A design method for hybrid power system architecture of distributed electric propulsion aircraft based on power flow matrix

Abstract

The present invention discloses a method for designing a hybrid power system architecture for a distributed electric propulsion aircraft based on a power flow matrix, belonging to the technical field of overall aircraft design. The present invention directly integrates the architecture design of a hybrid power system for a distributed electric propulsion aircraft with the overall design of the aircraft platform, fully considering the impact of hybrid power system architecture-related parameters on the overall parameters of the aircraft platform. This method establishes a direct link between the hybrid power system architecture, power allocation coefficient, and the efficiency of each component of the distributed electric propulsion aircraft and overall parameters such as the gross weight, speed, climb rate, and altitude of the aircraft. This method can then directly and efficiently iteratively optimize the hybrid power system architecture of the distributed electric propulsion aircraft based on the overall requirements of the aircraft, better addressing the design matching problem between complex hybrid power systems and distributed electric propulsion aircraft platforms.

Description

Technical Field

[0001] The present invention relates to the technical field of overall aircraft design, and more particularly to a method for designing a hybrid power system architecture for a distributed electric propulsion aircraft based on a power flow matrix. Background Art

[0002] The current overconsumption of traditional fossil fuel-based energy has led to serious global environmental pollution. To address this, the aviation industry is actively conducting feasibility studies on the use of renewable energy as the primary propulsion system for aircraft. With the advancement of new energy technologies such as lithium-ion batteries, power electronics, and generators / drive motors, aviation research and development institutions both domestically and internationally are actively exploring innovative distributed electric propulsion aircraft configurations for both military and civilian applications.

[0003] Due to the limitations of the performance of core components of electric propulsion aircraft, such as current battery energy density and motor power density, pure electric propulsion aircraft still have a large gap compared with traditional aircraft platforms in terms of platform payload, flight speed, range, and flight time, making the application scenarios of pure electric propulsion aircraft extremely limited.

[0004] By combining engines with energy storage devices (lithium batteries, generators), hybrid electric propulsion aircraft can adapt to aircraft design requirements for multiple engine states, meeting the aircraft's requirements for high power at takeoff and low fuel consumption at cruise. The engine's operating state is unaffected by environmental changes, maintaining stable operation within the optimal operating range. The engine also exhibits excellent economic and emission performance. Therefore, the overall performance of hybrid electric propulsion aircraft platforms far exceeds that of pure electric propulsion aircraft, and they also offer advantages in range, flight time, fuel economy, and environmental impact compared to traditional energy aircraft platforms.

[0005] Although hybrid electric propulsion aircraft can reduce fuel consumption and improve environmental impact, their energy and power systems are more complex, the overall architectural design has more elements, and the newly added power distribution system increases the total weight of the aircraft, all of which bring more challenges to the overall design of the aircraft.

[0006] Therefore, how to propose a design method for a hybrid power system architecture of a distributed electric propulsion aircraft and achieve design matching between a complex hybrid power system and a distributed electric propulsion aircraft platform is an urgent problem that technical personnel in this field need to solve. Summary of the Invention

[0007] In view of this, the present invention provides a method for designing a hybrid power system architecture for a distributed electric propulsion aircraft, establishing a direct connection between the hybrid power system architecture, power distribution coefficient, efficiency of each component and overall parameters such as the total weight, speed, climb rate, and altitude of the aircraft, and thus directly and efficiently iteratively optimizing the hybrid power system architecture of the distributed electric propulsion aircraft based on the overall requirements of the aircraft.

[0008] In order to achieve the above object, the present invention adopts the following technical solutions:

[0009] The present invention proposes a method for designing a hybrid power system architecture for a distributed electric propulsion aircraft based on a power flow matrix, comprising the following steps:

[0010] Define the typical mission profile of a distributed electric propulsion hybrid aircraft and decompose the typical mission profile into mission segments;

[0011] Calculate the required power matrix for each mission segment;

[0012] Constructing a hybrid power system architecture, defining an ideal power distribution matrix between components based on component connection relationships in the hybrid power system architecture, and defining a power efficiency transfer matrix for each component under different mission segments;

[0013] Calculating a real power distribution matrix between components based on the ideal power distribution matrix and the power efficiency transfer matrix between components;

[0014] The power flow matrix of each component is calculated using the required power matrix of each task segment and the actual power distribution matrix between the components.

[0015] Preferably, the required power matrix P of each task segment is m_segi,req The calculation formula is as follows:

[0016]

[0017] Where μ is the friction coefficient between the landing gear and the ground during takeoff and landing, m is the mass of the distributed electric propulsion hybrid aircraft, θ is the flight path angle of the distributed electric propulsion hybrid aircraft, v is the aircraft speed, h is the flight altitude, and C is the flight path angle of the distributed electric propulsion hybrid aircraft. L is the lift coefficient of the whole aircraft, C D is the drag coefficient of the whole aircraft, ρ is the atmospheric density, S is the wing area, W is the aircraft gravity, g is the acceleration of gravity, v takeoff is the takeoff roll speed, v climb is the climbing speed, v decent is the descent speed, v landing is the landing speed, h takeoff is the takeoff altitude relative to sea level, h curiseis the cruising altitude, h landing is the landing altitude relative to sea level.

[0018] Preferably, the hybrid system architecture includes an energy source, a power source and a thrust source; the energy source includes fuel and a battery; the power source includes an electric motor and a turboshaft engine, and the electric motor includes a generator and a drive motor connected in sequence; the thrust source is a propeller thrust source;

[0019] The propeller thrust source is driven to meet the power requirements of each mission segment; the propeller thrust source is driven by an electric motor respectively; the turboshaft engine and the battery drive the electric motor at the same time; and the fuel provides energy for the turboshaft engine.

[0020] Preferably, the ideal power distribution matrix between components includes: an ideal power distribution matrix between mission required power and thrust source, thrust source and power source, power source and power source, and energy source and power source;

[0021] The mission requires power - the ideal power distribution matrix B between thrust sources req,TS,m_segi The power distribution coefficient of the distributed electric propulsion hybrid aircraft driven by the thrust source is composed;

[0022] The ideal power distribution matrix B between the thrust source and the power source TS,PS,m_segi The elements in include the ideal power distribution coefficient of the motor-driven thrust source;

[0023] The ideal power distribution matrix B between the power sources PS,PS,m_segi The elements in include the ideal power distribution coefficient of the turboshaft engine drive motor;

[0024] The ideal power distribution matrix B between the energy source and the power source PS,ES,m_segi The elements in include the ideal power distribution coefficient for the battery-driven motor.

[0025] Preferably, the power efficiency transfer matrix η of each component under different task segments is component,m_segi The power transfer efficiency η of the turboshaft engine under the corresponding mission segment turboshaft,m_segi , generator power transfer efficiency η gen,m_segi , drive motor power transfer efficiency η motor,m_segi and propeller power transfer efficiency η prop,m_segi constitute.

[0026] Preferably, the power flow matrix of each component includes a thrust source power flow matrix, a direct power source power flow matrix, a total power source power flow matrix and an energy source power flow matrix;

[0027] The directly connected power source power flow matrix represents the power flow matrix of the power source directly connected to the thrust source;

[0028] The total power source power flow matrix represents the power source power flow matrix of all power sources in the hybrid powertrain architecture.

[0029] Preferably, the thrust source power flow matrix A TS,m_segi The calculation formula is:

[0030] A TS,m_segi (i) = P m_segi,req (i) B req,TS,m_segi (i);

[0031] Where A TS,m_segi (i) is the thrust source power flow matrix under the i-th mission segment, B req,TS,m_segi (i) is the ideal power distribution matrix between the mission power and thrust sources in the i-th mission segment, P m_segi_req (i) is the required power matrix under the i-th task segment.

[0032] Preferably, the direct-connected power source power flow matrix The calculation formula is as follows:

[0033]

[0034] Where, is the power flow matrix of the direct power source under the i-th task segment, is the real power distribution matrix between the thrust source and the power source in the i-th mission segment.

[0035] Preferably, the total power source power flow matrix A PS,m_segi It is expressed as follows:

[0036]

[0037] Where A PS,m_segi (i) is the total power source power flow matrix under the i-th task segment, is the real power allocation matrix between power sources under the i-th task segment.

[0038] Preferably, the energy source power flow matrix is ​​expressed as:

[0039]

[0040] Where A ES,m_segi (i) is the energy source power flow matrix under the i-th task segment, is the actual power allocation matrix between energy source and power source under the i-th task segment.

[0041] It can be seen from the above technical solutions that, compared with the prior art, the present invention discloses a method for designing a hybrid power system architecture for a distributed electric propulsion aircraft, which directly integrates the architecture design of the hybrid power system for a distributed electric propulsion aircraft with the overall design of the aircraft platform, fully considers the influence of the parameters related to the hybrid power system architecture on the overall parameters of the aircraft platform, and establishes a direct connection between the hybrid power system architecture, power distribution coefficient, efficiency of each component and overall parameters such as the total weight, speed, climb rate, and altitude of the aircraft. Furthermore, the hybrid power system architecture of the distributed electric propulsion aircraft can be directly and efficiently iteratively optimized according to the overall requirements of the aircraft, thereby better solving the design matching problem between the complex hybrid power system and the distributed electric propulsion aircraft platform. BRIEF DESCRIPTION OF THE DRAWINGS

[0042] In order to more clearly illustrate the embodiments of the present invention or the technical solutions in the prior art, the following briefly introduces the drawings required for use in the embodiments or the description of the prior art. Obviously, the drawings described below are merely embodiments of the present invention. For ordinary technicians in this field, other drawings can be obtained based on the provided drawings without paying any creative work.

[0043] Figure 1 A schematic flow chart of the design method provided by the present invention;

[0044] Figure 2 This is a schematic diagram of a typical mission profile of a distributed electric propulsion hybrid aircraft;

[0045] Figure 3 Schematic diagram of the hybrid power system architecture of a distributed electric propulsion aircraft. DETAILED DESCRIPTION

[0046] The following will clearly and completely describe the technical solutions in the embodiments of the present invention in conjunction with the accompanying drawings. Obviously, the described embodiments are only part of the embodiments of the present invention, not all of the embodiments. Based on the embodiments of the present invention, all other embodiments obtained by ordinary technicians in this field without making creative efforts are within the scope of protection of the present invention.

[0047] The embodiment of the present invention proposes a distributed electric propulsion aircraft hybrid power system architecture design method based on the power flow matrix, referring to Figure 1 , the present invention comprises the following steps:

[0048] S1. Define the typical mission profile of a distributed electric propulsion hybrid aircraft and decompose the typical mission profile into mission segments.

[0049] In this step, based on user requirements and combined with overall aircraft design experience, the core overall design requirement parameters of the distributed electric propulsion hybrid aircraft are preliminarily defined: design gross weight, flight speed, range, mission payload, and flight altitude.

[0050] like Figure 2 As shown in Figure 2, the mission profile of the distributed electric propulsion hybrid aircraft is decomposed and preliminarily defined as the takeoff and run phase, the climb phase, the cruise phase, the descent phase, and the landing phase. The expression is as follows:

[0051] Misson segs =['Takeoff"Climb"Cruise"Descent"Landing'] T (1)

[0052] Among them, Misson segs Indicates the mission segments of a typical mission profile. Takeoff represents the takeoff segment, Climb represents the climb segment, Cruise represents the cruise segment, Descent represents the descent segment, and Landing represents the landing segment.

[0053] According to formula (1) and Figure 2 , further defining the starting and ending heights of each mission stage under a typical mission profile, the function expression is:

[0054] Mission h_begin =[h takeoff h takeoff h cruise h cruise h landing ] T (2)

[0055] Mission h_end =[h takeoff h curise h cruise h landing h landing ] T (3)

[0056] In formula (2) and formula (3), Mission h_begin For each mission segment, the starting height matrix, Mission h_end The ending height matrix for each task segment, h takeoff is the takeoff altitude relative to sea level, h curise is the cruising altitude, h landing is the landing altitude relative to sea level.

[0057] According to formula (1) and Figure 1, further defining the starting and ending speeds of each mission phase under a typical mission profile, the expressions are as follows

[0058] Mission vel_begin =[0 v takeoff v cruise v cruise v landing ] T (4)

[0059] Mission vel_end =[v takeoff v climb v cruise v decent 0] T (5)

[0060] In formula (4) and formula (5), Mission vel_begin For each mission segment, the starting speed matrix is vel_end is the ending velocity matrix for each task segment, v takeoff is the takeoff roll speed, v climb is the climbing speed, v decent is the descent speed, v landing is the landing speed.

[0061] S2. Calculate the required power matrix for each task segment.

[0062] According to the law of conservation of energy, the power required for each mission segment of a distributed electric propulsion hybrid aircraft consists of three parts: the power to overcome flight resistance, the power to overcome flight gravity and climb to a specified altitude, and the power to change speed during flight. Its functional expression is:

[0063]

[0064] In formula (6), T is the total thrust, V is the aircraft speed, D is the drag, h is the flight altitude, W is the aircraft gravity, and g is the acceleration due to gravity.

[0065] According to the definition of formula (6), combined with formula (2)-formula (5), the required power of the distributed electric propulsion hybrid aircraft under different mission segments is given, which can be specifically expressed as:

[0066]

[0067] In formula (7), P m_segi,req is the required power matrix under each mission phase, μ is the friction coefficient between the landing gear and the ground during take-off and landing, m is the mass of the distributed electric propulsion hybrid aircraft, θ is the flight path angle of the distributed electric propulsion hybrid aircraft, v is the aircraft speed, CL is the lift coefficient of the whole aircraft, C D is the drag coefficient of the entire aircraft, ρ is the atmospheric density, and S is the wing area.

[0068] S3. Construct a hybrid power system architecture, define the ideal power distribution matrix between components based on the component connection relationship in the hybrid power system architecture, and define the power efficiency transfer matrix of each component under different mission segments.

[0069] This embodiment uses Figure 3 Taking the hybrid power system architecture of distributed electric propulsion aircraft as an example, a series hybrid distributed electric drive architecture with lithium batteries is defined. The hybrid distributed electric drive architecture includes 2 energy sources (batteries and fuel), 5 power sources (1 turboshaft engine, 4 sets of generators / drive motors), and 4 thrust sources (4 propulsion propellers).

[0070] The relationship between the various components of the hybrid power system of the distributed electric propulsion aircraft is as follows: the distributed electric propulsion hybrid power aircraft is driven by four propeller thrust sources to meet the power requirements of each mission segment; the four propellers are driven by four sets of electric motors respectively; the turboshaft engine and battery drive the four sets of electric motors at the same time; and the fuel provides energy for the turboshaft engine.

[0071] According to the relationship between the components of the hybrid power system of the distributed electric propulsion aircraft, the ideal power distribution matrix between the mission required power and the propeller thrust source is defined as follows:

[0072] B req,TS,m_segi =[λ req,TS1,m_segi λ req,TS2,m_segi λ req,TS3,m_segi λ req,TS4,m_segi ] (8)

[0073] In formula (8), B req,TS,m_segi is a 1×4×5 three-dimensional matrix, which represents the required power of the i-th mission segment and the ideal power distribution matrix between thrust sources, λ req,TS1,m_segi ,λ req,TS2,m_segi ,λ req,TS3,m_segi ,λ req,TS4,m_segi They respectively represent the power distribution coefficients of the first propeller thrust source, the second propeller thrust source, the third propeller thrust source and the fourth propeller thrust source driving the distributed electric propulsion hybrid aircraft under the same i-th mission segment.

[0074] Similarly, the ideal power distribution matrix between the thrust source-power source, power source-power source, and energy source-power source of the hybrid power system of the distributed electric propulsion aircraft can be defined. The specific expression is as follows

[0075]

[0076]

[0077] In formula (9)-formula (11), B TS,PS,m_segi 、B PS,PS,m_segi 、B PS,ES,m_segi are three-dimensional matrices of 4ⅹ5ⅹ5, 5ⅹ5ⅹ5, and 5ⅹ2ⅹ5, respectively, representing the ideal power allocation matrix of thrust source-power source, the ideal power allocation matrix of power source-power source, and the ideal power allocation matrix of energy source-power source for the i-th mission segment; λ TS1,PS2,m_segi ,λ TS2,PS3,m_segi ,λ TS3,PS4,m_segi ,λ TS4,PS5,m_segi They represent the ideal power allocation coefficients for the second power source driving the first thrust source, the third power source driving the second thrust source, the fourth power source driving the third thrust source, and the fifth power source driving the fourth thrust source in the same i-th mission segment; λ PS2,PS1,m_segi ,λ PS3,PS1,m_segi ,λ PS4,PS1,m_segi ,λ PS5,PS1,m_segi They represent the ideal power allocation coefficients of the first power source driving the second power source, the first power source driving the third power source, the first power source driving the fourth power source, and the first power source driving the fifth power source in the same i-th task segment; λ PS2,ES2,m_segi ,λ PS3,ES2,m_segi ,λ PS4,ES2,m_segi ,λ PS5,ES2,m_segi They respectively represent the ideal power distribution coefficients of the second energy source driving the second power source, the second energy source driving the third power source, the second energy source driving the fourth power source, and the second energy source driving the fifth power source in the same i-th task segment.

[0078] Under the same i-th task segment, the allocation coefficients of each power allocation matrix satisfy:

[0079]

[0080] In formula (12), since each thrust source is driven by a single power source, λ TS1,PS2,mission_segi ,λ TS2,PS3,m_segi ,λ TS3,PS4,m_segi ,λ TS4,PS5,m_segi Both are 1.

[0081] The efficiency of each component of the hybrid power system of a distributed electric propulsion aircraft needs to be considered during power transmission, and the efficiency in different mission segments is also different. Therefore, the efficiency matrix of each component of the hybrid power system of a distributed electric propulsion aircraft in different mission segments is defined as follows:

[0082]

[0083] In formula (13), ηcomponent,m_segi is the power transfer efficiency matrix of each component under the i-th task segment, η turboshaft,m_segi ,η gen,m_segi ,η motor,m_segi ,η prop,m_segi They represent the power transfer efficiency of the turboshaft engine, generator, drive motor and propeller in the i-th mission segment respectively.

[0084] S4. Calculate the actual power distribution matrix between the components based on the ideal power distribution matrix and the power efficiency transfer matrix between the components.

[0085] According to the power efficiency transfer matrix of each component of the distributed electric propulsion aircraft hybrid system in formula (13), combined with the ideal power distribution matrix between the thrust source-power source, power source-power source, and energy source-power source of the distributed electric propulsion aircraft hybrid system in formula (9)-formula (11), the real power distribution matrix of the distributed electric propulsion aircraft hybrid system is obtained:

[0086]

[0087]

[0088] In formula (14)-formula (16) They represent the thrust source-power source real power allocation matrix, the power source-power source real power allocation matrix, and the energy source-power source real power allocation matrix of the i-th mission segment respectively.

[0089] S5. Calculate the power flow matrix of each component using the required power matrix of each mission segment and the actual power distribution matrix between components. The power flow matrix of each component includes the thrust source power flow matrix, the direct power source power flow matrix, the total power source power flow matrix, and the energy source power flow matrix.

[0090] The required power of each mission segment of the distributed electric propulsion hybrid aircraft is transferred to the thrust source power flow matrix through the mission required power-propeller thrust source power allocation matrix. According to the power balance analysis, the thrust source power flow matrix A under the i-th mission segment can be obtained. TS,m_segi (i)

[0091] A TS,m_segi (i) = P m_segi,req (i) B req,TS,m_segi (i) (17)

[0092] B req,TS,m_segi (i) is the ideal power distribution matrix between the mission power and thrust sources in the i-th mission segment, P m_segi_req (i) is the required power matrix under the i-th task segment.

[0093] Substituting equations (7) and (8) into equation (17) and expanding them with equation (1) yields

[0094]

[0095]

[0096] In formula (18)-formula (22), A TS,takeoff 、A TS,climb 、A TS,cruise 、A TS,decent 、A TS,landing They represent the thrust source power flow matrices of the distributed electric propulsion aircraft hybrid system during takeoff, climb, cruise, descent, and landing phases.

[0097] The thrust source power of the distributed electric propulsion hybrid aircraft in each mission segment is transferred to the direct power source power flow matrix through the thrust source-power source real power allocation matrix. The direct power source power flow matrix represents the power flow matrix of the power source directly connected to the thrust source. According to the power balance analysis, the direct power source power flow matrix under the i-th mission segment can be obtained. for

[0098]

[0099] In formula (23), is the real power distribution matrix between the thrust source and the power source in the i-th mission segment.

[0100] The power of the direct-connected power source in each mission segment of the distributed electric propulsion hybrid aircraft is transferred to the total power source power flow matrix through the power source-power source real power allocation matrix. The total power source power flow matrix represents the power source power flow matrix of all power source components in the entire hybrid system architecture. According to the power balance analysis, the total power source power flow matrix A under the i-th mission segment can be obtained. PS,m_segi (i) can be expressed as:

[0101]

[0102] Where, is the real power allocation matrix between power sources under the i-th task segment.

[0103] Substitute equations (14), (17)-(22) into equation (23), and substitute the result together with equation (15) into equation (24) to obtain:

[0104]

[0105]

[0106] In formula (25) to formula (29), APS,takeoff 、A PS,climb 、A PS,cruise 、A PS,decent 、A PS,landing They are respectively represented as the total power source power flow matrix of the distributed electric propulsion aircraft hybrid system during takeoff, climb, cruise, descent and landing phases.

[0107] The total power source power of the distributed electric propulsion hybrid aircraft in each mission segment is transferred to the energy source power flow matrix through the power source-energy source real power allocation matrix. According to the power balance analysis, the energy source power flow matrix A under the i-th mission segment can be obtained. ES,m_segi (i) can be expressed as:

[0108]

[0109] Where, is the actual power allocation matrix between energy source and power source under the i-th task segment.

[0110] Substituting equations (16), (25)-(29) into equation (30), we can obtain

[0111]

[0112]

[0113] In formula (31)-formula (35), A ES,takeoff 、A ES,climb 、A ES,cruise 、A ES,decent 、A ES,landing They are respectively represented as the energy source power flow matrices of the distributed electric propulsion aircraft hybrid system during takeoff, climb, cruise, descent, and landing phases.

[0114] It can be seen from the expressions of the thrust source power flow matrix, power source power flow matrix and energy source power flow matrix under each mission segment that the present invention establishes a direct connection between the distributed electric propulsion aircraft hybrid power system architecture, power distribution coefficient, efficiency of each component and overall parameters such as the aircraft's total weight, speed, climb rate, altitude, etc., and thus can directly and efficiently iteratively optimize the distributed electric propulsion aircraft hybrid power system architecture through the overall requirements of the aircraft.

[0115] The various embodiments in this specification are described in a progressive manner, with each embodiment focusing on the differences from other embodiments. Reference can be made to the common and similar parts between the various embodiments. For the devices disclosed in the embodiments, since they correspond to the methods disclosed in the embodiments, the description is relatively simple, and the relevant parts can be referred to the method description.

[0116] The above description of the disclosed embodiments is intended to enable one skilled in the art to implement or use the present invention. Various modifications to these embodiments will be readily apparent to one skilled in the art, and the general principles defined herein may be implemented in other embodiments without departing from the spirit or scope of the present invention. Therefore, the present invention is not limited to the embodiments shown herein but is intended to conform to the widest scope consistent with the principles and novel features disclosed herein.

Claims

1. A method for designing a hybrid power system architecture for a distributed electric propulsion aircraft based on a power flow matrix, characterized in that: The following steps are involved: Define the typical mission profile of a distributed electric propulsion hybrid aircraft and decompose the typical mission profile into mission segments; Calculate the required power matrix for each mission segment; Constructing a hybrid power system architecture, defining an ideal power distribution matrix between components based on component connection relationships in the hybrid power system architecture, and defining a power efficiency transfer matrix for each component under different mission segments; Calculating a real power distribution matrix between components based on the ideal power distribution matrix and the power efficiency transfer matrix between components; The power flow matrix of each component is calculated using the required power matrix of each task segment and the actual power distribution matrix between the components.

2. The method for designing a hybrid power system architecture for a distributed electric propulsion aircraft based on a power flow matrix according to claim 1, characterized in that: The required power matrix P of each task segment m_segi,req The calculation formula is as follows: Where μ is the friction coefficient between the landing gear and the ground during takeoff and landing, m is the mass of the distributed electric propulsion hybrid aircraft, θ is the flight path angle of the distributed electric propulsion hybrid aircraft, v is the aircraft speed, h is the flight altitude, and C is the flight path angle of the distributed electric propulsion hybrid aircraft. L is the lift coefficient of the whole aircraft, C D is the drag coefficient of the whole aircraft, ρ is the atmospheric density, S is the wing area, W is the aircraft gravity, g is the acceleration of gravity, v takeoff is the takeoff roll speed, v climb is the climbing speed, v decent is the descent speed, v landing is the landing speed, h takeoff is the takeoff altitude relative to sea level, h curise is the cruising altitude, h landing is the landing altitude relative to sea level.

3. The method for designing a hybrid power system architecture for a distributed electric propulsion aircraft based on a power flow matrix according to claim 1, characterized in that: The hybrid system architecture includes an energy source, a power source and a thrust source; the energy source includes fuel and a battery; the power source includes an electric motor and a turboshaft engine, and the electric motor includes a generator and a drive motor connected in sequence; the thrust source is a propeller thrust source; The propeller thrust source is driven to meet the power requirements of each mission segment; the propeller thrust source is driven by an electric motor respectively; the turboshaft engine and the battery drive the electric motor at the same time; and the fuel provides energy for the turboshaft engine.

4. The method for designing a hybrid power system architecture for a distributed electric propulsion aircraft based on a power flow matrix according to claim 3, characterized in that: The ideal power distribution matrix between components includes: the ideal power distribution matrix between mission required power and thrust source, thrust source and power source, power source and power source, and energy source and power source; The mission requires power - the ideal power distribution matrix B between thrust sources req,TS,m_segi The power distribution coefficient of the distributed electric propulsion hybrid aircraft driven by the thrust source is composed; The ideal power distribution matrix B between the thrust source and the power source TS,PS,m_segi The elements in include the ideal power distribution coefficient of the motor-driven thrust source; The ideal power distribution matrix B between the power sources PS,PS,m_segi The elements in include the ideal power distribution coefficient of the turboshaft engine drive motor; The ideal power distribution matrix B between the energy source and the power source PS,ES,m_segi The elements in include the ideal power distribution coefficient for the battery-driven motor.

5. The method for designing a hybrid power system architecture for a distributed electric propulsion aircraft based on a power flow matrix according to claim 4, characterized in that: The power efficiency transfer matrix η of each component under different mission segments component,m_segi The power transfer efficiency η of the turboshaft engine under the corresponding mission segment turboshaft,m_segi , generator power transfer efficiency η gen,m_segi , drive motor power transfer efficiency η motor,m_segi and propeller power transfer efficiency η prop,m_segi constitute.

6. The method for designing a hybrid power system architecture for a distributed electric propulsion aircraft based on a power flow matrix according to claim 4, characterized in that: The power flow matrix of each component includes the thrust source power flow matrix, the direct power source power flow matrix, the total power source power flow matrix and the energy source power flow matrix; The directly connected power source power flow matrix represents the power flow matrix of the power source directly connected to the thrust source; The total power source power flow matrix represents the power source power flow matrix of all power sources in the hybrid powertrain architecture.

7. The method for designing a hybrid power system architecture for a distributed electric propulsion aircraft based on a power flow matrix according to claim 6, characterized in that: Thrust source power flow matrix A TS,m_segi The calculation formula is: A TS,m_segi (i)=P m_segi,req (i)·B req,TS,m_segi (i); Where A TS,m_segi (i) is the thrust source power flow matrix under the i-th mission segment, B req,TS,m_segi (i) is the ideal power distribution matrix between the mission power and thrust sources in the i-th mission segment, P m_segi_req (i) is the required power matrix under the i-th task segment.

8. The method for designing a hybrid power system architecture for a distributed electric propulsion aircraft based on a power flow matrix according to claim 7, characterized in that: Directly connected power source power flow matrix The calculation formula is as follows: Where, is the power flow matrix of the direct power source under the i-th task segment, is the real power distribution matrix between the thrust source and the power source in the i-th mission segment.

9. The method for designing a hybrid power system architecture for a distributed electric propulsion aircraft based on a power flow matrix according to claim 8, characterized in that: Total power source power flow matrix A PS,m_segi It is expressed as follows: Where A PS,m_segi (i) is the total power source power flow matrix under the i-th task segment, is the real power allocation matrix between power sources under the i-th task segment.

10. The method for designing a hybrid power system architecture for a distributed electric propulsion aircraft based on a power flow matrix according to claim 9, characterized in that: The energy source power flow matrix is ​​expressed as: Where A ES,m_segi (i) is the energy source power flow matrix under the i-th task segment, is the actual power allocation matrix between energy source and power source under the i-th task segment.

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