Hybrid power systems and control methods for aircraft

Abstract

This invention relates to the field of aircraft technology, specifically disclosing a hybrid power system for aircraft and its control method. The hybrid power system for aircraft includes an engine, a generator, a first inverter, a high-voltage energy storage module, a second inverter, and a low-voltage energy storage module. The output shaft of the engine is connected to the generator. The first stator winding of the generator is electrically connected to the first inverter. The first inverter is electrically connected to the high-voltage energy storage module, as well as the lift and thrust devices of the aircraft. The high-voltage energy storage module is electrically connected to the lift and thrust devices. The second stator winding of the generator is electrically connected to the second inverter. The second inverter is electrically connected to the low-voltage energy storage module and onboard electrical components. The low-voltage energy storage module can provide electrical energy to the generator through the second inverter. This hybrid power system for aircraft has a high range and simplifies components, improves system integration, and thus reduces control complexity.

Description

Technical Field

[0001] This invention relates to the field of aircraft technology, and in particular to a hybrid power system for aircraft and its control method. Background Technology

[0002] Currently, pure electric aircraft primarily rely on lithium-ion batteries for power. However, limited by the current energy density of lithium-ion batteries, these aircraft generally suffer from short range and insufficient flight time, severely restricting their application scenarios. To overcome these shortcomings, hybrid power systems have been proposed as a promising technological solution. This system retains the advantages of electric drive while introducing an internal combustion engine to generate electricity to meet the energy demands for long-range flight. Since the internal combustion engine can operate stably within its high-efficiency operating range, this solution can effectively reduce overall carbon emissions.

[0003] In a typical hybrid power system, the internal combustion engine starts and drives a generator to produce electricity, which powers the propulsion / lift system to enable flight. A separate battery pack powers the onboard electrical equipment, but its limited capacity makes it difficult to support long-duration flight missions. Furthermore, an additional dedicated starter motor is needed to start the internal combustion engine, increasing the number of system components, structural complexity, and control difficulty.

[0004] Therefore, there is an urgent need to propose a hybrid power system and its control method for aircraft to solve the above-mentioned technical problems. Summary of the Invention

[0005] According to one aspect of the present invention, a hybrid power system for an aircraft is provided, which has a high range to meet the needs of long-duration flight, and simplifies the components, improves the system integration, and thus reduces the difficulty of control.

[0006] To achieve this objective, the present invention adopts the following technical solution: A hybrid power system for an aircraft includes an engine, a generator, a first inverter, a high-voltage energy storage module, a second inverter, and a low-voltage energy storage module. The output shaft of the engine is connected to the generator, which includes a first stator winding and a second stator winding. The first stator winding is electrically connected to the first inverter, and the first inverter is electrically connected to the high-voltage energy storage module and the lift and thrust devices of the aircraft, respectively. The first stator winding is used to provide electrical energy to the high-voltage energy storage module, the lift device and the thrust device through the first inverter when the engine is running. The high-voltage energy storage module is electrically connected to the lift device and the thrust device. The second stator winding is electrically connected to the second inverter, which is electrically connected to the low-voltage energy storage module and the airborne electrical equipment of the aircraft. The second stator winding is used to provide electrical energy to the low-voltage energy storage module and the airborne electrical equipment through the second inverter when the engine is running. The low-voltage energy storage module can also provide electrical energy to the generator through the second inverter so that the generator can drive the engine to start. The low-voltage energy storage module is electrically connected to the airborne electrical equipment.

[0007] Optionally, the output terminals of the first inverter and the high-voltage energy storage module are both electrically connected to the input terminal of the first combiner module, and the output terminal of the first combiner module is electrically connected to the lift device and the thrust device, respectively.

[0008] Optionally, the hybrid power system further includes a bidirectional voltage regulator module, through which the second inverter is electrically connected to the low-voltage energy storage module and the onboard electrical appliances, respectively.

[0009] Optionally, both the second inverter and the low-voltage energy storage module are electrically connected to the input terminal of the multi-output transformer, and the multiple output terminals of the multi-output transformer are respectively electrically connected to multiple airborne electrical appliances, the multiple airborne electrical appliances having different operating voltages.

[0010] Optionally, both the second inverter and the low-voltage energy storage module are electrically connected to the input terminal of the second combiner module, and the output terminal of the second combiner module is electrically connected to the input terminal of the multi-output transformer.

[0011] Optionally, the hybrid power system for the aircraft further includes a first sampling module, a second sampling module, and a third sampling module. The first sampling module is used to collect the voltage signal and the first current signal at the output terminal of the first inverter and transmit the voltage signal and the first current signal to the flight control computer of the aircraft. The second sampling module is used to collect the second current signal at the input terminal of the lift device and transmit the second current signal to the flight control computer. The third sampling module is used to collect the third current signal at the input terminal of the thrust device and transmit the third current signal to the flight control computer. The flight control computer generates a control strategy based on the received voltage signal, the first current signal, the second current signal, and the third current signal, and transmits the control strategy to the engine controller. The engine controller controls the engine to operate according to the control strategy.

[0012] Optionally, the hybrid power system for the aircraft further includes a cooling system, which includes an expansion tank, a cooling water pump, a first water-cooled radiator, a second water-cooled radiator, a first temperature sensor, and a second temperature sensor. The outlet of the expansion tank is connected to the inlet of the cooling water pump, the outlet of the cooling water pump is connected to the inlet of the first water-cooled radiator, the outlet of the first water-cooled radiator is connected to the inlet of the second water-cooled radiator, the outlet of the second water-cooled radiator is connected to the inlet of the cooling water circuit of the first inverter, the outlet of the cooling water circuit of the first inverter is connected to the inlet of the cooling water circuit of the engine, and the outlet of the cooling water circuit of the engine is connected to the inlet of the expansion tank. The first temperature sensor is used to detect a first temperature signal of the first inverter and transmit the first temperature signal to the flight control computer of the aircraft. The second temperature sensor is used to detect a second temperature signal of the engine and transmit the second temperature signal to the flight control computer. The flight control computer controls the cooling water pump to operate based on the received first and second temperature signals.

[0013] Optionally, the first water-cooled radiator and the second water-cooled radiator are respectively located below the two propellers of the aircraft.

[0014] According to another aspect of the present invention, the present invention provides a control method for a hybrid power system for an aircraft, controlling the operation of the hybrid power system for an aircraft described in any of the above-mentioned technical solutions, the control method for the hybrid power system for an aircraft comprising the following steps: The voltage signal U and the first current signal A1 at the output terminal of the first inverter, the second current signal A2 at the input terminal of the lift device, and the third current signal A3 at the input terminal of the thrust device are detected in real time, and the voltage signal U, the first current signal A1, the second current signal A2, and the third current signal A3 are transmitted to the flight control computer of the aircraft. If the first current signal A1, the second current signal A2, and the third current signal A3 are all less than a preset current threshold, the flight control computer transmits a control strategy to the engine controller to maintain the working state. If any one of the first current signal A1, the second current signal A2, and the third current signal A3 is greater than or equal to the preset current threshold, and the voltage signal U satisfies: Umin-c≤U≤Umax-c, then the flight control computer transmits a control strategy to the engine controller to maintain the working state. If any one of the first current signal A1, the second current signal A2, and the third current signal A3 is greater than or equal to the preset current threshold, and the voltage signal U < Umin-c, then the flight control computer transmits a control strategy to the engine controller to gradually increase the throttle opening according to the preset opening increment, until the voltage signal U satisfies: Umid-c ≤ U ≤ Umax-c, then the engine maintains the current working state; If any one of the first current signal A1, the second current signal A2, and the third current signal A3 is greater than or equal to the preset current threshold, and the voltage signal U > Umax-c, then the flight control computer transmits a control strategy to the engine controller to gradually reduce the throttle opening according to the preset opening reduction, until the voltage signal U satisfies: Umid-c ≤ U ≤ Umax-c, then the engine maintains the current operating state. Among them, the minimum available voltage of the high-voltage energy storage module is Umin, and the maximum available voltage of the high-voltage energy storage module is Umax; Umin-c is the minimum value of the voltage adjustment threshold, and Umin-c / Umin=(1.02-1.03); Umax-c is the maximum value of the voltage adjustment threshold, and Umax-c / Umax=(0.97-0.98); Umid-c is the redundancy value of the voltage adjustment threshold, and Umid-c / Umax-c=(0.97-0.98).

[0015] Optionally, the control method for the hybrid power system of the aircraft further includes the following steps: The first temperature signal T1 of the first inverter and the second temperature signal T2 of the engine are detected in real time, and the first temperature signal T1 and the second temperature signal T2 are transmitted to the flight control computer. If the first temperature signal T1 > T1max or the second temperature signal T2 > T2max, the flight control computer controls the cooling water pump to gradually increase its speed according to a preset speed increment until the first temperature signal T1 < T1max, and the second temperature signal T2 satisfies: T2min ≤ T2 ≤ T2max, and the cooling water pump maintains the current speed. If the second temperature signal T2 < T2min, the flight control computer controls the cooling water pump to gradually reduce its speed according to a preset speed reduction amount until T2min ≤ T2 ≤ T2max, and the cooling water pump maintains the current speed. Wherein, T1min is the minimum operating temperature of the first inverter, and T1max is the maximum operating temperature of the first inverter; T2min is the minimum operating temperature of the engine, and T2max is the maximum operating temperature of the engine, and T2max > T1max.

[0016] The beneficial effects of this invention are: This invention provides a hybrid power system for aircraft, comprising an engine, a generator, a first inverter, a high-voltage energy storage module, a second inverter, and a low-voltage energy storage module. The generator includes a first stator winding and a second stator winding. During the engine-driven rotation of the generator, high-voltage and low-voltage currents are generated in the first and second stator windings, respectively. The high-voltage current powers the lift and thrust devices, while the low-voltage current powers the onboard electrical components. This configuration allows a single generator to simultaneously meet the high-voltage power requirements for propulsion and the low-voltage power requirements for flight control and steering mechanisms. Compared to existing technologies that use separate battery packs to power onboard equipment, this provides a longer power supply time, thereby improving the aircraft's range.

[0017] The high-voltage energy storage module in the high-voltage current power supply branch can store excess energy when the electrical energy output by the generator is greater than the electrical energy required by the lift and thrust devices, and release energy to supplement the electrical energy when the electrical energy output by the generator is less than the electrical energy required by the lift and thrust devices, thereby improving the stability of the aircraft's flight.

[0018] The low-voltage energy storage module in the low-voltage current supply branch can store energy when the generator output is greater than the energy required by the onboard electrical components, and release energy when the generator output is less than the voltage required by the onboard electrical components, thus improving the reliability of the onboard electrical components. Furthermore, during engine startup, the energy in the low-voltage energy storage module can be supplied to the generator via a second inverter to drive the generator and start the engine. Compared to existing technologies that require a separate starter motor, this simplifies components, improves system integration, and reduces control complexity. Additionally, using energy from the low-voltage energy storage module to drive engine startup significantly reduces the rate of current change in the generator windings, thereby reducing the risk of insulation breakdown and overheating, which is beneficial for generator protection. Simultaneously, the risk of electric shock is lower, which helps protect the personal safety of maintenance personnel.

[0019] The present invention also provides a control method for a hybrid power system for an aircraft, for controlling the operation of the aforementioned hybrid power system for an aircraft, which improves the safety and reliability of aircraft flight. Attached Figure Description

[0020] To more clearly illustrate the technical solutions in the embodiments of the present invention, the accompanying drawings used in the description of the embodiments of the present invention will be briefly introduced below. Obviously, the accompanying drawings described below are only some embodiments of the present invention. For those skilled in the art, other drawings can be obtained based on the content of the embodiments of the present invention and these drawings without creative effort.

[0021] Figure 1 This is a schematic diagram of a hybrid power system for an aircraft provided in an embodiment of the present invention; Figure 2 This is a schematic diagram of the cooling system provided in an embodiment of the present invention; Figure 3 This is a schematic diagram of the structure of the aircraft provided in an embodiment of the present invention; Figure 4 This is a flowchart of a control method for a hybrid power system for an aircraft provided in an embodiment of the present invention; Figure 5 This is a control flowchart of the cooling system provided in an embodiment of the present invention.

[0022] In the picture: 100. Lift device; 110. Propeller; 200. Thrust device; 300. Airborne electrical equipment; 400. Flight control computer; 500. Engine controller; 1. Engine; 2. Generator; 21. First stator winding; 22. Second stator winding; 3. First inverter; 4. High-voltage energy storage module; 5. Second inverter; 6. Low-voltage energy storage module; 7. First busbar module; 8. Bidirectional voltage regulator module; 9. Multi-output transformer; 10. Second busbar module; 11. First sampling module; 12. Second sampling module; 13. Third sampling module; 14. Multi-parameter transmitter; 15. Expansion tank; 16. Cooling water pump; 17. First water-cooled radiator; 18. Second water-cooled radiator; 19. First temperature sensor; 20. Second temperature sensor. Detailed Implementation

[0023] The technical solution of the present invention will now be clearly and completely described with reference to the accompanying drawings. Obviously, the described embodiments are only some, not all, of the embodiments of the present invention. Based on the embodiments of the present invention, all other embodiments obtained by those skilled in the art without creative effort are within the scope of protection of the present invention.

[0024] In the description of this invention, it should be noted that the terms "center," "upper," "lower," "left," "right," "vertical," "horizontal," "inner," and "outer," etc., indicate the orientation or positional relationship based on the orientation or positional relationship shown in the accompanying drawings. They are used only for the convenience of describing the invention and for simplifying the description, and do not indicate or imply that the device or element referred to must have a specific orientation, or be constructed and operated in a specific orientation. Therefore, they should not be construed as limitations on the invention. Furthermore, the terms "first" and "second" are used for descriptive purposes only and should not be construed as indicating or implying relative importance. The terms "first position" and "second position" refer to two different positions. Furthermore, "above," "on top of," and "over" the first feature in relation to the second feature includes the first feature directly above and diagonally above the second feature, or simply indicates that the first feature is at a higher horizontal level than the second feature. "Below," "under," and "below" the first feature in relation to the second feature includes the first feature directly below and diagonally below the second feature, or simply indicates that the first feature is at a lower horizontal level than the second feature.

[0025] In the description of this invention, it should be noted that, unless otherwise explicitly specified and limited, the terms "installation," "connection," and "linking" should be interpreted broadly. For example, they can refer to a fixed connection, a detachable connection, or an integral connection; they can refer to a mechanical connection or an electrical connection; they can refer to a direct connection or an indirect connection through an intermediate medium; and they can refer to the internal connection of two components. Those skilled in the art can understand the specific meaning of the above terms in this invention based on the specific circumstances.

[0026] Embodiments of the present invention are described in detail below. Examples of these embodiments are shown in the accompanying drawings, wherein the same or similar reference numerals denote the same or similar elements or elements having the same or similar functions throughout. The embodiments described below with reference to the accompanying drawings are exemplary and are only used to explain the present invention, and should not be construed as limiting the present invention.

[0027] This embodiment provides a hybrid power system for aircraft, which has a long range to meet the needs of long-duration flights. Furthermore, it simplifies components, improves system integration, and reduces control complexity.

[0028] Specifically, such as Figure 1 As shown, the hybrid power system for the aircraft includes an engine 1, a generator 2, a first inverter 3, a high-voltage energy storage module 4, a second inverter 5, a low-voltage energy storage module 6, a first sampling module 11, a second sampling module 12, and a third sampling module 13.

[0029] The output shaft of engine 1 is connected to generator 2, which includes a first stator winding 21 and a second stator winding 22. The first stator winding 21 is electrically connected to a first inverter 3, which is also electrically connected to a high-voltage energy storage module 4, the lift device 100, and the thrust device 200 of the aircraft. The first stator winding 21 provides electrical energy to the high-voltage energy storage module 4, the lift device 100, and the thrust device 200 via the first inverter 3 when engine 1 is running. The high-voltage energy storage module 4 is electrically connected to the lift device 100 and the thrust device 200. The second stator winding 22 is electrically connected to a second inverter 5, which is also electrically connected to a low-voltage energy storage module 6 and the aircraft's onboard electrical components 300. The second stator winding 22 is used to provide electrical energy to the low-voltage energy storage module 6 and the onboard electrical appliances 300 through the second inverter 5 when the engine 1 is running. The low-voltage energy storage module 6 can also provide electrical energy to the generator 2 through the second inverter 5, so that the generator 2 can drive the engine 1 to start. The low-voltage energy storage module 6 is electrically connected to the onboard electrical appliances 300.

[0030] This hybrid power system for aircraft can simultaneously meet the high-voltage power supply required for propulsion and the low-voltage power supply required for flight control and steering gear through a single generator 2. Compared with existing technologies that use separate battery packs to power onboard electrical equipment, this system provides a longer power supply time, thereby improving the aircraft's endurance. Furthermore, the high-voltage energy storage module 4 in the high-voltage current supply branch can store excess energy when the electrical energy output from generator 2 exceeds the energy required by the lift device 100 and thrust device 200, and release energy to supplement power when the electrical energy output from generator 2 is less than the energy required by the lift device 100 and thrust device 200, thus improving the stability of the aircraft's flight. The low-voltage energy storage module 6 in the low-voltage current supply branch can store energy when the electrical energy output from generator 2 exceeds the energy required by the onboard electrical equipment 300, and release energy when the electrical energy output from generator 2 is less than the energy required by the onboard electrical equipment 300, thus improving the reliability of the onboard electrical equipment 300's operation. Furthermore, during the engine 1 startup process, the electrical energy in the low-voltage energy storage module 6 can also be supplied to the generator 2 via the second inverter 5, driving the generator 2 to start the engine 1. Compared with the existing technology that requires an additional separate starter motor, this simplifies components, improves system integration, and thus reduces control complexity. Simultaneously, using the electrical energy in the low-voltage energy storage module 6 to drive the engine 1 startup significantly reduces the rate of current change in the generator 2 windings, thereby reducing the risk of insulation breakdown and overheating, which is beneficial for protecting the generator 2. Moreover, the risk of electric shock is lower, which is beneficial for the personal safety of maintenance personnel.

[0031] It is understandable that the first stator winding 21 and the second stator winding 22 in generator 2 share the same iron core and rotor magnet.

[0032] Optionally, see [link to relevant documentation] Figure 1 The output terminals of the first inverter 3 and the high-voltage energy storage module 4 are both electrically connected to the input terminal of the first combiner module 7. The output terminal of the first combiner module 7 is electrically connected to the lift device 100 and the thrust device 200, respectively. It is understood that the power of an aircraft is generally output from multiple sources, typically no fewer than four sources to the lift device 100 and no fewer than one source to the thrust device 200. Therefore, by setting up the first combiner module 7, the power distribution between the bus and the load can be better realized, reducing electromagnetic interference.

[0033] Furthermore, in one possible technical solution, the first bus module 7 can be a bus bar with one input connector and multiple output connectors.

[0034] Optionally, see [link to relevant documentation] Figure 1 The hybrid power system also includes a bidirectional voltage regulator module 8. The second inverter 5 is electrically connected to the low-voltage energy storage module 6 and the onboard electrical appliances 300 via the bidirectional voltage regulator module 8. That is, the second inverter 5 is electrically connected to the bidirectional voltage regulator module 8, and the bidirectional voltage regulator module 8 is electrically connected to both the low-voltage energy storage module 6 and the onboard electrical appliances 300. The bidirectional voltage regulator module 8 has both boost and buck functions. During startup, it can boost the voltage of the low-voltage energy storage module 6 to supply the generator 2. During power generation, it can buck the high-voltage electricity generated by the generator 2 and supply it to the low-voltage energy storage module 6, achieving "dual-purpose functionality" and simplifying the system structure. Furthermore, through its boost and buck functions, the bidirectional voltage regulator module 8 can stabilize the voltage output by the generator 2 within a certain range, ensuring the stability of the onboard electrical appliances 300.

[0035] Furthermore, in one possible technical solution, the bidirectional voltage regulator module 8 can be a bidirectional DC-DC converter.

[0036] Optionally, see [link to relevant documentation] Figure 1 The hybrid power system for the aircraft also includes a multi-output transformer 9. Specifically, the second inverter 5 and the low-voltage energy storage module 6 are both electrically connected to the input terminal of the multi-output transformer 9. The multiple output terminals of the multi-output transformer 9 are respectively electrically connected to multiple airborne electrical appliances 300, each operating at a different voltage. By using the multi-output transformer 9 to meet the power requirements of multiple airborne electrical appliances 300 operating at different voltages, compared to the existing technology where each airborne electrical appliance 300 is paired with a separate step-up or step-down transformer, the number of components is reduced, thus simplifying the system structure.

[0037] Optionally, see [link to relevant documentation] Figure 1The second inverter 5 and the low-voltage energy storage module 6 are both electrically connected to the input terminal of the second bus module 10, and the output terminal of the second bus module 10 is electrically connected to the input terminal of the multi-output transformer 9. By setting up the second bus module 10, the power distribution between the bus and the load can be better realized, and electromagnetic interference can be reduced.

[0038] Further, see also Figure 1 The hybrid power system for the aircraft also includes a first sampling module 11, a second sampling module 12, and a third sampling module 13.

[0039] The first sampling module 11 is used to collect the voltage signal and the first current signal at the output terminal of the first inverter 3, and transmit the voltage signal and the first current signal to the flight control computer 400 of the aircraft.

[0040] The second sampling module 12 is used to collect the second current signal at the input end of the lift device 100 and transmit the second current signal to the flight control computer 400.

[0041] The third sampling module 13 is used to collect the third current signal at the input end of the thrust device 200 and transmit the third current signal to the flight control computer 400.

[0042] The flight control computer 400 generates a control strategy based on the received voltage signal, first current signal, second current signal and third current signal, and sends the control strategy to the engine controller 500. The engine controller 500 controls the engine 1 to work according to the control strategy.

[0043] By setting up a first sampling module 11, a second sampling module 12, and a third sampling module 13 to detect the corresponding voltage and current values, and adjusting the operation of engine 1 based on the detected voltage and current values, it is possible to avoid engine 1 overload or insufficient power, thereby improving fuel economy; at the same time, it can quickly detect abnormal current or insulation faults, enhancing flight safety and system reliability.

[0044] Optionally, see [link to relevant documentation] Figure 1 The hybrid power system for the aircraft also includes a multi-parameter transmitter 14. The outputs of the first sampling module 11, the second sampling module 12, and the third sampling module 13 are all electrically connected to the input of the multi-parameter transmitter 14, which in turn is electrically connected to the input of the flight control computer 400. By using the multi-parameter transmitter 14, the voltage signal, the first current signal, the second current signal, and the third current signal are uniformly conditioned and standardized before being transmitted to the flight control computer 400, thereby significantly simplifying system wiring and improving signal anti-interference capability and measurement accuracy.

[0045] Furthermore, the hybrid power system for the aircraft also includes a cooling system for dissipating heat from the aircraft's first inverter 3 and engine 1 to ensure the reliability of the aircraft's operation.

[0046] Specifically, such as Figure 2 As shown, the cooling system includes an expansion tank 15, a cooling water pump 16, a first water-cooled radiator 17, a second water-cooled radiator 18, a first temperature sensor 19, and a second temperature sensor 20.

[0047] The expansion tank 15 has its outlet connected to the inlet of the cooling water pump 16, which in turn connects to the inlet of the first water-cooled radiator 17. The outlet of the first water-cooled radiator 17 connects to the inlet of the second water-cooled radiator 18, which in turn connects to the inlet of the cooling water circuit of the first inverter 3. The outlet of the cooling water circuit of the first inverter 3 connects to the inlet of the cooling water circuit of the engine 1, and the outlet of the cooling water circuit of the engine 1 connects to the inlet of the expansion tank 15. By using two radiators, dual cooling of the cooling water is achieved, ensuring that the cooling water is reduced to the required temperature while allowing the use of two smaller radiators, thus saving costs. Since high temperatures have a significant impact on the operation of the first inverter 3, while low temperatures have a significant impact on the operation of the engine 1, the engine 1 is placed in the downstream water channel of the first inverter 3. This not only meets the rapid cooling requirement of the first inverter 3, but also allows the cooling water that has been heated after heat exchange with the first inverter 3 to cool the engine 1, preventing the engine 1 from cooling down excessively, thus ensuring the reliability of the operation of the first inverter 3 and the engine 1.

[0048] The first temperature sensor 19 is used to detect the first temperature signal of the first inverter 3 and transmit the first temperature signal to the flight control computer 400 of the aircraft. The second temperature sensor 20 is used to detect the second temperature signal of the engine 1 and transmit the second temperature signal to the flight control computer 400. The flight control computer 400 controls the cooling water pump 16 to work according to the received first and second temperature signals.

[0049] By setting a first temperature sensor 19 and a second temperature sensor 20 to detect the temperature of the first inverter 3 and the engine 1 respectively, and enabling the flight control computer 400 to control the operation of the cooling water pump 16 according to the detected temperature values, the flow rate of cooling water can be dynamically adjusted according to the working requirements of the aircraft, thereby ensuring that the first inverter 3 and the engine 1 always operate at a suitable temperature, thus improving flight safety and system reliability.

[0050] Furthermore, such as Figure 3As shown, the first water-cooled radiator 17 and the second water-cooled radiator 18 are respectively disposed below the two propellers 110 of the aircraft. Arranging the first water-cooled radiator 17 and the second water-cooled radiator 18 below the propellers 110 allows the downwash airflow of the propellers 110 to dissipate heat from the first water-cooled radiator 17 and the second water-cooled radiator 18. This improves the operational reliability of the first water-cooled radiator 17 and the second water-cooled radiator 18 while eliminating the need for other heat dissipation components, thus simplifying the structure. By separately distributing the first water-cooled radiator 17 and the second water-cooled radiator 18 below the two propellers 110 of the aircraft, the impact of sudden changes in the downwash airflow from a single propeller 110 can be reduced, further improving the reliability of the cooling system.

[0051] This embodiment also provides a control method for a hybrid power system for an aircraft, used to control the operation of the aforementioned hybrid power system. This control method for a hybrid power system improves the safety and reliability of aircraft flight.

[0052] Specifically, such as Figure 4 As shown, the control method for the hybrid power system of an aircraft includes the following steps: The voltage signal U and the first current signal A1 at the output terminal of the first inverter 3, the second current signal A2 at the input terminal of the lift device 100, and the third current signal A3 at the input terminal of the thrust device 200 are detected in real time, and the voltage signal U, the first current signal A1, the second current signal A2, and the third current signal A3 are transmitted to the flight control computer 400 of the aircraft.

[0053] If the first current signal A1, the second current signal A2, and the third current signal A3 are all less than the preset current threshold, the flight control computer 400 transmits a control strategy to the engine controller 500 to maintain the working state.

[0054] If any one of the first current signal A1, the second current signal A2, and the third current signal A3 is greater than or equal to a preset current threshold, and the voltage signal U satisfies: Umin-c≤U≤Umax-c, then the flight control computer 400 transmits a control strategy to the engine controller 500 to maintain the working state.

[0055] If any one of the first current signal A1, the second current signal A2, and the third current signal A3 is greater than or equal to the preset current threshold, and the voltage signal U < Umin-c, then the flight control computer 400 transmits a control strategy to the engine controller 500 to gradually increase the throttle opening according to the preset opening increment, until the voltage signal U satisfies: Umid-c ≤ U ≤ Umax-c, then the engine 1 maintains the current working state.

[0056] If any one of the first current signal A1, the second current signal A2, and the third current signal A3 is greater than or equal to a preset current threshold, and the voltage signal U > Umax-c, then the flight control computer 400 transmits a control strategy to the engine controller 500 to gradually reduce the throttle opening according to a preset opening reduction, until the voltage signal U satisfies: Umid-c ≤ U ≤ Umax-c, then the engine 1 maintains its current operating state.

[0057] Among them, the minimum available voltage of the high-voltage energy storage module 4 is Umin, and the maximum available voltage of the high-voltage energy storage module 4 is Umax; Umin-c is the minimum value of the voltage adjustment threshold, and Umin-c / Umin=(1.02-1.03); Umax-c is the maximum value of the voltage adjustment threshold, and Umax-c / Umax=(0.97-0.98); Umid-c is the redundancy value of the voltage adjustment threshold, and Umid-c / Umax-c=(0.97-0.98).

[0058] Optionally, the preset current threshold is 2%-5% of the rated current of generator 2.

[0059] It is understandable that when the first current signal A1, the second current signal A2, and the third current signal A3 are all less than the preset current threshold, it indicates that the aircraft is in a stable flight state and all components of the aircraft are in a stable working state. Therefore, there is no need to adjust engine 1.

[0060] It is worth noting that in the aforementioned hybrid power system for aircraft, the high-voltage energy storage module 4 is connected in parallel with the lift device 100 and the thrust device 200. Therefore, the voltage at the output terminal of the first inverter 3 is equal to the voltage of the high-voltage energy storage module 4, and also equal to the voltages of the lift device 100 and the thrust device 200. When the required power (current) of the lift device 100 and the thrust device 200 increases instantaneously due to takeoff, climb, landing, or encountering extreme weather, the high-voltage energy storage module 4, acting as a buffer source, will experience an instantaneous increase in its output current I. This increase in current I leads to a greater voltage drop across the internal resistance of the high-voltage energy storage module 4 (the product of I and r), and consequently, a decrease in the terminal voltage U of the high-voltage energy storage module 4. 端 =EI×r, where E is the internal electromotive force of the high-voltage energy storage module 4. Therefore, the terminal voltage U output by the high-voltage energy storage module 4 is... 端 It will inevitably decrease, and vice versa.

[0061] Furthermore, if the voltage of the high-voltage energy storage module 4 is not between (Umin, Umax), it will damage the high-voltage energy storage module 4 and affect its service life.

[0062] Therefore, when any one of the first current signal A1, the second current signal A2, and the third current signal A3 is greater than or equal to the preset current threshold, and the voltage signal U satisfies: Umin-c≤U≤Umax-c, it indicates that although the current required by the lift device 100 and the thrust device 200 of the aircraft increases (such as during takeoff, crawling, landing, or extreme weather), causing the voltage of the high-voltage energy storage module 4 to decrease, the voltage value is still within the voltage adjustment threshold range of the high-voltage energy storage module 4. At this time, the voltage change will not damage the high-voltage energy storage module 4, and the operation of the engine 1 does not need to be adjusted.

[0063] When any one of the first current signal A1, the second current signal A2, and the third current signal A3 is greater than or equal to the preset current threshold, and the voltage signal U < Umin-c, it indicates that the lift device 100 and the thrust device 200 of the aircraft require a large current, which causes the voltage of the high-voltage energy storage module 4 to decrease too much. At this time, it is necessary to adjust the operation of the engine 1 to avoid the voltage of the high-voltage energy storage module 4 dropping below Umin, which would affect the service life of the high-voltage energy storage module 4.

[0064] When any one of the first current signal A1, the second current signal A2, and the third current signal A3 is greater than or equal to a preset current threshold, and the voltage signal U > Umax - c, it indicates that the current required by the lift device 100 and the thrust device 200 of the aircraft is too small, resulting in an excessive increase in the voltage of the high-voltage energy storage module 4. At this time, it is necessary to adjust the operation of the engine 1 to prevent the voltage of the high-voltage energy storage module 4 from increasing above Umax, thus affecting the service life of the high-voltage energy storage module 4. Since the execution of the control strategy of the engine 1 takes time, setting the voltage adjustment threshold range (between Umin - c and Umax - c) within the usable voltage range of the high-voltage energy storage module 4 (between Umin and Umax) allows the engine 1 to start responding before the actual voltage value of the high-voltage energy storage module 4 reaches the limit of the usable voltage, and shortens the operating time of the high-voltage energy storage module 4 under extreme voltage, thereby significantly extending the overall cycle life of the high-voltage energy storage module 4.

[0065] In addition, after each adjustment of the engine 1 throttle, the voltage value of the high-voltage energy storage module 4 needs to be adjusted to between Umid-c and Umax-c (Umid-c is a safety margin value close to Umax-c). This can reduce repeated adjustments caused by voltage fluctuations and enable the lift device 100 and thrust device 200 to work in a high-efficiency working state with higher voltage, thereby improving the operational stability and reliability of the aircraft.

[0066] Optionally, the high-voltage energy storage module 4 includes n individual battery cells, where n ≥ 1. If the available voltage range of the individual battery cells is (3V, 4.2V), then the available voltage range of the high-voltage energy storage module 4 is (3nV, 4.2nV).

[0067] It is worth noting that the preset throttle opening increment and preset throttle opening decrement can be calibrated experimentally. The control strategy of gradually increasing the throttle opening with the preset opening increment and gradually decreasing the throttle opening with the preset opening decrement can avoid over-adjustment and improve the reliability and efficiency of throttle opening adjustment.

[0068] Optionally, the preset opening increment can be 2%-3% of the current throttle opening, and the preset opening decrement can be 2%-3% of the current throttle opening.

[0069] Furthermore, such as Figure 2 and Figure 5 As shown, the control method for the hybrid power system of an aircraft also includes the following steps: The first temperature signal T1 of the first inverter 3 and the second temperature signal T2 of the engine 1 are detected in real time, and the first temperature signal T1 and the second temperature signal T2 are transmitted to the flight control computer 400.

[0070] If the first temperature signal T1 > T1max or the second temperature signal T2 > T2max, the flight control computer 400 controls the cooling water pump 16 to gradually increase its speed according to the preset speed increment until the first temperature signal T1 < T1max and the second temperature signal T2 satisfies: T2min ≤ T2 ≤ T2max, and the cooling water pump 16 maintains the current speed.

[0071] If the second temperature signal T2 < T2min, the flight control computer 400 controls the cooling water pump 16 to gradually reduce its speed according to the preset speed reduction amount until T2min ≤ T2 ≤ T2max, and the cooling water pump 16 maintains the current speed.

[0072] Wherein, T1min is the minimum operating temperature of the first inverter 3, and T1max is the maximum operating temperature of the first inverter 3; T2min is the minimum operating temperature of the engine 1, and T2max is the maximum operating temperature of the engine 1, and T2max > T1max.

[0073] It is worth noting that the first inverter 3 is more sensitive to high temperatures. Specifically, if the temperature of the first inverter 3 exceeds its maximum operating temperature T1max, it will significantly impact its operating status and lifespan. However, if the temperature of the first inverter 3 is slightly lower than its minimum operating temperature T1min, the impact on its operating status and lifespan will be minimal. The engine 1 is more sensitive to low temperatures, especially prohibiting operation at extremely low temperatures. Therefore, if the temperature of the engine 1 is lower than its minimum operating temperature T2min, it will significantly impact its operating status and lifespan. However, if the temperature of the engine 1 is slightly higher than its maximum operating temperature T2max, the impact on its operating status and lifespan will be minimal.

[0074] Therefore, when the first temperature signal T1 > T1max or the second temperature signal T2 > T2max occurs, in the process of adjusting the speed of the cooling water pump 16 to increase the flow rate of the cooling water and thereby cool down the first inverter 3 and the engine 1, it is necessary to always ensure that T2min ≤ T2 ≤ T2max, while the lower limit value of the first temperature signal T1 is not required.

[0075] Furthermore, when T1 < T1min, the cooling system can continue to operate in its current state without adjustment.

[0076] The above-mentioned scheme of adjusting the speed of the cooling water pump 16 based on the first temperature signal T1 and the second temperature signal T2, thereby cooling the first inverter 3 and the engine 1, not only enables the first inverter 3 and the engine 1 to operate at a suitable temperature, but also reduces the difficulty of control, avoids repeated adjustments, and helps to save energy.

[0077] It is worth noting that the preset speed increment and preset speed decrement can be calibrated experimentally. The control strategy of gradually increasing the power of the cooling water pump 16 with the preset speed increment and gradually decreasing the speed of the cooling water pump 16 with the preset speed decrement can avoid over-adjustment and improve the reliability and efficiency of the cooling water pump 16 speed regulation.

[0078] Optionally, the preset speed increment can be 2%-3% of the rated speed of the cooling water pump 16, and the preset speed reduction can be 2%-3% of the rated speed of the cooling water pump 16.

[0079] Obviously, the above embodiments of the present invention are merely examples for clearly illustrating the present invention, and are not intended to limit the implementation of the present invention. Those skilled in the art can make other variations or modifications based on the above description. It is neither necessary nor possible to exhaustively describe all embodiments here. Any modifications, equivalent substitutions, and improvements made within the spirit and principles of the present invention should be included within the scope of protection of the claims of the present invention.

Claims

1. A hybrid power system for aircraft, characterized in that, It includes an engine (1), a generator (2), a first inverter (3), a high-voltage energy storage module (4), a second inverter (5) and a low-voltage energy storage module (6). The output shaft of the engine (1) is connected to the generator (2). The generator (2) includes a first stator winding (21) and a second stator winding (22). The first stator winding (21) is electrically connected to the first inverter (3), and the first inverter (3) is electrically connected to the high-voltage energy storage module (4) and the lift device (100) and thrust device (200) of the aircraft, respectively. The first stator winding (21) is used to provide electrical energy to the high-voltage energy storage module (4), the lift device (100) and the thrust device (200) through the first inverter (3) when the engine (1) is running. The high-voltage energy storage module (4) is electrically connected to the lift device (100) and the thrust device (200). The second stator winding (22) is electrically connected to the second inverter (5). The second inverter (5) is electrically connected to the low-voltage energy storage module (6) and the airborne electrical appliances (300) of the aircraft. The second stator winding (22) is used to provide electrical energy to the low-voltage energy storage module (6) and the airborne electrical appliances (300) through the second inverter (5) when the engine (1) is running. The low-voltage energy storage module (6) can provide electrical energy to the generator (2) through the second inverter (5) so that the generator (2) drives the engine (1) to start. The low-voltage energy storage module (6) is electrically connected to the airborne electrical appliances (300).

2. The hybrid power system for an aircraft according to claim 1, characterized in that, The output terminals of the first inverter (3) and the high-voltage energy storage module (4) are both electrically connected to the input terminal of the first combiner module (7), and the output terminal of the first combiner module (7) is electrically connected to the lift device (100) and the thrust device (200) respectively.

3. The hybrid power system for an aircraft according to claim 1, characterized in that, The hybrid power system also includes a bidirectional voltage regulator module (8), and the second inverter (5) is electrically connected to the low-voltage energy storage module (6) and the airborne electrical appliances (300) through the bidirectional voltage regulator module (8).

4. The hybrid power system for an aircraft according to claim 1, characterized in that, The second inverter (5) and the low-voltage energy storage module (6) are both electrically connected to the input terminal of the multi-output transformer (9). The multiple output terminals of the multi-output transformer (9) are electrically connected to multiple airborne electrical appliances (300), and the multiple airborne electrical appliances (300) have different operating voltages.

5. The hybrid power system for an aircraft according to claim 4, characterized in that, The second inverter (5) and the low-voltage energy storage module (6) are both electrically connected to the input terminal of the second combiner module (10), and the output terminal of the second combiner module (10) is electrically connected to the input terminal of the multi-output transformer (9).

6. The hybrid power system for an aircraft according to claim 1, characterized in that, The hybrid power system for the aircraft further includes a first sampling module (11), a second sampling module (12), and a third sampling module (13). The first sampling module (11) is used to collect the voltage signal and the first current signal at the output terminal of the first inverter (3) and transmit the voltage signal and the first current signal to the flight control computer (400) of the aircraft. The second sampling module (12) is used to collect the second current signal at the input terminal of the lift device (100) and transmit the second current signal to the flight control computer (400). The third sampling module (13) is used to collect the third current signal at the input terminal of the thrust device (200) and transmit the third current signal to the flight control computer (400). The flight control computer (400) generates a control strategy based on the received voltage signal, the first current signal, the second current signal, and the third current signal, and transmits the control strategy to the engine controller (500). The engine controller (500) controls the engine (1) to work according to the control strategy.

7. The hybrid power system for an aircraft according to any one of claims 1-6, characterized in that, The hybrid power system for the aircraft also includes a cooling system comprising an expansion tank (15), a cooling water pump (16), a first water-cooled radiator (17), a second water-cooled radiator (18), a first temperature sensor (19), and a second temperature sensor (20). The outlet of the expansion tank (15) is connected to the inlet of the cooling water pump (16), the outlet of the cooling water pump (16) is connected to the inlet of the first water-cooled radiator (17), the outlet of the first water-cooled radiator (17) is connected to the inlet of the second water-cooled radiator (18), and the outlet of the second water-cooled radiator (18) is connected to the inlet of the cooling water path of the first inverter (3). The outlet of the cooling water path of the inverter (3) is connected to the inlet of the cooling water path of the engine (1), and the outlet of the cooling water path of the engine (1) is connected to the inlet of the expansion tank (15). The first temperature sensor (19) is used to detect the first temperature signal of the first inverter (3) and transmit the first temperature signal to the flight control computer (400) of the aircraft. The second temperature sensor (20) is used to detect the second temperature signal of the engine (1) and transmit the second temperature signal to the flight control computer (400). The flight control computer (400) controls the cooling water pump (16) to work according to the received first temperature signal and second temperature signal.

8. The hybrid power system for an aircraft according to claim 7, characterized in that, The first water-cooled radiator (17) and the second water-cooled radiator (18) are respectively located below the two propellers (110) of the aircraft.

9. A control method for a hybrid power system for an aircraft, controlling the operation of the hybrid power system for an aircraft according to any one of claims 1-8, characterized in that, The control method for the hybrid power system of the aircraft includes the following steps: The voltage signal U and the first current signal A1 at the output terminal of the first inverter (3), the second current signal A2 at the input terminal of the lift device (100), and the third current signal A3 at the input terminal of the thrust device (200) are detected in real time, and the voltage signal U, the first current signal A1, the second current signal A2 and the third current signal A3 are transmitted to the flight control computer (400) of the aircraft. If the first current signal A1, the second current signal A2 and the third current signal A3 are all less than the preset current threshold, the flight control computer (400) transmits a control strategy to the engine controller (500) to maintain the working state. If any one of the first current signal A1, the second current signal A2, and the third current signal A3 is greater than or equal to the preset current threshold, and the voltage signal U satisfies: Umin-c≤U≤Umax-c, then the flight control computer (400) transmits a control strategy to the engine controller (500) to maintain the working state. If any one of the first current signal A1, the second current signal A2, and the third current signal A3 is greater than or equal to the preset current threshold, and the voltage signal U < Umin-c, then the flight control computer (400) transmits a control strategy to the engine controller (500) to gradually increase the throttle opening according to the preset opening increment until the voltage signal U satisfies: Umid-c ≤ U ≤ Umax-c, then the engine (1) maintains the current working state; If any one of the first current signal A1, the second current signal A2, and the third current signal A3 is greater than or equal to the preset current threshold, and the voltage signal U > Umax-c, then the flight control computer (400) transmits a control strategy to the engine controller (500) to gradually reduce the throttle opening according to the preset opening reduction, until the voltage signal U satisfies: Umid-c ≤ U ≤ Umax-c, then the engine (1) maintains the current working state; Among them, the minimum available voltage of the high-voltage energy storage module (4) is Umin, and the maximum available voltage of the high-voltage energy storage module (4) is Umax; Umin-c is the minimum value of the voltage adjustment threshold, and Umin-c / Umin=(1.02-1.03); Umax-c is the maximum value of the voltage adjustment threshold, and Umax-c / Umax=(0.97-0.98); Umid-c is the redundancy value of the voltage adjustment threshold, and Umid-c / Umax-c=(0.97-0.98).

10. The control method for a hybrid power system for an aircraft according to claim 9, characterized in that, The control method for the hybrid power system of the aircraft further includes the following steps: The first temperature signal T1 of the first inverter (3) and the second temperature signal T2 of the engine (1) are detected in real time, and the first temperature signal T1 and the second temperature signal T2 are transmitted to the flight control computer (400). If the first temperature signal T1 > T1max or the second temperature signal T2 > T2max, the flight control computer (400) controls the cooling water pump (16) to gradually increase the speed according to the preset speed increment until the first temperature signal T1 < T1max, and the second temperature signal T2 satisfies: T2min ≤ T2 ≤ T2max, and the cooling water pump (16) maintains the current speed. If the second temperature signal T2 < T2min, the flight control computer (400) controls the cooling water pump (16) to gradually reduce the speed according to the preset speed reduction until T2min≤T2≤T2max, and the cooling water pump (16) maintains the current speed; Wherein, T1min is the minimum operating temperature of the first inverter (3), T1max is the maximum operating temperature of the first inverter (3); T2min is the minimum operating temperature of the engine (1), T2max is the maximum operating temperature of the engine (1), and T2max > T1max.

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