Aircraft optimization design method and system and aircraft

Abstract

The invention relates to the technical field of aviation, in particular to an aircraft optimization design method and system and an aircraft. Establishing a relation model of the actuator design parameters and the generalized control force; optimizing a target and a constraint condition; using an optimization algorithm to solve and obtain an optimal actuator parameter; and verifying and comparing the performance of the aircraft with the optimal design parameters and the non-optimal design parameters through a simulation test. By optimizing the parameters of the actuator, the maneuverability and the anti-interference capability during flight are improved. According to the method, the fault-tolerant capability is used as an optimization target at the initial stage of design, and dangerous situations possibly caused by insufficient control capability due to failure are found in advance; the optimization result is verified through a simulation system, and the test risk under the failure condition in the actual flight test is avoided in advance. By the adoption of the optimized configuration, after key faults are injected, the flight attitude is more stable, and therefore the capability of continuing safe flight and landing is remarkably improved.

Description

Technical Field

[0001] This invention relates to the field of aviation technology, and in particular to an aircraft optimization design method, system, and aircraft. Background Technology

[0002] With the development of technology, the design concepts and methods of flight are constantly evolving. Taking the electric vertical take-off and landing (eVTOL) aircraft as an example, as a new type of aircraft, its characteristic is the use of a distributed power system to achieve vertical take-off, landing, and cruise. Unlike traditional aircraft that use only a few actuators, eVTOLs typically carry multiple electrically driven rotors, ducted rotors, or aerodynamic control surfaces, potentially numbering as many as thirty or more.

[0003] On the one hand, this brings extremely high design freedom. Due to the extremely high redundancy of the actuators, designers can optimize the aircraft's control capabilities by adjusting parameters such as the installation position and angle of each actuator without significantly affecting the aircraft's weight and power consumption. For example, adjusting the installation angle of the eVTOL rotor can drastically change its ability to generate yaw moment during vertical takeoff and landing, with almost no impact on the aircraft's weight and power consumption. Because the number of adjustable parameters increases dramatically, relying on the design experience and intuition of traditional models often fails to fully tap the potential of flight performance. Therefore, how to utilize the high redundancy characteristics to coordinately optimize all parameters and achieve optimal flight performance is a technical challenge in this field.

[0004] On the other hand, aircraft safety is often placed at the highest level. Major international airworthiness certification bodies, such as the European Aviation Safety Agency (EASA), stipulate in their airworthiness standards that eVTOLs must be able to tolerate the failure of any single system and maintain safe flight and controlled landing afterward. Among these, "One-Engine-Inoperative" (OEI) is the most common failure mode. When a rotor fails, not only is some lift lost, but attitude oscillations also occur due to torque imbalance. At this time, the aircraft's controllability deteriorates, and the safety margin is significantly reduced. If this situation is not adequately considered in the initial design, the aircraft may suffer a disaster due to insufficient remaining capacity in the event of a failure. However, existing aircraft designs often only consider flight performance under fault-free conditions, and in the face of a failure, they can only maintain attitude balance as much as possible through onboard software. When the aircraft's own capabilities are compromised, this software compensation cannot provide absolute safety. Therefore, this failure mode should be fully considered during the aircraft design phase, ensuring that the aircraft still provides sufficient controllability and safety margin in the event of a failure.

[0005] In summary, there is an urgent need in this field for a design methodology that can proactively optimize a large number of parameters during the configuration design phase and incorporate the critical safety requirement of system failure into the optimization objectives, thereby ensuring that the aircraft possesses optimal control qualities and strong fault tolerance, and guaranteeing flight safety. Summary of the Invention

[0006] The technical problem to be solved by the present invention is to design an aircraft optimization design method, system and aircraft, which aims to solve the problems that the above-mentioned aircraft design methods fail to fully optimize actuator parameters and do not consider the impact of failure on flight control.

[0007] To address the aforementioned technical problems, this invention provides an aircraft optimization design method, specifically comprising the following steps:

[0008] Step S1: Define the actuators used by the aircraft and their design parameters.

[0009] Step S2: Establish a relationship model between actuator design parameters and generalized control force.

[0010] Step S3: Define optimization objectives and constraints based on actuator design parameters and relational models.

[0011] Step S4: Use a global optimization algorithm to solve for and obtain the optimal actuator parameters.

[0012] Step S5: Verify and compare the aircraft performance with optimal actuator design parameters and non-optimal actuator design parameters through simulation experiments.

[0013] Furthermore, in this invention, the control performance during the vertical take-off and landing phase is improved by optimizing the design parameters of the lift system. In step S1, the design parameters of the actuator are: the installation position and installation angle of the lift system, or the propeller size of the lift system.

[0014] Furthermore, in this invention, the actuator includes multiple lift systems, and the actuator parameters are specifically defined as follows:

[0015] Define the body coordinate system: with the aircraft's center of gravity as the origin. The axis is forward along the direction of the machine head. The axis is to the right along the wing direction. The axis is perpendicular to The axis is pointing downwards;

[0016] The installation location of each lift system is represented in the body coordinate system as follows: , The first individual lifting systems , , The position coordinates of the axis;

[0017] The installation angle of each lift system is defined as follows: , , The first A lifting system around shaft and Angle of the axis;

[0018] The diameter of the propeller blades in each lift system is defined as: , For the first The diameter of the propeller blades in the lift system.

[0019] Furthermore, in step S2 of this invention, the generalized control force represents the resultant force and resultant torque that control the motion of the entire aircraft within its airframe. During the vertical takeoff and landing phase, the resultant force and resultant torque that control the aircraft's motion are respectively the force controlling the aircraft's orbital motion. Rolling torque of shaft rotation Control the aircraft to orbit Pitching moment of shaft rotation Control the aircraft to orbit Yaw moment of shaft rotation and control the aircraft along Lift force of the axis moving up and down Generalized control force vector The system control input is Constructing a generalized control force vector With control input The relationship between the control matrix B and the efficiency matrix W, where:

[0020] , Indicates the first The roll, pitch, and yaw moments, as well as the lift generated by the lift system within the aircraft system. , Indicates the first The rotation direction of the lift system is along The positive axis is 1 for clockwise and -1 for counterclockwise. The conversion factor from tension to torque. The diameter of the propeller;

[0021] It is a diagonal matrix, and the numbers on the diagonal are... Indicates the first The efficiency of the power system is 1 at the maximum, indicating normal operation, and 0 at the minimum, indicating complete failure.

[0022] The number for the lift system. This represents the number of lift systems.

[0023] Furthermore, in step S3 of this invention, the design parameters of the actuator are based on the installation position of the lift system. With installation angle The system's control input Defined as a vector consisting of the tension forces independently generated by each lift system. , The tension in each lift system, , Let be the number of lift systems, then the generalized control force vector With control input The relationship is expressed as:

[0024] ;

[0025] in: , This represents the maximum tension of a single lift system.

[0026] Furthermore, in step S3 of this invention, the installation position of the lift system is optimized by design. With installation angle The optimal objective is to minimize the maximum tension of the lift system during grading after any single failure occurs.

[0027]

[0028] in, Indicates through design The value achieves the optimization goal. It is the objective function value. For the first The vector of trim results after the failure of a lift system corresponds to the optimization objective of minimizing the maximum value in the trim results under all single failure scenarios.

[0029] The constraints are:

[0030] (1) The aircraft can be trimmed after any single failure, and Each component in All meet ,Right now:

[0031] , For the first After the failure of the first lift system, the second The tension of the lifting system;

[0032] (2) The maximum yaw moment shall not be less than 10% of the maximum roll moment, that is: ,in:

[0033] This represents the maximum yaw moment that the entire lift system can produce. The third row of matrix B is the first... The summation symbol represents the summation of all elements in the third row. ;

[0034] This represents the maximum rolling torque that the entire lift system can produce. The first row of matrix B is the first... The summation symbol represents the summation of all elements in the first row. ;

[0035] (3) The installation position of the lift system is within ± 0.1 m of the initial value, that is: , These are the initial design values ​​for the installation location of the lift system;

[0036] (4) The installation angle of the lift system is within ±15° in the vertical direction, that is: .

[0037] Furthermore, in step S3 of this invention, the design parameters of the actuator are selected based on the propeller size of the lift system and the control input of the system. The square of the system's rotational speed. Then the generalized control force vector With control input The relationship is expressed as: ,in:

[0038] For each power system, the tensile force coefficient is used.

[0039] air density;

[0040] , For the first The diameter of the propeller blades in the propulsion system.

[0041] Furthermore, in step S3 of this invention, the propeller size is optimized with the trim power after each flight stage or failure as the optimization objective. The optimal objective is to minimize the power of the case with the highest power among all single failure scenarios, i.e.:

[0042] ;

[0043] in, Indicates through design The value achieves the optimization goal. It is the objective function value. For the first Power after the failure of a lift system The calculation method is as follows:

[0044]

[0045] This represents the propeller's torque coefficient. For the first In the event of failure of the first power system, the first The rotational speed of each power system is given by the summation symbol, which represents the summation of power over all power systems. ;

[0046] The constraints are:

[0047] (1) The aircraft can be trimmed after any single failure, and Each component in All meet ,Right now:

[0048] , For the first After the failure of the first lift system, the second The square of the rotational speed of the lift system The square of the maximum speed;

[0049] (2) The maximum yaw moment shall not be less than 10% of the maximum roll moment, that is: ,in:

[0050] This represents the maximum yaw moment that the entire lift system can produce. The third row of matrix B is the first... The summation symbol represents the summation of all elements in the third row. ;

[0051] This represents the maximum rolling torque that the entire lift system can produce. The first row of matrix B is the first... The summation symbol represents the summation of all elements in the first row. ;

[0052] (3) , and These represent the minimum and maximum values ​​of the propeller blade diameter, which are limited by the size of the aircraft.

[0053] Furthermore, in step S4 of this invention, the global optimization algorithm is any one of a genetic algorithm, a pattern search algorithm, or an alternative optimization algorithm.

[0054] The present invention also provides an aircraft optimization design system, employing the aforementioned aircraft optimization design method, comprising:

[0055] The actuator parameter definition module is used to define the body coordinate system and optimize design parameters. The optimized design parameters include at least one of the actuator installation position parameters, installation angle parameters, and propeller blade diameter parameters, and divide the actuator into variable design groups and restricted design groups.

[0056] Model building module: used to establish a generalized control force model with the effects of actuator failure based on optimized design parameters. The generalized control force model establishes a mapping relationship between the generalized control force vector and the system control input vector through the control matrix and efficiency matrix. The control matrix is ​​used to represent the roll moment, pitch moment, yaw moment and lift generated by the lift system within the machine system. The efficiency matrix is ​​used to quantify the normal operation or single failure state of the actuator.

[0057] The optimization configuration module is used to preset the optimization objective function and constraints. The objective function is to minimize the maximum tension or maximum power during balancing after a single actuator failure. The constraints include maneuverability constraints, actuator parameter adjustment range constraints, and balancing constraints after failure.

[0058] The optimal solution module uses a global optimization algorithm to solve the objective function and outputs the optimal actuator design parameters.

[0059] The results verification module is used to call the simulation system to verify the optimal actuator design parameters, compare the trim performance and maneuverability of the aircraft before and after optimization in normal state and single actuator failure state, and confirm the optimization results.

[0060] Furthermore, in this invention, the simulation system in the result verification module includes an aircraft mechanics and control model unit, a parameter selection unit, a failure injection unit, and a flight status display unit. The aircraft mechanics and control model unit includes actuator parameters and a generalized control force model. The parameter selection unit is used to switch actuator design parameters as needed. The failure injection unit is used to inject failure states as needed. The flight status display unit is used to display simulation results and provide verification of optimization results.

[0061] The present invention also provides an aircraft, including a fuselage, multiple lift systems, and multiple thrust systems. The fuselage is provided with wings and a tail, both of which are equipped with aerodynamic control surfaces. The tail is equipped with a lift system, and the wings are equipped with multiple motor arms, each mounted with a lift system. The tail section of the fuselage is equipped with a thrust system. The installation positions and angles of the lift systems are as described above for the installation positions of lift systems. With installation angle Setting the design optimization results of the aircraft optimization design method as actuator design parameters.

[0062] The present invention also provides another type of aircraft, including a fuselage, multiple lift systems and multiple thrust systems. The fuselage is provided with wings and a tail, and both the wings and the tail are provided with aerodynamic control surfaces. The tail is provided with a lift system, and the wings are provided with multiple motor arms. The motor arms are equipped with lift systems. The tail of the fuselage is provided with a thrust system. The diameter of the propeller blades of the lift system is set according to the design optimization results of the aforementioned aircraft optimization design method that uses the propeller size of the lift system as an actuator design parameter.

[0063] Compared with the prior art, the present invention has the following beneficial effects:

[0064] (1) Improved maneuverability: This invention improves maneuverability and anti-interference capability during flight by optimizing actuator parameters.

[0065] (2) Enhanced safety: Increased safety margin after encountering critical actuator failure, ensuring safe flight even in the event of actuator failure. With the optimized configuration of this invention, the flight attitude is more stable after a critical fault is injected, thereby significantly improving the ability to continue safe flight and landing.

[0066] (3) Pre-design: Fault tolerance is taken as the optimization target in the early stage of design, and dangerous situations that may be caused by insufficient control due to failure are discovered in advance; avoid relying on the dangerous situations and design changes caused by problems exposed in the later stage of testing, shorten the R&D cycle and save costs.

[0067] (4) Automated design: It provides an automated design process that replaces the design method that relies on experience or manual trial and error, and can find non-intuitive optimal solutions for highly redundant and complex aircraft configurations.

[0068] (5) Pre-testing: Verify the optimization results through the simulation system to avoid test risks in the event of failure during actual flight testing. Attached Figure Description

[0069] The specific embodiments of the present invention will be further explained below with reference to the accompanying drawings.

[0070] Figure 1 This is a flowchart of the aircraft optimization design method of the present invention.

[0071] Figure 2 This is a schematic diagram showing the coordinate system and lift system numbering in the aircraft optimization design method of the present invention.

[0072] Figure 3 This is a system block diagram of the aircraft optimization design system of the present invention.

[0073] Figure 4 This is a system block diagram of the simulation system in the result verification module of the aircraft optimization design system of the present invention.

[0074] Figure 5 This is a schematic diagram of the structure of the aircraft of the present invention. Detailed Implementation

[0075] Example 1

[0076] Combination Figure 1 As shown, the aircraft optimization design method in this embodiment specifically includes the following steps:

[0077] Step S1: Define the actuators used by the aircraft and their design parameters.

[0078] In this embodiment, the control performance during the vertical take-off and landing phase is improved by optimizing the design parameters of the lifting system. In step S1, the design parameters of the actuator are: the installation position and installation angle of the lifting system, or the propeller size of the lifting system.

[0079] In this embodiment, preferably, the actuator includes multiple lift systems, and the actuator parameters are specifically defined as follows:

[0080] Define the body coordinate system: such as Figure 2 As shown, with the aircraft's center of gravity as the origin, The axis is forward along the direction of the machine head. The axis is to the right along the wing direction. The axis is perpendicular to The axis is pointing downwards;

[0081] The installation location of each lift system is represented in the body coordinate system as follows: , The first individual lifting systems , , The position coordinates of the axis;

[0082] The installation angle of each lift system is defined as follows: , , The first A lifting system around shaft and Angle of the axis;

[0083] The diameter of the propeller blades in each lift system is defined as: , For the first The diameter of the propeller blades in the lift system.

[0084] For the aforementioned aircraft, it is for axis- The plane formed by the axes possesses symmetry; the dimensions of the wings and tail are constrained by aerodynamic design, therefore the various lift systems... axis, The position of the axis (i.e.) () is a fixed value. For example... Figure 2 As shown, this embodiment takes 12 lift systems as an example, numbered 1-12. The 11th and 12th lift systems installed on the tail fin are... The position along the axis is also fixed. Due to the limitations of the structural design in this embodiment, each lift system rotates around... The rotation angle of the shaft can only be 0, that is... .

[0085] Therefore, lift systems can be divided into two groups: the design variable group and the design constrained group. (Definition set) To allow along The set of lift system numbers for shaft adjustment positions, specific to this embodiment. Define a set To allow along The set of numbers for the lifting systems that rotate on the shaft, for this embodiment. .

[0086] Step S2: Establish a relationship model between actuator design parameters and generalized control force.

[0087] In this embodiment, preferably, in step S2, the generalized control force represents the resultant force and resultant torque that control the motion of the entire aircraft within its airframe. During the vertical takeoff and landing phase, the resultant force and resultant torque that control the aircraft's motion are respectively the force controlling the aircraft's orbital motion. Rolling torque of shaft rotation Control the aircraft to orbit Pitching moment of shaft rotation Control the aircraft to orbit Yaw moment of shaft rotation and control the aircraft along Lift force of the axis moving up and down Generalized control force vector The system control input is Constructing a generalized control force vector With control input The relationship between the control matrix B and the efficiency matrix W, where:

[0088] , Indicates the first The roll, pitch, and yaw moments, as well as the lift generated by the lift system within the aircraft system. , Indicates the first The rotation direction of the lift system is along The positive axis is 1 for clockwise and -1 for counterclockwise. The conversion factor from tension to torque. This is the propeller diameter. Of all the parameters, only... and This refers to the design freedom provided in this embodiment, or the parameters to be optimized. Because each lift system revolves around... Rotation angle of the shaft It can only be 0. It can be further simplified to:

[0089]

[0090] It is a diagonal matrix, and the numbers on the diagonal are... Indicates the first The efficiency of the power system is 1 at the maximum, indicating normal operation, and 0 at the minimum, indicating complete failure.

[0091] The number for the lift system. This represents the number of lift systems.

[0092] Control Matrix With efficiency matrix A relationship between the thrust of a single lift system and the generalized control force was established. This matrix is ​​influenced by actuator parameters; therefore, actuator parameters directly affect the magnitude of the generalized control force, which is the direct factor affecting the aircraft's control over its attitude and altitude changes. Thus, the aircraft's control capability can be improved by modifying actuator parameters.

[0093] Step S3: Define optimization objectives and constraints based on actuator design parameters and relational models.

[0094] In this embodiment, preferably, in step S3, the design parameters of the actuator are based on the installation position of the lift system. With installation angle The system's control input Defined as a vector consisting of the tension forces independently generated by each lift system. , The tension in each lift system, , Let be the number of lift systems, then the generalized control force vector With control input The relationship is expressed as:

[0095] ;

[0096] in: , This represents the maximum tension of a single lift system.

[0097] This embodiment improves control performance during the vertical takeoff and landing (VTOL) phase by optimizing the lift system parameters, with a particular focus on trim capability and maneuverability during this phase. Trim is defined as the lift provided by the aircraft counteracting gravity and maintaining zero torque in all directions. According to this definition, during trim, the aircraft can stably maintain its altitude and attitude. Based on the generalized control force model, its expression is:

[0098]

[0099] The fact that the aircraft can be trimmed means that the equation has a solution. Makes the equation true, and Each component in All meet Clearly, does the equation have a solution that satisfies the conditions? Received The influence of this matrix is ​​due to the design parameters and failure scenarios of the actuators. For example, in the extreme case where the entire lift system fails... Therefore, the equation has no solution. During aircraft leveling, it is crucial to ensure that the output of each lift system is uniform, avoiding situations where the thrust of one lift system is significantly higher than that of others, leading to overheating and failure. Especially when some lift systems fail, the loss of lift and the resulting unbalanced torque cause other lift systems to compensate, resulting in a significant increase in the thrust of those systems and further overheating failure. This situation should be avoided as much as possible when designing actuator parameters.

[0100] The maneuverability of an aircraft is defined as the maximum torque it can generate in all directions. For the aircraft in this embodiment, its ability to generate yaw moment is relatively weak; therefore, its yaw moment must be guaranteed to be no less than a certain threshold. Specifically, in this embodiment, the selected design constraint is that the yaw capability is no less than 10% of its roll capability. Furthermore, to ensure that the aircraft's center of gravity and strength requirements do not change significantly, the installation position and angle of each lift system should be subject to certain restrictions.

[0101] In this embodiment, preferably, in step S3, the installation position of the lift system is optimized. With installation angle The optimal objective is to minimize the maximum tension of the lift system during grading after any single failure occurs.

[0102]

[0103] in, Indicates through design The value achieves the optimization goal. It is the objective function value. For the first The vector of trim results after the failure of a lift system corresponds to the optimization objective of minimizing the maximum value in the trim results under all single failure scenarios.

[0104] The constraints are:

[0105] (1) The aircraft can be trimmed after any single failure, and Each component in All meet ,Right now:

[0106] , For the first After the failure of the first lift system, the second The tension of the lifting system;

[0107] (2) The maximum yaw moment shall not be less than 10% of the maximum roll moment, that is: ,in:

[0108] This represents the maximum yaw moment that the entire lift system can produce. The third row of matrix B is the first... The summation symbol represents the summation of all elements in the third row. ;

[0109] This represents the maximum rolling torque that the entire lift system can produce. The first row of matrix B is the first... The summation symbol represents the summation of all elements in the first row. ;

[0110] (3) The installation position of the lift system is within ± 0.1 m of the initial value, that is: , These are the initial design values ​​for the installation location of the lift system;

[0111] (4) The installation angle of the lift system is within ±15° in the vertical direction, that is: .

[0112] Step S4: Use a global optimization algorithm to solve for and obtain the optimal actuator parameters. Since the expressions for the optimization objective and constraints are complex and nonlinear, and the number of parameters to be optimized is extremely large, it is impossible to obtain the optimal solution through manual trial and error. Therefore, a global optimization algorithm should be used to solve for the optimal solution.

[0113] In this embodiment, preferably, in step S4, the global optimization algorithm is any one of a genetic algorithm, a pattern search algorithm, or an alternative optimization algorithm. Specifically, in this embodiment, a genetic algorithm is used to solve the problem based on the optimization toolbox provided by MATLAB.

[0114] Step S5: Verify and compare the aircraft performance with optimal actuator design parameters and non-optimal actuator design parameters through simulation experiments.

[0115] To confirm that the optimization results met expectations, simulation was used to verify the results after the optimization calculations were completed. During verification, to facilitate comparison of optimization effects, parameter switching and failure injection were performed, and the flight simulation results were directly compared based on the flight status display. Since flight parameters are directly reflected in changes to the control matrix B, parameter switching was performed by selecting the B matrix before and after optimization using signal selection. Failure injection was performed by changing the actuator efficiency by setting the diagonal value of the W matrix; for example, when there are no failures, the W matrix is... The identity matrix (all diagonal elements are 1, and all other elements are 0) can be completely disabled by setting the first element of the first row and first column of the W matrix to 0.

[0116] Example 2

[0117] The difference between the aircraft optimization design method in this embodiment and the aircraft optimization design method in Embodiment 1 lies in the different actuator design parameters. Correspondingly, the generalized control force model, optimization objective, and constraints also differ. The specific differences are as follows:

[0118] In this embodiment, preferably, in step S3, the design parameters of the actuator are selected to determine the propeller size of the lift system, and the system control input... The square of the system's rotational speed. Then the generalized control force vector With control input The relationship is expressed as:

[0119]

[0120] in:

[0121] For each power system, the tensile force coefficient is used.

[0122] air density;

[0123] , For the first The diameter of the propeller blades in the propulsion system .

[0124] The parameters to be optimized are Similar to Example 1, the set of designable optimizations should be defined based on the aerodynamic characteristics and structural design constraints of the aircraft. With The range of selectable values.

[0125] In this embodiment, preferably, in step S3, the propeller size is optimized with the trim power of each flight stage or after failure as the optimization objective. The optimal objective is to minimize the power of the case with the highest power among all single failure scenarios, that is:

[0126] ;

[0127] in, Indicates through design The value achieves the optimization goal. It is the objective function value. For the first Power after the failure of a lift system The calculation method is as follows:

[0128]

[0129] This represents the propeller's torque coefficient. For the first In the event of failure of the first power system, the first The rotational speed of each power system is given by the summation symbol, which represents the summation of power over all power systems. ;

[0130] The constraints are:

[0131] (1) The aircraft can be trimmed after any single failure, and Each component in All meet ,Right now:

[0132] , For the first After the failure of the first lift system, the second The square of the rotational speed of the lift system The square of the maximum speed;

[0133] (2) The maximum yaw moment shall not be less than 10% of the maximum roll moment, that is: ,in:

[0134] This represents the maximum yaw moment that the entire lift system can produce. The third row of matrix B is the first... The summation symbol represents the summation of all elements in the third row. ;

[0135] This represents the maximum rolling torque that the entire lift system can produce. The first row of matrix B is the first... The summation symbol represents the summation of all elements in the first row. ;

[0136] (3) , and These represent the minimum and maximum values ​​of the propeller blade diameter, which are limited by the size of the aircraft.

[0137] Specifically in this embodiment, the sum of the diameters of the propeller blades of the lift system on the side of the motor arm near the front of the aircraft on one side of the wing is no greater than 0.4 times the wingspan. For example... Figure 2 As shown, taking n=12 as an example, the sum of the diameters of the propeller blades of the three lift systems at the leading edge of the left wing and the three lift systems at the leading edge of the right wing is no greater than 40% of the wingspan. , , This refers to the wingspan.

[0138] The remaining technical features in this embodiment are the same as those in Embodiment 1, and will not be repeated here.

[0139] Example 3

[0140] Combination Figure 3 As shown, the aircraft optimization design system in this embodiment adopts the aircraft optimization design method as described in Embodiment 1, including:

[0141] The actuator parameter definition module is used to define the body coordinate system and optimize design parameters. The optimized design parameters include at least one of the actuator installation position parameters, installation angle parameters, and propeller blade diameter parameters, and divide the actuator into variable design groups and restricted design groups.

[0142] Model building module: used to establish a generalized control force model with the effects of actuator failure based on optimized design parameters. The generalized control force model establishes a mapping relationship between the generalized control force vector and the system control input vector through the control matrix and efficiency matrix. The control matrix is ​​used to represent the roll moment, pitch moment, yaw moment and lift generated by the lift system within the machine system. The efficiency matrix is ​​used to quantify the normal operation or single failure state of the actuator.

[0143] The optimization configuration module is used to preset the optimization objective function and constraints. The objective function is to minimize the maximum tension or maximum power during balancing after a single actuator failure. The constraints include maneuverability constraints, actuator parameter adjustment range constraints, and balancing constraints after failure.

[0144] The optimal solution module uses a global optimization algorithm to solve the objective function and outputs the optimal actuator design parameters.

[0145] The results verification module is used to call the simulation system to verify the optimal actuator design parameters, compare the trim performance and maneuverability of the aircraft before and after optimization in normal state and single actuator failure state, and confirm the optimization results.

[0146] In this embodiment, preferably, combined with Figure 4 As shown, the simulation system in the result verification module includes an aircraft mechanics and control model unit, a parameter selection unit, a failure injection unit, and a flight status display unit. The aircraft mechanics and control model unit includes actuator parameters and a generalized control force model. The parameter selection unit is used to switch actuator design parameters as needed. The failure injection unit is used to inject failure states as needed. The flight status display unit is used to display simulation results and provide verification of optimization results.

[0147] Example 4

[0148] Combination Figure 5As shown, the aircraft in this embodiment includes a fuselage 100, multiple lift systems 700, and multiple thrust systems 400. The fuselage 100 has wings 200 and a tail 300. Both the wings 200 and tail 300 have aerodynamic control surfaces 600. The tail 300 has a lift system 700. Multiple motor arms 500 are mounted on the wings 200, and each motor arm 500 has a lift system 700 installed on it. Specifically, in this embodiment, the two motor arms 500 on both sides of the fuselage 100 have a lift system 700 at one end closer to the forward direction of the fuselage 100, and a thrust system 400 at the end farther from the forward direction of the fuselage 100. The tail of the fuselage 100 has a thrust system 400, and the remaining motor arms 500 have lift systems 700 at both ends. The installation position and angle of the lift systems are set according to the design optimization results of the aircraft optimization design method in Embodiment 1, i.e., the installation position of the lift systems is adopted. With installation angle As a parameter for actuator design.

[0149] Example 5

[0150] The difference between the aircraft in this embodiment and the aircraft in embodiment 4 is that the aircraft in this embodiment uses the propeller size of the lift system as the actuator design parameter, that is, the diameter of the propeller blades of the lift system. The design optimization results are set according to the aircraft optimization design method in Example 2. The other technical features are the same and will not be repeated here.

[0151] Many specific details have been set forth in the foregoing description to provide a thorough understanding of the present invention. However, the above description is merely a preferred embodiment of the present invention, and the present invention can be implemented in many other ways different from those described herein. Therefore, the present invention is not limited to the specific embodiments disclosed above. Furthermore, any person skilled in the art can make many possible variations and modifications to the technical solutions of the present invention, or modify them into equivalent embodiments, using the methods and techniques disclosed above, without departing from the scope of the present invention. Any simple modifications, equivalent changes, and modifications made to the above embodiments based on the technical essence of the present invention, without departing from the content of the present invention, shall still fall within the protection scope of the present invention.

Claims

1. An aircraft optimization design method, characterized in that: Includes the following steps: Step S1: Define the actuators used by the aircraft and their design parameters; Step S2: Establish a model relating actuator design parameters to generalized control force; Step S3: Define optimization objectives and constraints based on actuator design parameters and relational models; Step S4: Use a global optimization algorithm to solve for and obtain the optimal actuator parameters; Step S5: Verify and compare the aircraft performance with optimal actuator design parameters and non-optimal actuator design parameters through simulation experiments.

2. The aircraft optimization design method according to claim 1, characterized in that: By optimizing the design parameters of the lift system to improve the control performance during vertical take-off and landing, in step S1, the design parameters of the actuator are: the installation position and installation angle of the lift system, or the propeller size of the lift system.

3. The aircraft optimization design method according to claim 2, characterized in that: The actuator includes multiple lift systems, and the actuator parameters are specifically defined as follows: Define the body coordinate system: with the aircraft's center of gravity as the origin. The axis is forward along the direction of the machine head. The axis is to the right along the wing direction. The axis is perpendicular to The axis is pointing downwards; The installation location of each lift system is represented in the body coordinate system as follows: , The first individual lifting systems , , The position coordinates of the axis; The installation angle of each lift system is defined as follows: , , The first A lifting system around shaft and Angle of the axis; The diameter of the propeller blades in each lift system is defined as: , For the first The diameter of the propeller blades in the lift system.

4. The aircraft optimization design method according to claim 3, characterized in that: In step S2, the generalized control force represents the resultant force and resultant torque that control the motion of the entire aircraft within its airframe. During the vertical takeoff and landing phase, the resultant force and resultant torque that control the aircraft's motion are respectively the force controlling the aircraft's orbital motion. Rolling torque of shaft rotation Control the aircraft to orbit Pitching moment of axis rotation Control the aircraft to orbit Yaw moment of shaft rotation and control the aircraft along Lift force of the axis moving up and down Generalized control force vector The system control input is Constructing a generalized control force vector With control input The relationship between the control matrix B and the efficiency matrix W, where: , Indicates the first The roll, pitch, and yaw moments, as well as the lift generated by the lift system within the aircraft system. , Indicates the first The rotation direction of the lift system is along The positive axis is 1 for clockwise and -1 for counterclockwise. The conversion factor from tension to torque. The diameter of the propeller; It is a diagonal matrix, and the numbers on the diagonal are... Indicates the first The efficiency of the power system is 1 at the maximum, indicating normal operation, and 0 at the minimum, indicating complete failure. The number for the lift system. This represents the number of lift systems.

5. The aircraft optimization design method according to claim 4, characterized in that: In step S3, the design parameters of the actuator are based on the installation position of the lift system. With installation angle The system's control input Defined as a vector consisting of the tension forces independently generated by each lift system. , The tension in each lift system, , Let be the number of lift systems, then the generalized control force vector With control input The relationship is expressed as: ； in: , This represents the maximum tension of a single lift system.

6. The aircraft optimization design method according to claim 5, characterized in that: In step S3, the installation position of the lift system is optimized. With installation angle The optimal objective is to minimize the maximum tension of the lift system during grading after any single failure occurs. in, Indicates through design The value achieves the optimization goal. It is the objective function value. For the first The vector of trim results after the failure of a lift system corresponds to the optimization objective of minimizing the maximum value in the trim results under all single failure scenarios. The constraints are: (1) The aircraft can be trimmed after any single failure, and Each component in All meet ,Right now: , For the first After the failure of the first lift system, the second The tension of the lifting system; (2) The maximum yaw moment shall not be less than 10% of the maximum roll moment, that is: ,in: This represents the maximum yaw moment that the entire lift system can produce. The third row of matrix B is the first... The summation symbol represents the summation of all elements in the third row. ; This represents the maximum rolling torque that the entire lift system can produce. The first row of matrix B is the first... The summation symbol represents the summation of all elements in the first row. ; (3) The installation position of the lift system is within ± 0.1 m of the initial value, that is: , These are the initial design values ​​for the installation location of the lift system; (4) The installation angle of the lift system is within ±15° in the vertical direction, that is: .

7. The aircraft optimization design method according to claim 4, characterized in that: In step S3, the design parameters of the actuator are selected to determine the propeller size of the lift system and the control input of the system. The square of the system's rotational speed. Then the generalized control force vector With control input The relationship is expressed as: ,in: For each power system, the tensile force coefficient is used. air density; , For the first The diameter of the propeller blades in the propulsion system.

8. The aircraft optimization design method according to claim 7, characterized in that: In step S3, the propeller size is optimized with the trim power after each flight stage or failure as the optimization objective. The optimal objective is to minimize the power of the case with the highest power among all single failure scenarios, i.e.: ； in, Indicates through design The value achieves the optimization goal. It is the objective function value. For the first Power after the failure of a lift system The calculation method is as follows: ； This represents the propeller's torque coefficient. For the first In the event of failure of the first power system, the first The rotational speed of each power system is given by the summation symbol, which represents the summation of power over all power systems. ; The constraints are: (1) The aircraft can be trimmed after any single failure, and Each component in All meet ,Right now: , For the first After the failure of the first lift system, the second The square of the rotational speed of the lift system The square of the maximum speed; (2) The maximum yaw moment shall not be less than 10% of the maximum roll moment, that is: ,in: This represents the maximum yaw moment that the entire lift system can produce. The third row of matrix B is the first... The summation symbol represents the summation of all elements in the third row. ; This represents the maximum rolling torque that the entire lift system can produce. The first row of matrix B is the first... The summation symbol represents the summation of all elements in the first row. ; (3) , and These represent the minimum and maximum values ​​of the propeller blade diameter, which are limited by the size of the aircraft.

9. The aircraft optimization design method according to claim 1, characterized in that: In step S4, the global optimization algorithm is any one of genetic algorithm, pattern search algorithm, or alternative optimization algorithm.

10. An aircraft optimization design system, characterized in that: The aircraft optimization design method as described in any one of claims 1-9 includes: The actuator parameter definition module is used to define the body coordinate system and optimize the design parameters. The optimized design parameters include at least one of the actuator installation position parameters, installation angle parameters and propeller blade diameter parameters, and divide the actuator into variable design groups and restricted design groups. Model building module: used to build a generalized control force model with the effects of actuator failure based on optimized design parameters. The generalized control force model establishes a mapping relationship between the generalized control force vector and the system control input vector through the control matrix and efficiency matrix. The control matrix is ​​used to represent the roll moment, pitch moment, yaw moment and lift generated by the lift system in the machine system. The efficiency matrix is ​​used to quantify the normal operation or single failure state of the actuator. The optimization configuration module is used to preset the optimization objective function and constraints. The objective function is to minimize the maximum tension or maximum power during the balancing process after a single actuator failure. The constraints include maneuverability constraints, actuator parameter adjustment range constraints, and balancing constraints after failure. The optimal solution module uses a global optimization algorithm to solve the objective function and outputs the optimal actuator design parameters; The results verification module is used to call the simulation system to verify the optimal actuator design parameters, compare the trim performance and maneuverability of the aircraft before and after optimization in normal state and single actuator failure state, and confirm the optimization results.

11. The aircraft optimization design system according to claim 10, characterized in that: The simulation system in the result verification module includes an aircraft mechanics and control model unit, a parameter selection unit, a failure injection unit, and a flight status display unit. The aircraft mechanics and control model unit includes actuator parameters and a generalized control force model. The parameter selection unit is used to switch actuator design parameters as needed. The failure injection unit is used to inject failure states as needed. The flight status display unit is used to display simulation results and provide verification of optimization results.

12. An aircraft, characterized in that: The aircraft includes a fuselage (100), multiple lift systems (700) and multiple thrust systems (400). The fuselage (100) is provided with wings (200) and tail (300). Both the wings (200) and tail (300) are provided with aerodynamic control surfaces (600). The tail (300) is provided with a lift system (700). The wings (200) are provided with multiple motor arms (500). The motor arms (500) are equipped with lift systems (700). The tail of the fuselage (100) is provided with a thrust system (400). The installation position and installation angle of the lift system (700) are set according to the design optimization results of the aircraft optimization design method according to any one of claims 1, 2, 3, 4, 5, 6 or 9.

13. An aircraft, characterized in that: The aircraft includes a fuselage (100), multiple lift systems (700) and multiple thrust systems (400). The fuselage (100) is provided with wings (200) and tail (300). Both the wings (200) and tail (300) are provided with aerodynamic control surfaces (600). The tail (300) is provided with a lift system (700). The wings (200) are provided with multiple motor arms (500). The motor arms (500) are equipped with lift systems (700). The tail of the fuselage (100) is provided with a thrust system (400). The diameter of the propeller blades of the lift system (700) is set according to the design optimization results of the aircraft optimization design method according to any one of claims 1, 2, 3, 4, 7, 8 or 9.

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