaller than the rate of increase in the flow rate of fuel discharged by the fuel pump 23 at the start of pre-ignition control T1.

[0049] In the range from the start of ignition T3 to the completion of ignition T4, the flow rate of fuel discharged by the fuel pump 23 may increase linearly, curvely, in a stepped manner, or at an irregularly fluctuating rate of increase.

[0050] As described above, since the fuel flow rate gradually increases during the ignition operation of the igniter 40, ignition failure due to fuel shortage is prevented, and the fuel supplied to the combustor 5 can be stably ignited when starting the aircraft engine 1.

[0051] In step S6, the control device 25 determines whether ignition of the fuel in the combustion chamber 20 has been completed. The control device 25 may determine that ignition is complete when the rate of increase of the rotational speed detected by the rotational speed sensor 27 exceeds a predetermined threshold. The control device 25 may also determine that ignition is complete when the rotational speed detected by the rotational speed sensor 27 exceeds a predetermined threshold. If the aircraft engine 1 is equipped with a temperature sensor that detects the temperature of the combustor 5 or the downstream side of the combustor 5, the control device 25 may determine that ignition is complete when the temperature detected by the temperature sensor exceeds a predetermined threshold.

[0052] When ignition is determined to be complete at time T4, in step S7, the control device 25 controls the fuel pump 23 so that the flow rate of fuel discharged by the fuel pump 23 decreases to a predetermined initial flow rate A1. The initial flow rate A1 is a value smaller than the ignition start flow rate A2.

[0053] At time T5, when the fuel flow rate discharged by the fuel pump 23 reaches the initial flow rate A1, in step S8, the control device 25 gradually increases the fuel flow rate discharged by the fuel pump 23 to the idling flow rate A3. The aircraft engine 1 is in an idling state when it has started and the aircraft remains stationary. The aircraft is also in an idling state when it has taken off with rapid acceleration and is cruising after the climb has finished. The idling state is the state in which the flow rate of fuel supplied from the fuel pump 23 to the combustor 5 is the minimum amount required to maintain the operating state of the aircraft engine 1, that is, the state in which the rotating shaft 2 is rotating. After entering the idling state, the control device 25 controls the fuel pump 23 in accordance with the output demands for the aircraft engine 1.

[0054] Figure 2 illustrates a configuration in which the combustor 5 is equipped with six fuel nozzles 30 and three igniters 40, but the configuration is not limited to this. Figure 7 is a cross-sectional view corresponding to Figure 2, showing a modified combustor 105. As shown in Figure 7, in the combustor 105, multiple fuel nozzles 30 are arranged at intervals from each other in the circumferential direction Z around the axis L of the combustion chamber 20. The number of fuel nozzles 30 in the combustor 105 is eight. For example, the fuel nozzles 30 include a first fuel nozzle 31, a second fuel nozzle 32, a third fuel nozzle 33, a fourth fuel nozzle 34, a fifth fuel nozzle 35, a sixth fuel nozzle 36, a seventh fuel nozzle 37, and an eighth fuel nozzle 38.

[0055] Multiple igniters 40 are arranged in the combustor 105. There are two igniters 40 in the combustor 105. The igniters 40 include a first igniter 41 and a second igniter 42. The first igniter 41 and the second igniter 42 are arranged point-symmetrically with respect to the axis L when viewed from the axial direction X. The first igniter 41 is located between the second fuel nozzle 32 and the third fuel nozzle 33 in the circumferential direction Z. The second igniter 42 is located between the sixth fuel nozzle 36 and the seventh fuel nozzle 37 in the circumferential direction Z. No igniters are located in the following regions in the circumferential direction Z: between the first fuel nozzle 31 and the second fuel nozzle 32, between the third fuel nozzle 33 and the fourth fuel nozzle 34, between the fourth fuel nozzle 34 and the fifth fuel nozzle 35, between the fifth fuel nozzle 35 and the sixth fuel nozzle 36, between the seventh fuel nozzle 37 and the eighth fuel nozzle 38, and between the eighth fuel nozzle 38 and the first fuel nozzle 31.

[0056] The first outlet 40b of the first igniter 41 injects a spark towards the fuel discharged from the second fuel nozzle 32, which is adjacent to one side of the first igniter 41 in the circumferential direction Z. When the fuel discharged from the second fuel nozzle 32 is ignited and a flame is generated, that flame ignites the fuel discharged from the first fuel nozzle 31. The second outlet 40c of the first igniter 41 injects a spark towards the fuel discharged from the third fuel nozzle 33, which is adjacent to the other side of the first igniter 41 in the circumferential direction Z. When the fuel discharged from the third fuel nozzle 33 is ignited and a flame is generated, that flame ignites the fuel discharged from the fourth fuel nozzle 34. In this way, the first igniter 41 has the role of igniting the fuel from the four fuel nozzles 31, 32, 33, and 34. Note that the ignition by the second igniter 42 is based on the same principle as ignition by the first igniter, so a detailed explanation is omitted.

[0057] It should be noted that the technology of this disclosure is not limited to the embodiments described above. For example, the controlled objects that constitute the fuel supply system 21 and are controlled by the control device 25 are not limited to the electric motor of the fuel pump 23. For example, the fuel supply system 21 may include a fuel supply valve. In that case, the control device 25 may control the fuel supply valve to adjust the opening degree of the fuel supply valve in order to adjust the flow rate of fuel supplied to the combustor 5. If the fuel pump 23 is a pressurizer that applies pressure to push fuel into a bag-shaped fuel tank 22, the control device 25 may control the pressurizer.

[0058] The flow sensor 26 may be omitted, and the fuel flow rate may be calculated from the rotational speed of the fuel pump 23. The rotational speed of the fuel pump 23 may be detected by a rotational speed sensor for the fuel pump 23, or it may be estimated from the command value output from the control device 25 to the electric fuel pump 23, the current value flowing through the fuel pump 23, etc. If the fuel pump 23 is a positive displacement pump such as a gear pump, it becomes easier to detect the fuel flow rate from the rotational speed of the fuel pump 23.

[0059] The igniter 40 may ignite the gunpowder P in the gunpowder chamber S by applying a physical shock to the gunpowder P instead of a heat source 40d. The igniter 40 may be positioned between two adjacent fuel nozzles 30 in the circumferential direction Z, and offset radially Y from the two fuel nozzles 30. The gunpowder chamber 40a of the igniter 40 may include a first gunpowder chamber communicating with a first outlet 40b and a second gunpowder chamber communicating with a second outlet 40c.

[0060] As described above, the embodiments have been explained as examples of the technology disclosed in this application. However, the technology in this disclosure is not limited to the embodiments described above and can be applied to embodiments that have been modified, replaced, added, or omitted as appropriate. Furthermore, it is possible to combine the components described in the embodiments to create new embodiments. For example, some components or methods in one embodiment may be applied to other embodiments, and some components in an embodiment can be separated from other components in that embodiment and extracted as appropriate. In addition, the components described in the attached drawings and detailed description include not only components that are essential for solving the problem, but also components that are not essential for solving the problem, in order to illustrate the technology.

[0061] The functions of the elements disclosed herein can be performed using circuits or processing circuits, including general-purpose processors, dedicated processors, integrated circuits, ASICs (Application Specific Integrated Circuits), FPGAs (Field Programmable Gate Arrays), conventional circuits, and / or combinations thereof, configured or programmed to perform the disclosed functions. A processor is considered a processing circuit or circuit because it includes transistors and other circuits. In this disclosure, a circuit, unit, or means is hardware that performs the enumerated functions, or hardware programmed to perform the enumerated functions. The hardware may be hardware disclosed herein, or other known hardware that is programmed or configured to perform the enumerated functions. If the hardware is a processor, which is considered a type of circuit, the circuit, means, or unit is a combination of hardware and software, and the software is used to configure the hardware and / or the processor.

[0062] [Aspect] The embodiments described above are specific examples of the following embodiments.

[0063] (Aspect 1) A control device for an aircraft engine comprising a fuel supply system that supplies fuel from a fuel pump to a combustor via a fuel supply passage, and an igniter that ignites the fuel in the combustor, The system includes a processing circuit that controls the fuel supply system, and the processing circuit is The igniter is instructed to perform an ignition operation, A control device for an aircraft engine, which controls the fuel supply system such that the flow rate of fuel flowing through the fuel supply line increases over time, from the time when the ignition operation is commanded to the ignition device until the completion of fuel ignition is detected in the combustor.

[0064] According to Embodiment 1, the fuel flow rate gradually increases during the ignition operation of the igniter, thus preventing ignition failure due to fuel shortage. Therefore, the fuel supplied to the combustor can be stably ignited when starting the aircraft engine.

[0065] (Aspect 2) The control device for an aircraft engine according to embodiment 1, wherein the processing circuit performs pre-ignition control, which drives the fuel pump before commanding the ignition operation to the igniter in order to push out any air that may be present in the fuel supply line.

[0066] According to embodiment 2, before the ignition operation of the igniter begins, air present in the fuel supply line is pushed into the combustor by the fuel flow. Therefore, when the igniter performs the ignition operation, it is prevented that air present in the fuel supply line is supplied to the combustor along with the fuel. Thus, ignition stability can be further improved.

[0067] (Aspect 3) The processing circuit performs pre-ignition control, which drives the fuel pump for a predetermined period of time before commanding the ignition operation to the igniter. The control device for an aircraft engine according to embodiment 1 or 2, wherein in the pre-ignition control, the fuel supply system is controlled so that the amount of fuel supplied to the combustor is greater than the volume of the fuel supply passage.

[0068] According to embodiment 3, any air that may be present in the fuel supply line can be effectively pushed into the combustor.

[0069] (Aspect 4) The processing circuit performs pre-ignition control to drive the fuel pump before commanding the ignition operation to the igniter. A control device for an aircraft engine according to any one of embodiments 1 to 3, wherein in the pre-ignition control, the fuel supply system is controlled such that the flow rate in the pre-ignition control is greater than the flow rate at the time the ignition operation is commanded to the igniter.

[0070] According to embodiment 4, the strong fuel flow makes it easier to push out the air present in the fuel supply passage.

[0071] (Aspect 5) The processing circuit performs pre-ignition control to drive the fuel pump before commanding the ignition operation to the igniter. The control device for an aircraft engine according to any one of embodiments 1 to 4, wherein in the pre-ignition control, the fuel supply system is controlled so that the flow rate in the pre-ignition control is greater than the maximum flow rate during idling of the aircraft engine or the maximum flow rate during flight of the aircraft engine.

[0072] According to embodiment 5, the strong fuel flow makes it easier to push out the air present in the fuel supply passage.

[0073] (Aspect 6) The control device for an aircraft engine according to any one of embodiments 1 to 5, wherein the fuel pump is an electric pump.

[0074] According to embodiment 6, fuel can be freely supplied to the combustor regardless of the engine's operating state.

[0075] (Aspect 7) The control device for an aircraft engine according to any one of embodiments 1 to 6, wherein the igniter is an ignition cartridge including a powder storage section containing explosives, a first outlet communicating with the powder storage section, and a second outlet communicating with the powder storage section.

[0076] According to embodiment 7, when using an ignition cartridge in which the ignition heat generated by the burning of gunpowder is emitted into the combustion chamber from the first and second outlets for several seconds, the ignition of the fuel can be completed within the output period of the ignition heat.

[0077] (Pattern 8) A control method for an aircraft engine comprising a fuel supply system that supplies fuel from a fuel pump to a combustor via a fuel supply passage, and an igniter that ignites the fuel in the combustor, The igniter is instructed to perform an ignition operation, A control method for an aircraft engine, comprising controlling the fuel supply system such that the flow rate of fuel flowing through the fuel supply line increases over time, from the time when the ignition operation is commanded to the ignition device until the completion of fuel ignition in the combustor is detected.

[0078] (Aspect 9) An aircraft engine control program that causes at least one processor to execute the control method described in Embodiment 8. [Explanation of symbols]

[0079] 1. Aircraft engine 5 Combustor 20 Combustion chamber 21 Fuel supply system 23 Fuel pump 24 Fuel supply path 25 Control device 30 Fuel Nozzles 31. No. 1 Fuel Nozzle 32. Second fuel nozzle 33 Third fuel nozzle 34. Fuel No. 4 No. 4 40 Igniter 40a Explosives storage section 40b Exit 1 40c 2nd exit 41 1st igniter 42 Second igniter 50 Processing Circuits 51 processors 53a Control Program F fuel L axis P Explosives R1, R2, R3, R4 discharge area S Explosives storage room VP virtual plane X-axis direction Y radial direction Z circumferential direction

Claims

1. A control device for an aircraft engine comprising a fuel supply system that supplies fuel from a fuel pump to a combustor via a fuel supply passage, and an igniter that ignites the fuel in the combustor, The system includes a processing circuit that controls the fuel supply system, and the processing circuit is The igniter is instructed to perform an ignition operation, From the time when the ignition operation is commanded to the ignition device until the completion of fuel ignition is detected in the combustor, the fuel supply system is controlled so that the flow rate of fuel flowing through the fuel supply path increases over time. A control device for an aircraft engine, which performs pre-ignition control by driving the fuel pump before commanding the ignition operation to the igniter in order to expel any air that may be present in the fuel supply line.

2. The aircraft engine control device according to claim 1, wherein in the pre-ignition control, the fuel supply system is controlled so that the amount of fuel supplied to the combustor is greater than the volume of the fuel supply passage.

3. The control device for an aircraft engine according to claim 1, wherein in the pre-ignition control, the fuel supply system is controlled such that the flow rate in the pre-ignition control is greater than the flow rate at the time the ignition operation is commanded to the igniter.

4. The control device for an aircraft engine according to claim 1, wherein in the pre-ignition control, the fuel supply system is controlled so that the flow rate in the pre-ignition control is greater than the maximum flow rate during idling of the aircraft engine or the maximum flow rate during flight of the aircraft engine.

5. The control device for an aircraft engine according to claim 1, wherein the fuel pump is an electric pump.

6. The control device for an aircraft engine according to any one of claims 1 to 5, wherein the igniter is an ignition cartridge including a powder storage section containing explosives, a first outlet communicating with the powder storage section, and a second outlet communicating with the powder storage section.

7. A control method for an aircraft engine comprising a fuel supply system that supplies fuel from a fuel pump to a combustor via a fuel supply passage, and an igniter that ignites the fuel in the combustor, The igniter is instructed to perform an ignition operation, From the time when the ignition operation is commanded to the ignition device until the completion of fuel ignition in the combustor is detected, the fuel supply system is controlled so that the flow rate of fuel flowing through the fuel supply path increases over time. A control method for an aircraft engine, comprising performing pre-ignition control to drive the fuel pump before commanding the ignition operation to the igniter in order to expel any air that may be present in the fuel supply line.

8. One or more processors, One or more storage media, A fuel supply system that supplies fuel from the fuel pump to the combustor via a fuel supply line, The combustor is further equipped with an igniter for igniting fuel, An aircraft engine, wherein the one or more storage mediums store a program which controls the fuel supply system such that the flow rate of fuel flowing through the fuel supply line increases over time from the time the ignition operation is commanded to the igniter until the completion of fuel ignition in the combustor is detected, and which drives the fuel pump before commanding the ignition operation to the igniter to expel any air present in the fuel supply line.

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