An electric aircraft frame stability adjustment system and method

Abstract

This invention discloses an electric aircraft frame stabilization adjustment system and method, comprising: an electric aircraft frame assembly having a limiting space; multiple arm assemblies, some of which extend into the limiting space and are movably connected to the electric aircraft frame assembly; an adjustment assembly disposed within the limiting space and connected to the electric aircraft frame assembly, with its actuator connected to one end of the arm assembly; an airframe attitude measurement unit connected to the electric aircraft frame assembly, used to acquire the current attitude data of a light electric helicopter; and multiple limiting assemblies, with their first ends movably connected to the electric aircraft frame assembly. This invention maintains frame balance by redistributing the lift center based on real-time data from the airframe attitude measurement unit through the adjustment assembly.

Description

Technical Field

[0001] This invention discloses an electric aircraft frame stability adjustment system and method, belonging to the field of electric aircraft technology. Background Technology

[0002] During flight, the stability of existing electric aircraft relies primarily on fixed structural design or passive adjustment mechanisms. Traditional frames often employ rigid connections and fixed arm positions, making dynamic adjustments based on real-time flight attitude impossible, resulting in weak anti-interference capabilities. Under uneven loads, attitude imbalance can easily occur due to center of gravity shift, while traditional counterweight adjustment devices can only achieve static balance and cannot cope with dynamic changes during flight.

[0003] Currently, attitude is controlled by rotor speed or control surfaces, but these methods rely on lift difference for indirect adjustment, resulting in slow response and susceptibility to power system performance limitations. Furthermore, traditional arm limiters often use fixed slots or mechanical blocks, which only limit the arm's range of motion and cannot dynamically optimize the limit boundaries based on real-time attitude, making them prone to structural damage from impacts. Summary of the Invention

[0004] The purpose of this invention is to solve the stability problem of existing electric aircraft in dynamic environments, and to propose an electric aircraft frame stability adjustment system and method.

[0005] The technical solution of the present invention is as follows:

[0006] According to a first aspect of the present invention, an electric aircraft frame stabilization adjustment system is provided, comprising: an electric aircraft frame assembly having a limiting space; an arm assembly, a portion of which extends into the limiting space, the arm assembly comprising multiple arms, each arm assembly being movably connected to the electric aircraft frame assembly; an adjustment assembly disposed within the limiting space and connected to the electric aircraft frame assembly, the adjustment assembly comprising multiple adjustments, each adjustment assembly corresponding to a plurality of arm assemblies, the actuator end of the adjustment assembly being connected to one end of an arm assembly; an airframe attitude measurement unit connected to the electric aircraft frame assembly, the airframe attitude measurement unit being spaced apart from the arm assemblies and the adjustment assembly, the airframe attitude measurement unit being used to acquire current attitude data of a light electric helicopter; and a limiting assembly comprising multiple limiting assemblies, each limiting assembly corresponding to a plurality of arm assemblies, the first end of the limiting assembly being connected to the arm assembly and movably connected to the electric aircraft frame assembly.

[0007] Furthermore, the electric aircraft frame assembly includes: an electric aircraft frame body having a limiting space, and multiple arm assemblies movably connected to the electric aircraft frame body; a first support leg, which is spaced apart from the arm assemblies, limiting assemblies, adjusting assemblies, and the aircraft attitude measurement unit, with one end of the first support leg connected to the electric aircraft frame body and the first support leg being set at an angle to the electric aircraft frame body; and a second support leg, which is spaced apart from the first support leg, arm assemblies, adjusting assemblies, and the aircraft attitude measurement unit, with one end of the second support leg connected to the electric aircraft frame body and the second support leg being set at an angle to the electric aircraft frame body, and the second support leg being set at an angle to the first support leg.

[0008] Furthermore, the electric aircraft frame body includes: an electric aircraft frame shell, an adjustment assembly connected to the electric aircraft frame shell; support columns, including multiple support columns, which are spaced apart on the electric aircraft frame shell; and an electric aircraft frame cover, which is fastened together to form a limiting space, with the multiple support columns connected to one of the electric aircraft frame shell and the electric aircraft frame cover.

[0009] Furthermore, at least one of the multiple arm assemblies includes: an arm rod, which is movably connected to the electric aircraft frame housing; and an arm connecting plate, which is connected to a first end of the arm rod, and the arm connecting plate is spaced apart from the adjustment assembly, the adjustment assembly and the airframe attitude measurement unit.

[0010] Furthermore, the adjustment assembly includes: a lead screw motor connected to the electric aircraft frame housing, the lead screw motor being spaced apart from the aircraft attitude measurement unit, the lead screw motor comprising multiple lead screw motors, each lead screw motor corresponding to a multiple arm rod, the actuator end of the lead screw motor being threadedly connected to the second end of the arm rod; and a positioning seat connected to the electric aircraft frame housing, the positioning seat comprising multiple positioning seats, each positioning seat corresponding to a multiple arm rod, each positioning seat being movably connected to its corresponding arm rod.

[0011] Furthermore, the limiting component includes: a positioning block, which is spaced apart from the lead screw motor and the positioning seat, and includes multiple positioning blocks, each of which is respectively set to correspond one-to-one with multiple boom rods, the positioning block is connected to the boom rod, and the positioning block is set adjacent to the boom connecting plate; and a first limiting rod, which includes multiple first limiting rods, the first end of which is hinged to the first end of the positioning block.

[0012] The second limiting rod, comprising multiple second limiting rods, has its first end hinged to the second end of the positioning block, and the second limiting rod and the first limiting rod are symmetrically arranged about the positioning block; a first slider, with the second end of the first limiting rod hinged to the first slider; and a second slider, with the second end of the second limiting rod hinged to the second slider; wherein, the electric aircraft frame body further includes: a first limiting post and a second limiting post, the first limiting post and the second limiting post being spaced apart on the electric aircraft frame shell, the first slider being movably connected to the second limiting post, and the second slider being movably connected to the first limiting post.

[0013] According to a second aspect of the present invention, an electric aircraft frame stabilization adjustment method is provided, applied to the electric aircraft frame stabilization adjustment system of the first aspect, comprising: in response to a flight start signal of a light electric helicopter, acquiring current attitude data; based on the current attitude data, obtaining a current center of gravity offset; based on the current center of gravity offset, making a judgment by a preset center of gravity offset to obtain a judgment result; in response to the judgment result that the current center of gravity offset is greater than the preset center of gravity offset, generating a control command set, the control command set being used to control multiple lead screw motors to perform corresponding target actions.

[0014] Furthermore, the electric aircraft frame stability adjustment method also includes: in response to the second judgment result that the current center of gravity offset is less than or equal to the preset center of gravity offset, repeatedly acquiring the current attitude data.

[0015] Furthermore, in response to the judgment result that the current center of gravity offset is greater than the preset center of gravity offset, a control instruction set is generated, including: in response to the judgment result that the current center of gravity offset is greater than the preset center of gravity offset, determining the arm extension stroke based on the current center of gravity offset; determining the number of motor rotations based on the arm extension stroke; and generating a control instruction set based on the number of motor rotations.

[0016] The beneficial effects of this invention are as follows:

[0017] This invention provides an electric aircraft frame stabilization adjustment system and method. By adjusting the components according to real-time data from the aircraft attitude measurement unit, the system dynamically adjusts the position of the arm components, adjusts the extension length of the corresponding arms, redistributes the center of lift, maintains frame balance, and achieves coordinated optimization of aircraft attitude and arm movement. This provides key support for the efficient and safe application of light electric helicopters in logistics, emergency rescue and other scenarios.

[0018] It should be understood that the above general description and the following detailed description are exemplary and explanatory only, and are not intended to limit the invention. Attached Figure Description

[0019] Figure 1This is a structural diagram illustrating a first embodiment of an electric aircraft frame stabilization adjustment system according to an exemplary embodiment.

[0020] Figure 2 This is a structural diagram illustrating a second embodiment of an electric aircraft frame stabilization adjustment system according to an exemplary embodiment.

[0021] Figure 3 This is a partial structural diagram of a third embodiment of an electric aircraft frame stabilization adjustment system according to an exemplary embodiment.

[0022] Figure 4 This is a flowchart illustrating a flight control method for a lightweight electric helicopter according to an exemplary embodiment. Detailed Implementation

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

[0024] In the description of this invention, it should be noted that the terms "center", "upper", "lower", "left", "right", "vertical", "horizontal", "inner", "outer", etc., indicate the orientation or positional relationship based on the orientation or positional relationship shown in the accompanying drawings. They are only for the convenience of describing this invention and simplifying the description, and do not indicate or imply that the device or element referred to must have a specific orientation, or be constructed and operated in a specific orientation. Therefore, they should not be construed as limitations on this invention.

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

[0026] Example 1: As Figure 1 , Figure 2 and Figure 3The diagram illustrates an electric aircraft frame stabilization and adjustment system according to an exemplary embodiment. The system includes: an electric aircraft frame assembly 20, an arm assembly 10, an adjustment assembly 50, an airframe attitude measurement unit 40, and a limiting assembly 30. The electric aircraft frame assembly 20 has a limiting space, and a portion of the arm assembly 10 extends into this limiting space. Multiple arm assemblies 10 are included, and each arm assembly 10 is movably connected to the electric aircraft frame assembly 20. The adjustment assembly 50 is disposed within the limiting space and connected to the electric aircraft frame assembly 20. Multiple adjustment assemblies 50 are also included. The 0 is set one-to-one with multiple arm assemblies 10. The actuator end of the adjustment assembly 50 is connected to one end of the arm assembly 10. The body attitude measurement unit 40 is connected to the electric aircraft frame assembly 20. The body attitude measurement unit 40 is set at intervals with the arm assembly 10 and the adjustment assembly 50 respectively. The body attitude measurement unit 40 is used to acquire the current attitude data of the light electric helicopter. The limit assembly 30 includes multiple components. The multiple limit assemblies 30 are set one-to-one with multiple arm assemblies 10. The first end of the limit assembly 30 is connected to the arm assembly 10. The first end of the limit assembly 30 is movably connected to the electric aircraft frame assembly 20.

[0027] In this embodiment, the limiting space provided by the electric aircraft frame assembly 20 allows some arm assemblies 10 to extend into and be movably connected. Combined with multiple adjustment components 50 corresponding to each arm assembly 10, precise adjustments can be made to each arm assembly 10. The airframe attitude measurement unit 40 is connected to the electric aircraft frame assembly 20 and spaced apart from the arm assemblies 10 and adjustment components 50, accurately acquiring current attitude data and providing a basis for adjustment, ensuring the timeliness and effectiveness of the adjustment. Simultaneously, multiple limiting components 30 corresponding to each arm assembly 10 have their first ends movably connected to both the arm assembly 10 and the electric aircraft frame assembly 20, ensuring both the flexibility of the arm assemblies 10 and providing reliable limiting protection. The overall structure is compact, and the components work collaboratively, significantly improving the system's stability, safety, and adjustment accuracy.

[0028] In this embodiment, the electric aircraft frame assembly 20 includes: an electric aircraft frame body, a first support leg 203, and a second support leg 204. The electric aircraft frame body has a limiting space. Multiple arm assemblies 10 are movably connected to the electric aircraft frame body. The first support leg 203 is spaced apart from the arm assemblies 10, the limiting assembly 30, the adjusting assembly 50, and the aircraft attitude measurement unit 40. One end of the first support leg 203 is connected to the electric aircraft frame body. The first support leg 203 is set at an angle to the electric aircraft frame body. The second support leg 204 is spaced apart from the first support leg 203, the arm assemblies 10, the adjusting assembly 50, and the aircraft attitude measurement unit 40. One end of the second support leg 204 is connected to the electric aircraft frame body. The second support leg 204 is set at an angle to the electric aircraft frame body and the first support leg 203.

[0029] In this embodiment, the limiting space provided by the electric aircraft frame assembly 20 provides a basis for the installation and movement of components such as the arm assembly 10. The first support leg 203 and the second support leg 204 are both set at an angle to the electric aircraft frame assembly and are also at an angle to each other, forming a stable support structure, which can effectively improve the overall load-bearing capacity and structural stability of the electric aircraft frame. At the same time, the first support leg 203 is spaced apart from the arm assembly 10, the limiting assembly 30, the adjustment assembly 50 and the aircraft attitude measurement unit 40, and the second support leg 204 is also spaced apart from the first support leg 203, the arm assembly 10, the adjustment assembly 50 and the aircraft attitude measurement unit 40, avoiding mutual interference between the components. This ensures the flexible movement of the arm assembly 10, the precise adjustment of the adjustment assembly 50 and the accuracy of data acquisition by the aircraft attitude measurement unit 40. Furthermore, the independent and angled support leg design enhances the overall stability of the electric aircraft during takeoff, landing and flight.

[0030] Furthermore, the electric aircraft frame body includes: an electric aircraft frame housing 202, support columns 205, and an electric aircraft frame cover 201. The adjustment assembly 50 is connected to the electric aircraft frame housing 202. The support columns 205 include multiple columns, which are spaced apart on the electric aircraft frame housing 202. The electric aircraft frame cover 201 is fastened together to form a limiting space. The multiple support columns 205 are connected to one of the electric aircraft frame housing 202 and the electric aircraft frame cover 201.

[0031] In this embodiment, the electric aircraft frame body forms a limiting space by fastening the electric aircraft frame housing 202 and the electric aircraft frame cover 201 together, providing a closed and stable installation environment for components such as the adjustment assembly 50 and ensuring the safe operation of internal components. Multiple spaced support columns 205 are connected to one of the electric aircraft frame housing 202 and the electric aircraft frame cover 201, which not only enhances the overall structural strength and deformation resistance of the electric aircraft frame housing 202 and the electric aircraft frame cover 201 after fastening, but also avoids interference with the installation and operation of components such as the adjustment assembly 50 through the spaced design. At the same time, the adjustment assembly 50 is directly connected to the electric aircraft frame housing 202, which improves the stability and response efficiency of the adjustment operation. The overall structural design is compact and stable, effectively optimizing the load-bearing performance and internal space utilization of the frame body.

[0032] In the above embodiments, at least one of the multiple arm assemblies 10 includes: an arm rod 101 and an arm connecting plate 102. The arm rod 101 is movably connected to the electric aircraft frame housing 202. The arm connecting plate 102 is connected to the first end of the arm rod 101. The arm connecting plate 102 is spaced apart from the adjustment assembly 50, the adjustment assembly 50 and the airframe attitude measurement unit 40. At least one of the multiple arm assemblies 10 includes an arm rod 101 that is movably connected to the electric aircraft frame housing 202, ensuring the overall flexibility of the arm assembly 1 and facilitating position adjustment according to flight requirements. The arm connecting plate 102 is connected to the first end of the arm rod 101 and can serve as an adapter structure for connecting the arm assembly 10 to other components (such as rotors), improving connection reliability. At the same time, the arm connecting plate 102 is spaced apart from the adjustment assembly 50 and the airframe attitude measurement unit 40, which can effectively avoid mutual interference between components, ensure the accuracy of the adjustment assembly 50 in adjusting the arm assembly 10 and the accuracy of the data acquired by the airframe attitude measurement unit 40. The overall design is reasonable and improves the stability of system operation.

[0033] In one exemplary embodiment, the adjustment assembly 50 includes a lead screw motor 502 and a positioning seat 501. The lead screw motor 502 is connected to the electric aircraft frame housing 202 and is spaced apart from the airframe attitude measurement unit 40. There are multiple lead screw motors 502, and each of the multiple lead screw motors 502 is correspondingly arranged with a multiple arm rod 101. The actuator end of the lead screw motor 502 is threadedly connected to the second end of the arm rod 101. The positioning seat 501 is connected to the electric aircraft frame housing 202 and includes multiple positioning seats 501. Each of the multiple positioning seats 501 is correspondingly arranged with a multiple arm rod 101, and each positioning seat 501 is movably connected to the corresponding arm rod 101.

[0034] In this embodiment, in the adjustment component 50, the lead screw motor 502 is connected to the electric aircraft frame housing 202 and spaced apart from the aircraft attitude measurement unit 40, which ensures stable installation and avoids interference with attitude measurement. Multiple lead screw motors 502 correspond one-to-one with multiple arm rods 101, and their actuators are threadedly connected to the second end of the arm rod 101. This allows for independent and precise position adjustment of each arm rod 101 through accurate threaded transmission, resulting in sensitive response and high adjustment accuracy. Simultaneously, multiple positioning seats 501 correspond one-to-one with multiple arm rods 101. Each positioning seat 501 is movably connected to and fixed to the electric aircraft frame housing 202, providing stable guidance for the movement of the arm rods 101 and preventing them from deviating or wobbling. This further improves the stability and reliability of the adjustment process. The overall structure exhibits strong synergy, effectively ensuring the accuracy and efficiency of the arm component attitude adjustment.

[0035] In this embodiment, the limiting component 30 includes: a positioning block 304, a second limiting rod 305, a first slider 302, a second slider 306, and a first limiting rod 303. The positioning block 304 is spaced apart from the lead screw motor 502 and the positioning seat 501. There are multiple positioning blocks 304, and each of the multiple positioning blocks 304 is respectively arranged in a one-to-one correspondence with a multiple arm rod 101. The positioning block 304 is connected to the arm rod 101 and is arranged adjacent to the arm connecting plate 102. There are multiple first limiting rods 303, and the first end of the first limiting rod 303 is hinged to the first end of the positioning block 304. There are multiple second limiting rods 305, and the first end of the second limiting rod 305 is hinged to the second end of the positioning block 304. The second limiting rods 305 and the first limiting rod 303 are arranged symmetrically about the positioning block 304. The second end of the first limiting rod 303 is hinged to the first slider 302, and the second end of the second limiting rod 305 is hinged to the second slider 306. The electric aircraft frame body also includes: a first limiting post 207 and a second limiting post 206, the first limiting post 207 and the second limiting post 206 being spaced apart on the electric aircraft frame housing 202, the first slider 302 being movably connected to the second limiting post 206, and the second slider 306 being movably connected to the first limiting post 207.

[0036] In this embodiment, the limiting component 30, through multiple positioning blocks 304 corresponding one-to-one with the arm rod 101, and in conjunction with the symmetrically arranged first limiting rod 303 and second limiting rod 305, forms a stable linkage structure. It can move synchronously with the positioning blocks 304 (connected to the arm rod 101), and is also movably connected to the second limiting post 206 and the first limiting post 207 on the electric aircraft frame housing 202 via the first slider 302 and the second slider 306, respectively, providing precise guidance and range constraint for the movement of the arm rod 101. The positioning blocks 304 are spaced apart from the lead screw motor 502 and the positioning seat 501, avoiding mutual interference between the adjusting components and the limiting components. Simultaneously, the symmetrical hinge design of the first limiting rod 303 and the second limiting rod 305 can evenly distribute the stress during the movement of the arm rod 101. Combined with the fixed support of the first limiting post 207 and the second limiting post 206, the stability and reliability of the limiting are greatly improved, ensuring that the arm assembly moves flexibly within a safe range, further guaranteeing the overall system's operational safety.

[0037] Example 2: Figure 4 This is an exemplary embodiment illustrating a method for stabilizing the frame of an electric aircraft, implemented by a terminal, which includes at least a CPU. The method includes:

[0038] Step S101: In response to the flight start signal of the light electric helicopter, acquire the current attitude data.

[0039] In step S101, when the light electric helicopter receives a flight start signal (such as a takeoff command issued by the operator, a remote control signal, or a start command triggered by a preset program), the system immediately triggers the airframe attitude measurement unit 40 to enter the working state. This unit collects the helicopter's current spatial attitude parameters in real time through built-in sensors (such as gyroscopes, accelerometers, magnetometers, etc.), including but not limited to pitch angle, roll angle, yaw angle, and other angular information, as well as linear acceleration, angular velocity, and other motion state data. This current attitude data will serve as an initial reference, providing the original basis for the subsequent dynamic adjustment of the arm assembly 10 by the adjustment component 50, ensuring that the helicopter can achieve stable control based on accurate attitude information from the start-up stage, laying the foundation for smooth takeoff and attitude correction during subsequent flight.

[0040] Step S102: Based on the current attitude data, obtain the current center of gravity offset.

[0041] In step S102, after acquiring the current attitude data (such as pitch angle, roll angle, angular velocity, etc.), the system will combine the preset structural parameters of the electric aircraft (including the length of each arm assembly 10, the weight distribution of the electric aircraft frame assembly 20, the preset position of the load, etc.) and analyze them through the center of gravity calculation model.

[0042] When the aircraft attitude measurement unit 40 detects a certain roll angle in the fuselage, it indicates a difference in force on both sides of the fuselage, possibly due to a shift in the center of gravity to one side. The system compares the real-time collected attitude angle data with the preset balance attitude parameters, calculates the attitude deviation value, and then, by combining the fuselage's moment of inertia, the weight parameters of each component, etc., it uses the mechanical balance formula to deduce the direction and specific amount of the current center of gravity offset relative to the preset reference center of gravity in three-dimensional space. This process provides a precise quantitative basis for the subsequent adjustment component 50 to adjust the position of the arm component 10, ensuring that the center of gravity offset can be effectively compensated.

[0043] Step S103: Based on the current center of gravity offset, a judgment is made using a preset center of gravity offset to obtain the judgment result.

[0044] In step S103, after obtaining the current center of gravity offset, the system will call the preset center of gravity offset. This parameter is a threshold range preset according to the safe flight standards, structural bearing limits and dynamic stability requirements of electric aircraft, such as ±5cm in the X-axis direction and ±5cm in the Y-axis direction, for comparison and judgment.

[0045] Specifically, the system quantitatively compares the current center of gravity offset distance in each axis direction in three-dimensional space with preset thresholds. If the current offset is within the preset range in all directions, the judgment result is "center of gravity offset is within the safe range, no adjustment action is needed," and the arm assembly 10 maintains its current state. If the offset in a certain direction (such as the X-axis) exceeds the preset threshold, the judgment result is "center of gravity offset exceeds the limit, targeted adjustment is required." The system will further clarify the direction of the excess and the specific difference, such as a 7cm offset in the positive X-axis direction exceeding the threshold by 2cm. If multiple directions exceed the limit simultaneously, the judgment result will integrate the offset situations in each direction, providing a basis for the subsequent adjustment assembly 50 to formulate a coordinated adjustment strategy. This judgment mechanism based on preset thresholds can avoid frequent adjustments caused by small offsets, reducing energy consumption and mechanical wear, while ensuring timely response when the center of gravity shifts significantly, thus ensuring flight safety.

[0046] Step S104: In response to the judgment result that the current center of gravity offset is greater than the preset center of gravity offset, a control instruction set is generated. The control instruction set is used to control multiple lead screw motors to perform corresponding target actions.

[0047] In step S104, when the judgment result is that the current center of gravity offset is greater than the preset center of gravity offset, the system will start a targeted adjustment process: First, based on the specific value of the current center of gravity offset (such as a 7cm offset in the positive X-axis direction) and direction, combined with the correlation model between the arm assembly 10 and the center of gravity adjustment (such as the quantitative relationship that the extension of the arm on one side can cause the center of gravity to shift to the opposite side), the required extension stroke of each corresponding arm rod 101 is calculated. If the offset in the positive X-axis direction exceeds the limit, it may be necessary to control the arm rod 101 in the negative X-axis direction to extend outward by 3cm, while the arm rod 101 in the positive X-axis direction should be appropriately contracted by 2cm, so as to offset the center of gravity offset through coordinated adjustment.

[0048] Subsequently, based on the pitch parameters of the lead screw motor 502 (e.g., each rotation corresponds to a 0.5cm movement of the boom 101), the extension and retraction stroke of the boom is converted into the precise number of motor rotations—the aforementioned 3cm extension stroke corresponds to 6 forward rotations of the lead screw motor 502, and the 2cm retraction stroke corresponds to 4 reverse rotations.

[0049] The extension and retraction stroke of the boom and the number of rotations of the motor are shown in formula (1):

[0050] (1)

[0051] in, Let λ be the extension / retraction stroke of the boom, and λ be the geometric amplification factor of the linkage or connecting rod. If the boom is connected to the lead screw and nut via connecting rods, levers, or other multi-bar mechanisms, geometric relationships such as "mechanism amplification" or "mechanism shortening" may also occur. For the transmission ratio, if the linear motor output shaft and lead screw do not have an additional speed reduction (or speed increase) mechanism, then it can be set as follows: If there are gears, synchronizer pulleys, or other transmission devices, the total transmission ratio needs to be factored into the lead screw rotation, making the total transmission ratio i (if it's a reduction, then...). If it is the growth rate, then ). is the number of motor rotations, and p is the pitch, determined by the lead screw motor 502.

[0052] Finally, based on the number of rotations of these motors, the system generates a set of control instructions that includes the target rotation direction, number of rotations, and execution timing, ensuring that each lead screw motor 502 operates synchronously according to the cooperative strategy and quickly pulls the center of gravity back to a safe range.

[0053] When the second judgment result is that the current center of gravity offset is less than or equal to the preset center of gravity offset, the system determines that no adjustment is needed. At this time, the body attitude measurement unit 40 will be triggered to continue to repeatedly acquire the current attitude data at a preset frequency (such as 50 times per second) to continuously monitor the changes in the center of gravity and form a dynamic closed-loop monitoring mechanism to ensure that it can respond immediately when the center of gravity offset just exceeds the threshold, thus avoiding the safety risks caused by lag.

[0054] In this embodiment, upon responding to the flight start signal, the airframe attitude measurement unit 40 can quickly collect attitude data, providing an initial benchmark for subsequent adjustments and ensuring accurate data for the adjustment of the arm assembly 10 from the start-up phase. Combining structural parameters such as the electric aircraft frame assembly 20, the current center of gravity offset is derived based on the data from the airframe attitude measurement unit 40, providing a precise quantitative reference for the adjustment assembly 50 to adjust the arm assembly 10. By using a preset threshold, frequent movements of the arm assembly 10 due to minor offsets can be avoided, reducing wear and ensuring timely response when the center of gravity shifts significantly. When adjustment is required, the number of rotations can be determined based on the required extension / retraction stroke of the arm lever 101 and parameters such as the pitch of the lead screw motor 502. A control command set is then generated to precisely control the movement of the lead screw motor 502, enabling coordinated adjustment of the arm assembly 10 and rapid correction of the center of gravity offset. When no adjustment is needed, the airframe attitude measurement unit 40 continuously monitors, forming a closed-loop mechanism. The entire process, through the collaboration of all components, balances adjustment accuracy, response timeliness, and operational safety.

[0055] Although embodiments of the present invention have been disclosed above, they are not limited to the applications listed in the specification and embodiments. It can be applied to various fields suitable for the present invention. Other modifications can be readily made by those skilled in the art. Therefore, without departing from the general concept defined by the claims and their equivalents, the present invention is not limited to the specific details and examples shown and described herein.

Claims

1. An electric aircraft frame stability adjustment system, characterized in that, include: Electric aircraft frame assembly (20), the electric aircraft frame assembly (20) includes: electric aircraft frame body, the electric aircraft frame body having a limiting space; the electric aircraft frame body includes: electric aircraft frame housing (202). Arm assembly (10), a portion of the arm assembly (10) extends into the limiting space, the arm assembly (10) includes a plurality of them, and the plurality of arm assemblies (10) are respectively movably connected to the electric aircraft frame body; Adjustment component (50), the adjustment component (50) is disposed in the limiting space, the adjustment component (50) is connected to the electric aircraft frame housing (202), the adjustment component (50) includes multiple, the multiple adjustment components (50) are respectively disposed in correspondence with multiple arm components (10), and the execution end of the adjustment component (50) is connected to one end of the arm component (10); Aircraft attitude measurement unit (40), the aircraft attitude measurement unit (40) is connected to the electric aircraft frame assembly (20), the aircraft attitude measurement unit (40) is spaced apart from the arm assembly (10) and the adjustment assembly (50), the aircraft attitude measurement unit (40) is used to acquire the current attitude data of the light electric helicopter; Limiting component (30), the limiting component (30) includes multiple, the multiple limiting components (30) are respectively arranged in correspondence with the multiple arm components (10), the first end of the limiting component (30) is connected to the arm component (10), and the first end of the limiting component (30) is movably connected to the electric aircraft frame component (20). At least one of the plurality of arm assemblies (10) includes: An arm rod (101) is movably connected to the electric aircraft frame housing (202); The arm connecting plate (102) is connected to the first end of the arm rod (101), and the arm connecting plate (102) is spaced apart from the adjustment component (50) and the body posture measurement unit (40). The adjustment component (50) includes: A lead screw motor (502) is connected to the electric aircraft frame housing (202). The lead screw motor (502) is spaced apart from the body attitude measurement unit (40). There are multiple lead screw motors (502). Each of the multiple lead screw motors (502) is correspondingly arranged with a multiple arm rod (101). The actuator end of the lead screw motor (502) is threadedly connected to the second end of the arm rod (101). Positioning seat (501), the positioning seat (501) is connected to the electric aircraft frame housing (202), the positioning seat (501) includes multiple, the multiple positioning seats (501) are respectively arranged in correspondence with multiple arm rods (101), and each positioning seat (501) is movably connected to the corresponding arm rod (101); The limiting component (30) includes: Positioning block (304), the positioning block (304) is spaced apart from the lead screw motor (502) and the positioning seat (501), the positioning block (304) includes multiple, the multiple positioning blocks (304) are respectively arranged in one-to-one correspondence with multiple arm rods (101), the positioning block (304) is connected to the arm rod (101), and the positioning block (304) is arranged adjacent to the arm connecting plate (102); The first limiting rod (303) includes multiple first limiting rods (303), and the first end of the first limiting rod (303) is hinged to the first end of the positioning block (304); The second limiting rod (305) includes multiple second limiting rods (305), the first end of the second limiting rod (305) is hinged to the second end of the positioning block (304), and the second limiting rod (305) and the first limiting rod (303) are symmetrically arranged about the positioning block (304). The first slider (302) has its second end hinged to the first limiting rod (303); The second slider (306) is hinged to the second end of the second limiting rod (305); The electric aircraft frame body further includes a first limiting post (207) and a second limiting post (206), the first limiting post (207) and the second limiting post (206) being spaced apart on the electric aircraft frame housing (202), the first slider (302) being movably connected to the second limiting post (206), and the second slider (306) being movably connected to the first limiting post (207).

2. The electric aircraft frame stability adjustment system according to claim 1, characterized in that, The electric aircraft frame assembly (20) includes: The first support foot (203) is spaced apart from the arm assembly (10), the limiting assembly (30), the adjusting assembly (50) and the body attitude measuring unit (40). One end of the first support foot (203) is connected to the electric aircraft frame body. The first support foot (203) is set at an angle to the electric aircraft frame body. The second support foot (204) is spaced apart from the first support foot (203), the arm assembly (10), the adjustment assembly (50) and the body attitude measurement unit (40). One end of the second support foot (204) is connected to the electric aircraft frame body. The second support foot (204) is set at an angle to the electric aircraft frame body and at an angle to the first support foot (203).

3. The electric aircraft frame stability adjustment system according to claim 2, characterized in that, The electric aircraft frame body includes: Support columns (205), the support columns (205) include a plurality of them, and the plurality of support columns (205) are arranged at intervals on the electric aircraft frame housing (202); An electric aircraft frame cover (201) is fastened to the electric aircraft frame housing (202) to form the limiting space, and a plurality of the support columns (205) are connected to one of the electric aircraft frame housing (202) and the electric aircraft frame cover (201).

4. A method for stabilizing and adjusting the frame of an electric aircraft, applied to the electric aircraft frame stabilization and adjustment system according to any one of claims 1-3, characterized in that, include: In response to the flight start signal of the light electric helicopter, acquire current attitude data; Based on the current attitude data, the current center of gravity offset is obtained; Based on the current center of gravity offset, a judgment is made using a preset center of gravity offset to obtain the judgment result; In response to the judgment result that the current center of gravity offset is greater than the preset center of gravity offset, a control instruction set is generated, which is used to control multiple lead screw motors to perform corresponding target actions.

5. The electric aircraft frame stability adjustment method according to claim 4, characterized in that, The electric aircraft frame stability adjustment method also includes: In response to the determination result that the current center of gravity offset is less than or equal to the preset center of gravity offset, the current attitude data is repeatedly acquired.

6. The electric aircraft frame stability adjustment method according to claim 4, characterized in that, In response to the determination result that the current center of gravity offset is greater than the preset center of gravity offset, the control command set is generated, including: In response to the judgment result that the current center of gravity offset is greater than the preset center of gravity offset, the extension and retraction stroke of the boom is determined based on the current center of gravity offset; The number of motor rotations is determined based on the arm's extension and retraction stroke. The control command set is generated based on the number of rotations of the motor.

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