Patent Application
20250187727
This application claims the benefit of Korean Patent Application No. 10-2023-0175392, filed on Dec. 6, 2023, which application is hereby incorporated herein by reference in its entirety.
The present disclosure relates to a flying vehicle having an auxiliary thruster and a method for controlling the auxiliary thruster.
A variety of emergency landing devices that can be used during an emergency of a flying vehicle have been proposed, and such emergency landing devices are intended to save the lives of passengers in the flying vehicle that is inoperable or face the risk of fire and explosion.
For example, a combination of parachutes and airbags deployed for an emergency landing of a flying vehicle is known. Properly deploying parachutes and airbags may ensure a gentle descent to the ground and attenuate an impact load generated when colliding with the ground as much as possible. This technology is being studied by reducing falling speed with a parachute and controlling a sink rate with an airbag in order to lower a passenger injury rate.
However, despite increased safety obtained when an emergency landing device in which the parachute and the airbag are combined are used in a flying vehicle, risks may remain. The reason is that due to the fact that the flying vehicle is inoperable and has fuel or other combustible materials, passengers must perform difficult and dangerous emergency landings along with the flying vehicle with a significant risk of explosion during an emergency landing process or during a ground collision.
The present disclosure relates to a flying vehicle having an auxiliary thruster and a method for controlling the auxiliary thruster. Particular embodiments relate to a flying vehicle having an auxiliary thruster that may protect passengers inside by reducing an impact load through auxiliary thrust force obtained using fuel from a fuel cell during an emergency and a method for controlling the auxiliary thruster, for example, in a flying vehicle using a fuel cell.
An embodiment of the present disclosure provides a flying vehicle having an auxiliary thruster that may protect passengers inside by reducing an impact load through auxiliary thrust force obtained using fuel from a fuel cell during an emergency and a method for controlling the auxiliary thruster, for example, in a flying vehicle using a fuel cell.
According to an embodiment of the present disclosure, a flying vehicle may include a fuel cell installed in a fuselage and providing power to the fuselage and an auxiliary thruster installed in the fuselage and providing auxiliary thrust force by ejecting combustion gas generated by burning at least a portion of fuel to be supplied to the fuel cell.
The fuselage may be provided with at least one rotor, and the fuel cell may provide thrust force to the fuselage by applying power to a drive motor of the rotor.
The fuel cell may include a fuel tank provided in the fuselage and configured to store the fuel, a compressor configured to pressurize air received from an air inlet of the fuselage to a predetermined pressure, and a fuel cell stack configured to electrochemically react the fuel with oxygen in the air to generate electricity.
The auxiliary thruster may include a combustion chamber configured to receive the fuel from the fuel tank and to receive air from the compressor, a nozzle connected to the combustion chamber to eject the combustion gas generated in the combustion chamber, and a plurality of vanes disposed at regular intervals from each other in the nozzle and configured to vary a posture thereof to control an ejection direction of the combustion gas.
A first fuel supply line may be connected between the fuel tank and the fuel cell stack, a second fuel supply line may be connected between the first fuel supply line and the combustion chamber, a first valve may be disposed between the first fuel supply line and the second fuel supply line, a first air supply line may be connected between the compressor and the fuel cell stack, a second air supply line may be connected between the first air supply line and the combustion chamber, and a second valve may be disposed between the first air supply line and the second air supply line.
The flying vehicle may further include a fuel pump installed in the second fuel supply line and configured to control a mass flow rate of the fuel supplied to the combustion chamber.
The flying vehicle may further include a controller configured to control an operation of the auxiliary thruster during an emergency and a sensing unit including a plurality of sensors electrically connected to the controller.
The sensing unit may include a sensor configured to sense a posture of the fuselage and a sensor configured to measure an altitude of the flying vehicle.
The controller may be configured to calculate a current sink rate from a current altitude value of the flying vehicle input from the sensing unit, compare the current sink rate with a predetermined necessary sink rate during an emergency, and calculate the auxiliary thrust force required for the vehicle from the difference between the current sink rate and the required sink rate.
The controller may be configured to determine an angle of the vane from a current posture of the flying vehicle input from the sensing unit and control an operation of a driver connected to the vane to control the ejection direction of the combustion gas discharged from the nozzle.
The ejection direction of the combustion gas may be maintained to have a right angle to the ground.
At least one of the auxiliary thrusters may be coupled to a lower portion of the fuselage, and the auxiliary thruster may be at least partially embedded in the fuselage.
The fuel may be hydrogen or a hydrogen compound.
According to another embodiment of the present disclosure, a method for controlling an auxiliary thruster may include measuring a current sink rate of a flying vehicle during an emergency of the flying vehicle, when the current sink rate is greater than or equal to a required sink rate, calculating auxiliary thrust force for reducing a sink rate, and commanding combustion gas to be generated and ejected by supplying at least a portion of fuel to be supplied to a fuel cell of the flying vehicle to an auxiliary thruster and burning the at least a portion of fuel.
Control of the auxiliary thruster may be performed when the flying vehicle is in a predetermined altitude.
Control of the auxiliary thruster may be terminated when the current sink rate is less than the required sink rate.
In the commanding, a mass flow rate of the fuel for obtaining the auxiliary thrust force may be determined, and an operation of a fuel pump disposed between a fuel tank of the fuel cell and a combustion chamber of the auxiliary thruster may be controlled to supply the fuel at a determined mass flow rate.
The method for controlling an auxiliary thruster may further include controlling an ejection direction of the combustion gas, and the ejection direction of the combustion gas may be maintained in a direction perpendicular to the ground.
According to an embodiment of the present disclosure, an impact load may be reduced by auxiliary thrust force obtained by ejecting combustion gas generated by burning fuel of a fuel cell into a nozzle, thereby improving the safety of passengers even during an emergency.
The above and other aspects, features, and advantages of embodiments of the present disclosure will be more clearly understood from the following detailed description, taken in conjunction with the accompanying drawings, in which:
Hereinafter, embodiments of the present disclosure will be described in detail with reference to the accompanying drawings. In adding reference numerals to elements of each of the drawings, although the same elements are illustrated in other drawings, like reference numerals may refer to like elements.
In the present specification, the terms such as “first,” “second,” or “third” may be used to describe various elements, but these elements are not limited in order, size, position, or importance by terms such as “first,” “second,” and “third,” and these terms may be used for distinguishing one component from other components.
First, in the present specification, a flying vehicle 1 may refer to a vehicle configured to fly and move over the air. In other words, the flying vehicle may refer to a rotorcraft, a drone, a tilt rotor aircraft, a vertical take-off and landing aircraft, a fixed-wing aircraft, and the like, and may also include a vehicle that may move on the ground using wheels and fly with the wheels separated from the ground.
Furthermore, the flying vehicle may also include a manned flying vehicle and an unmanned flying vehicle. The manned flying vehicle may include a fuselage that can operate by autonomous flight in addition to a fuselage controlled by a pilot.
For example, the flying vehicle 1 may include a fuselage 2 in which a boarding space is arranged, wings 3 mounted on the fuselage, and a plurality of rotors 4 provided on the wings 3. The plurality of rotors 4 may be rotated by respective drive motors thereof.
Some of the plurality of rotors 4 may be formed as tilting rotors capable of tilting upward or forward for lifting or cruising the fuselage 2, and the others thereof may be formed as lifting rotors for lifting the fuselage 2.
The rotors 4 may be configured so that at least two rotors may be disposed on the left and right wings 3 with respect to a center of the fuselage 2, respectively, but the position and number of rotors are not necessarily limited thereto. Additionally, in a wingless rotorcraft such as a helicopter, one rotor may be disposed on the fuselage.
The fuselage 2 may accommodate multiple passengers, and the fuselage 2 may allow vertical takeoff and landing and horizontal cruising due to a tilting rotor.
The fuel cell 10 may be installed in the fuselage 2 and may provide power to the fuselage 2. Furthermore, by applying power to the drive motor from the fuel cell 10, thrust force may be supplied to the fuselage 2 by the plurality of rotors 4.
The fuel cell 10 may electrochemically react fuel, such as hydrogen or a hydrogen compound, with oxygen in the air, and may directly convert chemical energy of fuel into electrical energy. Accordingly, the flying vehicle 1 according to an embodiment of the present disclosure may be an electric thrusting vehicle that is propelled by generating electricity through the fuel cell 10 during flight and applying power to the drive motor of the rotor 4.
More specifically, for example, in an anode or a fuel electrode of the fuel cell 10 using hydrogen and oxygen, hydrogen is decomposed into two hydrogen ions and two electrons. The electrons move through a conductor to a cathode or an air electrode, and hydrogen ions pass through an electrolyte and move to the cathode. In the cathode, the moved hydrogen ions, electrons, and oxygen react to generate liquid water, heat, and electrical energy. The fuel cell 10 uses such a movement of electrons as power.
For this purpose, the fuel cell 10 may include a fuel tank 11 provided in the fuselage 2 and formed as a pressure vessel to store fuel and a compressor 12 pressurizing the air received from an air inlet 5 of the fuselage 2 to a predetermined pressure.
High-purity fuel compressed and stored in the fuel tank 11 may be depressurized by a regulator 13 and may be supplied to a fuel cell stack 15 through a first fuel supply line FL1. Optionally, fuel remaining after being consumed for power generation of the fuel cell stack may be returned to the first fuel supply line FL1 and may then be provided back to the fuel cell stack 15.
Air may be supplied to the fuel cell stack 15 through the compressor 12 and the first air supply line OL1. The air remaining after being consumed in the power generation of the fuel cell stack may be discharged to the outside air.
The fuel cell stack 15 may be assembled by stacking a plurality of unit cells. Each unit cell may be provided with an electrolyte membrane that can move hydrogen ions and an anode and a cathode each sintered on both surfaces of the electrolyte membrane. Furthermore, the fuel cell stack 15 may further include a gas-impermeable separator having a flow path for providing gas, that is, hydrogen and oxygen, to each electrode.
For example, direct current electricity generated by the fuel cell stack 15 may be converted to alternating current electricity by an inverter and then output, and it may rotate the rotor 4 by driving the drive motor. Accordingly, thrust force may be supplied to the fuselage 2.
However, embodiments of the present disclosure are not necessarily limited thereto, and when a direct current motor is adopted as the drive motor, direct current electricity generated by the fuel cell stack 15 may be output without an inverter, and the rotor 4 may be rotated by driving the drive motor.
Optionally, the flying vehicle 1 according to an embodiment of the present disclosure may further include a secondary battery (not illustrated), and when the drive motor is not driven, electricity generated by the fuel cell stack 15 may be supplied to the secondary battery and stored. Moreover, as an application of stable power is required, a DC/DC converter may be adopted when voltage is boosted or decreased.
The auxiliary thruster 20 may provide auxiliary thrust force to the fuselage 2 by burning and ejecting the fuel supplied to the fuel cell 10. For this purpose, the auxiliary thruster may include a combustion chamber 21, a nozzle 22, and a plurality of vanes 23, and it may be configured and operated independently of a main thruster including a drive motor and a rotor 4.
On the other hand, the auxiliary thruster 20 may be connected to the first fuel supply line FL1 and the fuel tank 11 constituting the fuel cell 10 through a second fuel supply line FL2, and it may be connected to the first air supply line OL1 and the compressor 12 constituting the fuel cell 10 through a second air supply line OL2.
A first valve 25 may be disposed between the first fuel supply line FL1 and the second fuel supply line FL2, and a second valve 26 may be disposed between the first air supply line OL1 and the second air supply line OL2. In this case, a three-way valve may be adopted as the first valve 25 and the second valve 26, respectively.
A first port of the first valve 25 may be connected to the fuel tank 11, a second port may be connected to the fuel cell stack 15, and a third port may be connected to the combustion chamber 21 of the auxiliary thruster 20.
Furthermore, the first port of the second valve 26 may be connected to the compressor 12, the second port may be connected to the fuel cell stack 15, and the third port may be connected to the combustion chamber 21 of the auxiliary thruster 20.
When the first port and the third port communicate with each other in the first valve 25 and the first port and the third port communicate with each other in the second valve 26, in the combustion chamber 21, fuel supplied through the fuel tank 11 and the fuel supply lines FL1 and FL2 may be burned with an oxidizing agent supplied through the compressor 12 and the air supply lines OL1 and OL2, that is, oxygen in the air. The combustion chamber 21 may include an igniter 24 for igniting fuel and oxygen.
Additionally, a fuel pump 27 for adjusting the amount (mass flow rate) of fuel supplied to the combustion chamber 21 may be installed in the second fuel supply line FL2.
The combustion chamber 21 and the nozzle 22 configured in this manner may act like a kind of rocket. In other words, since the high-temperature and high-pressure combustion gas generated by combustion in the combustion chamber 21 has the characteristic of attempting to escape toward a lower pressure side, they are ejected outward through the nozzle 22, and as a reaction thereto, the fuselage 2 receives auxiliary thrust force for elevating the fuselage 2.
The plurality of vanes 23 may be disposed in the nozzle 22 at regular intervals from each other. Each of the vanes 23 may have a variable angle with respect to an axis or hypotenuse of the nozzle 22. To this end, a rotation axis of each vane 23 may be connected to a driver 28. The driver 28 may include a motor and a gear unit, and optionally, a bevel gear, etc. may be added to the gear unit as needed.
For example, the detailed configuration and operation of the vane 23 and the driver 28 applicable to the flying vehicle 1 according to an embodiment of the present disclosure are described in “Missile Jet Vane Control System and Method,” published in U.S. Pat. No. 5,806,791. This patent document may be incorporated and cited herein for reference.
In this manner, as the nozzle 22 is fixed and a posture of the plurality of vanes 23 arranged in the nozzle 22 is configured to vary, the effect of suppressing an increase in a weight of the flying vehicle 1 may be expected.
At least one auxiliary thruster 20 having the combustion chamber 21 and the nozzle 22 may be coupled to a lower portion of the fuselage 2, and each auxiliary thruster may be at least partially embedded in the fuselage 2. Each auxiliary thruster may protrude from the fuselage 2 to the minimum possible level.
The flying vehicle according to an embodiment of the present disclosure may include a controller 30 configured to control an operation of the auxiliary thruster 20 during an emergency and a sensing unit 31 equipped with a plurality of sensors and electrically connected to the controller 30. Here, power may be applied to the controller 30 and the sensing unit 31 from a separate battery 32.
First, the sensing unit 31 may monitor a posture of the fuselage 2 using, for example, an inertial navigation system (INS) installed within the fuselage 2.
The INS is a device that provides navigation information such as the position, speed, and posture of the flying vehicle 1 by integrating acceleration and angular velocity measured through sensors. The sensing unit 31 may calculate a roll angle, a pitch angle, and a yaw angle from an accelerometer output and a gyroscope output of the INS, and it may monitor the posture of the flying vehicle 1 by continuously sensing changes in these angles.
Additionally, the sensing unit 31 may measure an altitude of the flying vehicle 1 using at least one of an altimeter, an accelerometer, and the global positioning system (GPS) installed within the fuselage 2. The controller 30 may calculate a sink rate of the flying vehicle 1 using an altitude value input from the sensing unit 31. The sink rate may be defined as an altitude value that changes per unit time (e.g., 1 second). The sink rate may be required to calculate auxiliary thrust force.
Furthermore, the sensing unit 31 may optionally further include a flow sensor and a pressure sensor installed in the combustion chamber 21 of the auxiliary thruster 20.
The controller 30 may be implemented with various processing devices such as a microprocessor embedded with a semiconductor chip capable of performing various operations or commands, and it may control an overall operation of the auxiliary thruster 20, thus providing the auxiliary thrust force required for the flying vehicle 1 during an emergency.
For example, the controller 30 may be integrated into an upper control system of the flying vehicle 1 or may be used concurrently. Additionally, the controller 30 may be electrically connected to various sensors of the sensing unit 31.
Specifically, the controller 30 independently may determine the failure or falling of the flying vehicle 1, or it may generate an alarm according to a determination result input from the upper control system.
For example, based on an inoperable failure signal input from the drive motor for the rotor 4 of the flying vehicle 1, or a falling signal input from an altimeter, an accelerometer, GPS, INS, and the like, of the sensing unit 31 in consideration of a sink rate according to weight, the controller 30 may determine that the flying vehicle 1 is in an emergency situation.
The controller 30 may be electrically connected to an alarm unit provided in the fuselage 2. This alarm unit may generate a visual and/or auditory alarm to display and inform passengers in the fuselage of an emergency situation in the flying vehicle 1.
Accordingly, the controller 30 may command an operation of the first valve 25 and the second valve 26, the fuel pump 27, the igniter 24, and a motor in the driver 28 of the vane 23, for operation of the auxiliary thruster 20 in an emergency situation.
Additionally, the controller 30 may calculate a current sink rate from a current altitude value of the flying vehicle 1 input from an altimeter, an accelerometer, and GPS of the sensing unit 31. Next, the controller 30 may compare a current sink rate with a predetermined required sink rate that the flying vehicle 1 can handle during an emergency, and from a difference between the sink rates, the controller 30 may calculate the auxiliary thrust force required for the flying vehicle 1 for gentle falling. Moreover, the controller 30 may determine the amount of fuel and air for obtaining the required auxiliary thrust force.
The controller 30 may control an operation of the first valve 25 and the second valve 26, it may supply fuel from the fuel tank 11 and the first fuel supply line FL1 constituting the fuel cell 10 to the combustion chamber 21 of the auxiliary thruster 20 via the second fuel supply line FL2, and it may supply air from the compressor 12 and the first air supply line OL1 constituting the fuel cell to the combustion chamber 21 via the second air supply line OL2. Specifically, the controller 30 may control an operation of the fuel pump 27 and the compressor 12 to adjust the amount of fuel and oxygen in the air.
The controller 30 may be electrically connected to the igniter 24 of the combustion chamber 21 to control an operation thereof, and it may allow the igniter 24 in the combustion chamber 21 to ignite the fuel and oxygen, thus allowing for an occurrence of combustion. Optionally, the controller 30 may feedback control an operation of the fuel pump 27 and the compressor 12 until the flow rate and pressure of combustion gas in the combustion chamber 21 sensed by the flow sensor and the pressure sensor of the sensing unit 31 reach a target value.
Furthermore, the controller 30 may determine an angle of the vane 23 in the nozzle 22 from a current posture of the flying vehicle 1 input from the INS of the sensing unit 31, and it may control an ejection direction of combustion gas discharged from the nozzle 22 by controlling the operation of the motor in the driver 28 connected to the vane 23. The ejection direction of the combustion gas may be approximately perpendicular to the ground.
Additionally, the controller 30 may include a timer or be electrically connected to the timer to obtain time data from the timer.
The flying vehicle 1 according to an embodiment of the present disclosure, configured as described above, may drive the first valve 25 during a normal operation to allow fuel, that is, hydrogen or a hydrogen compound, to flow into the fuel cell stack 15 of the fuel cell 10, and it may drive the second valve 26 together with the compressor 12 to allow oxygen in the air to flow into the fuel cell stack 15 of the fuel cell 10.
In this manner, the flying vehicle 1 according to an embodiment of the present disclosure may generate electricity by the fuel cell 10 during flight and may allow power to be applied to the drive motor of the rotor 4, thereby providing thrust force and enabling navigation.
For example, when the flying vehicle 1 according to an embodiment of the present disclosure falls rather rapidly due to reasons such as failure, under the control of the controller 30, a flow path may be changed by the first valve 25 to allow the fuel to flow into the combustion chamber 21 of the auxiliary thruster 20, and the flow path may be changed by the second valve 26 to allow oxygen in the air to flow into the combustion chamber of the auxiliary thruster 20.
The igniter 24 in the combustion chamber 21 may ignite fuel and oxygen to generate combustion, the high-temperature and high-pressure combustion gas generated by combustion may be ejected into the outside air through the nozzle 22, and as a reaction thereto, the fuselage 2 may receive auxiliary thrust force that elevates the fuselage and alleviates rapid descent.
In the case of an emergency, when the flying vehicle 1 receives the auxiliary thrust force, this helps to maintain operating speed and an altitude, and furthermore, there is an advantage of being able to extend the flight time necessary to select a safe landing site.
In this manner, the flying vehicle 1 according to an embodiment of the present disclosure may ensure a gentle descent to the ground by the auxiliary thrust force that elevates the fuselage 2, and it may maximally attenuate an impact load generated in the event of a collision with the ground, thereby increasing a chance of a smooth and safe landing and protecting passengers in the flying vehicle 1.
The method for controlling the auxiliary thruster 20 according to an embodiment of the present disclosure may include an operation of measuring a current sink rate of a flying vehicle 1 during an emergency (S10), an operation of calculating auxiliary thrust force for reducing a sink rate when the current sink rate is greater than or equal to a required sink rate (S30), and an operation of commanding combustion gas to be generated and ejected by supplying at least a portion of fuel to be supplied to a fuel cell 10 of the flying vehicle 1 to an auxiliary thruster 20 and burning the at least a portion of fuel (S40).
The controller 30 itself may determine the failure or fall of the flying vehicle 1, and it may generate an alarm informing an emergency situation of the flying vehicle according to a determination result input from an upper control system (S00).
For example, based on an inoperable failure signal input from the drive motor for the rotor 4 of the flying vehicle 1 or a falling signal input from an altimeter, an accelerometer, GPS, INS, and the like of the sensing unit in consideration of a sink rate according to weight, the controller 30 may determine that the flying vehicle 1 is in an emergency situation.
Accordingly, the controller 30 may drive an alarm unit, and the alarm unit may generate a visual and/or audible alarm to inform passengers in the fuselage 2 of an emergency situation of the flying vehicle 1.
Actually, the control of the auxiliary thruster 20 according to an embodiment of the present disclosure may be performed when an altitude value input from the altimeter, the accelerometer, the GPS, and the like of the sensing unit 31 in an emergency situation of the flying vehicle 1 is in, for example, at least 100 m.
In an emergency of the flying vehicle 1, the controller 30 may measure the current sink rate of the flying vehicle (S10). The sensing unit 31 may measure the altitude of the flying vehicle 1 using at least one of the altimeter, the accelerometer, and the GPS installed in the fuselage 2. The controller 30 may calculate a sink rate of the fuselage 2 using an altitude value input from the sensing unit 31. The sink rate may be defined as an altitude value that changes per unit time (e.g., 1 second).
Then, the controller 30 may compare the current sink rate with a predetermined required sink rate that the flying vehicle 1 can handle during an emergency (S20).
When the current sink rate is greater than or equal to the required sink rate, the controller 30 may calculate the auxiliary thrust force required by the flying vehicle 1 so as to induce a gentle descent by reducing the sink rate (S30). The auxiliary thrust force determined by the difference between the current sink rate and the required sink rate may be obtained by Equation 1 below.
Here, T is auxiliary thrust force (N), m is a mass (kg) of the flying vehicle 1 including passenger(s), VZ1 is a current sink rate (m/s), VZ2 is a required sink rate (m/s), and t is the time(s). The required sink rate VZ2 may be determined in advance as a predetermined value, and it may be, for example, approximately 13 m/s.
When the current sink rate is less than the required sink rate, the controller 30 may terminate the control of the auxiliary thruster 20 for reducing the sink rate.
Next, the controller 30 may supply at least a portion of the fuel to be supplied to the fuel cell 10 of the flying vehicle 1 to the auxiliary thruster 20 and burn the at least a portion of the fuel, thus generating and ejecting combustion gas to provide the auxiliary thrust force (S40).
To this end, the controller 30 may transmit a control command to the first valve 25 and the second valve 26, and it may specifically transmit the control command to the fuel pump 27 and the compressor 12, thus controlling the amount (mass flow rate) of fuel and air.
Furthermore, the controller 30 may determine a mass flow rate of oxygen in the air and the fuel for obtaining the required auxiliary thrust force.
The auxiliary driving force may be expressed by Equation 2 below.
Here, {dot over (m)} is a mass flow rate (kg/s) of the combustion gas, Ve is the nozzle outlet speed (m/s) of the combustion gas, Pe is the nozzle outlet pressure (bar) of the combustion gas, Pa is atmospheric pressure (about 1 bar), and Ae is a nozzle outlet area (m2).
For example, when fuel, oxygen, and a pressure ratio (pressure in the combustion chamber) are determined, these variables may be obtained through the Chemical Equilibrium with Applications (CEA) provided by NASA in the United States. Furthermore, when only an area of a nozzle throat is determined in addition to these variables, the mass flow rate of the combustion gas may also be determined.
Accordingly, the controller 30 may determine and adjust the mass flow rate of oxygen in the air and fuel so as to form the mass flow rate of the combustion gas for obtaining the required auxiliary thrust force.
For example, an appropriate mass flow of fuel and oxygen corresponding to the required auxiliary thrust force may be determined by being extracted from a database obtained in advance numerically or experimentally, along with a mixing ratio (F/A) which is a mass ratio of fuel to oxygen. Then, the controller may control an operation of the fuel pump 27 and the compressor 12 and control the amount (mass flow) of fuel and air by a determined value.
Furthermore, the controller 30 may transmit a control command for controlling an operation of the igniter 24 in the combustion chamber 21 to the igniter 24, and it may allow the igniter 24 in the combustion chamber 21 to ignite the fuel and oxygen, thus allowing for an occurrence of combustion. Optionally, the controller 30 may feedback control an operation of the fuel pump 27 and the compressor 12 until the flow rate and pressure of combustion gas in the combustion chamber sensed by the flow sensor and the pressure sensor of the sensing unit 31 reach a target value.
Additionally, the controller 30 may determine an angle of the vane 23 in the nozzle 22 from a current posture of the flying vehicle 1 input from the INS, and it may control an operation of the motor in the driver 28 connected to the vane to control the ejection direction of the combustion gas discharged from the nozzle (S50). A posture of each vane 23 may be changed so that the ejection direction of the combustion gas is maintained in a direction substantially perpendicular to the ground.
In this manner, the flying vehicle according to an embodiment of the present disclosure may ensure the gentle descent to the ground by the auxiliary thrust force for elevating the fuselage, and it may maximally attenuate the impact load generated in the event of a collision with the ground, thereby increasing a chance of a smooth and safe landing and protecting passengers in the flying vehicle. Accordingly, embodiments of the present disclosure may have the effect of improving the safety of passengers even during an emergency.
The aforementioned description merely illustrates the technical concept of the present disclosure, and a person skilled in the art to which the present invention pertains may make various modifications and modifications without departing from the essential characteristics of the present disclosure.
Therefore, the example embodiments disclosed in this specification and drawings are not intended to limit but to explain the technical concept of the present disclosure, and the scope of the technical idea of the present disclosure is not limited by these example embodiments. The scope of protection of the present disclosure should be interpreted by the following claims, and all technical ideas within the scope equivalent thereto should be interpreted as being included in the scope of the present disclosure.
| Number | Date | Country | Kind |
|---|---|---|---|
| 10-2023-0175392 | Dec 2023 | KR | national |
