WO2023202448A1

WO2023202448A1 - Distributed hybrid aircraft

Info

Publication number: WO2023202448A1
Application number: PCT/CN2023/087928
Authority: WO
Inventors: 向杰
Original assignee: 向杰
Priority date: 2022-04-22
Filing date: 2023-04-12
Publication date: 2023-10-26

Abstract

A distributed hybrid aircraft: the hybrid aircraft comprises distributed flight units (109) and distributed first power supply apparatuses (110, 201), and further comprises a second power supply apparatus (107, 204), etc. A configuration scheme comprises: the second power supply apparatus comprises an electric power source which has a high capacity and high energy density, such as power supply in the form of hydrogen power, internal combustion engine power and the like; the second power supply apparatus serves as the preferred power source to provide endurance main power energy; the first power supply apparatuses comprise small-capacity and large-current rechargeable batteries; and the first power supply apparatuses serve as standby power sources and provide electric energy requirements in emergency states such as rapid acceleration, emergency steering and preferred power source failure. The hybrid aircraft further comprises an operation mode scheme, an electric energy regulation scheme, a charging scheme and the like, wherein a threshold is set to enable the aircraft to perform operation mode switching; an electric energy regulation mode is run to regulate battery consumption in each first power supply apparatus; and the charging scheme is run to promptly restore battery levels in each first power supply apparatus. On the basis of the characteristics of the present hybrid scheme, the present invention further relates to a detailed aircraft construction scheme, for example, comprising a precursor fan (502), an inclinable unit, a leading edge lift structure (503), a suspended wing (504), etc.

Description

A distributed hybrid aircraft

Technical field

This patent relates to the field of aircraft, mainly to a distributed hybrid aircraft.

Background technique

In recent years, electric-driven aircraft have developed faster and faster. Compared with fuel-powered aircraft, electric-driven aircraft have many advantages, including faster response speed and better environmental protection. At the same time, because electric energy transmission is very flexible and convenient, it can be easily Enable distributed electric drive for greater safety. Since the current battery energy density is not high, simply using distributed batteries to drive an aircraft can easily lead to short endurance, and battery management is also difficult. If centralized power generation equipment is used as the power source, once the power source fails, the aircraft will be lost. It will go on strike and is not very safe. Currently, we use the characteristics of aircraft electric drives to carry out directional designs for designated usage scenarios or specific needs, and can obtain a variety of new aircraft products that meet various application needs.

Technical solutions

In response to the above situation, this patent proposes a distributed hybrid electric drive aircraft solution.

A basic hybrid solution is:

The aircraft is equipped with multiple distributed flight units and multiple first power supply devices, where the flight unit includes one or two electric-driven lift fans; the first power supply device includes a rechargeable battery and a power controller; each first power supply device, Through the first power supply cable, it is connected to at least one adjacent flight unit, and the connected power supply drives the lift fan therein to rotate; each flight unit is connected to an adjacent first power supply device, which provides power.

A second power supply device is installed, which contains a power source, and the power source form includes but is not limited to: hydrogen-to-electricity conversion equipment, or fuel-to-electricity conversion equipment, or rechargeable batteries; the second power supply device, through the second power supply cable, is connected to each The second power supply device is connected to a first power supply device to provide power; the second power supply device is installed near the center of the aircraft.

The power controller can obtain control information and associated power information, where the control information includes control instructions for the power controller and/or speed management instructions for the associated lift fan.

The power controller manages the associated power according to the power information and control information. The management actions include at least actions A and B among actions A, B, C, D, and E, among which:

Action A: Charge the rechargeable battery in the first power supply device with the electric energy connected to the second power supply cable.

Action B: Output the electric energy of the rechargeable battery in the first power supply device to the connected lift fan.

Action C: Directly output the power connected to the second power supply cable to the connected lift fan.

Action D: Charge the electric energy connected to the second power supply cable to the rechargeable battery in the corresponding first power supply device, and output it to the connected lift fan at the same time.

Action E: Output the electric energy connected to the second power supply cable and the electric energy of the rechargeable battery to the connected lift fan at the same time.

A solution for the distributed multi-channel design of the second power supply cable is: the first power supply device has two or three second power supply cables, which are directly connected to input power; all the first power supply devices obtain power from the second power supply cable in a parallel manner electricity.

An optimization method is: the transmission voltage of the second power supply cable is higher than the transmission voltage of the first power supply cable. The power controller includes a voltage reduction function for reducing the high voltage of the second power supply cable. The reduced voltage value is close to The voltage value of the battery in the first power supply device is either close to the voltage value required for charging the battery in the first power supply device, or close to the standard voltage value required to drive the lift fan; and/or the power controller further includes It has a boost function, which is used to boost the power of the rechargeable battery and output it. The boosted voltage value is close to the standard voltage value required to drive the lift fan. When the ratio of the transmission voltage of the second power supply cable to the transmission voltage of the first power supply cable is not less than 3, the performance optimization of the aircraft is obviously reflected. The voltage is close, which can be understood as being as equal as possible, or that circuits with two voltage values can operate normally after being directly connected, or that the two voltage values are equal after a loss value is reserved.

Another optimization method is: the second power supply device and the second power supply cable are designed with reference to the standard power requirements of the lift fan connected to the power supply; the output power of the first power supply device and the transmission power of the first power supply cable are designed with reference to the lift fan of the connected power supply. The maximum power requirement of the fan is designed; the battery output power of the first power supply device is designed with reference to the maximum power requirement of the lift fan connected to the power supply, or with reference to the difference between the maximum power and the standard power. When the ratio of the maximum power of the lift fan to the standard power is no less than 2, the aircraft performance is significantly optimized.

One way to drive a lift fan is: the electric energy controller can obtain the rotation speed management value of the associated lift fan, adjust the value based on the control information or the associated electric energy information, and then send it to the associated lift fan.

The aircraft can contain one operating mode: A operating mode or B operating mode, specifically:

A operation mode: the second power supply device is the preferred power supply, providing power to the lift fan through the power supply link; the battery power of the first power supply device is the backup power supply, providing supplementary power to the lift fan in case of temporary emergency needs, and Provides full power to the lift fan in the event of a preferred power failure;

B operation mode: The battery in the first power supply device is the primary power supply to provide power to the lift fan; the second power supply device is a backup power supply. In case of temporary emergency needs, it provides supplementary power to the lift fan through the power supply link, or provides power to the charging fan. Charges the battery and provides full power to the lift fan in the event of a primary power source failure.

One specific form of the temporary emergency demand is that when the speed management value of the lift fan of the aircraft is greater than the specified value, it is a temporary emergency demand; when it is less than or equal to the specified value, it is a normal flight demand. The specified value refers to the standard power value setting.

For the above two operating modes, the recommended power configuration solutions are:

Including operating mode A, the ratio of the sum of the maximum battery power in all first power supply devices to the maximum available power of the second power supply device is 5%-35%; or including operating mode B, the maximum available power of the second power supply device is The ratio of power supply to the sum of the maximum battery power in all first power supply devices is 5%-35%. The maximum power supply refers to the power supply when the power source is at full resource load. For example, when the power source is a fuel-to-electricity conversion equipment, the power that can be generated when the fuel volume is 100%. The maximum battery power refers to the power of the battery when it is fully charged.

Based on the above two operating modes, one combination solution is: including both operating modes A and B. The sum of the maximum battery power in all first power supply devices is less than the maximum power supply of the second power supply device. When detecting When the power supply of the second power supply device is higher than the set threshold, the A operation mode is adopted. When the power supply is lower than the set threshold, the B operation mode is adopted.

The aircraft may also include a power regulation mode, that is, when the battery power in a certain first power supply device is lower than the average power level setting value, the A or B power regulation mode is operated, where:

A power regulation mode: maintain the output current of the power supply device, reduce the battery output current to the lift fan, increase the current obtained from the second power supply cable to the lift fan; or increase the current from the second power supply cable to charge the battery.

B power regulation mode: reduce the output current of the power supply device, thereby reducing the lift of the connected lift fan, and increase the output current of other first power supply devices, causing the increase of the lift of other connected lift fans to compensate.

Based on the above two power adjustment modes, one combination scheme is to include both power adjustment modes A and B. When it is detected that the power supply of the second power supply device is higher than the set threshold, power adjustment mode A is adopted; When it is lower than the set threshold, the B power regulation mode is adopted.

The aircraft may also include a charging management method. The steps are: detect the output power of the second power supply device; when the output power is less than the standard power, detect the transmission power of the second power supply cable one by one; when the transmission power is less than the standard power, detect the transmission power one by one. Detect the rechargeable battery in the first power supply device connected to the second power supply cable; when the voltage of the detected rechargeable battery is lower than the set value and is not in a charged state, find the battery with the lowest voltage among them; the second power supply The device charges the battery with the lowest voltage through the power supply link.

An electric energy information processing method is: each first power supply device can independently communicate with each other and exchange electric energy information, and then obtain the electric energy information of all first power supply devices, perform electric power calculation and analysis, and complete its own electric energy management actions.

An additional power information processing method includes: each first power supply device can obtain the power supply information of the second power supply device.

A specific aircraft solution based on the hybrid energy solution is: the flight units are arranged in a honeycomb array; a lift fan is placed at the center of the honeycomb unit, and there are 2 or 3 angular distances between the center position and the peripheral angular positions of the honeycomb unit of 120 degree of the first connection fixing device, one of the first connection fixing devices is equipped with a first power supply cable, and the angular position corresponding to the first connection fixing device has or is paired with a first power supply device, where pairing refers to The first power supply device matches and supplies power to the flight unit and does not supply power to other flight units; adjacent flight units are connected and fixed at the same angular position using the first connection fixing device; the lift fans are divided into numbers as close as possible a forward-rotating fan and a counter-rotating fan; the second power supply cable is arranged in the first connection fixture, and or the second power supply cable is arranged in the second connection fixture, and the second connection fixture is located around the honeycomb unit face position.

A more detailed aircraft structure plan is: one or more honeycomb units in the middle are configured as loading bins, and no flying units are configured; the aircraft is arranged with 2 to 4 layers of honeycomb units from the inside to the outside as flying units.

An additional solution is: there is a suspended wing structure above the rear, and there is at least one connection and fixing device between the suspended wing and the main body of the aircraft. The distance between the suspended wing and the main body of the aircraft can prevent downward airflow from entering the flight unit during forward flight; The suspended wings generate lift as the aircraft flies forward.

An additional solution is that the aircraft has a perimeter structure on its forward edge. The perimeter of the perimeter structure has an arc-shaped structure that guides the front airflow downwards and backwards. This structure generates lift when the aircraft flies forward.

An additional solution is that at least one flight unit at the rear of the aircraft can tilt forward to push the airflow backward to generate forward thrust.

An additional configuration is to include a front-drive fan with its rotating surface facing forward, pushing the airflow to the rear and driving the aircraft to fly forward. Its energy comes from the direct drive of the power source of the second power supply device; the front-drive fan is located close to the second power supply device. Two power supply devices are located in the middle or lower part of the honeycomb array surface and at the rear.

A flight control method based on a front-driven fan or a tiltable unit and a leading-edge lift structure is to adjust the pitch and roll attitude of the aircraft by adjusting the lift of the flight units at different positions, and achieve slow forward, forward, left, and right flight; by adjusting the flight The speed of the forward and reverse lift fans in the unit adjusts the heading of the aircraft; in take-off, landing, hovering and slow flight states, the lift is completely or mainly provided by the flight unit; during fast forward flight, the main forward power is provided by The front drive fan or the fan of the tilting unit is provided, and the main lift is provided by the leading lift structure, the rear suspended wing and the flight unit.

An electronic device is characterized in that it includes a memory, and the memory stores program instructions. The program instructions are stored in any of the aforementioned aircraft and can complete any of the aforementioned management and control functions of the aircraft.

beneficial effects

This patent document proposes a distributed hybrid solution for aircraft, which has the advantages of high safety and high flexibility, and also solves the problem of endurance. Its specific effects will be described in conjunction with the implementation details in the following implementation plans.

Description of the drawings

Figure 1 is a schematic diagram of a hybrid module according to an embodiment of the invention.

Figure 2 is a schematic diagram of electric energy operation according to the embodiment of the invention.

Figure 3 is a schematic diagram of the distribution of an aircraft with 24 cellular flight units.

Figure 4 is a detailed structural implementation diagram of a flight unit.

Figure 5 is a schematic diagram of the distribution of another aircraft with 24 cellular flight units.

Figure 6 is a schematic diagram of an aircraft electrical energy network with 24 cellular flight units.

Figure 7 is a distributed charging flow chart according to the embodiment of the invention.

Figure 8 is a schematic distribution diagram of an aircraft with a tilting device according to an embodiment of the invention.

Figure 9 is a schematic distribution diagram of 56 cellular flight units according to the embodiment of the invention.

Figure 10 is a side view of an aircraft with 56 cellular flight units according to the embodiment of the invention.

Embodiments of the invention

The specific implementation of this patent will now be described with reference to the accompanying drawings.

In this patent document: "electricity" and "electrical energy" have the same meaning, and both refer to the electrical energy mainly used for lift fans on aircraft. Batteries all refer to rechargeable batteries. Terms such as "standard power" and "maximum power" can be understood with reference to this implementation manual. The angle size is described in the angle system, and a circumferential angle is described as 360 degrees. The result value of each calculation process may be the result of rounding, and the left and right sides of the equal sign in the calculation expression are not strictly equal. A power supply cable consists of two power wires forming a current loop.

An electric lift fan refers to a device that uses electricity as energy and can control the speed. Including ESC, motor, fan blades and other components. The throttle signal is input to the ESC, and the ESC drives the motor. The motor drives the fan blades to rotate, driving the airflow and generating pulling force, which is referred to as a lift fan.

Association refers to components that are directly connected to it or belong to its internal structure. The associated electric energy, for the first electric energy controller, refers to the electric energy contained in these three components: the directly connected first power supply cable, the directly connected second power supply cable, and the internal rechargeable battery. An associated lift fan refers to a lift fan that is directly connected to it, or is supplied with electrical energy by it, or is controlled by it. The associated battery refers to the internal rechargeable battery of the first power supply device.

The uniform flight state refers to a flight state close to the same speed value or a state close to hovering. The speed changes are very small and the passengers have basically no feeling. It is not limited to no change in speed or a complete hovering state. For example: during a period of flight, if the speed change is not greater than 5% of the average value, it is judged to be flying at a constant speed.

Normal flight status refers to a state where the aircraft is flying at a constant speed, external airflow and other interferences are basically unchanged, and the required power power is relatively stable. For example: during a period of flight, if the power supply change is not greater than 5% of the average value, it is judged to be a normal flight state.

An aircraft implementation, the hybrid module structure is shown in Figure 1:

It has multiple distributed flight units and multiple distributed first power supply devices. The flight unit contains an electrically driven lift fan. The first power supply device includes a rechargeable battery and an electric energy controller. Each first power supply device is connected to a flight unit through a first power supply cable to provide power to it. Each flight unit is connected to an adjacent first power supply device to provide power.

The aircraft is also equipped with a second power supply device, which contains a power source. The power source is a gasoline power generation device. The gasoline engine drives a generator to generate electricity. The power is connected to each first power supply device through a second power supply cable.

The second power supply device detects the remaining amount of gasoline in the gasoline power generation equipment, obtains the remaining power that the second power supply device can provide, and sends this information to all electric energy controllers. At the same time, the power controller can detect the voltage of the rechargeable battery in the first power supply device, and can detect the voltage and current of the directly connected first power supply cable and the second power supply cable.

The power controller can use management actions to manage the associated power, as shown in Figure 2. The management actions are as follows:

Action A: Open the K301 power channel and charge the rechargeable battery in the first power supply device with the power connected by the second power supply cable; close the K302 and K303 power channels.

Action B: Open the K302 power channel and output the power from the rechargeable battery in the first power supply device to the connected lift fan; close the K301 and K303 power channels.

Action C: Open the K303 power channel and directly output the power connected to the second power supply cable to the connected lift fan; close the K301 and K302 power channels.

Action D: Open the K301 and K303 power channels, charge the power connected by the second power supply cable to the rechargeable battery in the corresponding first power supply device, and output it to the connected lift fan at the same time; close the K302 power channel.

Action E: Open the K302 and K303 power channels, output the power from the second power supply cable and the rechargeable battery to the connected lift fan at the same time, and close the K301 power channel.

The aircraft also contains a flight control system, which generates speed management instructions for each lift fan, which are throttle signals. The throttle signal is transmitted in the form of digital communication protocol, such as DSHOT600 protocol. The throttle signal is transmitted to the electric energy controller in the first power supply device, and the electric energy controller can adjust the value of the throttle signal and then send it to the associated lift fan.

Based on the implementation description of the module composition, an implementation plan of the structural position is shown in Figure 3: the number of the flight unit and the first power supply device of the aircraft is 24, arranged and fixed according to the honeycomb arrangement, and designed in two layers: inside and outside around the aircraft. The aircraft presents six aspects, and the outermost layer of each aspect is composed of 3 or 4 flight units, which are arranged alternately. There is a second connection and fixing device 101 around each flying unit, and the second connection and fixing device can connect and fix adjacent units. In the middle of the flight unit, there are two coaxial lift fans that rotate forward and backward, including fan blades 103, motor 102, and electric control (not shown). In the middle of the flight unit, three of the honeycomb-shaped hexagons are connected and fixed with first connection fixing devices 104 with equal angular spacing. The first connection fixing device is hollow in design, and the first power supply cable and the second power supply cable can be placed in the hollow part. Power supply cable. Adjacent flight units are connected and fixed at the same angular position using the first connection and fixing device 104. This angular position presents a triangular space 105. This space does not coincide with the rotating surface of the lift fan and does not affect the downward airflow driven by the lift fan. Flow, used to place the first power supply device. There are three honeycomb positions in the middle of the aircraft without flight units. There are loading bays 106 designed to load people and cargo. A second power supply device 107 is installed in the loading bay position. The second power supply device includes gasoline power generation equipment.

In this implementation, the double-layer dense honeycomb flight unit is arranged to increase the upward lift of the entire aircraft; compared with traditional helicopter blades, the lift fan blades are much smaller in size, so the inertia is much smaller, so the control response will be Very fast; the numerous flight units on the aircraft divide the lift surface into multiple small units, so that differential lift control can be performed for each angle and each small position, so that the aircraft can cope with very complex airflow environments.

The detailed structure of a flight unit is shown in Figure 4: the three first connection and fixing devices 104 of the flight unit are combined in the middle. There are motors 102 and fan blades 103 in the upper and lower directions of the joint. The first connection and fixing devices 104 are in A second connection and fixing device 101 is connected downward at the peripheral position. The second connection and fixing device 101 is annular, and its plane position is substantially the same as the rotation plane of the lower fan blade. The second connection and fixation device 101 can not only connect and fix adjacent flight units, but also function as a duct wall for the lower fan blades, including guiding the airflow downward and reducing noise.

Another structural position implementation is shown in Figure 5: the aircraft is not equipped with a second connection fixture, and adjacent flight units are connected and fixed using the first connection fixture, and the first power supply cable and the second power supply cable are designed at the A connection fixture. There are four outer flight units 108 on one side of the aircraft. There are only two first connection fixing devices connecting the middle position of each flight unit and the peripheral angular position of the flight unit, and no first connection fixation device is provided toward the outward direction. The aircraft has a total of 3 such side settings.

As shown in Figure 3, the electrical energy network design of the aircraft is shown in Figure 6, including: a first power supply device 201, a trifurcated second power supply cable 202 and a V-shaped second power supply cable 203, a second power supply device 204, a first Power supply cable 205, node 206. The second power supply cables connect all the first power supply devices and are also connected in a honeycomb shape. Each first power supply device has two or three second power supply cables directly connected. All the second power supply cables are connected. Although they pass through the first power supply device, they are also connected inside the first power supply device. The connection means that there are no semiconductor switches or any other components other than the second power supply cable. connect. The second power supply device 204 outputs electric energy through two second power supply cables. The first power supply device 201 is connected to the node 206 through the first power supply cable 205, and each first power supply device 201 is connected to and only one node 206. Figure 6 corresponds to Figure 3, in which: the first power supply device 201 corresponds to the position of the triangular space 105; the second power supply cables 202 and 203 and the first power supply cable 205 correspond to the position of the first connection fixing device 104; the second power supply device 204 It is the same component as the second power supply device 107; the node 206 corresponds to the electric control position in the middle of the flight unit and serves as a connection point to provide power to the corresponding lift fan in Figure 3. Features of this design include: The honeycomb multi-channel design of the second power supply cable improves safety.

The aircraft contains two voltage systems, high and low. The high-voltage 380V system mainly refers to gasoline power generation equipment and the second power supply cable, and their voltage setting values are both 380V. The low voltage is a 48V system, in which the rechargeable battery in the first power supply device is set to 12S lithium battery, the maximum working voltage is 4.2V*12=50.4V, and the minimum working voltage is set to 2.7V*12=32.4V. The working standard of the lift fan The voltage value is 44.4V, and its operating voltage range is 44.4V-50.4V.

The power controller includes a voltage reduction function. When the battery needs to be charged, the 380V electric energy of the second power supply cable is stepped down to the charging voltage value required for the 12S lithium battery. When the second power supply cable and the rechargeable battery need to supply power to the lift fan at the same time, and the current voltage of the rechargeable battery is not lower than 44.4V, the 380V electric energy of the second power supply cable is reduced to the current voltage value of the 12S battery, so that the two electric energy At the same time, the output is combined to the lift fan. When the second power supply cable needs to independently supply power to the lift fan, the 380V electric energy of the second power supply cable is stepped down to 44.4V and output to the lift fan. The voltage reducing circuit can also: detect the voltage and current input by the second power supply cable; detect the output voltage and current of the first power supply device.

The power controller also includes a voltage boosting function. When the second power supply cable and the rechargeable battery need to supply power to the lift fan at the same time, and the voltage of the rechargeable battery is lower than 44.4V, the 380V power of the second power supply cable will be stepped down to 44.4V. The electric energy of the rechargeable battery is boosted to 44.4V, and the two electric energies are combined and output to the lift fan at the same time. The boost circuit also has the switching function of the battery output.

An implementation method of adjusting the throttle signal includes: the electric energy controller receives the throttle signal value sent from the flight control system. This value corresponds to a voltage of 44.4V, which is converted into the actual working voltage value of the current lift fan. If the received throttle signal value is 20% and the current actual working voltage value is 50V, the converted throttle signal value is 20%*44.4V/50V=17.76%, and then the converted throttle value 17.76% is sent to the associated Lift fan.

When the aircraft is in normal flight status, the power requirement of the lift fan in each flight unit is set to 2KW, which is called the standard power of the lift fan in a flight unit. The power parameters output from the second power supply cable to each first power supply device are set with reference to this value: working voltage 380V, maximum current 2KW/380V=5.26A. When one second power supply cable is connected to supply multiple first power supply devices, the maximum transmittable current of the cable is calculated by superposition. For example, if there are two, the standard power is 4KW, that is: the operating voltage is 380V, and the maximum current is 5.26A*2 =10.52A. When the first power supply device has more than one second power supply cable to supply power, calculate the maximum transmittable current of each second power supply cable equally. For example, if there are two second power supply cables, the standard power of each second power supply cable is 1KW, that is: working voltage 380V, maximum current 5.26A/2=2.63A. Set the maximum power value of the lift fan in each flight unit to 4.5KW, then the maximum transmission power of the first power supply cable, based on this value, is set to: maximum voltage 50.4V, maximum current 4.5KW/44.4V=101A. The maximum output power of the battery in the first power supply device is set according to the difference, that is, 4.5KW-2KW=2.5KW. The specific settings are: maximum voltage 50.4V, maximum current 2.5KW/44.4V=56A. The power requirement of the entire aircraft under normal flight conditions is calculated based on the standard power of the flight unit, that is: 2KW*24=48KW, adding 5% redundancy, the value is 48KW*(1+5%)=50.4KW, this value As the standard power of the second power supply device, the specific parameters are: working voltage 380V, working current 50.4KW/380V=133A. There are two second power supply cables directly connected to the second power supply device. The maximum power that each cable can withstand is 50.4KW/2=25.2KW, that is, the working voltage is 380V and the maximum current is 25.2KW/380=66A. When the second power supply device and all the batteries in the first power supply device output electric energy at the same time, the maximum power output of the aircraft is 4.5KW*24=108KW.

For distributed electric-driven aircraft, the reliability and weight of the power cables have a great impact on the aircraft. During a flight of the aircraft, it is in a normal flight state most of the time. At this time, a small and stable power supply is required for a long time. In this embodiment, it is designed to be the above 48KW. During the period of rapid acceleration, a large amount of temporary burst electric energy is required. At this time, a high-power power supply is required. In this embodiment, it is designed to reach the maximum 108KW mentioned above, which exceeds the dual-engine power performance of a simple fuel engine.

The first power supply cable designed according to the maximum power requirement is used to transmit the increased current after voltage reduction and sudden large current. Its length is limited to a short length between the lift fan and the first power supply device, which reduces the weight. , also improves reliability.

The second power supply cable as the main line only needs to be designed according to the standard power requirements, and the current is smaller, which reduces the size and weight of the power cable. The transmission voltage of the second power supply cable is 380V, which is significantly higher than the transmission voltage of the first power supply cable of 48V, which is 7.92 times. With the same power, the transmission current is reduced, and the size and weight of the second power supply cable are further reduced.

As the second power supply device of the main energy source, it works in the standard power state most of the time and does not need to cope with sudden high power demands. Therefore, it can fully optimize the narrower standard power state and improve efficiency, such as using Atkinson cycle engine.

In this implementation, when the gasoline power generation equipment is fully charged, it uses an output power of 50.4KW and is configured to run for one hour, that is, the maximum power supply of the second power supply device is 50.4KWH. When the rechargeable battery of a single first power supply device is fully charged, the power configuration is 200WH, the maximum output power is 2.5KW, and the output magnification is 12.5. Therefore, the sum of the maximum battery power in all the first power supply devices is 200WH*24=4800WH, and the ratio to the maximum power supply of the second power supply device is 4800WH/50.4KWH=9.52%. When the second power supply device is exhausted or damaged and the batteries of all first power supply devices are fully charged, only the batteries are used to output power, and the maximum usage time is 200WH/2KW*60=6 minutes. In an emergency state with large current demand, the first power supply device and the second power supply device output electric energy at the same time, and the shortest usage time is 200WH/2.5KW*60=4.8 minutes. Assuming that the aircraft speed is 120 kilometers per hour and the emergency time is about 5 minutes, it can fly over 10 kilometers. Most special situations can be dealt with, thus greatly improving safety.

The aircraft uses a power controller to manage how to output power to the lift fan. It only uses one mode for management at a time. It has two operating modes, A and B, specifically:

A operation mode: C action is the first choice, B action and E action are the backup. After receiving the throttle signal, the power controller detects the remaining power supply of the second power supply device and the voltage of the second power supply cable. If the power of the second power supply device can be provided normally, under normal flight conditions, use action C to provide power for lift. For fans, when there is a temporary emergency demand for high power, action E is used to provide it; if the power of the second power supply device cannot be supplied normally, action B is used to provide it.

B operation mode: B action is the first choice, C action, D action and E action are the backup. After receiving the throttle signal, the power controller detects the associated battery. If the power can be supplied normally, action B is used to provide power to the lift fan under normal flight conditions. Action E is used to provide power for temporary high-power emergency needs. If the associated battery is damaged, If it fails and the voltage is 0, use action C to provide it. If it is just low battery, use action D to provide it.

The way to identify the temporary emergency demand for high power is to check whether the throttle signal value of the connected lift fan is greater than the set threshold. The threshold is set with reference to the standard power value, that is, the power value of the flight unit corresponding to the throttle signal. Greater than the standard power value of 2KW.

The method of identifying high-power temporary emergency demand can also be: detect the output current and voltage of the directly connected first power supply device, and calculate the power value. If it is greater than 2KW, it is judged that there is a high-power emergency demand.

In addition to the operating mode, the aircraft also includes a static charging mode. That is, when it is stationary on the ground, the lift fan is completely stationary and there is no throttle signal. After the power controller receives the charging command, it performs action A to charge.

The ratio of the remaining power supply to the maximum power supply of the second power supply device. When this value is 10%, that is, when the remaining power supply is 50.4KWH*10%=5.04KWH, it is used as the setting threshold. The aircraft can only be in one operating mode at the same time. It is set that when the power supply of the second power supply device is detected to be higher than the threshold, operating mode A is used, and when it is lower than the threshold, operating mode B is used.

The battery capacity of the first power supply device is set to be small, and a rechargeable battery that can be charged and discharged with high current and frequently charged and discharged is used as a backup power supply. As the main power source, the gasoline engine power generation equipment has a capacity configured to cover the entire voyage. In actual implementation scenarios, it can just fill up the gasoline and charge the battery in a short time, or it can quickly put it into use without charging the battery at all. This enables the aircraft to adapt to rapid application scenarios. As for the battery of the first power supply device used as a backup power source in this implementation, due to its small capacity, it can be quickly charged before takeoff. Its specific capacity can be enough to allow the aircraft to safely descend to the ground at a regular flight height, with a certain margin. When designing, for example, the reserve fuel capacity for flights on international routes has a reference data of approximately 10% of the route's flight time.

An implementation method to improve security: when each power controller performs various functions, it does not belong to each other and operates in a distributed manner. Specifically, each power controller independently communicates data with each other and regularly sends its associated power information. , receiving the associated electric energy information of other electric energy controllers, thereby obtaining the electric energy information of all first power supply devices, and at the same time obtaining the power supply information of the second power supply device, performing calculation and analysis, and completing respective electric energy actions.

During the use of the aircraft, the power controller can monitor the voltage value of the rechargeable battery, execute a charging mode, and charge in time. The implementation is shown in Figure 7. The specific steps are detailed as follows:

S01, the power controller starts the charging task regularly;

S02, detect whether the output power of the second power supply device is less than the standard power of 50.4KW;

S03, detect whether the input power of the directly connected second power supply cable is less than 2KW;

S04, detect whether the voltage of the associated rechargeable battery is lower than 48V;

S05, detect whether the associated rechargeable battery is in a non-charging state;

S06, detect whether the associated rechargeable battery is the rechargeable battery with the lowest voltage to be charged;

S07, if steps S02-S06 are all yes, start charging the associated battery;

S08, the charging information is formed and sent to other power controllers, and the step ends.

In the above steps, the information to be charged includes: whether it is a rechargeable battery to be charged, and the voltage value of the associated battery. The judgment method of the rechargeable battery to be charged is: if all steps S02-S05 are yes, it is judged to be a rechargeable battery to be charged, otherwise it is not a rechargeable battery to be charged. To receive charging information, the power controller is set to passively receive updates. Note that steps S02-S05 do not have to be performed in the above order and can be in any order; if any of steps S02-S06 is negative, jump directly to step S08.

In actual usage scenarios, the batteries of each first power supply device often have inconsistent capacities. Some batteries are in a fully charged state, while some batteries may be close to being exhausted. For this situation, the power controller includes A or B power regulation mode, where:

A power regulation mode: When the power management action of the power controller is action E, the current output current is maintained, which reduces the battery output current to the lift fan and increases the current obtained from the second power supply cable to the lift fan; when the power management action is Action C increases current from the second power supply cable to charge the battery, which changes to action D.

B power regulation mode: Based on the current power management action, the output current of the power supply device is reduced, thereby reducing the lift of the connected lift fan, and the adjacent first power supply device increases the output current, bringing the connected The lift of other lift fans is increased to compensate. The method of reduction is to reduce the value of the throttle signal output to the associated lift fan and send information to the adjacent first power supply device to increase the value of the throttle signal of the adjacent flight unit.

The power controller can use the following methods for power regulation mode and charging mode:

When the power supply of the second power supply device can be provided normally, then: run the charging mode; or detect the current associated battery voltage, and when it is lower than 48V, run the A power adjustment mode. If it cannot be provided normally, then: detect the battery voltage in all first power supply devices. When the battery associated with this power controller has the lowest voltage and is lower than 10% of the average, run the B power regulation mode.

The power adjustment mode obeys the channel rules of the charging mode, that is, when it is necessary to increase the output power of the second power supply device, the second power supply device must still have power margin, and the power channel from the second power supply device to the first power supply device must still have Power headroom.

Based on the implementation shown in Figure 3, an expanded implementation shown in Figure 8 is: the three honeycomb units in the middle are used for loading bins, and their length and width are designed with reference to the seating space for one person, For example, when a person is riding, the left and right width is about 70cm, and the front and rear length is about 100cm. The diameter of the honeycomb unit is designed to be slightly larger than 70cm. Since the front and rear length of the three honeycomb units is close to 130cm at this time, the length and width are both sufficient for riding a person. human needs. A tiltable device is added at the rear of the aircraft, including three flight units 109 and two first power supply devices 110 . The tilting device and the non-tilting main unit are connected using two tilting connection devices 111, and a second power supply cable is flexibly connected at this position to provide power to the tilting device. The tilt connection device 111 can be implemented using any disclosed angular tilt connection technology, such as using loose-leaf hinge technology. When the tiltable device tilts forward, it can drive the airflow to the rear without changing the current attitude of the aircraft, generating forward thrust, thereby increasing the riding comfort of the aircraft.

Other embodiments of the invention

Based on the first embodiment, a second embodiment with a different configuration is: the power source in the second power supply device is a rechargeable battery, which has an independent and quick replacement function and also serves as the main energy source of the aircraft; the first power supply device The battery configuration is the same as in the first embodiment. In the application scenario of aircraft, an aircraft is equipped with multiple main energy batteries, which are replaced immediately after the aircraft lands. Under the current situation of limited battery energy density, although the flight distance of the aircraft is limited, it can immediately go back and fly again, which is very efficient.

On the basis of the first embodiment, a third embodiment with a different configuration is: the battery in the first power supply device serves as the main energy source of the aircraft, and its capacity is configured to cover the entire voyage demand; the power source in the second power supply device It is a rechargeable battery, but its capacity is relatively small. It is used to: adjust the imbalance of battery power usage in each first power supply device during use; provide supplementary sudden power demand for the lift fan in the flight unit; provide a unified An interface for charging the batteries in each first power supply device. In this implementation, the aircraft requires a certain period of charging before use, but during the entire flight, the interdependence of electrical energy between the flight units is lower and the safety is higher.

Distributed cellular aircraft can be flexibly formed into various shapes. The fourth embodiment is shown in Figure 9. A total of 56 flight units are arranged in a honeycomb arrangement, and the overall shape is a symmetrical polygon. The flight units of the entire aircraft are symmetrical left and right, with the top left side of the aircraft being symmetrical. The fans on the first floor rotate in the forward direction clockwise, the upper one on the right rotates in the counterclockwise direction, the lower one on the left rotates in the opposite direction, and the lower one on the right rotates in the forward direction, achieving a fully symmetrical left and right arrangement.

The overall side view of the aircraft is shown in Figure 10: The aircraft has a leading edge lift structure 503. This structure refers to the arc-shaped structure on the front edge of the aircraft that guides the front airflow to flow downward and backward. This structure is used when the aircraft flies forward. Generate lift; the aircraft has a loading bay 505 in the middle lower part, and a second power supply device at the rear of the loading bay. The power source of the second power supply device is a gasoline generator 501. The gasoline generator includes a gasoline engine and a generator. There is a front-driven fan 502 at the rear of the gasoline engine. The front-driven fan is driven by the mechanical connection of the gasoline engine. Its rotating surface faces forward, pushing the airflow to the rear and driving the aircraft to fly forward. There is a clutch device in the mechanical connection device, and the fan pitch is adjustable. , used to start the front drive fan and adjust the pulling force. As shown in Figure 9, there are two honeycomb units above the loading bin without a lift fan in the middle, which are used to place and connect the loading bin; there is a suspended wing structure 504 above the rear of the aircraft, and there are two connections between the suspended wing and the main body of the aircraft. The distance between the fixed device and the suspended wing and the main body of the aircraft can prevent downward airflow from entering the flight unit during forward flight. The suspended wing generates lift when the aircraft flies forward, and its lift curve is close to the leading edge lift structure, so that the leading edge lift can be maintained during flight. The structure forms a rough front-to-back balance.

In this embodiment, the lift fan downwash flow can easily interfere with the front drive fan and its inlet and outlet airflow, reducing efficiency. One way to improve this situation includes: the front drive fan is directly connected to the second power supply device located in the middle. The front drive fan can use a culvert The duct form is used to reduce the windward area of the fan, reduce the number of flight units affected by the airflow under the influence of the ducted fan, and introduce the airflow source of the ducted fan from the top of the loading bin through the arc-shaped drainage structure; other methods can also be used Methods include reducing the lift fans in some flight units in the rear direction, replacing them with ducted fans, and integrating the ducted fans and honeycomb units on one plane.

The flight control of the distributed hybrid aircraft in this embodiment can be achieved by the following methods: adjusting the pitch and roll attitude of the aircraft by adjusting the lift of the flight units at different positions, and achieving slow forward, backward, left, and right flight; The heading of the aircraft is adjusted by adjusting the speed of the forward and reverse lift fans in the flight unit; during takeoff, landing, hovering and slow flight, the flight unit provides lift entirely or mainly; during fast forward flight, the aircraft's heading is mainly Forward power is provided by the front-drive fan, and the main lift is provided by the leading lift structure, the rear suspended wing and the lift fan. When cruising in normal fast forward flight, compared with the hovering state, the leading edge lift structure and suspended wings share part of the lift, and the lift fan takes on the other part of the lift. Because the lift fan bears part of the lift, it maintains the attitude control function of the aircraft. At the same time, because it only bears part of the lift, its pulling efficiency is also greatly improved. For example, the calculation example: a 4112 brushless motor, equipped with 14*4.7 propellers, power supply is 25V, when the current is 13A, the pulling force is 2KG. When the pulling force is shared half, that is, when only 1kg pulling force is needed, the current only requires 4.7A. The efficiency of the former is 6.2kG/kw, and the efficiency of the latter is increased to 8.5kg/W.

Persons of ordinary skill in the art can understand the aircraft structures and process methods of the above embodiments, wherein a software program can be written through a computer, and then the software program can be written into the processor of the aircraft to instruct the relevant hardware to complete various functions of the aircraft. Functional action. The software program can be stored in a memory, and the memory can be a magnetic disk, an optical disk, a read-only memory or a random access memory, etc.

The calculation of electric energy values in the above embodiments does not fully consider the loss, which needs to be considered depending on the actual implementation. For example, in terms of voltage reduction, as far as the current technical level is concerned, there is a situation where 2KW is only 1.9KW after voltage reduction, and the efficiency is only 95%. Circuit loss is a well-known thing to those skilled in the industry. It does not affect the authenticity of this patent and cannot be used as a reason to hinder the implementation of this patent.

Claims

A distributed hybrid aircraft characterized by:

Multiple distributed flight units and multiple first power supply devices are installed, where the flight unit includes one or two electric-driven lift fans; the first power supply device includes a rechargeable battery and an electric energy controller; each first power supply device, through A first power supply cable is connected to at least one adjacent flight unit, and is connected to the power supply to drive the lift fan therein to rotate; each flight unit is connected to an adjacent first power supply device, which provides power;

A second power supply device is installed, which contains a power source, and the power source form includes but is not limited to: hydrogen-to-electricity conversion equipment, or fuel-to-electricity conversion equipment, or rechargeable batteries; the second power supply device, through the second power supply cable, is connected to each is connected to a first power supply device and can provide power to it;

The power controller can obtain control information and associated power information, where the control information includes control instructions for the power controller and/or speed management instructions for the associated lift fan;

The power controller manages the associated power according to the power information and control information. The management actions include at least actions A and B among actions A, B, C, D, and E, among which:

Action A: Charge the rechargeable battery in the first power supply device with the electric energy connected to the second power supply cable;

Action B: Output the electric energy of the rechargeable battery in the first power supply device to the connected lift fan;

Action C: Directly output the electric energy connected to the second power supply cable to the connected lift fan;

Action D: Charge the electric energy connected by the second power supply cable to the rechargeable battery in the corresponding first power supply device, and output it to the connected lift fan at the same time;

Action E: Output the electric energy connected to the second power supply cable and the electric energy of the rechargeable battery to the connected lift fan at the same time.
The aircraft according to claim 1, characterized in that: the first power supply device has two or three second power supply cables, which are directly connected to input power; and all the first power supply devices obtain power from the second power supply cable in a parallel manner.
The aircraft according to claim 1 or 2, characterized in that: the transmission voltage of the second power supply cable is higher than the transmission voltage of the first power supply cable, and the power controller includes a voltage reduction function for reducing the high voltage of the second power supply cable. The reduced voltage value is close to the battery voltage value in the first power supply device, or close to the voltage value required for charging the battery in the first power supply device, or close to the standard voltage value required to drive the lift fan; and /or; the electric energy controller also includes a voltage boosting function, which is used to boost the electric energy of the rechargeable battery and output it. The boosted voltage value is close to the standard voltage value required to drive the lift fan.
The aircraft according to any one of claims 1 to 3, characterized in that: the second power supply device and the second power supply cable are designed with reference to the standard power requirements of the connected power supply lift fan; the output power of the first power supply device, the first power supply cable The transmission power is designed with reference to the maximum power requirement of the connected power supply lift fan; the battery output power of the first power supply device is designed with reference to the maximum power requirement of the connected power supply lift fan, or the difference between the maximum power and the standard power.
The aircraft according to claim 4, characterized in that: the ratio of the transmission voltage of the second power supply cable to the transmission voltage of the first power supply cable is not less than 3; the ratio of the maximum power of the lift fan to the standard power is not less than 2.
The aircraft according to any one of claims 1 to 5, characterized in that the electric energy controller can obtain the rotational speed management value of the associated lift fan, adjust the value according to the control information or the associated electric energy information, and then send it to the associated lift fan.
The aircraft according to any one of claims 1 to 6, characterized by including A operating mode or B operating mode, wherein:

A operation mode: the second power supply device is the preferred power supply, providing power to the lift fan through the power supply link; the battery power of the first power supply device is the backup power supply, providing supplementary power to the lift fan in case of temporary emergency needs, and Provides full power to the lift fan in the event of a preferred power failure;

B operation mode: The battery in the first power supply device is the primary power supply to provide power to the lift fan; the second power supply device is a backup power supply. In case of temporary emergency needs, it provides supplementary power to the lift fan through the power supply link, or provides power to the charging fan. Charges the battery and provides full power to the lift fan in the event of a primary power source failure.
The aircraft according to claim 7, characterized in that: including operation mode A, the ratio of the sum of the maximum battery power in all first power supply devices to the maximum available power of the second power supply device is 5%-35%; or Including B operation mode, the ratio of the maximum power supply of the second power supply device to the sum of the maximum battery power of all first power supply devices is 5%-35%.
The aircraft according to claim 7, characterized in that: it includes two operating modes A and B at the same time, and the sum of the maximum battery power in all first power supply devices is less than the maximum power supply of the second power supply device. When detecting When the power supply of the second power supply device is higher than the set threshold, the A operation mode is used; when the power supply is lower than the set threshold, the B operation mode is used.
The aircraft according to any one of claims 1 to 9, characterized in that: when the battery power in a certain first power supply device is lower than the average power level setting value, the A or B power adjustment mode is operated, wherein:

A power regulation mode: maintain the output current of the power supply device, reduce the battery output current to the lift fan, increase the current obtained from the second power supply cable to the lift fan; or increase the current from the second power supply cable to charge the battery;

B power regulation mode: reduce the output current of the power supply device, thereby reducing the lift of the connected lift fan, and increase the output current of other first power supply devices, causing the increase of the lift of other connected lift fans to compensate.
The aircraft according to any one of claims 1 to 10, characterized by comprising a charging management method that: detects the output power of the second power supply device; when the output power is less than the standard power, detects the transmission power of the second power supply cable one by one ; When the transmission power is less than the standard power, detect the rechargeable batteries in the first power supply device connected to the second power supply cable one by one; when the voltage of the detected rechargeable battery is lower than the set value and is not in a charged state , looking for the battery with the lowest voltage among them; the second power supply device charges the battery with the lowest voltage through the power supply link.
The aircraft according to any one of claims 1 to 11, characterized in that: each first power supply device can independently communicate with each other and exchange power information, thereby acquiring the power information of all first power supply devices, and simultaneously acquiring the second power supply device. The power supply information can be calculated and analyzed to complete its own electric energy actions.
The aircraft according to any one of claims 1 to 12, characterized in that: the flight units are arranged in a honeycomb array; a lift fan is placed at the center of the honeycomb unit, and there are 2 or 3 angles between the center position and the peripheral angular position of the honeycomb unit. The distance between the first connection fixtures is 120 degrees, and a first power supply cable is arranged in one of the first connection fixtures. The corresponding angular position of the first connection fixtures has or is paired with a first power supply device; Adjacent flight units are connected and fixed at the same angular position using the first connection fixture; the lift fans are divided into forward-rotating fans and counter-rotating fans whose numbers are as close as possible; the second power supply cable is arranged in the first connection fixture, Or, the second power supply cable is arranged in a second connection fixture located at a peripheral surface position of the honeycomb unit.
The aircraft according to claim 13, characterized in that: one or more honeycomb units in the middle are configured as loading bins, and no flying units are configured; and 2 to 4 layers of honeycomb units are arranged from the inside to the outside of the aircraft as flying units.
The aircraft according to claim 13 or 14, characterized in that: there is a suspended wing structure above the rear, there is at least one connecting and fixing device between the suspended wing and the main body of the aircraft, and the distance between the suspended wing and the main body of the aircraft can avoid falling when flying forward. The airflow enters the flight unit; the suspended wings generate lift when the aircraft flies forward.
The aircraft according to any one of claims 13 to 15, characterized by: including a leading edge lift structure, which means that the forward edge of the aircraft has a surrounding structure, and the periphery of the surrounding structure has an arc that guides the forward airflow to flow downward and backward. structure, which generates lift when the aircraft flies forward.
The aircraft according to any one of claims 13 to 16, characterized by: including a tiltable unit, which means that at least one rearmost flight unit can tilt forward to push the airflow backward to generate forward thrust.
The aircraft according to any one of claims 13 to 16, characterized by: including a front-driven fan, the rotating surface of the fan faces forward, pushes the airflow to the rear, and drives the aircraft to fly forward, and its energy comes from the power of the second power supply device. The front-drive fan is located close to the second power supply device, in the middle or lower part of the honeycomb array surface, and is located at the rear.
The aircraft according to claim 17 or 18, characterized by comprising a flight control method, which is: adjusting the pitch and roll attitude of the aircraft by adjusting the lift of the flight units at different positions, and achieving slow forward, backward, left and right flight; by adjusting the flight The speed of the forward and reverse lift fans in the unit adjusts the heading of the aircraft; in take-off, landing, hovering and slow flight states, the lift is completely or mainly provided by the flight unit; during fast forward flight, the main forward power is provided by The front drive fan or the fan of the tilting unit is provided, and the main lift is provided by the leading lift structure, the rear suspended wing and the flight unit.
An electronic device includes a memory, and the memory stores program instructions, characterized in that: the program instructions are stored in the aircraft described in any one of claims 1 to 19, and can complete the aircraft management function described in the claim.