A flight propeller safety control system and method

Abstract

The application belongs to the technical field of aircraft overall design, and is a flight propeller safety control system and method, which comprises a flight control computer, an ignition controller, an unlocking and ignition device, an airborne battery and a propelling device. The propelling device and the ignition controller are provided with a first ignition power supply circuit and a second ignition power supply circuit arranged in parallel, the ignition power supply circuits have two groups and are respectively connected with two groups of safety state switching actuators, the first ignition power supply circuit and the second ignition power supply circuit are both provided with mechanical safety pins fixed on the aircraft hanger, forming ignition circuit switches; after the aircraft is released from the aircraft, the mechanical safety pins are separated from the aircraft hanger, and the first power supply ignition circuit and the second power supply ignition circuit are simultaneously connected. After the separation switch and the inertial navigation assembly respectively send the "separated" instruction and the flight attitude information and send the ignition controller, the program automatic ignition is carried out, so that the stability of ignition is effectively realized.

Description

Technical Field

[0001] This application belongs to the field of aircraft overall design technology, and specifically relates to a flight propulsion safety control system and method. Background Technology

[0002] In the aviation field, thrusters are widely used for launching combat equipment, offering advantages such as simple structure, high thrust, and ease of use and maintenance. In recent years, small technology demonstrator aircraft equipped with thrusters have seen increasingly widespread application in supersonic and hypersonic technology research. Small aircraft using thrusters may take off in the following ways: First, in combination with a turbine engine, using the turbine engine for takeoff and the thruster for flight; second, using a large carrier aircraft to carry (boost) the small aircraft to high altitude for release (separation), with the thruster igniting in the air to propel the aircraft in continuous flight.

[0003] For air-launched missile systems like those launched from carrier aircraft, the propulsion system and onboard systems must possess extremely high safety standards to prevent accidental ignition during ground transport and carrier aircraft flight. Currently, the common approach for airborne attack equipment is as follows: First, the attack equipment system is powered down during ground transport and flight. Before launch, the carrier aircraft controls the system to power on and activate it, minimizing the powered flight time. Second, after launch, timing control methods can be used to ensure separation before ignition. This approach is widely used and prevents accidental triggering on the ground; however, even after the missile system is powered on, there is still a possibility of accidental triggering before launch.

[0004] Therefore, how to prevent small aircraft with propulsion systems from accidentally igniting their engines during takeoff is a problem that needs to be solved. Summary of the Invention

[0005] The purpose of this application is to provide a flight propulsion safety control system and method to solve the problem of accidental ignition in small aircraft with propulsion systems during takeoff.

[0006] The technical solution of this application is: a flight propulsion safety control system, including a flight control computer, an ignition controller, an unlocking and ignition device, an airborne battery, and a propulsion device; the flight control computer is electrically connected to the ignition controller, and the flight control computer is also electrically connected to an inertial navigation system and a disconnect switch; the airborne battery is simultaneously electrically connected to and supplies power to the inertial navigation system, the flight control computer, and the ignition controller; the unlocking and ignition device includes a safety status switch, a safety status switching actuator, and an ignition power-on circuit; the safety status switch is electrically connected to the flight control computer; a first ignition power supply circuit and a second ignition power supply circuit are arranged in parallel between the propulsion device and the ignition controller; there are two sets of ignition power-on circuits, each connected to one of the two sets of safety status switching actuators; both the first and second ignition power supply circuits are provided with mechanical safety pins fixed on the aircraft pylon, forming an ignition circuit switch; and the two sets of ignition power-on circuits are electrically connected to different igniters on the propulsion device.

[0007] When the aircraft is in the mounting state, the mechanical safety pin is inserted into the carrier pylon, controlling the first and second ignition power supply circuits to be in the open state; when the aircraft is released from the carrier, the mechanical safety pin separates from the carrier pylon, and the first and second power supply ignition circuits are simultaneously connected; the ignition controller can supply power to the unlocking and ignition devices, and the safety status switch can collect the status parameters of the two sets of safety status switching actuators and send them to the flight control computer in real time.

[0008] Preferably, the onboard battery is a constant voltage 28V.

[0009] As one specific implementation method, a flight thruster safety control method includes the following steps:

[0010] Real-time judgment of aircraft status; when the aircraft is in the hang-on state, keep the mechanical safety pin inserted into the aircraft pylon, and keep both ignition circuit switches in the open state at the same time, controlling the first ignition power supply circuit and the second ignition power supply circuit to be in the disconnected state.

[0011] After the aircraft is released from the carrier, the mechanical safety pin automatically separates from the carrier rack, the two ignition circuit switches are simultaneously in the off state, and the first ignition power supply circuit and the second ignition power supply circuit are simultaneously connected.

[0012] After the mechanical safety pin separates from the aircraft pylon, the separation switch is triggered, and the flight control computer receives the "separated" command with low latency, forming the necessary condition 1 for the "propeller ignition" command;

[0013] After the aircraft is released from the carrier, the flight control computer receives the aircraft attitude information transmitted from the inertial navigation system in real time, which is a necessary condition for forming the 'thrust ignition' command.

[0014] When necessary conditions 1 and 2 are met simultaneously, the flight control computer sends a "unlock" command to the ignition controller. The ignition controller supplies power to the fuse state switching actuator until both ignition power circuits are connected simultaneously. After the ignition power circuits are connected, the fuse state switch sends a "unlocked" message back to the flight control computer, forming necessary condition 3 for the "thruster ignition" command.

[0015] After receiving the 'separation' command, the flight control computer starts timing. The predetermined flight time after separation is a necessary condition for generating the 'thrust ignition' command.

[0016] After all the necessary conditions 1-4 for thruster ignition are met, the flight control computer sends an "ignition" command to the ignition controller. The ignition controller continuously outputs constant voltage and current to the two sets of ignition circuits, and the two sets of igniters detonate and ignite the thrusters under the action of the current.

[0017] Preferably, when necessary conditions 1 and 2 are met simultaneously, the ignition controller continuously supplies power to the safety state switching actuator for 1.5s; after necessary conditions 1-4 for thruster ignition are met, the ignition controller continuously outputs constant voltage and current to the two sets of ignition power circuits for a duration of not less than 0.2s.

[0018] Preferably, before the aircraft separates from the propulsion unit, the flight control computer detects the safety status signal in real time to determine whether an "unlocked" command has been issued. If so, the flight control computer immediately sends a "activate safety" command to the ignition controller.

[0019] Preferably, after the aircraft separates from the thruster, the inertial navigation system can acquire the aircraft's acceleration information in real time and use the aircraft's acceleration information to determine whether the thruster is working. If it is not working, the flight control computer repeatedly issues the 'unlock' and 'ignition' commands in sequence. If it is still not working, the flight control computer continuously issues the 'lock' command to the ignition controller and causes the aircraft to enter the return and recovery process.

[0020] Preferably, after the igniter detonates, the ignition controller collects information from the ignition power circuit and determines in real time whether a short circuit has occurred in the ignition power circuit. If so, the ignition controller is automatically disconnected.

[0021] This application discloses a flight propulsion safety control system and method, including a flight control computer, an ignition controller, an unlocking and ignition device, an onboard battery, and a propulsion device. A first ignition power supply circuit and a second ignition power supply circuit are arranged in parallel between the propulsion device and the ignition controller. There are two sets of ignition power supply circuits, each connected to a safety state switching actuator. Both the first and second ignition power supply circuits are equipped with mechanical safety pins fixed to the aircraft pylon, forming an ignition circuit switch. When the aircraft is launched from the aircraft pylon, the mechanical safety pins separate from the pylon, and the first and second power supply ignition circuits are simultaneously activated. After the separation switch and the inertial navigation system (INS) issue a "separated" command and flight attitude information to the ignition controller, automatic ignition is performed, effectively achieving ignition stability. Attached Figure Description

[0022] To more clearly illustrate the technical solutions provided in this application, the accompanying drawings will be briefly described below. Obviously, the drawings described below are merely some embodiments of this application.

[0023] Figure 1 This is a schematic diagram of the overall structure of this application.

[0024] 1. Flight control computer; 2. Ignition controller; 3. Unlocking and ignition device; 4. Onboard battery; 5. Propulsion system; 6. Inertial navigation system; 7. Disconnect switch; 8. Safety switch; 9. Safety switch actuator; 10. Mechanical safety pin; 11. First ignition power supply circuit; 12. Second ignition power supply circuit; 13. Ignition power supply circuit; 14. Ignition device. Detailed Implementation

[0025] The technical solutions of the embodiments of the present invention will be clearly and completely described below with reference to the accompanying drawings. Obviously, the described embodiments are only some embodiments of the present invention, and not all embodiments. Based on the embodiments of the present invention, all other embodiments obtained by those skilled in the art without creative effort are within the scope of protection of the present invention.

[0026] A flight propulsion safety control system, such as Figure 1 As shown, it includes a flight control computer 1, an ignition controller 2, an unlocking and ignition device 3, an onboard battery 4, and a propulsion device 5.

[0027] Flight control computer 1 is electrically connected to ignition controller 2. Inertial navigation system 6 and disconnect switch 7 are also electrically connected to flight control computer 1. Onboard battery 4 is electrically connected to and supplies power to inertial navigation system 6, flight control computer 1 and ignition controller 2. Unlocking and ignition device 3 includes safety status switch 8, safety status switching actuator 9 and ignition power circuit 13. Safety status switch 8 is electrically connected to flight control computer 1.

[0028] The propulsion device 5 and the ignition controller 2 are provided with a first ignition power supply circuit 11 and a second ignition power supply circuit 12 arranged in parallel. There are two sets of ignition power supply circuits 13, which are respectively connected to two sets of safety state switching actuators 9. The first ignition power supply circuit 11 and the second ignition power supply circuit 12 are each provided with a mechanical safety pin 10 fixed on the carrier frame to form an ignition circuit switch. The two sets of ignition power supply circuits 13 are respectively electrically connected to different igniters 14 on the propulsion device 5.

[0029] When the aircraft is in the flight-mounted state, it is mounted on the carrier aircraft, which is the carrier aircraft. The mechanical safety pin 10 is inserted into the carrier aircraft pylon, controlling the first ignition power supply circuit 11 and the second ignition power supply circuit 12 to be in the open state. When the aircraft is released from the propulsion device 5, the mechanical safety pin 10 is separated from the carrier aircraft pylon, and the first and second power supply ignition circuits are simultaneously connected. The ignition controller 2 can supply power to the unlocking and ignition device 3, and the safety status switch 8 can collect the status parameters of the two sets of safety status switching actuators 9 and send them to the flight control computer 1 in real time.

[0030] By using a mechanical safety pin 10 as the ignition circuit switch for the first and second power supply ignition circuits, and taking the aircraft's loaded flight state and the release from the propulsion device 5 as the triggering node for the switch, it is possible to accurately determine whether the aircraft has been released from the propulsion device 5.

[0031] In this way, when the aircraft is in flight with the load on, the possibility of accidental ignition is completely prevented by judging the state of the mechanical safety pin 10 in the first and second power supply ignition circuits.

[0032] After receiving the status of the aircraft being released from the propulsion device 5, the separation switch 7 and the inertial navigation component 6 respectively issue a "separated" command and flight attitude information; after receiving the "separated" command and flight attitude information, the ignition controller 2 controls the ignition power circuit 13 to ignite, and performs thruster ignition at the predetermined flight time after separation. In this way, the automatic ignition is achieved through multiple safety measures, thereby effectively realizing the stability of ignition.

[0033] Furthermore, by employing two sets of ignition circuits controlled by dual switches, even if a minor abnormality occurs in one ignition circuit, the system can still stably achieve automatic ignition by adjusting the power distribution, thereby improving reliability and safety.

[0034] Preferably, the onboard battery 4 is a constant voltage of 28V to ensure stable power supply.

[0035] As one specific implementation, a flight propulsion safety control method is also included, which specifically includes the following steps:

[0036] Step S100: Real-time judgment of the aircraft status. When the aircraft is in the hang-up state, keep the mechanical safety pin 10 inserted into the aircraft hangar, and keep both ignition circuit switches in the open state at the same time to control the first ignition power supply circuit 11 and the second ignition power supply circuit 12 to be in the disconnected state.

[0037] In step S200, after the aircraft is released from the carrier, the mechanical safety pin 10 automatically separates from the carrier rack, the two ignition circuit switches are simultaneously turned off, and the first ignition power supply circuit 11 and the second ignition power supply circuit 12 are simultaneously turned on.

[0038] In step S300, after the mechanical safety pin 10 separates from the aircraft pylon, the separation switch 7 is triggered, and the flight control computer 1 receives the "separated" command with low delay, forming the necessary condition 1 for the "thruster ignition" command;

[0039] Step S400: After the aircraft is released from the carrier aircraft, the flight control computer 1 receives the aircraft attitude information transmitted from the inertial navigation component 6 in real time, forming the necessary condition 2 for the 'thrust ignition' command.

[0040] In step S500, when necessary conditions 1 and 2 are met simultaneously, the flight control computer 1 sends a "unlock" command to the ignition controller 2. The ignition controller 2 supplies power to the fuse state switching actuator 9 until both sets of ignition power-on circuits 13 are connected simultaneously. After the ignition power-on circuit 13 is connected, the fuse state switch 8 sends "unlocked" back to the flight control computer 1, forming necessary condition 3 for the "thruster ignition" command.

[0041] In step S600, after receiving the 'separation' command, the flight control computer 1 starts timing. The predetermined flight time after separation is a necessary condition 4 for generating the 'thrust ignition' command.

[0042] In step S700, after all the necessary conditions 1-4 for thruster ignition are met, the flight control computer 1 sends an "ignition" command to the ignition controller 2. The ignition controller 2 continuously outputs constant voltage and current to the two sets of ignition power circuits 13, and the two sets of igniters 14 detonate and ignite the thruster under the action of the current.

[0043] By setting four necessary conditions as ignition requirements, the thruster cannot be ignited if any one of the necessary conditions is not met, thereby effectively improving safety and reliability.

[0044] Preferably, when necessary conditions 1 and 2 are simultaneously met, the ignition controller 2 continuously supplies power to the safety state switching actuator 9 for 1.5 seconds to ensure sufficient power supply. After necessary conditions 1-4 for thruster ignition are all met, the ignition controller 2 continuously outputs a constant voltage and current to the two sets of ignition power circuits 13 for a duration of not less than 0.2 seconds to ensure efficient ignition.

[0045] Preferably, before the aircraft separates from the propulsion unit, the flight control computer 1 detects the safety status signal in real time to determine whether an "unlocked" command has appeared. This command indicates that the mechanical safety pin 10 has disengaged from the fuselage, which means that an abnormality has occurred. If so, the flight control computer 1 immediately sends a "safety engaged" command to the ignition controller 2, thereby effectively preventing the propulsion unit from igniting accidentally.

[0046] Preferably, after the aircraft separates from the thruster, the inertial navigation system 6 can acquire the aircraft's acceleration information in real time and determine whether the thruster is working based on the aircraft's acceleration information. If it is not working, the flight control computer 1 repeatedly issues the 'disarm' and 'ignition' commands sequentially. If it is still not working, the flight control computer 1 continuously issues the 'arm' command to the ignition controller 2 and initiates the return and recovery process. By detecting the thruster's working status through a dual-determination method, the ignition process can be stopped in time when the thruster malfunctions, preventing further losses.

[0047] Preferably, after the igniter 14 is detonated, the ignition controller 2 collects information from the ignition power circuit 13 and determines in real time whether a short circuit has occurred in the ignition power circuit 13. If so, the ignition controller 2 is automatically disconnected to prevent further damage after the short circuit.

[0048] Finally, the following points should be noted: First, in the description of this application, it should be noted that, unless otherwise specified and limited, the terms "installation", "connection", and "linkage" should be interpreted broadly, and can be mechanical or electrical connections, or internal connections between two components, or direct connections. "Up", "down", "left", "right", etc. are only used to indicate relative positional relationships. When the absolute position of the described object changes, the relative positional relationship may change.

[0049] Secondly: The accompanying drawings of the embodiments disclosed in this invention only involve the structures involved in the embodiments disclosed in this invention. Other structures can refer to the general design. In the absence of conflict, the same embodiment and different embodiments of this invention can be combined with each other.

[0050] In conclusion, the above description is only a preferred embodiment of the present invention and is not intended to limit the present invention. Any modifications, equivalent substitutions, improvements, etc., made within the spirit and principles of the present invention should be included within the protection scope of the present invention.

Claims

1. A flight propulsion safety control system, characterized in that: The system includes a flight control computer (1), an ignition controller (2), an unlocking and ignition device (3), an onboard battery (4), and a propulsion device (5). The flight control computer (1) is electrically connected to the ignition controller (2). The flight control computer (1) is also electrically connected to an inertial navigation system (6) and a disconnect switch (7). The onboard battery (4) is electrically connected to and supplies power to the inertial navigation system (6), the flight control computer (1), and the ignition controller (2). The unlocking and ignition device (3) includes a safety status switch (8), a safety status switching actuator (9), and an ignition power-on circuit (13). (8) Electrically connected to the flight control computer (1), the propulsion device (5) and the ignition controller (2) are provided with a first ignition power supply circuit (11) and a second ignition power supply circuit (12) arranged in parallel, the ignition power supply circuit (13) has two sets and is respectively connected to two sets of safety state switching actuators (9), the first ignition power supply circuit (11) and the second ignition power supply circuit (12) are both provided with mechanical safety pins (10) fixed on the carrier hangar to form an ignition circuit switch, and the two sets of ignition power supply circuits (13) are respectively electrically connected to different igniters (14) on the propulsion device (5); When the aircraft is in the hang-on state, the mechanical safety pin (10) is inserted into the carrier rack, controlling the first ignition power supply circuit (11) and the second ignition power supply circuit (12) to be in the disconnected state; when the aircraft is released from the carrier, the mechanical safety pin (10) is separated from the carrier rack, and the first power supply ignition circuit and the second power supply ignition circuit are simultaneously connected; the ignition controller (2) can supply power to the unlocking and ignition device (3), and the safety status switch (8) can collect the status parameters of the two sets of safety status switching actuators (9) and send them to the flight control computer (1) in real time.

2. The flight propulsion safety control system as described in claim 1, characterized in that: The onboard battery (4) is a constant voltage 28V.

3. A safety control method for a flight propulsion system, characterized in that, Includes the following steps: The aircraft status is judged in real time. When the aircraft is in the hang-up state, the mechanical safety pin (10) is kept inserted into the aircraft hangar and the two ignition circuit switches are in the open state at the same time, controlling the first ignition power supply circuit (11) and the second ignition power supply circuit (12) to be in the disconnected state. When the aircraft is released from the carrier, the mechanical safety pin (10) automatically separates from the carrier rack, and the two ignition circuit switches are simultaneously turned off, while the first ignition power supply circuit (11) and the second ignition power supply circuit (12) are simultaneously turned on. After the mechanical safety pin (10) separates from the aircraft pylon, the separation switch (7) is triggered, and the flight control computer (1) receives the "separated" command with low delay, forming the necessary condition 1 for the "propeller ignition" command; After the aircraft is released from the carrier, the flight control computer (1) receives the aircraft attitude information transmitted by the inertial navigation component (6) in real time, forming the necessary condition 2 for the 'propeller ignition' command; When necessary conditions 1 and 2 are met simultaneously, the flight control computer (1) sends a "unlock" command to the ignition controller (2), and the ignition controller (2) supplies power to the fuse state switching actuator (9) until the two sets of ignition power circuits (13) are connected at the same time. After the ignition power circuit (13) is connected, the fuse state switch (8) sends "unlocked" back to the flight control computer (1), forming necessary condition 3 for the 'thruster ignition' command. After receiving the 'separation' command, the flight control computer (1) starts timing. The predetermined flight time after separation is a necessary condition for generating the 'thrust ignition' command. After all the necessary conditions 1-4 for thruster ignition are met, the flight control computer (1) sends an "ignition" command to the ignition controller (2). The ignition controller (2) continuously outputs constant voltage and current to the two sets of ignition power circuits (13), and the two sets of igniters (14) ignite and ignite the thruster under the action of the current.

4. The flight propulsion safety control method as described in claim 3, characterized in that: When necessary conditions 1 and 2 are met simultaneously, the ignition controller (2) continuously supplies power to the safety state switching actuator (9) for 1.5s; after necessary conditions 1-4 for thruster ignition are met, the ignition controller (2) continuously outputs constant voltage and current to the two sets of ignition power circuits (13) for a duration of not less than 0.2s.

5. The flight propulsion safety control method as described in claim 3, characterized in that: Before the aircraft separates from the propulsion system, the flight control computer (1) detects the safety status signal in real time and determines whether an "unlocked" command has appeared. If so, the flight control computer (1) immediately sends a "safety engaged" command to the ignition controller (2).

6. The flight propulsion safety control method as described in claim 3, characterized in that: After the aircraft separates from the thruster, the inertial navigation component (6) can acquire the aircraft acceleration information in real time and determine whether the thruster is working through the aircraft acceleration information. If it is not working, the flight control computer (1) repeatedly issues the 'unlock' and 'ignition' commands in sequence. If it is still not working, the flight control computer (1) continuously issues the 'lock' command to the ignition controller (2) and causes the aircraft to enter the return and recovery process.

7. The flight propulsion safety control method as described in claim 3, characterized in that: After the igniter (14) is detonated, the ignition controller (2) collects information from the ignition power circuit (13) and determines in real time whether there is a short circuit in the ignition power circuit (13). If so, the ignition controller (2) is automatically cut off.

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