1. A multi-dimensional vibration control method for a strut tail-supported aircraft model, wherein the method is through the arrangement in the aircraft model on the center of mass of pitch and yaw acceleration sensor measuring aircraft model, the main vibration acceleration of two component, calculate a main vibration vector and determine a strut real-time aircraft model plane, introduction of inertial force to solve a multidimensional active vibration damper on an active section by dynamic bending moment, and then obtain an initiative stress distribution on the active section, through the real-time vibration plane space position relation of multidimensional vibration damper to participate in the work of piezoelectric ceramic actuator serial number; a vibration control force is calculated in real time according to the stress on the active section of a piezoelectric ceramic actuator, and then the dynamic bending moment is generated in the process of a reverse bending moment resisting the vibration of the aircraft model; the multi-dimensional vibration active control system based on the piezoelectric ceramic actuator is adopted to control the multi-dimensional vibration; the specific steps are as follows:step 1: establish an absolute coordinate system of the aircraft model support system

the absolute coordinate system OXYZ (E) is established on an aircraft tail strut (4), and the coordinate origin is established in the equilibrium position at the intersection of the active section (F) and the axis of the aircraft tail strut (4), which is defined as O; the direction of the X axis coincides with the balance position of the axis of the aircraft tail strut (4) and points to the aircraft model (5),the direction of the Y axis is that the intersection of the active section (F) and the pitching plane points upward; the Z axis is determined by the right manipulation; a vibration measurement coordinate system OAXAYAZA(A) is established on the aircraft model (5), whose origin is established at the intersection of the centroid of the aircraft model (5) and the X axis of the absolute coordinate system OXYZ (E), which is defined as that the direction of the OA;XA coordinate axis coincides with the X axis of the absolute coordinate system OXYZ (E), the YA coordinate axis and the Y axis of the absolute coordinate system OXYZ (E) point upward, and the ZA coordinate axis is determined by right manipulation;step 2: obtain the components of the main vibration acceleration in the pitch plane and yaw plane in real time

using a pitching accelerometer (6) and a yaw accelerometer (7) at the centroid of the aircraft model (5) to measure the acceleration of the main vibration in the pitch plane and yaw plane perpendicular to each other, the acceleration of the main vibration is fed back to a real-time controller (8) controlled by an upper computer (9), and a plurality of acceleration sampling values of the pitch plane and the yaw plane are collected in each vibration control cycle, the acceleration components of the main vibration acceleration in the pitch direction and yaw direction in a vibration control cycle are calculated by formulas (1) and (2) respectively:

among them, the acceleration component of the apith(t) main vibration acceleration in the pitch direction, the acceleration component of the ayaw(t) main vibration acceleration in the yaw direction, apithi(t) , ayawi(t) is the acceleration sampling value of the aircraft model (5) in the pitching plane and the yaw plane at the i sampling time, and N is the number of acceleration sampling values in each vibration control cycle, wherein i=1,2. . . N;

step 3: solve the main vibration acceleration vector in real time

the main vibration acceleration is obtained by combining the acceleration components in the pitching direction and yaw direction; the main vibration acceleration consists of magnitude and direction; the main vibration acceleration vector is constructed by solving the magnitude and direction of the main vibration acceleration vector in each vibration control cycle in real time by using formulas (3) and (4):

among them, a(t) is the main vibration acceleration vector, |a(t) | is the magnitude of the a(t) vibration acceleration vector, and ?a(t) is the main vibration acceleration vector in a(t) directions;

step 4: establish a real-time vibration active control coordinate system of the aircraft model and determine the real-time vibration plane of the strut

the real-time active vibration control coordinate system O?X?Y?Z?(D) is established on the active section (F), and its origin coincides with the origin O of the absolute coordinate system OXYZ (E); it is defined that the direction of the O?; X?axis coincides with the direction of the X coordinate of the absolute coordinate OXYZ (E), the Y?axis coincides with the a(t) direction of the main vibration acceleration vector, and the Z60 coordinate axis is determined by the right manipulation; plane X?O?Y?is the real-time vibration plane X?O?Y?(C) of the strut; because of the randomness of the vibration of the aircraft model (5), the real-time active vibration control coordinate system O?X?Y?Z?(D) changes with time, and the real-time vibration plane X?O?Y?(C) changes with time;

step 5: the real-time inertial force and the stress distribution on the active section of the support system are solved

on the real time vibration plane X?O?Y?(C) of the aircraft tail strut (4), the inertia force is solved in real time by formula (5)

FI(t)=?meqa(t) (5)

formula (6) was used to calculate the dynamic bending moment on the active section (F) in real time

M(t)=FI(t)·L (6)

a dynamic stress distribution on the active section (F) was solved in real time by formula (7)

among them, meq is the equivalent mass of the support system, FI(t) is the real-time inertia force acting on the aircraft model (5), M(t) is the dynamic bending moment on the active section (F) of the aircraft model (5) during vibration, L is the distance from the centroid of the aircraft model (5) to the active section (F), ?(ya, za, t) is the dynamic stress at the length of the active section (F) inner distance X?coordinate axis ya, and Iz ?(t) is the real time inertia moment of the active section (F) to the Z?coordinate axis;

step 6: determine the serial number of the piezoelectric ceramic actuator in real time and calculate the vibration control force

a number of the piezoelectric ceramic actuators (3-1) are uniformly arranged in the circumferential direction of the multi-dimensional vibration damper (3) at the active section (F), the uniformly distributed circumferential radius is R, and the piezoelectric ceramic which coincides with the Z axis of the absolute coordinate system OXYZ (E) is set as No. 0 piezoelectric ceramic actuator (3-1), the No. 1 piezoelectric ceramic actuator (3-1), the No. 2 piezoelectric ceramic actuator (3-1), . . . , the No. n piezoelectric ceramic actuator (3-1), are arranged in a counterclockwise circular array in turn; the array angle between two adjacent piezoelectric ceramic actuators (3-1) is

in the real-time active vibration control coordinate system O?X?Y?Z?(D), the piezoelectric ceramic actuator (3-1) above the Z?axis participates in the vibration control, and the serial number of the piezoelectric ceramic actuator (3-1) participating in the work is:among them,represent the rounding of the calculated values ofrespectively, ?(t) is an angle between the main vibration acceleration vector a(t) and the Z axis of the absolute coordinate system OXYZ (E), and then the real-time coordinates of the center of the piezoelectric ceramic actuator (3-1) in the real-time active vibration control coordinate system O?X?Y?Z?(D) are determined as follows:

where, ?n c, is the angle between participating piezoelectric ceramic actuator (3-1) and the Z axis direction of the absolute coordinate system OXYZ (E), and the resultant force on the active section (F) of the participating piezoelectric ceramic actuator (3-1) is:

where,

is the contact area between no. nc participating piezoelectric ceramic actuator (3-1) and the active section (F), and resistance required by the no. nc participating piezoelectric ceramic actuator (3-1) is:

finally, all the participating piezoelectric ceramic actuators (3-1) generate the reverse bending moment MR (t) to resist the dynamic bending moment M(t) generated during the vibration of the aircraft model (5).