An electromagnetic drive disc-shaped aircraft control design method

Abstract

The invention relates to the field of aerospace technology, and particularly to a control design method for an electromagnetic-driven disc-shaped aircraft, which includes an internal stationary structure and an external rotating structure, and the external rotating structure is driven by electromagnetic driving force. A universal joint structure is arranged between the internal stationary structure and the external rotating structure, and there is a certain interval among the three, and they are sequentially connected in pairs through a transverse connecting pin shaft and a longitudinal connecting pin shaft. Two mutually perpendicular actuators are arranged between the internal stationary structure and the external rotating structure to realize the heading and elevation control of the aircraft. When the actuators extend or contract, the internal stationary structure and the external rotating structure deflect in opposite directions around the central axis. Through the cooperation of relevant structures, the invention solves the problem that for an electromagnetic-driven disc-shaped aircraft, due to its unconventional aerodynamic shape and special internal stationary and external rotating structure settings, the conventional control design method cannot well control it.

Description

Technical Field

[0001] The invention relates to the field of aerospace technology, and particularly to a method for designing the control of a disc-shaped aircraft driven by electromagnetic force. Background Art

[0002] With the increasing development demand of new energy aircraft and the rapid development of electromagnetic drive technology, the disc-shaped aircraft driven by electromagnetic force as the power has become a potential emerging research topic of new energy aircraft.

[0003] The control of conventional aircraft, such as the motion modes of rotary-wing aircraft mainly include ascending and descending, translation, and heading motion, and the motion modes of fixed-wing aircraft mainly include pitching, rolling, yawing, and accelerating and decelerating motion. The control mechanisms of rotary-wing aircraft mainly consist of parts such as the collective pitch control lever, the foot pedal, and the throttle lever; the control mechanisms of fixed-wing aircraft mainly consist of parts such as the two-axis side stick, the central stick or central disk, the foot pedal, and the throttle lever.

[0004] However, different from conventional aircraft, due to its unconventional aerodynamic shape and special internal static and external rotating structural configuration, the conventional control design method cannot well control the disc-shaped aircraft driven by electromagnetic force. Summary of the Invention

[0005] In view of this, the purpose of the invention is to propose a method for designing the control of a disc-shaped aircraft driven by electromagnetic force to solve the problem that the conventional control design method cannot well control the disc-shaped aircraft driven by electromagnetic force due to its unconventional aerodynamic shape and special internal static and external rotating structural configuration.

[0006] Based on the above purpose, the invention provides a method for designing the control of a disc-shaped aircraft driven by electromagnetic force, including an internal static structure and an external rotating structure, and the external rotating structure is driven by electromagnetic driving force. A universal joint structure is arranged between the internal static structure and the external rotating structure, and there is a certain interval among the three, and they are connected in pairs in sequence through a transverse connecting pin shaft and a longitudinal connecting pin shaft. Two mutually perpendicular actuators are arranged between the internal static structure and the external rotating structure to realize the heading and ascending and descending control of the aircraft. Among them, the method for designing the control of the disc-shaped aircraft is as follows;

[0007] Step 1: When the actuator extends or contracts, the internal static structure and the external rotating structure deflect in opposite directions around the central axis. An auxiliary actuator is arranged on the other side symmetrical to each actuator.

[0008] Step 2: The actuator located longitudinally is the longitudinal actuator, and the actuator located transversely is the transverse actuator. When the longitudinal actuator and the transverse actuator are operated simultaneously, the aircraft deflects in the skewed direction synthesized in the transverse and longitudinal directions.

[0009] Step 3: The actuator applies control loads to the internal stationary structure and the external rotating structure according to the determined control principle, and the flight attitude deflection and lifting operations of the aircraft follow the control principle and are controlled by the actuator and electromagnetic drive respectively.

[0010] Preferably, the control principle of the actuator includes:

[0011] Step 3.1: Establish the o-xyz geodetic coordinate system;

[0012] Step 3.2: Establish the internal stationary structure balance equation in the xoy plane;

[0013] Step 3.3: Establish the external rotating structure balance equation in the xoy plane;

[0014] Step 3.4: Combine the internal and external structure balance equations obtained in Steps 3.2 and 3.3 to obtain the overall structure balance equation, and perform functional characterization on the aerodynamic force perpendicular to the rotation plane generated by the rotating blades, the aerodynamic lift perpendicular to the course generated by the airflow, and the aerodynamic drag in the course direction generated by the airflow;

[0015] Step 3.5: Based on Step 3.4, deduce the relationship between the blade rotation speed and the aircraft navigation speed, and then determine all the parameters of the stable flight state;

[0016] Step 3.6: Take the xoz and yoz planes, repeat Steps 3.2 to 3.5, and obtain all the parameters of the flight state in the xoz and yoz planes.

[0017] Preferably, in Step 3.2, the internal stationary structure balance equation is:

[0018]

[0019]

[0020]

[0021] θ in1 and θ out1 are the rotation angles of the centers of gravity of the internal stationary structure and the external rotating structure around the transverse pin axis respectively, h in and h B1 are the distances from the center of gravity of the internal stationary structure and the action point B1 of the actuator to the transverse pin axis respectively, F B1 is the force applied at both ends of the longitudinally arranged actuator to the support points B1 in and B1 out of the internal stationary structure and the external rotating structure, G in is the weight of the internal stationary structure, FRx1 , F Ry1 are respectively the reaction forces of the horizontal pin shaft on the internal stationary structure and the external rotating structure in the longitudinal and vertical directions.

[0022] Preferably, in step 3.3, the equilibrium equation of the external rotating structure is:

[0023]

[0024]

[0025]

[0026] h out is the distance from the center of gravity of the external structure to the axis of the horizontal pin shaft. The distance from the action point of the actuator at this point to the axis of the horizontal pin shaft is equal to the distance from the action point of the actuator on the internal structure to the axis of the horizontal pin shaft; F D1 is the aerodynamic drag generated by the airflow in the course direction; h a is the distance from the aerodynamic center of the course airflow to the axis of the horizontal pin shaft; F L1 is the aerodynamic lift perpendicular to the course generated by the airflow; F u1 is the aerodynamic force perpendicular to the rotating plane generated by the rotating blade; G out1 is the weight of the external structure.

[0027] Preferably, in step 3.4, the equilibrium equation of the overall structure is:

[0028] G out h out sinθ out1 -G in h in sinθ in1 =F L1 h a sinθ out1 -F D1 h a cosθ out1

[0029] F u1 sinθ out1 =F D1

[0030] F u1 cosθ out1 =G out1 +G in +F L1

[0031] In the above equations, F u1 , F L1 , F D1It can be expressed in the following form:

[0032] F u1 = k u (s, V)ω 2

[0033] F L1 = k L (s, θ out1 , ω)V 2

[0034] F D1 = k D (s, θ out1 , ω)V 2

[0035] ω and V are respectively the rotational angular velocity of the blade and the velocity of the air flow in the course direction, and k u (s, V), k L (s, θ out1 , ω) and k D (s, θ out1 , ω) are respectively the coefficients related to the structural feature s, the inclination angle θ of the aircraft out1 and ω, and their characteristic curves are determined through calculation and experiment.

[0036] Preferably, in step 3.5, the relationship between the blade rotation speed and the aircraft navigation speed is deduced as follows:

[0037] k u (s, V)ω 2 = k L (s, θ out 1 , ω)V 2 + G out1 + G in

[0038] It can be seen from this that when ω increases, if θ out1 remains unchanged, the aircraft navigation speed V increases at this time; when the rotational angular velocity ω of the blade increases, the inclination angle θ is changed simultaneously out1 , at this time, due to the change of the inclination angle, the vertical balance of the aircraft structure is broken, so vertical lifting occurs to reach a new navigation state, and the navigation speed V also changes accordingly.

[0039] Preferably, in step 3.5, for the relationship between the blade rotation speed and the aircraft navigation speed, substituting it into the two groups of equations obtained in step 3.4 and step 3.2, the following flight state parameters can be obtained:

[0040]

[0041]

[0042] F Ry1 = G in + F B1 (θ out1 - θ in1 )

[0043] Advantages of the invention: A method for controlling the design of a disc-shaped aircraft driven by electromagnetic force is provided. By setting the dimensional linear actuator, the disc-shaped aircraft can be simply and efficiently controlled. The external structure is divided into two parts, upper and lower or inner and outer. The two parts rotate around the common central axis in opposite directions and maintain the total angular momentum to be zero. Since both parts have a large moment of inertia, each part has gyro-like attitude stability relative to its central axis, that is, its attitude is not easily affected by external disturbances, thus ensuring that the entire aircraft has high stability, that is, it is not easily affected by external disturbance airflows. BRIEF DESCRIPTION OF THE DRAWINGS

[0044] In order to more clearly illustrate the technical solutions in the embodiments of the invention or the prior art, the following will briefly introduce the drawings required for the description of the embodiments or the prior art. Obviously, the drawings in the following description are only the embodiments of the invention. For those of ordinary skill in the art, other drawings can be obtained based on these drawings without creative efforts.

[0045] Figure 1 It is the design drawing of the embodiment of the invention;

[0046] Figure 2 It is for the Figure 1 internal structure schematic diagram of the embodiment of the invention;

[0047] Figure 3 It is the design drawing of the external rotating structure and fixed frame of the embodiment of the invention;

[0048] Figure 4 It is the X-Y central sectional view of the gimbal connection structure and actuator arrangement design of the embodiment of the invention;

[0049] Figure 5 It is the Z-Y sectional view of the gimbal connection structure and actuator arrangement design of the embodiment of the invention;

[0050] Figure 6 It is the sectional view of E-E of the embodiment of the invention;

[0051] Figure 7 It is the sectional view of F-F of the embodiment of the invention;

[0052] Figure 8 It is the vertical x-y sectional view after the flight attitude control of the embodiment of the invention;

[0053] Figure 9Schematic diagram of force analysis for flight drive and control in an invention embodiment.

[0054] In the figure: 1. Internal stationary structure; 2. External rotating structure; 3. Gimbal connection structure; R1. Transverse connection pin; R2. Longitudinal connection pin; D1. Longitudinal actuator; D2. Transverse actuator; D3. Longitudinal auxiliary actuator; D4. Transverse auxiliary actuator. Detailed implementation manners

[0055] To make the purpose, technical solutions and advantages of the invention clearer and more understandable, the following further elaborates on the invention in detail with reference to specific embodiments and the accompanying drawings.

[0056] It should be noted that unless otherwise defined, the technical terms or scientific terms used in the invention embodiments should have the ordinary meanings understood by those with ordinary skills in the field to which the invention belongs. The "first", "second" and similar terms used in the invention do not indicate any order, quantity or importance, but are only used to distinguish different components. Words such as "including" or "comprising" mean that the elements or objects appearing before this word cover the elements or objects listed after this word and their equivalents, without excluding other elements or objects. Words such as "connected" or "linked" are not limited to physical or mechanical connections, but may include electrical connections, whether direct or indirect. "Upper", "lower", "left", "right", etc. are only used to represent relative position relationships, and when the absolute position of the object being described changes, the relative position relationship may also change accordingly.

[0057] As Figure 1 、 Figure 2 、 Figure 3 、 Figure 4 、 Figure 5 、 Figure 6 、 Figure 7 、 Figure 8 、 Figure 9 shown, a control design method for an electromagnetic-driven disc-shaped aircraft includes an internal stationary structure 1 and an external rotating structure 2. The internal stationary structure is a non-rotating structure, serving as a passenger and cargo cabin, cockpit, landing gear compartment, and upper and lower connection channels, etc. And the external rotating structure 2 is driven by electromagnetic driving force. A gimbal connection structure 3 is provided between the internal stationary structure 1 and the external rotating structure 2, and there is a certain interval among the three, and they are connected in pairs in sequence by a transverse connection pin R1 and a longitudinal connection pin R2. The external rotating structure 2 consists of two symmetrically arranged sub-rotating structures, and both sub-rotating structures can rotate circumferentially, and the rotation directions of the two sub-rotating structures are opposite. At this time, the total angular momentum of the system is zero. Two mutually perpendicular actuators are provided between the internal stationary structure 1 and the external rotating structure 2 to achieve the heading and elevation control of the aircraft. Among them, the control design method of the disc-shaped aircraft is as follows:

[0058] Step 1: When the actuator extends or contracts, the internal stationary structure 1 and the external rotating structure 2 deflect in opposite directions around the central axis. Auxiliary actuators are provided on the other side of each actuator symmetrically. The longitudinally arranged one is the longitudinal auxiliary actuator D3, and the transversely arranged one is the transverse auxiliary actuator D4.

[0059] Step 2: The longitudinally arranged actuator is the longitudinal actuator D1, and the transversely arranged actuator is the transverse actuator D2. When the longitudinal actuator D1 and the transverse actuator D2 are operated simultaneously, the aircraft deflects in the skewed direction synthesized in the transverse and longitudinal directions.

[0060] Step 3: The actuator applies control loads to the internal stationary structure 1 and the external rotating structure 2 according to a determined control principle. The deflection and lifting operations of the flight attitude of the aircraft follow the control principle and are controlled by the actuator and electromagnetic drive respectively.

[0061] In a preferred embodiment of the invention, the control principle of the actuator includes:

[0062] Step 3.1: Establish the o-xyz geodetic coordinate system.

[0063] Step 3.2: Establish the balance equation of the internal stationary structure in the xoy plane.

[0064] In Step 3.2, the balance equation of the internal stationary structure is:

[0065]

[0066]

[0067]

[0068] θ in1 and θ out1 are the rotation angles of the centers of gravity of the internal stationary structure and the external rotating structure around the transverse pin axis respectively. h in and h B1 are the distances from the center of gravity of the internal stationary structure and the action point B1 of the actuator to the transverse pin axis respectively. F B1 is the force applied at both ends of the longitudinally arranged actuator to the support points B1 in and B1 out on the internal stationary structure and the external rotating structure. G in is the weight of the internal stationary structure. F Rx1 and F Ry1 are the reaction forces of the transverse pin on the internal stationary structure and the external rotating structure in the longitudinal and vertical directions respectively.

[0069] Step 3.3, establish the external rotation structure equilibrium equation in the xoy plane;

[0070] In step 3.3, the equilibrium equation of the external rotating structure is:

[0071]

[0072]

[0073]

[0074] h out F is the distance from the center of gravity of the external structure to the axis of the transverse pin. The distance from the actuator action point to the axis of the transverse pin is equal to the distance from the actuator action point on the internal structure to the axis of the transverse pin. D1 is the aerodynamic resistance in the heading direction caused by the airflow; h a F is the distance from the aerodynamic center of the heading airflow to the axis of the transverse pin; L1 F is the aerodynamic lift generated by the airflow perpendicular to the heading direction; u1 G is the aerodynamic force perpendicular to the rotation plane generated by the rotating blades; out1 The weight of the external structure.

[0075] Step 3.4: Combine the internal and external structural balance equations obtained in Steps 3.2 and 3.3 to obtain the overall structural balance equation, and provide functional representations for the aerodynamic force perpendicular to the plane of rotation generated by the rotating blades, the aerodynamic lift perpendicular to the heading generated by the airflow, and the aerodynamic drag in the heading direction generated by the airflow;

[0076] In step 3.4, the overall structural equilibrium equation is:

[0077] G out h out sinθ out1 -G in h in sinθ in1 =F L1 h a sinθ out1 -F D1 h a cosθ out1

[0078] F u1 sinθ out1 =F D1

[0079] F u1 cosθ out1 =G out1 +G in +FL1

[0080] In the above equation, F u1 , F L1 , F D1 can be expressed in the following form:

[0081] F u1 = k u (s, V)ω 2

[0082] F L1 = k L (s, θ out1 , ω)V 2

[0083] F D1 = k D (s, θ out1 , ω)V 2

[0084] ω and V are the rotational angular velocity of the blade and the velocity of the air flow in the course direction respectively. k u (s, V), k L (s, θ out1 , ω) and k D (s, θ out1 , ω) are coefficients related to the structural feature s, the inclination angle θ out1 of the aircraft and ω respectively, and their characteristic curves are determined by calculation and experiment.

[0085] In step 3.5, the relationship between the blade rotation speed and the aircraft navigation speed is deduced as follows:

[0086] k u (s, V)ω 2 = k L (s, θ out 1 , ω)V 2 + G out1 + G in

[0087] It can be seen from this that when ω increases, if θ out1 remains unchanged, the aircraft navigation speed V increases at this time; when the rotational angular velocity ω of the blade increases, the inclination angle θ out1 is changed at the same time. At this time, due to the change of the inclination angle, the vertical balance of the aircraft structure is broken, so vertical lifting occurs to reach a new navigation state, and the navigation speed V also changes accordingly.

[0088] In step 3.5, for the relationship between the blade rotation speed and the aircraft navigation speed, substituting it into the two sets of equations obtained in step 3.4 and step 3.2, the following flight state parameters can be obtained:

[0089]

[0090]

[0091] F Ry1 = G in + F B1 (θ out1 - θ in1 )

[0092] It can be seen from this that: (1) Given the structural weight of the internal and external structures and the thrust applied to the actuator, all references of the flight states such as the blade rotation speed, the flight speed of the aircraft, the inclination angle of the aircraft, and the reaction force at the pin shaft can be calculated; (2) As can be seen from the flight state parameter function derived in this step. In the operation scheme designed in this application, the thrust of the actuator is directly proportional to the inclination angle of the internal structure and also shows a proportional trend with the inclination angle of the external structure, which is beneficial to the design of the flight operation system.

[0093] Step 3.6, take the xoz and yoz planes, and repeat Steps 3.2 to 3.5 to obtain all the parameters of the flight states in the xoz and yoz planes.

[0094] Those of ordinary skill in the art should understand that: The discussion of any embodiment above is only exemplary and is not intended to imply that the scope of the invention is limited to these examples; Under the concept of the invention, the technical features in the above embodiments or different embodiments can also be combined, and the steps can be implemented in any order, and there are many other variations in different aspects of the invention as described above, which are not provided in detail for the sake of brevity.

[0095] The embodiments of the invention are intended to cover all such substitutions, modifications, and variations that fall within the broad scope of the appended claims. Therefore, any omission, modification, equivalent substitution, improvement, etc. made within the spirit and principle of the invention shall be included within the protection scope of the invention.

Claims

1. An electromagnetic-driven control design method for a disc-shaped aircraft, comprising an internal stationary structure (1) and an external rotating structure (2), and the external rotating structure (2) is driven by electromagnetic driving force. A universal joint connection structure (3) is arranged between the internal stationary structure (1) and the external rotating structure (2), and there is a certain interval among the three, and they are sequentially connected in pairs by a transverse connection pin shaft (R1) and a longitudinal connection pin shaft (R2). It is characterized in that, The external rotation structure (2) is composed of two sub-rotation structures that are symmetrically arranged up and down. Both sub-rotation structures can perform circumferential rotation, and the rotation directions of the two sub-rotation structures are opposite. Two mutually perpendicular actuators are arranged between the internal stationary structure (1) and the external rotation structure (2) to achieve the heading and elevation control of the aircraft. Among them, the control design method of the disc-shaped aircraft is as follows: Step 1: When the actuator extends or contracts, the internal stationary structure (1) and the external rotation structure (2) deflect in opposite directions around the central axis. Auxiliary actuators are arranged on the other side symmetrical to each actuator. Step 2: The actuator located longitudinally is the longitudinal actuator (D1), and the actuator located transversely is the transverse actuator (D2). When the longitudinal actuator (D1) and the transverse actuator (D2) are operated simultaneously, the aircraft deflects in the skewed direction synthesized in the transverse and longitudinal directions. Step 3: The actuator applies control loads to the internal stationary structure (1) and the external rotation structure (2) according to the determined control principle. The flight attitude deflection and elevation operation of the aircraft follow the control principle and are controlled by the actuator and electromagnetic drive respectively. The control principle of the actuator includes: Step 3.1: Establish the o-xyz geodetic coordinate system. Step 3.2: Establish the internal stationary structure balance equation in the xoy plane. Step 3.3: Establish the external rotation structure balance equation in the xoy plane. Step 3.4: Combine the internal and external structure balance equations obtained in Step 3.2 and Step 3.3 to obtain the overall structure balance equation, and characterize the aerodynamic force perpendicular to the rotation plane generated by the rotating blades, the aerodynamic lift perpendicular to the heading generated by the airflow, and the aerodynamic drag in the heading direction generated by the airflow as functions. Step 3.5: Based on Step 3.4, deduce the relationship between the blade rotation speed and the aircraft navigation speed, and then determine all the parameters of the stable flight state. Step 3.6: Take the xoz and yoz planes, and repeat Steps 3.2 to 3.5 to obtain all the parameters of the flight state in the xoz and yoz planes.

2. The electromagnetic drive disc-type aircraft control design method according to claim 1, characterized in that In Step 3.2, the internal stationary structure balance equation is: θ in1 and θ out1 are the angles of rotation of the centers of gravity of the internal stationary structure and the external rotating structure about the axis of the transverse pin, respectively. h in and h B1 are the distances from the center of gravity of the internal stationary structure and the action point B1 of the actuator to the axis of the transverse pin, respectively. F B1 is the force applied at both ends of the longitudinally arranged actuator to the support points B1 in and B1 out on the internal stationary structure and the external rotating structure, respectively. G in is the weight of the internal stationary structure. F Rx1 and F Ry1 are the reaction forces of the transverse pin on the internal stationary structure and the external rotating structure in the longitudinal and vertical directions, respectively.

3. A method for controlling the design of a disc-shaped flying vehicle driven by electromagnetic force according to claim 2, characterized in that, In Step 3.3, the external rotation structure balance equation is: h out is the distance from the center of gravity of the external structure to the axis of the transverse pin, where the distance from the actuator action point to the axis of the transverse pin is equal to the distance from the actuator action point on the internal structure to the axis of the transverse pin; F D1 is the aerodynamic drag in the course direction generated by the airflow; h a is the distance from the aerodynamic center of the course airflow to the axis of the transverse pin; F L1 is the aerodynamic lift perpendicular to the course generated by the airflow; F u1 is the aerodynamic force perpendicular to the rotation plane generated by the rotating blade; G out1 is the weight of the external structure.

4. A method for controlling the design of a disc-shaped flying vehicle driven by electromagnetic force according to claim 3, characterized in that In Step 3.4, the overall structure balance equation is: G out h out sinθ out1 -G in h in sinθ in1 = F L1 h a sinθ out1 -F D1 h a cosθ out1 F u1 sinθ out1 = F D1 F u1 cosθ out1 = G out1 + G in + F L1 In the above equation, F u1 , F L1 , F D1 can be expressed in the following form: F u1 = k u (s, V)ω 2 F L1 = k L (s, θ out1 , ω)V 2 F D1 = k D (s, θ out1 , ω)V 2 ω and V are the rotational angular velocity of the blade and the velocity of the air flow in the course direction respectively, k u (s, V), k L (s, θ out1 , ω) and k D (s, θ out1 , ω) are coefficients related to the structural feature s, the inclination angle θ of the aircraft out1 and ω respectively, and their characteristic curves are determined by calculation and experiment.

5. A method for designing the control of a disc-shaped aircraft driven by electromagnetic force according to claim 4, characterized in that, In Step 3.5, the relationship between the blade rotation speed and the aircraft navigation speed is deduced as follows: k u (s,V)ω 2 = k L (s,θ out1 ,ω)V 2 + G out1 + G in It can be seen from this that when ω increases, if θ out1 remains unchanged, the flight speed V of the aircraft increases at this time; when the rotational angular velocity ω of the blade increases, the inclination angle θ out1 is changed at the same time. At this time, due to the change of the inclination angle, the vertical balance of the aircraft structure is broken, resulting in vertical lifting to reach a new flight state, and the flight speed V also changes accordingly.

6. A method for designing the control of a disc-shaped flying vehicle driven by electromagnetic force according to claim 5, characterized in that, In Step 3.5, substitute the relationship between the blade rotation speed and the aircraft navigation speed into the two sets of equations obtained in Step 3.4 and Step 3.2, and the following flight state parameters can be obtained: F Ry1 = G in + F B1 (θ out1 - θ in1 )。

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