Parallel oil-electric hybrid aircraft power system control method

Abstract

The application provides a parallel oil-electric hybrid aircraft power system control method and relates to the technical field of aircraft power system control. The application is suitable for a power system composed of a high power-to-weight ratio piston engine, a bidirectional output motor, a bidirectional generator and a high energy density battery pack. An aircraft operating state, a power system working state and a battery pack state of charge are obtained through an onboard sensor, and a unified demand power calculation and power mode determination mechanism is constructed in a flight management computer. The method determines an engine high efficiency output interval based on an engine universal characteristic curve, combines a flight task, a battery pack state of charge and a demand power relationship, autonomously selects an optimal mode from a pure electric, direct drive, hybrid, power generation and kinetic energy recovery mode and implements power distribution, so that the engine is kept in a high efficiency interval as much as possible, thereby improving power system efficiency, reducing fuel consumption and realizing safe, efficient and economic collaborative control in all working conditions.

Description

Technical Field

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

[0002] With the widespread application of drones and light general aviation aircraft in low-altitude reconnaissance, environmental monitoring, logistics transportation, and emergency rescue, hybrid power technology is gradually becoming an important development direction for aircraft power systems due to its ability to combine the high energy density of fuel with the high efficiency and fast response of electric motors. Currently, parallel hybrid electric vehicle architectures, with their high parallel power output efficiency and high safety redundancy of dual power sources, are widely used in small aircraft piston engine power systems. However, existing parallel hybrid electric vehicle power systems lack suitable engine control methods. They often borrow control logic from traditional engines, failing to fully adapt to the complex flight conditions of aircraft. Furthermore, they have significant shortcomings in global system efficiency optimization, dynamic switching of power modes, and battery balance control, making it difficult to achieve the synergistic goal of high efficiency and low fuel consumption.

[0003] The core advantage of hybrid-powered aircraft lies in the coordinated control of the engine and the electric motor. Its dual power source design allows the output torque of the engine and the electric motor to complement each other, so that the engine always operates in the high-efficiency output range, thereby improving engine output efficiency and reducing engine fuel consumption.

[0004] Existing parallel hybrid electric vehicle (HEV) control methods mostly focus on controlling the engine or motor individually. Even those methods that consider dual power systems generally suffer from local optimization rather than global optimization: most systems prioritize engine or motor efficiency optimization while neglecting the synergistic effects of engine status, high-energy-density battery pack state of charge (SOC), and other conditions. In summary, existing parallel HEV aircraft power system control methods, lacking a global efficiency optimization mindset, exhibiting poor adaptability to operating conditions, and insufficient dynamic SOC adaptation, fail to meet the aircraft's demand for overall system efficiency. Therefore, a control method that can accurately define the high-efficiency direct-drive range and dynamically adapt to the overall system and flight conditions is urgently needed to address these technical pain points. Summary of the Invention

[0005] To overcome the shortcomings of existing technologies, the purpose of this invention is to provide a control method for a parallel hybrid electric vehicle power system. By incorporating the engine's high-efficiency output range and the battery's state of charge into the power mode determination, the hybrid electric vehicle can achieve more efficient, safer, and more economical power coordination control under all operating conditions.

[0006] To achieve the above objectives, the present invention provides the following solution:

[0007] A control method for a parallel hybrid electric vehicle propulsion system, applied to the propulsion system of an aircraft including a high power-to-weight ratio piston engine, a bidirectional output motor, a bidirectional generator, and a high energy density battery pack, comprising:

[0008] The aircraft's operating status, the power system's working status, and the high-energy-density battery pack's state of charge are collected by airborne sensors, and the operating status, working status, and state of charge are input into the flight management computer as the basic data for control calculations.

[0009] Based on the universal characteristic curve of the high power-to-weight ratio piston engine, the high-efficiency output range of the engine is determined in the flight management computer. Within the high-efficiency output range, the engine's minimum drive curve and maximum drive curve are determined, and the high-efficiency output range, the minimum drive curve, and the maximum drive curve are used as parameters for power mode determination.

[0010] Based on the aforementioned basic data, and combined with the aircraft's current flight mission and flight conditions, the aircraft's power demand at the current moment is calculated in the flight management computer. The power demand is then compared with the power range corresponding to the engine's minimum drive curve and maximum drive curve to obtain the relationship data between the power demand and the engine's high-efficiency output range.

[0011] Based on the state of charge of the high-energy-density battery pack and the relationship between the required power and the engine's high-efficiency output range, at least one power mode is selected from pure electric drive mode, engine direct drive mode, hybrid electric drive mode, power generation mode and kinetic energy recovery mode, and a corresponding power mode decision command is generated.

[0012] According to the power mode decision command, a power allocation command is generated in the flight management computer and sent to the high power-to-weight ratio piston engine, the bidirectional output motor, and the bidirectional generator to adjust the output power of the high power-to-weight ratio piston engine, the output power of the bidirectional output motor, and the power generation power of the bidirectional generator. Under the premise of meeting the power requirements of the aircraft and the state of charge constraints of the high energy density battery pack, the high power-to-weight ratio piston engine operates within the high-efficiency output range.

[0013] Preferably, the airborne sensors include: an aircraft status sensor and a power system sensor; the aircraft status sensor is used to collect aircraft status data and input the collected aircraft status data into the flight management computer; the aircraft status data includes: flight speed, altitude, surface pressure, and ambient temperature and humidity; the power system sensor is used to collect power system status data and input the collected power system status data into the flight management computer as one of the basic data sources for control calculation; the power system status data includes: high power-to-weight ratio piston engine speed, engine fuel consumption rate, engine output power, bidirectional output motor speed, bidirectional output motor torque, temperature of key components of the power system, and voltage of the high energy density battery pack.

[0014] Preferably, the calculation of the aircraft's power demand at the current moment in the flight management computer includes:

[0015] Based on the aircraft's operating status in the aforementioned basic data, parameters related to flight dynamics are extracted; these parameters include aircraft load, air density, altitude, atmospheric conditions, and wing angle of attack.

[0016] Based on the parameters related to flight power, the power requirements for overcoming aerodynamic drag and maintaining flight during flight are calculated according to aerodynamic relationships, thus obtaining the initial power requirement data for the flight process.

[0017] The initial power demand data is corrected based on an empirical correction model and compared with the power range corresponding to the current flight profile. If the comparison is successful, the corrected power demand is taken as the required power of the aircraft.

[0018] Preferably, the high-efficiency output range of the engine is determined in the flight management computer, and the minimum drive curve and the maximum drive curve of the engine are determined within the high-efficiency output range, including:

[0019] Based on the universal characteristic curves provided by the engine manufacturer, the effective fuel consumption rate of the high power-to-weight ratio piston engine under different speed and torque combinations is analyzed, and the region with low effective fuel consumption rate and high efficiency is determined as the high-efficiency output range of the engine.

[0020] Within the engine's high-efficiency output range, a continuous power trajectory corresponding to the lower limit of output power along the direction of engine speed change is taken as the engine's minimum drive curve, and a continuous power trajectory corresponding to the upper limit of output power along the direction of engine speed change is taken as the engine's maximum drive curve. The operating power range of the high power-to-weight ratio piston engine is defined by the engine's minimum drive curve and the engine's maximum drive curve.

[0021] Preferably, during the power mode determination process, the setting methods for the state of charge range of the high-energy-density battery pack and the engine output power range include:

[0022] The state of charge (SOC) of the high-energy-density battery pack is divided into a low-charge range, an intermediate-charge range, and a high-charge range. When the SOC of the high-energy-density battery pack is less than 15% of its rated capacity, it is considered to be in the low-charge range. When the SOC of the high-energy-density battery pack is greater than or equal to 15% of its rated capacity but less than 60% of its rated capacity, it is considered to be in the intermediate-charge range. When the SOC of the high-energy-density battery pack is greater than or equal to 60% of its rated capacity, it is considered to be in the high-charge range.

[0023] The output power corresponding to the lowest drive curve of the engine is defined as the lowest drive power of the engine, and the output power corresponding to the highest drive curve of the engine is defined as the highest drive power of the engine. The lowest drive power and the highest drive power of the engine are used to compare the power required by the aircraft in segments when determining the power mode.

[0024] Preferably, the underlying control logic of the aircraft includes:

[0025] In the pure electric drive mode, the high power-to-weight ratio piston engine does not output power, and the high energy density battery pack supplies power to the bidirectional output motor. The bidirectional output motor outputs power and drives the propeller through the transmission mechanism to meet the power requirements of the aircraft.

[0026] In the engine direct drive mode, the bidirectional output motor does not output power, but supplies power to the high power-to-weight ratio piston engine by fuel. The high power-to-weight ratio piston engine outputs power and drives the propeller through the transmission mechanism.

[0027] In the hybrid electric drive mode, the high power-to-weight ratio piston engine and the bidirectional output motor simultaneously output power to drive the propeller. The high power-to-weight ratio piston engine provides the base power, and the bidirectional output motor provides power compensation.

[0028] In the power generation mode, a portion of the output power of the high power-to-weight ratio piston engine is used to drive the propeller, and another portion is converted into electrical energy by the bidirectional generator and stored in the high energy density battery pack to improve the state of charge of the high energy density battery pack.

[0029] In the kinetic energy recovery mode, when the power demand of the aircraft is less than zero, the bidirectional output motor switches to power generation mode, recovers excess kinetic energy during flight through the propeller and transmission system, converts the recovered mechanical energy into electrical energy and stores it in the high-energy-density battery pack.

[0030] Preferably, the upper-level control logic of the aircraft includes at least:

[0031] When the aircraft is gliding or in a condition where propulsion power is not required, causing the power system to be in a non-output state, the power system is set to kinetic energy recovery mode, so that the bidirectional output motor can recover kinetic energy to improve the state of charge of the high energy density battery pack.

[0032] When the power system is in output mode and the high power-to-weight ratio piston engine is in an unsuitable output mode or the high power-to-weight ratio piston engine cannot output power, the power system is set to pure electric drive mode, and the high energy density battery pack supplies power to the bidirectional output motor. The bidirectional output motor independently drives the propeller to meet the power requirements of the aircraft.

[0033] Preferably, when the state of charge of the high-energy-density battery pack is in the low-charge range, the upper-level control logic includes:

[0034] When the power demand of the aircraft is less than the minimum drive power of the engine, the power system is set to a combination of the engine direct drive mode and the power generation mode, so that the high power-to-weight ratio piston engine works in the high efficiency output range of the engine, and at the same time, the bidirectional generator charges the high energy density battery pack to prioritize the restoration of the state of charge of the high energy density battery pack.

[0035] When the power required by the aircraft is between the minimum drive power and the maximum drive power of the engine, the power system is set to the direct drive mode of the engine, so that the high power-to-weight ratio piston engine can independently drive the propeller within the high efficiency output range of the engine.

[0036] When the power required by the aircraft exceeds the maximum driving power of the engine, the power system is set to the direct drive mode of the engine acceleration, so that the high power-to-weight ratio piston engine outputs power at the upper boundary of the engine's high-efficiency output range to meet the propulsion power requirements of high-power flight conditions.

[0037] Preferably, when the state of charge of the high-energy-density battery pack is in the intermediate charge range, the upper-level control logic includes:

[0038] When the power required by the aircraft is less than the minimum driving power of the engine, the power system is set to a combination of the engine direct drive mode and the power generation mode, so that the high power-to-weight ratio piston engine operates within the high efficiency output range of the engine, and charges the high energy density battery pack through the bidirectional generator.

[0039] When the power required by the aircraft is between the minimum drive power and the maximum drive power of the engine, the power system is set to the direct drive mode of the engine, so that the high power-to-weight ratio piston engine can independently drive the propeller within the high efficiency output range of the engine.

[0040] When the power demand of the aircraft is greater than the maximum drive power of the engine and the power demand of the aircraft after deducting the maximum output power of the bidirectional output motor is still greater than the minimum drive power of the engine, the power system is set to the hybrid electric drive mode, and the output power of the high power-to-weight ratio piston engine is controlled within the high efficiency output range of the engine, with the bidirectional output motor providing residual power compensation.

[0041] When the power demand of the aircraft is greater than the maximum drive power of the engine and the power demand of the aircraft minus the maximum output power of the bidirectional output motor is less than or equal to the minimum drive power of the engine, the power system is set to the hybrid electric drive mode, and the output power of the high power-to-weight ratio piston engine is controlled to the minimum drive power of the engine, with the remaining power provided by the bidirectional output motor, so as to balance the efficient operation of the high power-to-weight ratio piston engine with the power demand of the aircraft.

[0042] Preferably, when the state of charge of the high-energy-density battery pack is in the high-charge range, the upper-level control logic includes:

[0043] When the power demand of the aircraft is less than or equal to the maximum output power of the bidirectional output motor, the power system is set to the pure electric drive mode, and the high energy density battery pack supplies power to the bidirectional output motor. The bidirectional output motor independently drives the propeller to reduce the number of start-stop cycles and fuel consumption of the high power-to-weight ratio piston engine.

[0044] When the power demand of the aircraft is greater than the maximum output power of the bidirectional output motor and the power demand of the aircraft minus the maximum output power of the bidirectional output motor is greater than the minimum drive power of the engine, the power system is set to the hybrid electric drive mode, and the output power of the high power-to-weight ratio piston engine is controlled within the high-efficiency output range of the engine, with the bidirectional output motor providing the differential power.

[0045] When the power demand of the aircraft is greater than the maximum output power of the bidirectional output motor and the power demand of the aircraft minus the maximum output power of the bidirectional output motor is less than or equal to the minimum drive power of the engine, the power system is set to the hybrid electric drive mode, and the output power of the high power-to-weight ratio piston engine is controlled to the minimum drive power of the engine. The bidirectional output motor provides residual power compensation, thereby ensuring the power demand of the aircraft while further improving the working economy of the high power-to-weight ratio piston engine.

[0046] The present invention discloses the following technical effects:

[0047] This invention introduces the engine's universal characteristic curve and the determination of its high-efficiency output range into the flight management computer. By consistently constraining power mode selection and power allocation within this high-efficiency output range, the high power-to-weight ratio piston engine operates within this range under most flight conditions. Compared to prior art methods that roughly control the engine based solely on immediate power requirements or simple speed ranges, this invention uses the engine's minimum and maximum drive curves as explicit power boundaries for comparison and calculation. This prevents the engine from prolonged operation in fuel-inefficient regions, reducing fuel consumption per unit of propulsion power at the system level and improving the range of hybrid electric vehicles.

[0048] This invention utilizes onboard sensors and a flight management computer to collaboratively collect and process data on the aircraft's operational status and the power system's working status. It then calculates the aircraft's required power based on the flight mission and operating conditions, using this as the input for power mode determination. Unlike prior art techniques that rely on only a few operating parameters or fixed assumptions for coarse control, this invention calculates required power from comprehensive data under actual flight conditions. This accurately reflects the drag, load, and handling requirements of the current flight profile, enabling precise matching between pure electric drive mode, engine direct drive mode, hybrid electric drive mode, power generation mode, and kinetic energy recovery mode, thus improving the responsiveness and adaptability of power output.

[0049] This invention explicitly incorporates the state of charge (SBC) of the high-energy-density battery pack into the power mode selection decision, ensuring that the power mode is no longer determined solely by engine operating conditions or power demand. Instead, it is comprehensively determined by a combination of three factors: aircraft power demand, engine high-efficiency output range, and battery SBC. Compared to prior art strategies that simply prohibit pure electric mode when the battery is too low or arbitrarily use the motor when the battery is fully charged, this invention can rationally arrange the use of electrical energy and the timing of recharging while ensuring the safety and lifespan of the high-energy-density battery pack. During long-endurance missions, it balances fuel consumption, battery SBC, and redundant power reserves, thereby improving the energy utilization efficiency and reliability of the entire hybrid power system.

[0050] This invention generates unified power mode decision commands and power allocation commands in the flight management computer, enabling coordinated control of the high power-to-weight ratio piston engine, bidirectional output motor, and bidirectional generator. This achieves orderly switching and linkage between various operating modes, including pure electric drive, engine direct drive, hybrid electric drive, power generation, and kinetic energy recovery. Compared to prior art schemes where each sub-component is controlled separately by independent controllers and mode switching relies on manual intervention or simple threshold logic, this invention can perform mode judgment and power allocation based on the same set of basic data within a single control framework. This avoids lag and conflicts during mode switching, reduces the impact of sudden torque changes and repeated energy flow switching, and thus improves the smoothness and comfort of the power system operation.

[0051] This invention prioritizes the operation of a high power-to-weight ratio piston engine within its efficient output range, while meeting the power requirements of the aircraft and the state of charge constraints of the high-energy-density battery pack. This not only improves fuel economy but also reduces the thermal load and mechanical stress on the engine under long-term high-load or low-efficiency conditions, thus extending engine life and maintenance cycles. Furthermore, this invention retains the ability of the electric motor and generator to participate in all operating conditions, providing additional power redundancy and energy regulation capabilities for the aircraft. Compared to the single power source or simple superposition solutions in the prior art, this invention is more suitable for low-altitude aircraft with complex flight missions and multi-segment flight profiles, contributing to improved mission safety and reliability. Attached Figure Description

[0052] To more clearly illustrate the technical solutions in the embodiments of the present invention or the prior art, the drawings used in the embodiments will be briefly introduced below. Obviously, the drawings described below are only some embodiments of the present invention. For those skilled in the art, other drawings can be obtained based on these drawings without creative effort.

[0053] Figure 1 A flowchart of the method provided in an embodiment of the present invention;

[0054] Figure 2 This is a schematic diagram of the technical route provided in the embodiments of the present invention;

[0055] Figure 3 This is a flowchart of the engine upper-level control method for determining performance, provided in an embodiment of the present invention.

[0056] Figure 4 This is a schematic diagram of a parallel hybrid electric power system provided in an embodiment of the present invention. Detailed Implementation

[0057] The technical solutions of the embodiments of the present invention will be clearly and completely described below with reference to the accompanying drawings. Obviously, the described embodiments are only some embodiments of the present invention, and not all embodiments. Based on the embodiments of the present invention, all other embodiments obtained by those skilled in the art without creative effort are within the scope of protection of the present invention.

[0058] The purpose of this invention is to provide a control method for the power system of a parallel hybrid electric vehicle. By coordinating the scheduling of the engine, motor and battery pack through a unified flight management computing framework, the method can achieve smooth switching between multiple modes and optimal energy allocation, thereby significantly improving the overall power performance and mission adaptability of the aircraft.

[0059] To make the above-mentioned objects, features and advantages of the present invention more apparent and understandable, the present invention will be further described in detail below with reference to the accompanying drawings and specific embodiments.

[0060] Figure 1 The method flowchart provided in the embodiments of the present invention is as follows: Figure 1 As shown, this invention provides a control method for a parallel hybrid electric vehicle power system, applicable to the power system of an aircraft including a high power-to-weight ratio piston engine, a bidirectional output motor, a bidirectional generator, and a high energy density battery pack, comprising:

[0061] The aircraft's operational status, the power system's working status, and the state of charge of the high-energy-density battery pack are collected by airborne sensors. The operational status, working status, and state of charge are then input into the flight management computer as the basic data for control calculations.

[0062] Based on the universal characteristic curve of the high power-to-weight ratio piston engine, the high-efficiency output range of the engine is determined in the flight management computer. Within the high-efficiency output range, the engine's minimum drive curve and maximum drive curve are determined, and the high-efficiency output range, minimum drive curve, and maximum drive curve are used as the basis parameters for power mode determination.

[0063] Based on the basic data, combined with the aircraft's current flight mission and flight conditions, the aircraft's power demand at the current moment is calculated in the flight management computer, and the power demand is compared with the power range corresponding to the engine's minimum drive curve and maximum drive curve to obtain the relationship data between the power demand and the engine's high-efficiency output range.

[0064] Based on the state of charge of the high-energy-density battery pack and the relationship between the required power and the engine's efficient output range, at least one power mode is selected from pure electric drive mode, engine direct drive mode, hybrid drive mode, power generation mode and kinetic energy recovery mode, and corresponding power mode decision instructions are generated.

[0065] Based on the power mode decision command, a power allocation command is generated in the flight management computer and sent to the high power-to-weight ratio piston engine, the bidirectional output motor, and the bidirectional generator to adjust the output power of the high power-to-weight ratio piston engine, the output power of the bidirectional output motor, and the power generation power of the bidirectional generator. Under the premise of meeting the power requirements of the aircraft and the state of charge constraints of the high energy density battery pack, the high power-to-weight ratio piston engine can operate in the high-efficiency output range.

[0066] In this invention, the engine drive curve and high-efficiency output range are determined. Based on the universal characteristic curves provided by the engine manufacturer, the high-efficiency output range and drive curve can be defined by analyzing key parameters such as the engine's effective fuel consumption rate, speed, and torque. The high-efficiency output range corresponds to a region of low fuel consumption and high efficiency. The lowest drive curve corresponds to the minimum output power curve within the engine's high-efficiency output range, and the highest drive curve corresponds to the maximum output power curve within the same range. These two drive curves define the high-efficiency and stable operating range of the engine under its output state.

[0067] like Figures 2 to 4 As shown, this embodiment includes the following steps:

[0068] S1. Collect the real-time status and operating mode of the aircraft and power system through airborne sensors;

[0069] S2. The Flight Management Computer (FMC) identifies the current operating environment and status of the engine based on the key operating status parameters of the aircraft (flight speed, altitude, aerodynamic drag, and load, etc.) collected in real time, and calculates the power required by the aircraft.

[0070] S3. Establish the engine drive curve and the optimal system output mode that changes with actual conditions;

[0071] S4 and FMC determine the required power mode of the power system based on the current operating status and the overall machine control model, and output the pre-set power mode decision to ensure optimal output efficiency.

[0072] S5. Construct the upper-level control logic and lower-level control logic of the parallel hybrid electric vehicle.

[0073] Among them, the airborne sensors mentioned in S1 include: aircraft-shaped sensors, including a flight speed measurement unit, an altitude measurement unit, a surface pressure measurement unit, an environmental condition (temperature, humidity, etc.) collection unit, etc.; and power system sensors, including an engine speed measurement unit, an engine fuel consumption rate measurement unit, an engine output power measurement unit, a motor speed measurement unit, a motor torque measurement unit, a temperature monitoring unit, and a power supply voltage monitoring unit.

[0074] Among them, such as Figure 3 As shown, the upper-level control logic of the parallel hybrid electric vehicle in S5 includes:

[0075] S500: When the aircraft's power system is in non-output mode, the M1 motor enters kinetic energy recovery mode.

[0076] S501. When the aircraft's power system is in output mode, but the engine is in an unsuitable output state or the engine is unable to output power, it enters pure electric mode. When the aircraft's power system is in output mode and the engine is in a suitable output state, the following judgment conditions are included:

[0077] S502. When the SOC is less than the minimum SOC limit and the actual power demand of the aircraft is less than the minimum drive curve of the engine, it enters the direct drive mode and the power generation mode.

[0078] S503. When the SOC is less than the minimum SOC limit, and the actual power demand of the aircraft is greater than the minimum drive curve of the engine but less than the maximum drive curve of the engine, the aircraft enters the direct drive mode of the engine.

[0079] S504. When the SOC is less than the minimum SOC limit and the actual power demand of the aircraft is greater than the engine's maximum drive curve, the aircraft enters the engine acceleration direct drive mode.

[0080] S505. When the SOC is greater than the minimum SOC limit, less than the intermediate SOC reference value, and the actual power demand of the aircraft is less than the minimum drive curve of the engine, the aircraft enters the engine direct drive mode and power generation mode.

[0081] S506. When the SOC is greater than the minimum SOC limit but less than the intermediate SOC reference value, and the actual power demand of the aircraft is greater than the minimum drive curve of the engine but less than the maximum drive curve of the engine, the aircraft enters the direct drive mode of the engine.

[0082] S507. When the SOC is greater than the minimum SOC limit and less than the intermediate SOC reference value, and the actual power demand of the aircraft is greater than the maximum drive curve of the engine, if the actual power demand of the aircraft minus the maximum output power of the output motor is greater than the minimum drive power of the engine, then the aircraft enters the hybrid mode, in which the engine output power is in the optimal output range.

[0083] S508. When the SOC is greater than the minimum SOC limit and less than the intermediate SOC reference value, and the actual power demand of the aircraft is greater than the maximum drive curve of the engine, if the actual power demand of the aircraft minus the maximum output power of the output motor is less than the minimum drive power of the engine, then the aircraft enters the hybrid mode, where the engine output power is the minimum drive power of the engine.

[0084] S509. When the SOC is greater than the intermediate reference value of SOC, if the actual power demand of the aircraft is greater than the maximum output power of the motor, and if the actual power demand of the aircraft minus the maximum output power of the motor is greater than the minimum drive power of the engine, then the aircraft enters the hybrid mode, in which the engine output power is in the optimal output range.

[0085] S510. When the SOC is greater than the intermediate reference value, if the actual power demand of the aircraft minus the maximum output power of the motor is less than the minimum drive power of the engine, then the aircraft enters hybrid mode, where the engine output power is the minimum drive power of the engine.

[0086] The SOC reference value of the power supply (high-energy-density battery pack) and the engine output power in the upper control of the S5 parallel hybrid electric aircraft are as follows:

[0087] S521 and SOClow are equal to 15% of the power supply rating.

[0088] S522 and SOCmid are equal to 60% of the power supply rating.

[0089] S523 and Gmin are the power corresponding to the engine's lowest drive curve;

[0090] S524 and Gmax represent the maximum power outputs corresponding to the engine's highest drive curves. This information can be adjusted according to the actual conditions of the aircraft.

[0091] The underlying control logic of the parallel hybrid electric vehicle propulsion system in the S5 includes:

[0092] S520, Pure Electric Mode: In this mode, the engine output is 0, and the power supply is used to provide energy to the output motor, which in turn drives the propeller.

[0093] S521, Direct Drive Mode: In this mode, the output motor output is 0, and the engine relies solely on fuel to provide energy. The engine outputs and drives the propeller.

[0094] S522, Hybrid Mode: In this mode, the engine and the output motor work together to drive the propeller;

[0095] S523, Generating Mode: This mode applies only to the generator. In this mode, a portion of the engine's output power is allocated to the generator to generate electricity. The generator converts mechanical energy into electrical energy and stores it in the power source.

[0096] S524, Kinetic Energy Recovery Mode: In this mode, the aircraft's power demand is less than 0, and the output motor M1 is converted into a generator to generate electricity by recovering energy and storing it in the power supply.

[0097] The method for calculating the power requirement of an aircraft in S2:

[0098] S21. First, the FMC receives information from various sensors, including aircraft load, air density, altitude, atmospheric conditions, wing angle of attack, etc.

[0099] S22. The specific formula for calculating the power requirement during flight is as follows:

[0100]

[0101] In the formula: Total takeoff mass; Wing area; This is the fuel consumption coefficient, which is the ratio of the current fuel mass to the initial fuel mass. For flight pressure; Zero-lift drag coefficient; The drag coefficient is the coefficient of elevation, where For the wing aspect ratio, Oswald factor; For flight altitude, Rise rate; For flight speed; For propeller efficiency; It is the acceleration due to gravity;

[0102] S23. Correct the required power calculated in the previous step according to the corresponding empirical formula, and then compare it with the power under the current flight profile to ensure the reliability of the calculation.

[0103] The optimal engine output mode selected based on actual conditions in S3 is as follows:

[0104] S331. When the engine output power is lower than the engine's minimum drive curve, the engine's direct drive efficiency is low, and the engine should increase its output power. However, the actual selection is constrained by the power supply SOC value and the maximum output power of the output motor (M1). If the power supply SOC value is lower than the intermediate reference value, the engine output state is adjusted to the optimal output range, with part of the power used to drive the propeller and part of the power used to generate electricity, and the generated power B = Gmin - R. At this time, the system is in direct drive mode + power generation mode. If the power supply SOC value is higher than the intermediate reference value, the specific output state of the generator is constrained by the rated power of the output motor. If the aircraft's power demand is less than the rated power of the output motor, the generator stops, the output is 0, and the power supply provides energy to the motor, which drives the propeller. At this time, the system is in pure electric mode. If the aircraft's power demand is higher than the rated power of the output motor, the engine output state is adjusted to the optimal output range, and the engine output is used entirely to drive the propeller. At this time, the system is in hybrid mode.

[0105] S332. When the engine output efficiency is higher than the engine's highest drive curve, the engine's fuel consumption is higher and the output efficiency decreases. The engine should then reduce its output power and select a more efficient output mode. However, the actual selection is constrained by the power supply's SOC. If the power supply's SOC value is less than the minimum reference value, the engine will still maintain an inefficient output state, and the system will remain in direct drive mode. At this time, the system is in direct drive mode. If the power supply's SOC value is greater than the minimum reference value, the engine's output state will be adjusted to the optimal output range, and the output motor will supplement the remaining required power. At this time, the system is in hybrid mode.

[0106] The specific determination and output decision of the power mode in S4 are as follows: The airborne FMC determines the power mode required by the power system based on the current operating status of the aircraft and the overall control model, and outputs a decision according to the pre-set power mode.

[0107] The advantages of this parallel hybrid electric vehicle power system control method are as follows: the high-efficiency output range of the engine is determined based on the universal characteristic curve of the engine, and the minimum drive curve and the maximum drive curve are defined. The optimal mode of the power system is determined based on the engine output state and the power supply SOC value, so that the engine is always in the optimal output state under each operating condition, thereby improving the engine output efficiency and reducing fuel consumption.

[0108] The various embodiments in this specification are described in a progressive manner, with each embodiment focusing on the differences from other embodiments. The same or similar parts between the various embodiments can be referred to each other.

[0109] This document uses specific examples to illustrate the principles and implementation methods of the present invention. The descriptions of the above embodiments are only for the purpose of helping to understand the method and core ideas of the present invention. Furthermore, those skilled in the art will recognize that, based on the ideas of the present invention, there will be changes in the specific implementation methods and application scope. Therefore, the content of this specification should not be construed as a limitation of the present invention.

Claims

1. A control method for a parallel hybrid electric vehicle power system, applied to the power system of an aircraft including a high power-to-weight ratio piston engine, a bidirectional output motor, a bidirectional generator, and a high energy density battery pack, characterized in that, include: The aircraft's operating status, the power system's working status, and the high-energy-density battery pack's state of charge are collected by airborne sensors, and the operating status, working status, and state of charge are input into the flight management computer as the basic data for control calculations. Based on the universal characteristic curve of the high power-to-weight ratio piston engine, the high-efficiency output range of the engine is determined in the flight management computer. Within the high-efficiency output range, the engine's minimum drive curve and maximum drive curve are determined, and the high-efficiency output range, the minimum drive curve, and the maximum drive curve are used as parameters for power mode determination. Based on the aforementioned basic data, and combined with the aircraft's current flight mission and flight conditions, the aircraft's power demand at the current moment is calculated in the flight management computer. The power demand is then compared with the power range corresponding to the engine's minimum drive curve and maximum drive curve to obtain the relationship data between the power demand and the engine's high-efficiency output range. Based on the state of charge of the high-energy-density battery pack and the relationship between the required power and the engine's high-efficiency output range, at least one power mode is selected from pure electric drive mode, engine direct drive mode, hybrid electric drive mode, power generation mode and kinetic energy recovery mode, and a corresponding power mode decision command is generated. According to the power mode decision command, a power allocation command is generated in the flight management computer and sent to the high power-to-weight ratio piston engine, the bidirectional output motor and the bidirectional generator to adjust the output power of the high power-to-weight ratio piston engine, the output power of the bidirectional output motor and the power generation power of the bidirectional generator. Under the premise of meeting the power requirements of the aircraft and the state of charge constraints of the high energy density battery pack, the high power-to-weight ratio piston engine operates within the high-efficiency output range. The high-efficiency output range of the engine is determined in the flight management computer, and the minimum and maximum drive curves of the engine are determined within the high-efficiency output range, including: Based on the universal characteristic curves provided by the engine manufacturer, the effective fuel consumption rate of the high power-to-weight ratio piston engine under different speed and torque combinations is analyzed, and the region with low effective fuel consumption rate and high efficiency is determined as the high-efficiency output range of the engine. Within the engine's high-efficiency output range, a continuous power trajectory corresponding to the lower limit of output power along the direction of engine speed change is taken as the engine's minimum drive curve, and a continuous power trajectory corresponding to the upper limit of output power along the direction of engine speed change is taken as the engine's maximum drive curve. The operating power range of the high power-to-weight ratio piston engine is defined by the engine's minimum drive curve and the engine's maximum drive curve.

2. The parallel hybrid-electric aircraft power system control method of claim 1, wherein, The airborne sensors include: an aircraft status sensor and a power system sensor; the aircraft status sensor is used to collect aircraft status data and input the collected aircraft status data into the flight management computer; the aircraft status data includes: flight speed, altitude, surface pressure, and ambient temperature and humidity; the power system sensor is used to collect power system status data and input the collected power system status data into the flight management computer as one of the basic data sources for control calculations; the power system status data includes: high power-to-weight ratio piston engine speed, engine fuel consumption rate, engine output power, bidirectional output motor speed, bidirectional output motor torque, temperature of key components of the power system, and voltage of the high energy density battery pack.

3. The control method of claim 1, wherein, The flight management computer calculates the aircraft's power demand at the current moment, including: Based on the aircraft's operating status in the aforementioned basic data, parameters related to flight dynamics are extracted; these parameters include aircraft load, air density, altitude, atmospheric conditions, and wing angle of attack. Based on the parameters related to flight power, the power requirements for overcoming aerodynamic drag and maintaining flight during flight are calculated according to aerodynamic relationships, thus obtaining the initial power requirement data for the flight process. The initial power demand data is corrected based on an empirical correction model and compared with the power range corresponding to the current flight profile. If the comparison is successful, the corrected power demand is taken as the required power of the aircraft.

4. The control method for the parallel hybrid electric vehicle power system according to claim 1, characterized in that, During the power mode determination process, the setting methods for the state of charge range of the high-energy-density battery pack and the engine output power range include: The state of charge (SOC) of the high-energy-density battery pack is divided into a low-charge range, an intermediate-charge range, and a high-charge range. When the SOC of the high-energy-density battery pack is less than 15% of its rated capacity, it is considered to be in the low-charge range. When the SOC of the high-energy-density battery pack is greater than or equal to 15% of its rated capacity but less than 60% of its rated capacity, it is considered to be in the intermediate-charge range. When the SOC of the high-energy-density battery pack is greater than or equal to 60% of its rated capacity, it is considered to be in the high-charge range. The output power corresponding to the lowest drive curve of the engine is defined as the lowest drive power of the engine, and the output power corresponding to the highest drive curve of the engine is defined as the highest drive power of the engine. The lowest drive power and the highest drive power of the engine are used to compare the power required by the aircraft in segments when determining the power mode.

5. The control method of claim 1, wherein, The underlying control logic of the aircraft includes: In the pure electric drive mode, the high power-to-weight ratio piston engine does not output power, and the high energy density battery pack supplies power to the bidirectional output motor. The bidirectional output motor outputs power and drives the propeller through the transmission mechanism to meet the power requirements of the aircraft. In the engine direct drive mode, the bidirectional output motor does not output power, but supplies power to the high power-to-weight ratio piston engine by fuel. The high power-to-weight ratio piston engine outputs power and drives the propeller through the transmission mechanism. In the hybrid electric drive mode, the high power-to-weight ratio piston engine and the bidirectional output motor simultaneously output power to drive the propeller. The high power-to-weight ratio piston engine provides the base power, and the bidirectional output motor provides power compensation. In the power generation mode, a portion of the output power of the high power-to-weight ratio piston engine is used to drive the propeller, and another portion is converted into electrical energy by the bidirectional generator and stored in the high energy density battery pack to improve the state of charge of the high energy density battery pack. In the kinetic energy recovery mode, when the power demand of the aircraft is less than zero, the bidirectional output motor switches to power generation mode, recovers excess kinetic energy during flight through the propeller and transmission system, converts the recovered mechanical energy into electrical energy and stores it in the high-energy-density battery pack.

6. The parallel hybrid-electric aircraft power system control method of claim 1, wherein, The upper-level control logic of the aircraft includes at least: When the aircraft is gliding or in a condition where propulsion power is not required, causing the power system to be in a non-output state, the power system is set to kinetic energy recovery mode, so that the bidirectional output motor can recover kinetic energy to improve the state of charge of the high energy density battery pack. When the power system is in output mode and the high power-to-weight ratio piston engine is in an unsuitable output mode or the high power-to-weight ratio piston engine cannot output power, the power system is set to pure electric drive mode, and the high energy density battery pack supplies power to the bidirectional output motor. The bidirectional output motor independently drives the propeller to meet the power requirements of the aircraft.

7. The control method of claim 4, wherein, When the state of charge of the high-energy-density battery pack is in the low-charge range, the upper-level control logic includes: When the power demand of the aircraft is less than the minimum drive power of the engine, the power system is set to a combination of the engine direct drive mode and the power generation mode, so that the high power-to-weight ratio piston engine operates within the engine's high-efficiency output range, and the high energy density battery pack is charged through the bidirectional generator to prioritize the restoration of the high energy density battery pack's state of charge. When the power required by the aircraft is between the minimum drive power and the maximum drive power of the engine, the power system is set to the direct drive mode of the engine, so that the high power-to-weight ratio piston engine can independently drive the propeller within the high efficiency output range of the engine. When the power required by the aircraft exceeds the maximum driving power of the engine, the power system is set to the direct drive mode of the engine acceleration, so that the high power-to-weight ratio piston engine outputs power at the upper boundary of the engine's high-efficiency output range to meet the propulsion power requirements of high-power flight conditions.

8. The control method for the parallel hybrid electric vehicle power system according to claim 4, characterized in that, When the state of charge of the high-energy-density battery pack is in the intermediate charge range, the upper-level control logic includes: When the power required by the aircraft is less than the minimum driving power of the engine, the power system is set to a combination of the engine direct drive mode and the power generation mode, so that the high power-to-weight ratio piston engine operates within the high efficiency output range of the engine, and charges the high energy density battery pack through the bidirectional generator; When the power required by the aircraft is between the minimum drive power and the maximum drive power of the engine, the power system is set to the direct drive mode of the engine, so that the high power-to-weight ratio piston engine can independently drive the propeller within the high efficiency output range of the engine. When the power demand of the aircraft is greater than the maximum drive power of the engine and the power demand of the aircraft after deducting the maximum output power of the bidirectional output motor is still greater than the minimum drive power of the engine, the power system is set to the hybrid electric drive mode, and the output power of the high power-to-weight ratio piston engine is controlled within the high efficiency output range of the engine, with the bidirectional output motor providing residual power compensation. When the power demand of the aircraft is greater than the maximum drive power of the engine and the power demand of the aircraft minus the maximum output power of the bidirectional output motor is less than or equal to the minimum drive power of the engine, the power system is set to the hybrid electric drive mode, and the output power of the high power-to-weight ratio piston engine is controlled to the minimum drive power of the engine, with the remaining power provided by the bidirectional output motor, so as to balance the efficient operation of the high power-to-weight ratio piston engine with the power demand of the aircraft.

9. The control method of claim 4, wherein, When the state of charge of the high-energy-density battery pack is in the high-charge range, the upper-level control logic includes: When the power demand of the aircraft is less than or equal to the maximum output power of the bidirectional output motor, the power system is set to the pure electric drive mode, and the high energy density battery pack supplies power to the bidirectional output motor. The bidirectional output motor independently drives the propeller to reduce the number of start-stop cycles and fuel consumption of the high power-to-weight ratio piston engine. When the power demand of the aircraft is greater than the maximum output power of the bidirectional output motor and the power demand of the aircraft minus the maximum output power of the bidirectional output motor is greater than the minimum drive power of the engine, the power system is set to the hybrid electric drive mode, and the output power of the high power-to-weight ratio piston engine is controlled within the high-efficiency output range of the engine, with the bidirectional output motor providing the differential power. When the power demand of the aircraft is greater than the maximum output power of the bidirectional output motor and the power demand of the aircraft minus the maximum output power of the bidirectional output motor is less than or equal to the minimum drive power of the engine, the power system is set to the hybrid electric drive mode, and the output power of the high power-to-weight ratio piston engine is controlled to the minimum drive power of the engine. The bidirectional output motor provides residual power compensation, thereby ensuring the power demand of the aircraft while further improving the working economy of the high power-to-weight ratio piston engine.

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