Performance optimization method for hybrid propulsion system of propeller tip jet based on thermal cycle planning

Abstract

This invention discloses a performance optimization method for a blade-tip jet hybrid propulsion system based on thermodynamic cycle planning, belonging to the field of aerospace hybrid propulsion technology. The method decomposes the design goals of high power-to-weight ratio and low fuel consumption of an aircraft propulsion system into optimization goals for the system's thermodynamic cycle parameters; identifies key parameters affecting cycle power and thermal efficiency, including compressor pressure ratio, turbine pressure ratio, rotor centrifugal pressure ratio, and blade-tip jet Mach number; collaboratively designs these strongly coupled key parameters and calculates system performance based on component models and energy conservation relationships; compares the calculated system power-to-weight ratio and fuel consumption with the optimization goals, and iteratively optimizes and adjusts component parameters until the optimization goals are met. This invention overcomes the limitations of traditional unidirectional engine design, achieving multi-level utilization and full recovery of system energy, and effectively solves the problem that traditional propulsion systems cannot simultaneously achieve high power-to-weight ratio and low fuel consumption.

Description

Technical Field

[0001] This invention relates to the field of aerospace composite propulsion technology, and in particular to a method for performance optimization of a tip-jet composite propulsion system based on thermodynamic cycle planning. Background Technology

[0002] The development history of propeller-driven jet aircraft is nearly 80 years. It is usually equipped with an aero engine as the air source / gas generator, and uses the reaction force generated by the jet from the propeller tip to drive the rotor to rotate and generate lift. The power-to-weight ratio and fuel consumption rate of the aero engine directly affect the effective payload capacity and flight time of the aircraft. However, the principle configuration of using a single power system as the air source generator to provide power to the rotor has always failed to meet the requirements of high power-to-weight ratio and low fuel consumption rate of aircraft. The main reasons include: (1) Traditional aero power systems can hardly meet the requirements of high effective payload and long flight time; (2) In terms of power-to-weight ratio, most existing helicopters use piston engines or turboshaft engines as the main power source. With the continuous advancement of engine design, although the power-to-weight ratio of the two types of power systems has made great progress, a heavy reduction gear is still required, so the power-to-weight ratio of the powertrain still cannot meet the high load requirements of future shipborne helicopters; (3) In terms of fuel consumption, the turboshaft engine with a relatively high power-to-weight ratio also has a high fuel consumption rate, which cannot meet the requirements of low-speed long-duration cruise; although the piston engine has a low fuel consumption rate, its power-to-weight ratio is also low, and the two cannot be taken into account at the same time.

[0003] The performance of a propeller-driven jet aircraft largely depends on the level of integrated design of its power system thermodynamic cycle and propulsion system. In traditional aero-engine design, the thermodynamic cycle is usually based on the Brayton cycle, which converts thermal energy into mechanical energy through the combination of compressor, combustion chamber and turbine. However, in the power configuration of a propeller-driven jet aircraft, this classic cycle mode and cycle optimization method are no longer applicable. This is mainly reflected in the following aspects: (1) The power system of a propeller-driven jet aircraft is deeply coupled with the propeller. Its key working components include not only the compressor, combustion chamber and turbine of a traditional aero-engine, but also the propeller and its built-in gas pipeline. After the turbine outlet gas flows into the built-in gas pipeline of the propeller, it will still undergo centrifugal pressurization during rotation, which greatly increases the complexity of the power system thermodynamic cycle, making it no longer a simple classic Brayton cycle; (2) The performance optimization and parameter iterative design of traditional aero-engines are usually "unidirectional calculation" according to the flow direction of the gas working fluid, and the design parameters are iteratively optimized in multiple rounds based on the calculation results of system performance. However, the propulsion system of a propeller-driven jet aircraft needs to be designed based on high power-to-weight ratio and low fuel consumption. First, the cycle boundary must be determined. Then, with the compressor inlet and propeller inlet as the starting points for calculation and the turbine outlet as the calculation coupling point, the energy matching scheme and performance of the system are designed and optimized. The cycle parameters involved in this design process are strongly coupled and cannot be achieved through traditional "one-way calculation".

[0004] Therefore, in order to further optimize the performance of the propulsion system of the tip jet aircraft, it is necessary to design a completely new thermodynamic cycle design method and parameter iterative optimization logic, and to match and iteratively optimize key parameters such as pressure ratio, turbine inlet temperature, mass flow rate, propeller speed, tip nozzle type and jet speed, rather than simply using the existing engine design methods and calculation logic. Summary of the Invention

[0005] To address the aforementioned technical challenges, this invention aims to propose a performance optimization method for a tip-jet hybrid propulsion system based on thermodynamic cycle planning. Specifically, considering the design goals of high power-to-weight ratio and low fuel consumption of the propulsion system, and effectively combining the performance characteristics of traditional piston engines and gas turbines, a thermodynamic cycle planning and performance optimization method for a tip-jet hybrid propulsion system is proposed. This method breaks through the single-link energy transfer mode of engines in traditional tip-jet aircraft. Under the premise of relatively low component temperature limits and thermal loads, it allocates, organizes, and coordinates the energy of the hybrid propulsion system based on thermodynamic cycle planning, achieving multi-level utilization and full recovery of system energy, thus solving the problem that traditional propulsion systems cannot simultaneously achieve high power-to-weight ratio and low fuel consumption.

[0006] This invention provides a performance optimization method for a blade tip jet hybrid propulsion system based on thermodynamic cycle planning. The hybrid propulsion system includes at least a gas turbine and a jet rotor. The gas turbine includes a compressor, a combustion chamber, and a turbine. The jet rotor includes internal ducts and a blade tip nozzle. The specific steps are as follows: S1. Based on the performance requirements of the aircraft, determine the design targets for the available power and maximum fuel consumption rate of the composite power system; S2. Transform the design objective into an optimization objective for the thermodynamic cycle parameters; S3. Based on the design level of aero-engines and the temperature resistance limit of materials, set the maximum operating temperature of the thermodynamic cycle; under the constraint of the maximum operating temperature of the thermodynamic cycle, identify the key factors affecting the useful work of the cycle; the key factors affecting the useful work of the cycle include the compressor pressure ratio and the turbine pressure ratio; S4. Based on the fuel consumption requirements of the composite power system, the tip nozzle form is determined to be a converging nozzle, and the inlet gas conditions of the tip nozzle are designed. It is also identified that the centrifugal pressure ratio of the jet rotor and the tip jet Mach number are the key parameters affecting the thermal efficiency and fuel consumption of the system. S5. Collaboratively design the key factors and key parameters, and sequentially calculate the inlet and outlet gas state parameters of each component of the composite power system. S6. Based on the calculation results of step S5, calculate the current power-to-weight ratio and current fuel consumption rate of the composite power system; S7. Determine whether the current power-to-weight ratio and current fuel consumption rate meet the optimization target; if not, adjust the component design parameters and return to step S5 for iterative optimization until the optimization target is met, thereby obtaining the optimized thermodynamic cycle key parameters.

[0007] Optionally, in step S5, the collaborative design includes: under the condition that the compressor and the turbine meet the preset work balance relationship, through the collaborative design of the turbine outlet gas parameters, rotor speed, wing internal pipe size and the blade tip nozzle size, the reaction thrust and power generated by the blade tip nozzle meet the system requirements.

[0008] Optionally, the work balance relationship is expressed by the following formula:

[0009] in, and These represent compressor work and turbine work, respectively. and These represent the inlet gas flow rates of the compressor and turbine, respectively. and These represent the adiabatic indices for the gas compression process in the compressor and the gas expansion process in the turbine, respectively. Represents the gas constant; and These represent the inlet gas temperatures of the compressor and turbine, respectively. and These represent the compressor boost ratio and the turbine pressure drop ratio, respectively. , and These represent the compressor efficiency, turbine efficiency, and rotor system mechanical efficiency, respectively.

[0010] Optionally, the centrifugal pressure ratio is the ratio of the total pressure in front of the propeller tip nozzle to the total pressure at the inlet of the jet rotor, and its value is determined by the rotor speed, the cross-sectional design of the duct inside the wing, and the flow loss in the duct inside the wing.

[0011] Optionally, the composite power system further includes a piston engine connected to the rotor main shaft; in step S6, the current power-to-weight ratio is calculated as the ratio of the total output power of the system to the total weight of the piston engine and the gas turbine; the total output power is the sum of the power of the piston engine and the power of the jet rotor.

[0012] Optionally, in step S5, the step of calculating the gas flow rate required for the propeller tip jet includes: subtracting the piston engine power from the total output power of the system to obtain the effective driving power of the jet rotor, and then solving the problem based on the energy conservation relationship between the effective driving power of the jet rotor, the jet power of the propeller tip nozzle, and the power consumed by the gas in the rotor for centrifugal pressurization.

[0013] Optionally, in step S5, the total gas pressure and total temperature at the propeller tip nozzle inlet are calculated based on the centrifugal pressurization ratio; the jet velocity at the propeller tip nozzle is obtained based on the total gas pressure, total temperature and critical velocity at the nozzle inlet.

[0014] Optionally, in step S3, the upper temperature limit is the highest temperature of the turbine's inlet gas, and its value is set based on the material's temperature resistance limit and the level of cooling technology.

[0015] Optionally, in step S7, the design parameters of the adjustment component include: adjusting one or more of the following: compressor boost ratio, turbine pressure ratio, rotor speed, cross-sectional area of ​​the wing duct, and throat area of ​​the blade tip nozzle.

[0016] Compared with the prior art, the present invention has at least the following beneficial effects: 1. In similar propeller-jet hybrid propulsion system configurations, the propeller-jet hybrid propulsion system performance optimization method based on thermodynamic cycle planning proposed in this invention can quickly identify and calculate the thermodynamic cycle process and key parameters.

[0017] 2. The performance optimization method proposed in this invention can be directly decomposed from the performance requirements of the flight platform to the design requirements of the power system, and the system flow, pressure and temperature can be iteratively calculated through an efficient parameter optimization algorithm.

[0018] 3. The power system optimized based on this invention has a significant improvement in power-to-weight ratio and fuel consumption rate compared to traditional aero piston engines and traditional aero turboshaft engines.

[0019] 4. The power-to-weight ratio of the optimized power system obtained based on the present invention is higher than 2.5 kW / kg, which is higher than that of traditional aviation piston engines and traditional aviation turboshaft engines; the system's minimum fuel consumption rate is roughly equivalent to that of traditional aviation piston engines and lower than that of traditional aviation gas turbines (0.3 kg / (kW·h)). Attached Figure Description

[0020] Figure 1 Analysis of thermodynamic cycle mode and identification of key parameters for the propeller tip jet combined propulsion system of this invention. Figure 2 This invention provides the logic and process for iterative optimization of the performance of the propeller tip jet hybrid propulsion system. Figure 3 This is a flowchart of the performance optimization method for a tip-jet hybrid propulsion system based on thermodynamic cycle planning, as described in this invention. Detailed Implementation

[0021] To better understand the above-described objectives, features, and advantages of the present invention, the invention will be further described in detail below with reference to the accompanying drawings and specific embodiments. It should be noted that, unless otherwise specified, the embodiments of the present invention and the features thereof can be combined with each other. Furthermore, the present invention can be implemented in other ways different from those described herein; therefore, the scope of protection of the present invention is not limited to the specific embodiments disclosed below.

[0022] A specific embodiment of the present invention, such as Figure 3 This paper discloses a performance optimization method for a tip-jet hybrid propulsion system based on thermodynamic cycle planning. The method consists of three parts: thermodynamic cycle mode analysis and key parameter identification, thermodynamic cycle parameter co-design, and system performance iterative optimization. The specific technical solution is as follows: Part 1: Thermodynamic Cycle Model Analysis and Key Parameter Identification. Through thermodynamic cycle analysis, the design goals of high power-to-weight ratio and low fuel consumption of the power system are transformed into optimization goals for thermodynamic cycle parameters, specifically including the upper limits of pressure and temperature during the cycle. To maximize the useful work of the cycle, under the constraint of the highest operating temperature of the cycle, i.e., the turbine inlet gas temperature, key factors affecting the useful work of the cycle are identified, mainly including the compressor boost ratio and turbine pressure drop ratio. Based on this, to minimize the system's fuel consumption, the high-temperature gas from the turbine outlet is delivered to the rotor. The gas further expands in the contracting nozzle located at the rotor tip to generate thrust, identifying the rotor's centrifugal boost ratio as a key parameter affecting the system's thermal efficiency and fuel consumption rate.

[0023] The thermodynamic cycle model analysis of the first step is attached. Figure 1 As shown.

[0024] Part Two: Coordinated Design of Thermodynamic Cycle Parameters: To achieve efficient energy utilization of the system, the cycle parameters of the gas turbine and the jet rotor are designed in a coordinated manner. The compressor pressure ratio and turbine expansion ratio of the gas turbine, the rotor centrifugal pressure ratio, and the jet Mach number are key cycle parameters. During the design process, it is necessary not only to ensure that the compressor and turbine components meet the work balance relationship, but also to coordinate the design of turbine outlet parameters, rotor speed, internal piping, and blade tip nozzles to ensure that the reaction thrust and power generated by the blade tip jet meet the system design requirements.

[0025] Part Three: Iterative Optimization of System Performance. The performance requirements of the power system include a high power-to-weight ratio and low fuel consumption. These are decomposed into the power system's operating parameters, namely, tip jet thrust and fuel consumption. Further decomposition leads to the power system's thermodynamic cycle parameters, namely, cycle useful work and cycle heating. Based on the power system's inlet conditions, component operating characteristics, and preliminary design parameters, the inlet and outlet states of each component are calculated sequentially, ultimately obtaining the gas pressure, temperature, velocity, and flow rate of the tip jet. Based on these parameters, the reaction thrust and converted power are calculated. Based on the converted power, the system's power-to-weight ratio and fuel consumption rate are further calculated and compared with the design targets. Based on the comparison results, the component design parameters are adjusted until the power system's design targets are met.

[0026] Further, see Figure 1-3 A performance optimization method for a propeller-tip jet hybrid propulsion system based on thermodynamic cycle planning is presented. The hybrid propulsion system includes a piston engine, a gas turbine, and a jet rotor. The gas turbine includes a compressor, a combustion chamber, and a turbine. The jet rotor includes internal ducts and a tip nozzle. The piston engine is connected to the rotor shaft (e.g., via a clutch). The gas turbine supplies gas to the rotor. The compressor and turbine of the gas turbine are coaxially arranged (i.e., rotating at the same speed). The combustion chamber is located between the compressor and turbine, with its inlet connected to the compressor outlet and its outlet connected to the turbine inlet. The specific steps are as follows: Step 1: Based on the usage scenarios and functional requirements of the propeller-tip jet aircraft, as well as the design requirements for a high power-to-weight ratio and low fuel consumption rate of the composite propulsion system, determine the available power and maximum fuel consumption rate of the composite propulsion system.

[0027] Step 2: Transform the design requirements of high power-to-weight ratio and low fuel consumption rate of the composite power system into optimization targets for thermodynamic cycle parameters through thermodynamic cycle analysis.

[0028] Specifically, the optimization objectives of thermodynamic cycle parameters include the upper limits of pressure and temperature in the cycle process, which are the optimization boundaries of the thermodynamic cycle parameters.

[0029] Step 3: Based on the design level of the aero-engine and the temperature resistance limit of the materials, set the maximum operating temperature of the thermodynamic cycle; under the constraint of the maximum operating temperature of the thermodynamic cycle, identify the key factors affecting the useful work of the cycle.

[0030] Specifically, the highest operating temperature of the thermodynamic cycle is the highest temperature of the gas at the turbine inlet; key factors affecting the useful work of the cycle include the compressor boost ratio and the turbine pressure drop ratio.

[0031] Step 4: Based on the fuel consumption requirements of the composite power system, determine that the tip nozzle is a converging nozzle, design the inlet gas conditions of the tip nozzle, and identify the centrifugal pressure ratio of the jet rotor and the tip jet Mach number as key parameters affecting the system's thermal efficiency and fuel consumption.

[0032] The centrifugal pressure ratio of a jet rotor is the total gas pressure in front of the rotor tip nozzle divided by the total pressure at the rotor inlet.

[0033] Step 5: Collaboratively design the key parameters that determine the energy utilization rate and thermal efficiency of the composite power system.

[0034] The key cycle parameters include the compressor pressure ratio and turbine expansion ratio of the gas turbine, as well as the centrifugal pressure ratio and tip jet Mach number of the jet rotor.

[0035] Step 6: Under the set operating environment, given the compressor boost ratio and efficiency design values, calculate the compressor specific work, outlet gas pressure, temperature and density and other parameters in combination with the intake conditions of the operating environment.

[0036] Step 7: Based on the highest operating temperature of the thermodynamic cycle in Step 3, calculate the heating amount and fuel flow rate of the gas turbine combustion chamber to obtain the turbine inlet gas conditions.

[0037] Step 8: Obtain the design values ​​of turbine pressure ratio and efficiency, as well as turbine outlet gas conditions, based on the compressor and turbine work balance relationship shown in formula (1).

[0038] The work balance relationship satisfied by the compressor and turbine is shown in equation (1): (1) in, and These represent compressor work and turbine work, respectively. and These represent the inlet gas flow rates of the compressor and turbine, respectively. and These represent the adiabatic indices for the gas compression process in the compressor and the gas expansion process in the turbine, respectively. Represents the gas constant; and These represent the inlet gas temperatures of the compressor and turbine, respectively. and These represent the compressor boost ratio and the turbine pressure drop ratio, respectively. , and These represent the compressor efficiency, turbine efficiency, and rotor system mechanical efficiency, respectively.

[0039] Step 9: Through the coordinated design of turbine outlet parameters, rotor speed, wing internal piping and blade tip nozzle, obtain a reasonable design value for rotor rotation centrifugal pressure ratio, so that the reaction thrust and power generated by the blade tip nozzle meet the system design requirements.

[0040] Specifically, turbine outlet parameters include the pressure, temperature, velocity, and flow rate of the gas exiting the turbine.

[0041] Understandably, the turbine outlet temperature determines the rotor inlet gas temperature; the rotor speed, the design of the internal duct cross-section, and the design of the blade tip nozzle jointly affect the gas flow velocity and dynamic pressure, thereby determining the total gas pressure in front of the blade tip nozzle and the pressure ratio of the rotor centrifugal pressurization.

[0042] Specifically, when the rotor rotates, it does work on a unit flow rate of gas, generating the centrifugal boost work required for the rotor rotation centrifugal boost ratio. The calculation method is shown in formula (2): (2) in, The adiabatic index represents the pressure increase of gas by centrifugal force in the rotor. Indicates the temperature of the gas entering the rotor; This indicates the efficiency of gas centrifugal pressurization in the rotor; This indicates the centrifugal pressure boost ratio of the gas in the rotor.

[0043] Step 10: Subtract the piston engine power from the available power of the composite power system obtained in Step 1 to obtain the effective driving power of the jet rotor; obtain the gas flow rate required for the jet tip jet based on the energy conservation relationship between the effective driving power of the jet rotor, the jet power of the blade tip nozzle, and the power consumed by centrifugal pressurization.

[0044] Furthermore, the calculation method for the propeller tip jet power is shown in formula (3).

[0045] (3) in, Indicates the effective driving power of the jet rotor; This indicates the velocity of the airflow ejected from the propeller tip nozzle; Indicates the rotor speed; This indicates the distance between the propeller tip nozzle and the rotor shaft; The adiabatic index represents the pressure increase of gas by centrifugal force in the rotor. Indicates the temperature of the gas entering the rotor; This indicates the efficiency of gas centrifugal pressurization in the rotor; This indicates the pressure ratio by which the gas is centrifugally pressurized in the rotor; m T This indicates the gas flow rate at the turbine outlet.

[0046] Furthermore, the power consumed by centrifugal boosting is the product of the specific work of centrifugal boosting and the gas flow rate.

[0047] Furthermore, the power of the power system = piston engine power + effective drive power of the jet rotor = jet power from the propeller tip nozzle - power consumed by centrifugal pressurization = jet power from the propeller tip nozzle - specific work of centrifugal pressurization × gas flow rate.

[0048] Furthermore, the expression for the energy conservation relationship between the effective driving power of the jet rotor, the airflow velocity at the rotor tip nozzle, and the power consumed by centrifugal pressurization is as follows: (4) in, Indicates the effective driving power of the jet rotor; This indicates the jet velocity at the propeller tip nozzle. This indicates the gas flow rate required for propeller tip ejection; Indicates the rotor speed; This indicates the distance between the propeller tip jet orifice and the main shaft.

[0049] Furthermore, the gas flow rate required for jet propulsion from the propeller tip nozzle. The expression is: (5) Step 11: Calculate the gas conditions at the inlet of the rotor tip nozzle based on the pressure ratio of the rotor's centrifugal rotation. The gas conditions include gas pressure and gas temperature.

[0050] Furthermore, the expression for the gas pressure at the propeller tip nozzle inlet is: (6) in, This represents the gas pressure at the propeller tip nozzle inlet; This represents the rotor inlet gas pressure; This represents the additional gas dynamic pressure caused by the centrifugal force of the rotor rotation; The total pressure recovery coefficient representing the flow in the rotor duct is preferably taken as 0.9.

[0051] Furthermore, the expression for the gas temperature at the propeller tip nozzle inlet is: (7) in, This represents the gas temperature at the inlet of the propeller tip nozzle; This represents the temperature of the gas entering the rotor. This represents the temperature drop of the gas inside the rotor due to heat dissipation; This represents the specific heat capacity of a gas.

[0052] Furthermore, the additional gas dynamic pressure caused by the centrifugal force of the rotor rotation The expression is: (8) in, Represents gas density.

[0053] Step 12: Calculate the jet velocity at the propeller tip nozzle. The relationship between the inlet gas pressure and the jet velocity is shown in formula (9): (9) in, Represents the total gas pressure at the propeller tip nozzle inlet; The total gas pressure at the nozzle tip is typically taken as the ambient atmospheric pressure. Represents the jet velocity at the nozzle; This represents the critical velocity of the nozzle.

[0054] Step 13: Obtain the power-to-weight ratio and fuel consumption rate of the power system based on the power of the composite power system and the power of the piston engine, and compare them with the design target. Based on the comparison results, adjust the component design parameters until the design target of the power system is met.

[0055] Furthermore, the expression for the power-to-weight ratio of the composite dynamic system is: (10) in, Represents the power output of a piston engine; Represents the total weight of the composite power system, including piston engines and gas turbines.

[0056] Furthermore, the expression for the fuel consumption rate of the hybrid power system is: (11) in, This represents the fuel flow rate of the gas turbine. This represents the fuel flow rate of the piston engine (obtained from step 7).

[0057] Step 14: Determine whether the power-to-weight ratio and fuel consumption rate of the composite power system meet the optimization objectives of the thermodynamic cycle parameters designed in Step 2. If they do, obtain the key parameters of the thermodynamic cycle. If they do not, return to Step 9.

[0058] The tip-jet hybrid propulsion system performance optimization method based on thermodynamic cycle planning disclosed in this invention exhibits significant advantages over traditional piston and turboshaft engines used in helicopters and rotorcraft in terms of power-to-weight ratio and system fuel consumption. Specifically, the power-to-weight ratio of the optimized propulsion system obtained using this invention is higher than 2.5 kW / kg, exceeding that of traditional aero-piston and turboshaft engines. The system's minimum fuel consumption is roughly equivalent to that of a traditional aero-piston engine and lower than the 0.3 kg / (kW·h) of a traditional aero-gas turbine.

[0059] The above description is only a preferred embodiment of the present invention, but the scope of protection of the present invention is not limited thereto. Any changes or substitutions that can be easily conceived by those skilled in the art within the scope of the technology disclosed in the present invention should be included within the scope of protection of the present invention.

Claims

1. A performance optimization method for a tip-jet hybrid propulsion system based on thermodynamic cycle planning, characterized in that, The composite power system includes at least a gas turbine and a jet rotor. The gas turbine includes a compressor, a combustion chamber, and a turbine. The jet rotor includes internal ducts and tip nozzles. The specific steps are as follows: S1. Based on the performance requirements of the aircraft, determine the design targets for the available power and maximum fuel consumption rate of the composite power system; S2. Transform the design objective into an optimization objective for the thermodynamic cycle parameters; S3. Based on the design level of aero-engines and the temperature resistance limit of materials, set the maximum operating temperature of the thermodynamic cycle; under the constraint of the maximum operating temperature of the thermodynamic cycle, identify the key factors affecting the useful work of the cycle; the key factors affecting the useful work of the cycle include the compressor pressure ratio and the turbine pressure ratio; S4. Based on the fuel consumption requirements of the composite power system, the tip nozzle form is determined to be a converging nozzle, and the inlet gas conditions of the tip nozzle are designed. It is also identified that the centrifugal pressure ratio of the jet rotor and the tip jet Mach number are the key parameters affecting the thermal efficiency and fuel consumption of the system. S5. Collaboratively design the key factors and key parameters, and sequentially calculate the inlet and outlet gas state parameters of each component of the composite power system. S6. Based on the calculation results of step S5, calculate the current power-to-weight ratio and current fuel consumption rate of the composite power system; S7. Determine whether the current power-to-weight ratio and current fuel consumption rate meet the optimization objective; If the requirements are not met, the component design parameters are adjusted and the process returns to step S5 for iterative optimization until the optimization objective is met, thereby obtaining the optimized key parameters of the thermodynamic cycle.

2. The method according to claim 1, characterized in that, In step S5, the collaborative design includes: under the condition that the compressor and the turbine meet the preset work balance relationship, through the collaborative design of turbine outlet gas parameters, rotor speed, wing internal pipe size and blade tip nozzle size, the reaction thrust and power generated by the blade tip nozzle meet the system requirements.

3. The method according to claim 2, characterized in that, The work balance relationship is expressed by the following formula: in, and These represent compressor work and turbine work, respectively. and These represent the inlet gas flow rates of the compressor and turbine, respectively. and These represent the adiabatic indices for the gas compression process in the compressor and the gas expansion process in the turbine, respectively. Represents the gas constant; and These represent the inlet gas temperatures of the compressor and turbine, respectively. and These represent the compressor boost ratio and the turbine pressure drop ratio, respectively. , and These represent the compressor efficiency, turbine efficiency, and rotor system mechanical efficiency, respectively.

4. The method according to claim 1, characterized in that, The centrifugal pressure ratio is the ratio of the total pressure at the propeller tip nozzle to the total pressure at the jet rotor inlet, and its value is determined by the rotor speed, the cross-sectional design of the duct inside the wing, and the flow loss in the duct inside the wing.

5. The method according to claim 1, characterized in that, The composite power system also includes a piston engine connected to the rotor main shaft; in step S6, the current power-to-weight ratio is calculated as the ratio of the total output power of the system to the total weight of the piston engine and the gas turbine; the total output power is the sum of the power of the piston engine and the power of the jet rotor.

6. The method according to claim 1, characterized in that, In step S5, the inlet and outlet gas state parameters of each component of the composite power system include the gas flow rate required for the tip jet. The step of calculating the gas flow rate required for the tip jet includes: subtracting the piston engine power from the available power of the composite power system obtained in step 1 to obtain the effective driving power of the jet rotor, and then solving the problem based on the energy conservation relationship between the effective driving power of the jet rotor, the jet power of the tip nozzle, and the power consumed by the gas in the rotor for centrifugal pressurization.

7. The method according to claim 1, characterized in that, In step S5, the inlet and outlet gas state parameters of each component of the composite power system include the total gas pressure and total temperature at the inlet of the propeller tip nozzle, which are calculated based on the centrifugal boost ratio; the jet velocity of the propeller tip nozzle is obtained based on the total gas pressure, total temperature and critical velocity at the nozzle inlet.

8. The method according to claim 1, characterized in that, In step S3, the highest operating temperature of the thermodynamic cycle is the highest temperature of the turbine inlet gas.

9. The method according to claim 1, characterized in that, In step S7, the design parameters of the adjustment components include: adjusting one or more of the following: compressor boost ratio, turbine pressure ratio, rotor speed, cross-sectional area of ​​the wing duct, and throat area of ​​the blade tip nozzle.

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