Helicopter autorotation landing control method based on normal acceleration and related device

Abstract

The invention discloses a helicopter autorotation landing control method based on normal acceleration and a related device, and relates to the technical field of autorotation landing, the method comprises the following steps: in a deceleration leveling stage, determining a first normal acceleration expected value according to a residual hang time estimation value of the deceleration leveling stage, and utilizing a total distance control law of the deceleration leveling stage to determine a second normal acceleration expected value; the normal motion of the target helicopter is controlled by tracking the first normal acceleration expected value; and in a landing stage, determining a second normal acceleration expected value according to the total hang time expected value in the landing stage, and controlling the normal motion of the target helicopter by using the total distance control law in the landing stage to track the second normal acceleration expected value as a target. The motion of the helicopter in the normal direction in the autorotation landing process is precisely controlled, impact generated when the helicopter touches the ground is reduced, and the success rate of autorotation landing is increased.

Description

Technical Field

[0001] This application relates to the technical field of autorotation landing, and particularly to a helicopter autorotation landing control method and related device based on normal acceleration. Background Art

[0002] The in-flight engine shutdown of a helicopter is one of the common fault types of helicopters. In order to improve the safety of helicopters, it is necessary to design a control strategy for the autorotation landing process of a helicopter after losing power. At present, there are two control strategies proposed in the patent document CN106873617A for the domestic research on the autorotation landing process: 1. Autorotate and glide under the conditions of the engine failure and tail rotor failure of an unmanned helicopter. When the height of the unmanned helicopter is greater than 25 meters, keep the forward speed of the unmanned helicopter = 27 m / s; the vertical descent speed = 8 m / s; when the height of the unmanned helicopter is greater than 10 meters and less than or equal to 25 meters, level off and decelerate the unmanned helicopter to reduce the forward speed and vertical speed; when the height of the unmanned helicopter is less than 10 meters, level off the pitch angle to the hover attitude; increase the collective pitch to reduce the touchdown speed, and complete the landing with the descent rate controlled within 3 m / s. 2. Autorotate and glide under the condition of the tail rotor failure of an unmanned helicopter, and stabilize the helicopter attitude. When the height of the unmanned helicopter is greater than 75 meters, keep the vertical descent speed of the unmanned helicopter = 8 m / s; when the height of the unmanned helicopter is greater than 25 meters, keep the vertical descent speed of the unmanned helicopter = 4 m / s; when the height of the unmanned helicopter is less than or equal to 25 meters, keep the vertical descent speed of the unmanned helicopter = 1 m / s to complete the landing.

[0003] The patent document CN106873617A divides the entire autorotation landing process into 3 stages according to the height. However, in the deceleration and leveling stage and the landing stage below 25 m, the first control method only mentions that a deceleration and leveling action needs to be performed, and the second control method only gives the expected vertical descent speed of the helicopter; the applicability is not strong, and the specific description of the deceleration and leveling process is lacking. Therefore, there is an urgent need for a control scheme that can ensure the helicopter to complete the autorotation landing safely and without collision. Summary of the Invention

[0004] The purpose of this application is to provide a helicopter autorotation landing control method and related device based on normal acceleration, which can ensure the helicopter to complete the autorotation landing safely and without collision.

[0005] To achieve the above purpose, this application provides the following solutions: In the first aspect, this application provides a helicopter autorotation landing control method based on normal acceleration, including: Determine the autorotation landing control stage of the target helicopter according to the height of the target helicopter; the target helicopter is a helicopter that has lost power.

[0006] When the autorotation landing control phase is the steady autorotation phase, the normal motion of the target helicopter is controlled with the aim of tracking the expected value of the normal velocity; the expected value of the normal velocity is determined according to the expected value of the longitudinal velocity and the trim descent rate; the normal direction is the direction perpendicular to the ground.

[0007] When the autorotation landing control phase is the deceleration and flare phase, according to the estimated remaining hovering time value of the deceleration and flare phase, the expected value of the first normal acceleration is determined, and the total pitch control law of the deceleration and flare phase is used to control the normal motion of the target helicopter with the aim of tracking the expected value of the first normal acceleration; the estimated remaining hovering time value of the deceleration and flare phase is determined according to the current altitude, the current longitudinal position and the current first motion state of the target helicopter; the first motion state includes the normal velocity, the normal acceleration, the longitudinal velocity and the longitudinal acceleration; the longitudinal direction is the forward flight direction of the target helicopter.

[0008] When the autorotation landing control phase is the landing phase, according to the expected value of the total hovering time of the landing phase, the expected value of the second normal acceleration is determined, and the total pitch control law of the landing phase is used to control the normal motion of the target helicopter with the aim of tracking the expected value of the second normal acceleration.

[0009] In a second aspect, the present application provides a computer device, including: a memory, a processor, and a computer program stored on the memory and executable on the processor, where the processor executes the computer program to implement the method for helicopter autorotation landing control based on normal acceleration described in the first aspect.

[0010] In a third aspect, the present application provides a computer-readable storage medium, on which a computer program is stored, and when the computer program is executed by a processor, it implements the method for helicopter autorotation landing control based on normal acceleration described in the first aspect.

[0011] In a fourth aspect, the present application provides a computer program product, including a computer program, and when the computer program is executed by a processor, it implements the method for helicopter autorotation landing control based on normal acceleration described in the first aspect.

[0012] According to the specific embodiments provided by the present application, the present application has the following technical effects: The present application provides a method and related device for helicopter autorotation landing based on normal acceleration. First, according to the altitude of the target helicopter, the autorotation landing control phase of the target helicopter is determined. When the autorotation landing control phase is the steady autorotation phase, the normal motion of the target helicopter is controlled with the goal of tracking the expected value of the normal velocity. When the autorotation landing control phase is the deceleration and flare phase, according to the estimated remaining hang time in the deceleration and flare phase, the first expected value of the normal acceleration is determined, and the total pitch control law in the deceleration and flare phase is used to control the normal motion of the target helicopter with the goal of tracking the first expected value of the normal acceleration. When the autorotation landing control phase is the landing phase, according to the expected total hang time in the landing phase, the second expected value of the normal acceleration is determined, and the total pitch control law in the landing phase is used to control the normal motion of the target helicopter with the goal of tracking the second expected value of the normal acceleration. Through the above solutions, the present application achieves fine control of the helicopter's motion in the normal direction during autorotation landing (especially in the deceleration and flare phase and the landing phase), thereby ensuring that the helicopter can complete autorotation landing safely and without collision. BRIEF DESCRIPTION OF THE DRAWINGS

[0013] In order to more clearly illustrate the technical solutions in the embodiments of the present application or in the prior art, the following will briefly introduce the drawings required to be used in the embodiments. Obviously, the drawings described below are only some embodiments of the present application. For those of ordinary skill in the art, other drawings can be obtained based on these drawings without creative efforts.

[0014] Figure 1 It is a schematic flowchart of a method for helicopter autorotation landing based on normal acceleration provided by an embodiment of the present application; Figure 2 It is a schematic flowchart of a method for helicopter autorotation landing based on normal acceleration provided by another embodiment of the present application; Figure 3 It is a schematic structural diagram of a computer device provided by an embodiment of the present application. DETAILED DESCRIPTION OF THE EMBODIMENTS

[0015] The following will clearly and completely describe the technical solutions in the embodiments of the present application with reference to the drawings in the embodiments of the present application. Obviously, the described embodiments are only some embodiments of the present application, rather than all embodiments. Based on the embodiments in the present application, all other embodiments obtained by those of ordinary skill in the art without creative efforts belong to the scope of protection of the present application.

[0016] Autorotative landing is a unique power-off landing process for rotorcraft. In this state, the helicopter relies on the upward inflow to drive the rotor. When the helicopter engine fails in flight, the available energy of the helicopter is limited. Therefore, the pilot needs to perform the correct maneuvers at the right time to ensure the safe landing of the helicopter. This application provides a control strategy for helicopter autorotative landing, which helps the helicopter to complete the autorotative landing process autonomously.

[0017] To make the above objects, features, and advantages of this application more obvious and understandable, the following further detailed description of this application will be given in conjunction with the accompanying drawings and specific embodiments.

[0018] In an exemplary embodiment, as Figure 1 shown, a method for controlling helicopter autorotative landing based on normal acceleration is provided, including the following steps 101 to 104. Among them: Step 101, determine the autorotative landing control stage of the target helicopter according to the height of the target helicopter; the target helicopter is a helicopter that has lost power.

[0019] According to the change in height from the ground, the autorotative gliding process of the target helicopter can be divided into three stages: steady-state autorotation, decelerated flare, and landing. Accordingly, the autorotative landing control stage is correspondingly divided into a steady-state autorotation stage, a decelerated flare stage, and a landing stage. For manned helicopters, the height range from which the steady-state autorotation stage starts is above 35m, the height range from which the decelerated flare stage starts is 15m - 35m, and the height range from which the landing stage starts is 0 - 2m; that is, steady-state autorotation starts and is maintained above 35m, decelerated flare starts and is maintained at 15 - 35m, and landing starts and is maintained at 0 - 2m.

[0020] For example, when the height of the target helicopter is above 35m, determine that the current autorotative landing control stage is the steady-state autorotation stage; when the height of the target helicopter is below 35m, determine that the current autorotative landing control stage is the decelerated flare stage; 35m is both the end height of the steady-state autorotation stage and the initial height of the decelerated flare stage; when the height of the target helicopter is below 2m, determine that the current autorotative landing control stage is the landing stage; 2m is both the end height of the decelerated flare stage and the initial height of the landing stage. Suppose the height of the target helicopter is 31m and the engine suddenly stops, then it directly enters the decelerated flare stage, skipping the steady-state autorotation stage.

[0021] Step 102, when the autorotative landing control stage is the steady-state autorotation stage, control the normal motion of the target helicopter with the goal of tracking the normal velocity expectation value; the normal velocity expectation value is determined according to the longitudinal velocity expectation value and the trim descent rate; the normal direction is perpendicular to the ground.

[0022] Step 103, when the autorotation landing control phase is the deceleration and flare phase, determine the expected value of the first normal acceleration according to the estimated remaining hovering time value in the deceleration and flare phase, and use the total pitch control law in the deceleration and flare phase to control the normal movement of the target helicopter with the goal of tracking the expected value of the first normal acceleration; the estimated remaining hovering time value in the deceleration and flare phase is determined according to the current altitude, current longitudinal position and current first motion state of the target helicopter; the first motion state includes normal velocity, normal acceleration, longitudinal velocity and longitudinal acceleration; the longitudinal direction is the forward flight direction of the target helicopter, parallel to the ground.

[0023] In addition, the lateral direction in this article is the direction perpendicular to the longitudinal direction and parallel to the ground.

[0024] Step 104, when the autorotation landing control phase is the landing phase, determine the expected value of the second normal acceleration according to the expected total hovering time value in the landing phase, and use the total pitch control law in the landing phase to control the normal movement of the target helicopter with the goal of tracking the expected value of the second normal acceleration.

[0025] As an optional implementation manner, in Step 103, the mathematical expression of the total pitch control law in the deceleration and flare phase is as follows: ; Where, is the total pitch; is the first-order differential of the total pitch; is the difference between the expected value of the first normal acceleration and the actual normal jerk; is 's first-order differential; , and are the first PID controller parameters; is the first total pitch lift speed, used to ensure the rationality of the expected value of the first normal acceleration; is the estimated remaining hovering time value in the deceleration and flare phase; is the current altitude; is the current normal velocity.

[0026] Furthermore, specifically determining the expected value of the first normal acceleration according to the estimated remaining hovering time value in the deceleration and flare phase includes: Determine the expected value of the first normal acceleration according to the estimated remaining hovering time value in the deceleration and flare phase by using the calculation formula for the expected value of the first normal acceleration; the calculation formula for the expected value of the first normal acceleration is: ; Where, is the expected value of the initial height in the landing phase; for example, the expected value of the initial height in the landing phase can be set to 2 m. is the current height; is the current normal velocity; is the expected value of the first normal acceleration.

[0027] The principle is as follows: for a given current height and descent speed (i.e., normal velocity), there is a constant normal acceleration that satisfies the following formula: ; where is the height (current height) at time, is the normal velocity (current normal velocity) at represents the height reached after uniformly accelerated linear motion for a certain time with a constant acceleration . Set the left end of the above formula to the expected value of the starting (initial) height of the landing phase set by the user , and calculate the constant acceleration , which is the current expected value of the normal acceleration (i.e., the first expected value of the normal acceleration ). Moreover, to ensure the rationality of the obtained constant normal acceleration , the current height and descent speed need to satisfy the following inequality. Otherwise, the target helicopter height change time curve obtained according to this constant normal acceleration will reach the expected height before (i.e., there is a root less than ), ; For the case where the above inequality is not satisfied, that is, when the obtained first expected value of the normal acceleration is unreasonable, a signal for quickly increasing the collective pitch (also referred to as a large-speed collective pitch signal in this article) needs to be applied to the collective pitch, that is . By applying a signal for quickly increasing the collective pitch to the collective pitch, the motion state of the target helicopter is changed to make it satisfy the above inequality, that is, to make the calculated satisfy the above inequality, thereby ensuring the rationality of the first expected value of the normal acceleration .

[0028] Furthermore, the determination process of the estimated remaining hover time in the deceleration and flare-out phase specifically includes: Based on the current altitude, current longitudinal position, and current first motion state of the target helicopter, use the formula for the remaining hovering time in the deceleration and flare phase to determine the estimated value of the remaining hovering time in the deceleration and flare phase; the formula for the remaining hovering time in the deceleration and flare phase is: ; Where, is transition value of; is a safety value set by the user and is greater than or equal to zero, making the estimated more in line with the actual hovering time during deceleration motion; is the maximum value of.

[0029] The calculation formula of is as follows: ; Where, is the current normal acceleration, is the current longitudinal position; is the current longitudinal velocity; is the current longitudinal acceleration; is the expected value of the normal initial velocity in the landing phase; is the expected value of the touchdown longitudinal velocity, set by the user. That is, the estimation method is to divide the expected velocity change by the current acceleration and then add a safety value.

[0030] The calculation formula of is as follows: ; Where, is the available energy, obtained by adding the rotational kinetic energy and longitudinal kinetic energy of the current target helicopter's rotor; is the ideal available energy, obtained by adding the rotational kinetic energy and longitudinal kinetic energy of the ideal helicopter rotor, that is, the best available energy that is expected to be achieved; is the maximum duration of the deceleration and flare phase.

[0031] As an alternative implementation, the mathematical expression of the total pitch control law in the landing phase is as follows: ; Where, is the total pitch; is the first-order differential of the total pitch; is the difference between the second normal acceleration expectation value and the actual normal jerk; is the first-order differential of; , and is the parameter of the second PID controller; is the pitch rate of the collective pitch, which is used to ensure the rationality of the expected value of the second normal acceleration; is the expected value of the total hang time during the landing phase; is the current altitude; is the current normal velocity.

[0032] Further, according to the expected value of the total hang time during the landing phase, the expected value of the second normal acceleration is determined, specifically including: According to the expected value of the total hang time during the landing phase, using the calculation formula of the expected value of the second normal acceleration, the expected value of the second normal acceleration is determined; the calculation formula of the expected value of the second normal acceleration is: ; wherein, is the touchdown altitude during the landing phase, equal to 0; is the current altitude; is the current normal velocity; is the expected value of the second normal acceleration; is the expected value of the total hang time during the landing phase.

[0033] As an optional implementation manner, the expected value of the total hang time during the landing phase is determined according to the expected value of the initial altitude during the landing phase and its corresponding expected value of the normal initial velocity during the landing phase determined.

[0034] In this embodiment, the determination of the expected value of the normal acceleration during the landing phase is similar to the determination of the expected value of the normal acceleration during the deceleration and flare phase. The difference is that the calculated during the deceleration and flare phase is refreshed in real time as the state of the target helicopter changes, while the during the landing phase is a constant value determined by the user according to and (constant and unchanged during the calculation of the entire landing phase), so as to ensure that the expected normal velocity of the target helicopter decays exponentially.

[0035] Referring to Figure 2 , in another exemplary embodiment, the control process during the steady-state autorotation phase further includes the control of longitudinal motion and lateral motion, so that the control target in step 101 can be set as follows: ; wherein: represents the expected value of the longitudinal velocity, represents the expected value of the lateral velocity, Denote the expected normal velocity. The three represent the expected earth axis velocities in the three axial directions respectively; is the expected forward flight velocity (i.e., the longitudinal velocity), which can be adjusted manually or according to the trajectory planning algorithm; is the expected steady-state descent rate (when describing the flight process of an aircraft, the descent rate is generally used to represent the descending velocity quantity), which can be adjusted according to and the trimmed descent rate in the steady-state autorotation state.

[0036] In the steady-state autorotation stage, the tracking of the longitudinal velocity and the normal velocity can be achieved through the speed following system of the helicopter.

[0037] As an alternative implementation, for the control objectives in the above steady-state autorotation stage, the expected value of the corresponding velocity can be used to subtract the feedback actual velocity signal and perform PID (Proportional, Integral, Derivative) parameter tuning to achieve the velocity tracking of each axis (longitudinal, lateral, normal).

[0038] The collective pitch control law is also called the collective pitch lever control law.

[0039] In another exemplary embodiment, in step 103 (deceleration and flare stage), the following control objectives are further included: ; Where: is the expected forward flight velocity before touchdown (i.e., the expected longitudinal velocity at touchdown), , are the pitch angle and the roll angle respectively, represents the minimum pitch angle in the deceleration and flare stage; represents the maximum pitch angle in the deceleration and flare stage; represents the minimum roll angle in the deceleration and flare stage; represents the maximum roll angle in the deceleration and flare stage; The minimum values and the maximum values can be given with reference to the flight capabilities and geometric dimensions of the target helicopter.

[0040] For the above control objectives, the expected longitudinal velocity at touchdown can be tracked by designing a PID controller, and the inner loop attitude limit is added to achieve it.

[0041] That is to say, when the target helicopter is in the deceleration and flare stage, the longitudinal motion and the lateral motion of the target helicopter are achieved through the speed following system; for the normal motion, first, based on the estimated remaining airborne time in the deceleration and flare stage at the current moment, the time history of the expected normal acceleration is obtained, and then the tracking of the expected normal acceleration is achieved based on the PID collective pitch control law of the acceleration error; for the situation that does not meet , a large velocity lift signal needs to be applied to the collective pitch ( ) Avoidance is carried out.

[0042] In this paper, the total pitch control law in the deceleration and flare phase and the total pitch control law in the landing phase are collectively referred to as the PID total pitch control law based on acceleration error.

[0043] In another exemplary embodiment, in step 104 (landing phase), the following control objectives are further included: ; Where: represents the minimum pitch angle in the landing phase; represents the maximum pitch angle in the landing phase; represents the minimum roll angle in the landing phase; represents the maximum roll angle in the landing phase; the subscript land represents the landing phase limit, which is convenient for distinguishing from the limit values in the deceleration and flare phase. The selection of the maximum and minimum values of the above attitude angles can be set by the user to apply to helicopters with different landing capabilities.

[0044] The above embodiments of the present application clarify the control objectives of each stage of the autorotation landing of the helicopter, and give the specific total pitch control law design ideas in the deceleration and flare phase and the landing phase. This design idea takes the collision time as the basis for the control law design, which is intuitive and conforms to physical laws. At the same time, a number of artificially set parameters are introduced to apply to different models of helicopters.

[0045] In an exemplary embodiment, a computer device is provided, including a memory and a processor. A computer program is stored in the memory, and when the processor executes the computer program, the steps in the above method embodiments are implemented. As Figure 3 shown.

[0046] In an exemplary embodiment, a computer-readable storage medium is provided, storing a computer program, and when the computer program is executed by a processor, the steps in the above method embodiments are implemented.

[0047] In an exemplary embodiment, a computer program product is provided, including a computer program, and when the computer program is executed by a processor, the steps in the above method embodiments are implemented.

[0048] Those of ordinary skill in the art can understand that all or part of the processes in the methods of the above embodiments can be completed by instructing relevant hardware through a computer program. The computer program can be stored in a non-volatile computer-readable storage medium. When the computer program is executed, it can include the processes of the embodiments of the above methods. Among them, any reference to a memory, database, or other medium used in the embodiments provided in the present application can include at least one of non-volatile and volatile memories. Non-volatile memories can include read-only memory (ROM), magnetic tapes, floppy disks, flash memories, optical memories, high-density embedded non-volatile memories, resistive random access memories (ReRAM), magnetoresistive random access memories (MRAM), ferroelectric random access memories (FRAM), phase change memories (PCM), graphene memories, etc. Volatile memories can include random access memory (RAM) or external cache memories, etc. By way of illustration and not limitation, RAM can be in various forms, such as static random access memory (SRAM) or dynamic random access memory (DRAM), etc.

[0049] The processors involved in the embodiments provided in the present application can be general-purpose processors, central processors, graphics processors, digital signal processors, programmable logics, data processing logics based on quantum computing, etc., without limitation.

[0050] The technical features of the above embodiments can be combined arbitrarily. For the sake of brevity in description, not all possible combinations of the technical features in the above embodiments are described. However, as long as there is no contradiction in the combination of these technical features, it should be considered as the scope recorded in this specification.

[0051] Specific examples are used in this article to elaborate on the principles and implementation manners of the present application. The description of the above embodiments is only used to help understand the method and its core idea of the present application; at the same time, for those of ordinary skill in the art, according to the idea of the present application, there will be changes in the specific implementation manners and application scopes. In summary, the content of this specification should not be construed as a limitation to the present application.

Claims

1. A helicopter autorotation landing control method based on normal acceleration, characterized in that, The helicopter autorotation landing control method based on normal acceleration includes: Determine the autorotation landing control phase of the target helicopter according to the altitude of the target helicopter; the target helicopter is a helicopter that has lost power; When the autorotation landing control phase is the steady autorotation phase, control the normal motion of the target helicopter with the goal of tracking the expected value of the normal velocity; the expected value of the normal velocity is determined according to the expected value of the longitudinal velocity and the trim descent rate; the normal direction is the direction perpendicular to the ground; When the autorotation landing control phase is the deceleration and flare phase, determine the first expected value of the normal acceleration according to the estimated remaining hovering time of the deceleration and flare phase, and use the total pitch control law in the deceleration and flare phase to control the normal motion of the target helicopter with the goal of tracking the first expected value of the normal acceleration; the estimated remaining hovering time of the deceleration and flare phase is determined according to the current altitude, current longitudinal position and current first motion state of the target helicopter; the first motion state includes normal velocity, normal acceleration, longitudinal velocity and longitudinal acceleration; the longitudinal direction is the forward flight direction of the target helicopter; When the autorotation landing control phase is the landing phase, determine the second expected value of the normal acceleration according to the expected total hovering time in the landing phase, and use the total pitch control law in the landing phase to control the normal motion of the target helicopter with the goal of tracking the second expected value of the normal acceleration.

2. The method for controlling autorotation landing of a helicopter based on normal acceleration according to claim 1, wherein The mathematical expression of the total pitch control law in the deceleration and flare phase is as follows: ； Among them, is the collective pitch; is the first-order differential of the collective pitch; is the difference between the expected value of the first normal acceleration and the actual normal jerk; is the first-order differential of; , and are the parameters of the first PID controller; is the first collective pitch raising speed; is the estimated remaining hovering time in the deceleration and flare phase; is the current altitude; is the current normal velocity.

3. The helicopter autorotation landing control method based on normal acceleration according to claim 2, characterized in that, The specific process of determining the first expected value of the normal acceleration according to the estimated remaining hovering time of the deceleration and flare phase includes: Determine the first expected value of the normal acceleration according to the estimated remaining hovering time of the deceleration and flare phase by using the calculation formula for the first expected value of the normal acceleration; the calculation formula for the first expected value of the normal acceleration is: ； wherein, is the expected value of the initial height in the landing phase; is the current height; is the current normal velocity; is the expected value of the first normal acceleration.

4. The helicopter autorotation landing control method based on normal acceleration according to claim 3, wherein The determination process of the estimated remaining hovering time of the deceleration and flare phase specifically includes: Determine the estimated remaining hovering time of the deceleration and flare phase according to the current altitude, current longitudinal position and current first motion state of the target helicopter by using the calculation formula for the remaining hovering time of the deceleration and flare phase; the calculation formula for the remaining hovering time of the deceleration and flare phase is: ； Among them, is transition value; is an insurance value greater than or equal to zero; is maximum value; The calculation formula is as follows: ； wherein, is the current normal acceleration, is the current longitudinal position; is the current longitudinal velocity; is the current longitudinal acceleration; is the expected value of the normal initial velocity in the landing phase; is the expected value of the touchdown longitudinal velocity; The calculation formula is as follows: ； Among them, is the available energy, which is obtained by adding the rotational kinetic energy and the longitudinal kinetic energy of the current target helicopter rotor; is the ideal available energy, which is obtained by adding the rotational kinetic energy and the longitudinal kinetic energy of the ideal helicopter rotor; is the maximum duration of the deceleration and flare phase.

5. The helicopter autorotation landing control method based on normal acceleration according to claim 1, characterized in that The mathematical expression of the total pitch control law in the landing phase is as follows: ； Among them, is the collective pitch; is the first-order differential of the collective pitch; is the difference between the expected value of the second normal acceleration and the actual normal jerk; is 's first-order differential; , , are the parameters of the second PID controller respectively; is the second collective pitch increasing speed; is the expected value of the total hovering time during the landing phase; is the current altitude; is the current normal velocity.

6. The method for controlling autorotation landing of a helicopter based on normal acceleration according to claim 5, characterized in that, The specific process of determining the second expected value of the normal acceleration according to the expected total hovering time in the landing phase includes: Determine the second expected value of the normal acceleration according to the expected total hovering time in the landing phase by using the calculation formula for the second expected value of the normal acceleration; the calculation formula for the second expected value of the normal acceleration is: ； Among them, is the touchdown height in the landing phase, equal to 0; is the expected value of the second normal acceleration is the current height; is the current normal velocity; is the expected value of the total airtime in the landing phase.

7. The helicopter autorotation landing control method based on normal acceleration according to claim 5, characterized in that The expected value of the total hang time in the landing phase is determined based on the expected value of the normal initial velocity in the landing phase and the expected value of the initial height in the landing phase and is determined accordingly.

8. A computer device, comprising: A memory, a processor, and a computer program stored on the memory and executable on the processor, wherein the processor executes the computer program to implement the helicopter autorotation landing control method based on normal acceleration according to any one of claims 1-7.

9. A computer-readable storage medium, having a computer program stored thereon, characterized in that, When the computer program is executed by the processor, it implements the helicopter autorotation landing control method based on normal acceleration according to any one of claims 1-7.

10. A computer program product, comprising a computer program, characterized in that, When the computer program is executed by a processor, it implements the normal-acceleration-based autorotation landing control method for a helicopter according to any one of claims 1-7.

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