Aircraft variable-core excitation forced angular oscillation wash time difference derivative testing device

Abstract

The application discloses a kind of aircraft variable turning center excitation forced angular oscillation wash time difference derivative testing device, belong to high-speed wind tunnel dynamic aerodynamic derivative testing technical field, comprising: transmission shaft, its end portion is fixedly installed with driving motor;Measuring balance, one end is connected with driving motor by motion conversion mechanism, the other end is installed with aircraft model;Tail support rod, it is sleeved in the outside of transmission shaft, and the measuring balance is provided with a pintle pivot, and the end of tail support rod is rotatably connected with measuring balance by pintle pivot.The application accurately adjusts the oscillation turning center position of aircraft model in high-speed wind tunnel by changing the length of tail support rod, so that it is located before or after the moment reference center, and then the pitch moment response under different turning center conditions is obtained.The application combines theoretical formula derivation, separates the pure wash time difference derivative from two or more groups of variable turning center forced oscillation test data through linear superposition and difference operation.

Description

Technical Field

[0001] This invention belongs to the field of high-speed wind tunnel dynamic aerodynamic derivative testing technology. More specifically, this invention relates to a device for testing the derivative of the forced angle oscillation wash flow time difference of an aircraft with a variable center of rotation, used to separate the derivative of the rate of change of angle of attack (the derivative of the wash flow time difference) and the derivative of the rotational angular rate in the forced angle oscillation test of an aircraft. Background Technology

[0002] In the design and research of aircraft, especially high-speed aircraft, the accurate acquisition of dynamic aerodynamic derivatives is crucial, as they directly relate to the aircraft's dynamic stability, maneuverability, and control law design. Among these, the airflow time-delay characteristics reflected by unsteady aerodynamic effects have a significant impact on flight dynamics, and the "washout time-delay derivative" is one of the core parameters for quantifying this unsteady aerodynamic torque (particularly critical in the pitch direction). Traditional forced oscillation experimental techniques are the standard method for measuring dynamic derivatives in wind tunnels. This involves forcing the model to undergo single-degree-of-freedom harmonic oscillations (such as pitch oscillations) around a fixed axis, measuring the unsteady aerodynamic response during the oscillation process, and then using an identification algorithm to obtain the derivatives.

[0003] However, existing technologies have a significant limitation: the center of rotation of the oscillation is usually fixed at the moment reference center (usually the center of mass or aerodynamic center) of the aircraft model in experiments. This setting inevitably couples two physical effects when measuring the derivative of the pitch moment about this point: (1) the derivative component caused by the change in rotational angular rate (angular velocity); and (2) the derivative component reflecting the fluid time delay (washout time difference) effect caused by the rate of change of angle of attack (angular acceleration) (i.e., the washout time difference derivative itself). This coupling makes it impossible to effectively separate the pure washout time difference derivative from the results of a single standard forced oscillation test. Since the washout time difference derivative is directly related to the physical nature of unsteady aerodynamics, its identification accuracy is insufficient or it cannot be obtained independently, which seriously restricts the depth of research on the unsteady aerodynamic characteristics of aircraft and also brings uncertainty and risk to the design of control systems that rely on precise aerodynamic parameters. Therefore, there is an urgent need to develop a new experimental technique and device that can decouple the measurement and accurately identify the time difference derivative of the wash flow to meet the urgent needs of the refined design and safety assessment of modern high-performance aircraft. Summary of the Invention

[0004] One object of the present invention is to solve at least the above-mentioned problems and / or defects, and to provide at least the advantages described below.

[0005] This invention provides a device for testing the derivative of the time difference of oscillation washing flow under the excitation forced angle of a variable rotation center of gravity of an aircraft, comprising: A drive shaft, with a drive motor fixedly mounted at its end; A measuring balance, one end of which is connected to the drive motor via a motion conversion mechanism, and an aircraft model is mounted on the other end of the measuring balance; The tail support rod is sleeved on the outside of the drive shaft. The measuring balance is provided with a pin shaft. The end of the tail support rod is rotatably connected to the measuring balance through the pin shaft.

[0006] Preferably, the structure of the motion conversion mechanism includes: A cam, which is slidably connected to the output shaft of the drive motor; A spherical bearing, the inner ring of which is slidably connected to the cam.

[0007] Preferably, the measuring balance is a five-component strain balance, strain gauges are attached to the measuring beam of the measuring balance, the strain gauges are connected by a Wheatstone bridge, and the signal output terminals of the strain gauges are connected to a dynamic data acquisition system.

[0008] Preferably, the structure of the measuring balance includes: The model connection segment is connected in series with the middle structural segment and the eccentric shaft connection segment at its rear end; The model connection segment is provided with a model connection cone at its end; The central structural segment includes: A pin connector, with a gap reserved between one end and the connecting section of the model; The pin connector is provided with a pin shaft connection hole for inserting the pin shaft. The balance measuring element is symmetrically arranged on the upper and lower sides of the pin connector. The balance measuring element is connected to the model connecting section, and a gap is reserved between the balance measuring element and the pin connector. The eccentric shaft connecting section has a U-shaped structure, and symmetrical joint bearing grooves are provided on the inner side of the eccentric shaft connecting section. The joint bearing is slidably disposed in the joint bearing groove.

[0009] Preferably, the single-sided balance measuring element includes three measuring beams arranged in parallel in a "three" shape, with strain gauges attached to the measuring beams and the strain gauges connected by a Wheatstone bridge.

[0010] Preferably, the device further includes an angular displacement measurement and compensation module, which includes a position recording encoder mounted on the output shaft of the drive motor.

[0011] Preferably, the method for separating the time difference derivative of the washing flow using an aircraft variable rotation center excitation forced angle oscillation washing flow time difference derivative testing device includes the following steps: Step 1: Connect the aircraft model to the measuring balance, and install the aircraft variable rotation center excitation forced angle oscillation wash flow time difference derivative test device as a whole in the wind tunnel test section; Step 2: Adjust the rotation center of the aircraft model's oscillation to the first axial position by changing the length of the tail strut. x 1. Start the wind tunnel to the specified Mach number, and drive the motion conversion mechanism to apply a frequency of [frequency value missing] to the aircraft model. f Amplitude is i Forced pitch angle oscillation of 0, recording the time history of pitch moment coefficient. C m1 ( t )ω; Step 3: Adjust the rotation center of the aircraft model's oscillation to the second axial position by changing the length of the tail strut. x 2. Maintaining the same wind tunnel conditions, oscillation frequency, and amplitude, record the time history of the pitch moment coefficient. C m2 ( t ); Step 4: Use the angular displacement measurement and compensation module to perform inertial compensation on the recorded torque signal of the aircraft model, and deduct the inertial torque of the aircraft model itself; Step 5: Based on the principle of linear superposition of unsteady aerodynamic forces, establish the pitch moment coefficient and rotational rate derivative of the aircraft model at different rotation centers. and the derivative of the washing flow time difference A system of linear equations; Step 6: Solve the system of linear equations to obtain the pure time difference derivative of the washing flow. .

[0012] Preferably, in step five, the pitch moment coefficient and rotational angular rate derivative of the aircraft model at different rotation centers are established. and the derivative of the washing flow time difference Specific methods for solving linear equation systems include: Let the distance from the aircraft model's head be... x s The pin-shaped shaft undergoes small-amplitude forced pitch oscillations, with the following oscillation pattern: i ( t )= i 0sin( ωt ) in i 0 represents amplitude. oh It is the angular frequency; The reference center of an aircraft model, usually the center of gravity, is located at a distance from the head of the aircraft model. x cg At this location, the distance between the reference center and the pivot shaft is: Δx = x cg - x s When the pin shaft makes a pitching motion, the angle of attack at the reference center is... α ( t It includes not only the geometric angle of attack at the pivot point. i ( t This also includes the induced angle of attack caused by the velocity component perpendicular to the airflow from the reference center; under the assumption of small amplitude and low frequency, the frequency is reduced. k = oh / 2 V Instantaneous angle of attack at the reference center of the aircraft model α ( t )for: in, For geometric angle of attack i ( t The first derivative of ) Differentiating with respect to time yields the rate of change of angle of attack at the reference center. : For geometric angle of attack i ( t The second derivative of ). In the frequency domain, for a small-amplitude linear system, the pitching moment coefficient C m This can be expressed as the rate of change of angle of attack. and pitch rate q A linear combination, ignoring higher-order terms, pitching moment coefficient C m Represented as: in c For reference chord length, V The incoming flow velocity; Will Substitute the values ​​and consider simple harmonic motion. , After simplification, the in-phase and quadrature components of the total torque coefficient can be obtained, where the in-phase component of the total torque coefficient is... i ( t In phase ， Reflecting the stiffness term, the orthogonal component of the total moment coefficient and the angular velocity In phase, reflecting the damping term; focus on the orthogonal component of the total moment coefficient, i.e., the component related to the pitch rate. q The in-phase portion corresponds to the dynamic derivative; At a certain turning point x s The orthogonal components measured below The relationship with the true derivative is: make , Then we have: By changing the length of the tail boom, two different pivot positions can be selected. x s1 and x s2 , corresponding to Δ x 1 and Δ x 2. Measure the corresponding orthogonal components. and Establish a system of linear equations: Solving for the given information, we get: , Furthermore, the following can be separated: , That is, the derivative of the time difference of the washing flow is completed. and the derivative of the rotational angular rate C mq Decoupling.

[0013] This invention offers at least the following advantages: It addresses the problem in existing forced oscillation dynamic derivative testing techniques where, because the rotation center is fixed at the aircraft's torque reference center, the measured dynamic derivative is coupled with the derivative of the rotational angular rate and the derivative of the rate of change of angle of attack (washout time difference derivative), making it impossible to independently evaluate unsteady aerodynamic effects. This invention provides a test device for the washout time difference derivative of aircraft variable-rotation-excited forced oscillation, comprising a motion conversion mechanism, a measuring balance, an angular displacement measurement and compensation module, and a tail support rod. By changing the length of the tail support rod, the oscillation rotation center position of the aircraft model in a high-speed wind tunnel can be precisely adjusted, placing it before or after the torque reference center, thereby obtaining the pitch torque response under different rotation center conditions. This invention, combined with theoretical formula derivation, uses linear superposition and differential operations to separate the pure washout time difference derivative from two or more sets of variable-rotation-excited forced oscillation test data. The invention features an ingenious structural design and a clear separation principle, effectively improving the identification accuracy of the aircraft's dynamic aerodynamic derivatives and providing crucial data support for the study of unsteady aerodynamic characteristics and control law design of aircraft.

[0014] This invention transforms the originally coupled derivatives of the washing flow time difference and the rotation angular rate into a linear relationship by physically changing the rotation center, thus avoiding the errors caused by the traditional method's reliance on empirical formulas for estimation.

[0015] All tests in this invention use the same aircraft model, measuring balance, and dynamic data acquisition system, eliminating the influence of model manufacturing errors and measuring balance calibration errors in different test batches, and ensuring the internal consistency of the separation effect.

[0016] This invention is not only applicable to high-speed wind tunnels, but can also be applied, with modifications, to the study of unsteady aerodynamics / hydraulics in low-speed wind tunnels or water tunnels.

[0017] Other advantages, objectives and features of the present invention will become apparent in part from the following description, and in part from those skilled in the art through study and practice of the invention. Attached Figure Description

[0018] Figure 1 This is a schematic diagram of the overall structure of the aircraft variable rotation center excitation forced angle oscillation washing flow time difference derivative test device in Example 1; Figure 2 A schematic diagram of the structure of the test device for the derivative of the time difference of the oscillation of the forced angle of the aircraft's variable center of rotation after removing the tail support rod; Figure 3 This is a schematic diagram of the structure of a measuring balance; Figure 4 This is a schematic diagram of the "three centers in one" position for measuring the balance.

[0019] The corresponding labels for each structure in the diagram are as follows: Model connection section 1, Middle structure section 2, Eccentric shaft connection section 3, Pin connector 21, Point O 4, Model connection cone 5, Bolt hole 51, Pin shaft connection hole 6, Balance measuring element 7, Measuring beam 71, Joint bearing groove 8, Aircraft model 9, Cam 10, Joint bearing 11, Tail support rod 12, Center hole 121, Drive shaft 13, Drive motor 14, Pin shaft 15. Detailed Implementation

[0020] The present invention will now be described in further detail with reference to the accompanying drawings, so that those skilled in the art can implement it based on the description.

[0021] It should be understood that terms such as “having,” “comprising,” and “including” as used herein do not exclude the presence or addition of one or more other elements or combinations thereof. Example 1 like Figure 1-Figure 4 As shown, a device for testing the derivative of the time difference of an aircraft's variable rotation center excitation forced angle oscillation washing flow includes: A drive motor 14 is fixedly mounted at the end of a drive shaft 13. A measuring balance, one end of which is connected to the drive motor 14 via a motion conversion mechanism, and an aircraft model 9 is mounted on the other end of the measuring balance; Tail support rod 12 is sleeved on the outside of the drive shaft 13. A pin shaft 15 is provided on the measuring balance. The end of the tail support rod 12 is rotatably connected to the measuring balance through the pin shaft 15. By replacing the tail support rod 12 of different lengths, the rotation center of the aircraft model 9 can be changed, and the steering can be positioned at a preset axial position in front of and behind the torque reference center of the aircraft model 9.

[0022] The structure of the motion conversion mechanism includes: Cam 10 is slidably connected to the output shaft of the drive motor 14; The spherical bearing 11 has its inner ring slidably connected to the cam 10. When the drive motor 14 drives the output shaft to rotate, it simultaneously drives the cam 10 and the spherical bearing 11 to rotate circumferentially. Due to the action of the cam 10, the rotational motion of the drive motor 14 is converted into a forced pitch angle oscillation motion of the measuring balance and the aircraft model 9. The cam 10 and the spherical bearing 11 can slide axially along the axis of the output shaft of the drive motor 14, which facilitates the installation of the cam 10 and the spherical bearing 11 without affecting their circumferential rotation.

[0023] The measuring balance is a five-component strain balance. The structure of the measuring balance includes: Model connecting segment 1, with the middle structural segment 2 and the eccentric shaft connecting segment 3 connected in series at its rear end; The model connecting section 1 is provided with a model connecting cone 5 at its end. The model connecting cone 5 is used to install the aircraft model 9. The model connecting cone 5 is provided with a bolt hole 51 at its end. The aircraft model 9 is tightened and fixed to the model connecting cone 5 by screwing bolts into the bolt hole 51. The middle structural segment 2 includes: The pin connector 21 has a gap reserved between one end and the model connecting section 1; The pin connector 21 is provided with a pin shaft connection hole for inserting the pin shaft. The balance measuring element 7 is symmetrically arranged on the upper and lower sides of the pin connector 21. The balance measuring element 7 is connected to the model connecting section 1, and a gap is reserved between the balance measuring element 7 and the pin connector 21. The center of the balance measuring element 7 and the center of the pin shaft connecting hole 6 are set to the same point, which is designated as point O4. The center of the aircraft model is adjusted to coincide with point O4. In this way, the rotation support center, the center of the balance measuring element 7, and the center of the aircraft model are precisely coincident at the same spatial structural point O4, realizing the "three centers in one" structural design. The eccentric shaft connecting section 3 has a U-shaped structure. Symmetrical spherical bearing grooves 8 are provided on the inner side of the eccentric shaft connecting section 3. The spherical bearing 11 is slidably disposed in the spherical bearing groove 8, allowing it to slide laterally within the groove. The eccentric shaft connecting section 3 is used to connect the cam 10 via the spherical bearing 11. In actual operation, under the action of the drive motor 14, the spherical bearing 11 drives the measuring balance to reciprocate around the "three-center convergence" position O point 4 with a small amplitude. The single-sided balance measuring element 7 includes three measuring beams 71 arranged in parallel in a "three" shape. Strain gauges are attached to the measuring beams 71 and are connected by a Wheatstone bridge. The signal output terminals of the strain gauges are connected to a dynamic data acquisition system.

[0024] The tail support rod 12 is provided with a central hole 121 that matches the pin shaft connection hole 6. The pin shaft 15 is inserted into the central hole 121 and the pin shaft connection hole 6 to realize the rotational connection between the measuring balance and the tail support rod 12. The two sides of the pin shaft connection hole 6 are in sliding fit with the contact surfaces of the central hole 121 of the tail support rod 12, and the balance measuring element 7 does not come into contact with the tail support rod 12. The pin shaft connection hole 6 coincides with the geometric center of the balance measuring element 7.

[0025] The aircraft variable rotation center excitation forced angle oscillation wash flow time difference derivative test device also includes an angular displacement measurement and compensation module. The angular displacement measurement and compensation module includes a position recording encoder, which is installed on the output shaft of the drive motor 14 and is used to measure the angular displacement of the output cam 10 in real time and record the position of the cam 10. The angular displacement measurement and compensation module collects the elastic deformation data of the tail support rod 12 and converts it for real-time recording of the oscillation angle.

[0026] The working principle of the aircraft variable rotation center excitation forced angle oscillation wash flow time difference derivative testing device in this embodiment is as follows: Connect the aircraft model to the measuring balance, and install the entire aircraft variable rotation center excitation forced angle oscillation wash flow time difference derivative testing device in the wind tunnel test section; adjust the rotation center of the aircraft model oscillation to the first axial position through the tail support rod, start the wind tunnel to the specified Mach number, drive the motion conversion mechanism to apply forced pitch angle oscillation to the aircraft model, and record the pitch moment coefficient time history; adjust the rotation center of the aircraft model oscillation to the second axial position through the tail support rod, maintain the same wind tunnel state and oscillation frequency and amplitude, and record the pitch moment coefficient time history; use the angular displacement measurement and compensation module to perform inertial compensation on the recorded aircraft model torque signal, and deduct the aircraft model's own inertial torque; based on the principle of linear superposition of unsteady aerodynamic forces, establish a linear equation system of the aircraft model pitch moment coefficient, rotational angular rate derivative, and wash flow time difference derivative at different rotation centers, solve the linear equation system, and separate the pure wash flow time difference derivative.

[0027] Example 2 A method for separating the derivative of the time difference of an aircraft's variable rotation center excitation forced angle oscillation wash flow, based on the aircraft variable rotation center excitation forced angle oscillation wash flow time difference derivative testing device in Example 1, includes the following steps: Step 1: Connect the aircraft model to the measuring balance, and install the aircraft variable rotation center excitation forced angle oscillation wash flow time difference derivative test device as a whole in the wind tunnel test section; Step 2: Adjust the rotation center of the aircraft model's oscillation to the first axial position by changing the length of the tail strut. x 1 (e.g., at 0.5 times the chord length in front of the reference center), start the wind tunnel to the specified Mach number, and drive the motion conversion mechanism to apply a frequency of 1 to the aircraft model. f Amplitude is i Forced pitch angle oscillations of 0 (≤±1°) were recorded, and the time history of the pitch moment coefficient was recorded. C m1 ( t )ω; Step 3: Adjust the rotation center of the aircraft model's oscillation to the second axial position by changing the length of the tail strut. x 2. Maintaining the same wind tunnel conditions, oscillation frequency, and amplitude, record the time history of the pitch moment coefficient. C m2 ( t ); Step 4: Use the angular displacement measurement and compensation module to perform inertial compensation on the recorded torque signal of the aircraft model, and deduct the inertial torque of the aircraft model itself; Step 5: Based on the principle of linear superposition of unsteady aerodynamic forces, establish the pitch moment coefficient and rotational rate derivative of the aircraft model at different rotation centers. and the derivative of the washing flow time difference The linear equation system can be solved using the following methods: Let the distance from the aircraft model's head be... x s The pin-shaped shaft undergoes small-amplitude forced pitch oscillations, with the following oscillation pattern: i ( t )= i 0sin( ωt ) in i 0 represents amplitude. oh It is the angular frequency; The reference center of an aircraft model, usually the center of gravity, is located at a distance from the head of the aircraft model. x cg At this location, the distance between the reference center and the pivot shaft is: Δ x = x cg - x s When the pin shaft makes a pitching motion, the angle of attack at the reference center is... α ( t It includes not only the geometric angle of attack at the pivot point. i ( t This also includes the induced angle of attack caused by the velocity component perpendicular to the airflow from the reference center; under the assumption of small amplitude and low frequency, the frequency is reduced. k = oh / 2 V Instantaneous angle of attack at the reference center of the aircraft model α ( t )for: in, For geometric angle of attack i ( t The first derivative of ) Differentiating with respect to time yields the rate of change of angle of attack at the reference center. : For geometric angle of attack i ( t The second derivative of ). In the frequency domain, for a small-amplitude linear system, the pitching moment coefficient C m This can be expressed as the rate of change of angle of attack. and pitch rate q A linear combination, ignoring higher-order terms, pitching moment coefficient C m Represented as: in c For reference chord length, V The incoming flow velocity; Will Substitute the values ​​and consider simple harmonic motion. , After simplification, the in-phase and quadrature components of the total torque coefficient can be obtained, where the in-phase component of the total torque coefficient is... i ( t In phase ， Reflecting the stiffness term, the orthogonal component of the total moment coefficient and the angular velocity In phase, reflecting the damping term; focus on the orthogonal component of the total moment coefficient, i.e., the component related to the pitch rate. q The in-phase portion corresponds to the dynamic derivative; At a certain turning point x s The orthogonal components measured below The relationship with the true derivative is: make , Then we have: By changing the length of the tail boom, two different pivot positions can be selected. x s1 and x s2 , corresponding to Δ x 1 and Δ x 2. Measure the corresponding orthogonal components. and Establish a system of linear equations: Solving for the given information, we get: , Furthermore, the following can be separated: , That is, the derivative of the time difference of the washing flow is completed. and the derivative of the rotational angular rate C mq Decoupling; Step 6: Solve the system of linear equations to obtain the pure time difference derivative of the washing flow. .

[0028] The number of devices and processing scale described herein are for the purpose of simplifying the description of the invention. Applications, modifications, and variations of the invention will be readily apparent to those skilled in the art.

[0029] Although embodiments of the present invention have been disclosed above, they are not limited to the applications listed in the specification and embodiments. They can be applied to various fields suitable for the present invention. For those skilled in the art, other modifications can be easily made. 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 illustrations shown and described herein.

Claims

1. A device for testing the derivative of the time difference of oscillation washing current under varying rotation center excitation forced angle of an aircraft, characterized in that, include: A drive shaft, with a drive motor fixedly mounted at its end; A measuring balance, one end of which is connected to the drive motor via a motion conversion mechanism, and an aircraft model is mounted on the other end of the measuring balance; A tail support rod is sleeved on the outside of the drive shaft. A pin shaft is provided on the measuring balance. The end of the tail support rod is rotatably connected to the measuring balance through the pin shaft. The method for separating the derivative of the washing flow time difference using a device for measuring the derivative of the washing flow time difference of an aircraft with a variable rotation center of gravity-excited forced angle oscillation includes the following steps: Step 1: Connect the aircraft model to the measuring balance, and install the aircraft variable rotation center excitation forced angle oscillation wash flow time difference derivative test device as a whole in the wind tunnel test section; Step 2: Adjust the rotation center of the aircraft model's oscillation to the first axial position by changing the length of the tail strut. x 1. Start the wind tunnel to the specified Mach number, and drive the motion conversion mechanism to apply a frequency of [frequency value missing] to the aircraft model. f Amplitude is θ Forced pitch angle oscillation of 0, recording the time history of pitch moment coefficient. C m1 ( t ); Step 3: Adjust the rotation center of the aircraft model's oscillation to the second axial position by changing the length of the tail strut. x 2. Maintaining the same wind tunnel conditions, oscillation frequency, and amplitude, record the time history of the pitch moment coefficient. C m2 ( t ); Step 4: Use the angular displacement measurement and compensation module to perform inertial compensation on the recorded torque signal of the aircraft model, and deduct the inertial torque of the aircraft model itself; Step 5: Based on the principle of linear superposition of unsteady aerodynamic forces, establish the pitch moment coefficient and rotational rate derivative of the aircraft model at different rotation centers. C mq and the derivative of the washing flow time difference A system of linear equations; Step 6: Solve the system of linear equations to obtain the pure time difference derivative of the washing flow. ; In step five, the pitch moment coefficient and rotational angular rate derivative of the aircraft model at different rotation centers are established. C mq and the derivative of the washing flow time difference Specific methods for solving linear equation systems include: Let the distance from the aircraft model's head be... x s The pin-shaped shaft undergoes small-amplitude forced pitch oscillations, with the following oscillation pattern: θ ( t )= θ 0sin( ωt ) in θ 0 represents amplitude. ω It is the angular frequency; The reference center of an aircraft model, usually the center of gravity, is located at a distance from the head of the aircraft model. x cg At this location, the distance between the reference center and the pivot shaft is: D x = x cg - x s When the pin shaft makes a pitching motion, the angle of attack at the reference center is... α ( t It includes not only the geometric angle of attack at the pivot point. θ ( t This also includes the induced angle of attack caused by the velocity component perpendicular to the airflow from the reference center; under the assumption of small amplitude and low frequency, the frequency is reduced. k = ωc / 2 V ,in c For reference chord length, V The instantaneous angle of attack at the reference center of the aircraft model is the incoming flow velocity. α ( t )for: in, For geometric angle of attack θ ( t The first derivative of ) Differentiating with respect to time yields the rate of change of angle of attack at the reference center. : For geometric angle of attack θ ( t The second derivative of ). In the frequency domain, for a small-amplitude linear system, the pitching moment coefficient C m This can be expressed as the rate of change of angle of attack. and pitch rate q A linear combination, ignoring higher-order terms, pitching moment coefficient C m Represented as: Will Substitute the values ​​and consider simple harmonic motion. , After simplification, the in-phase and quadrature components of the total torque coefficient can be obtained, where the in-phase component of the total torque coefficient is... θ ( t In phase ， Reflecting the stiffness term, the orthogonal component of the total moment coefficient and the angular velocity In phase, reflecting the damping term; focus on the orthogonal component of the total moment coefficient, i.e., the component related to the pitch rate. q The in-phase portion corresponds to the dynamic derivative; At a certain turning point x s The orthogonal components measured below The relationship with the true derivative is: make , Then we have: By changing the length of the tail boom, two different pivot positions can be selected. x s1 and x s2 , corresponding to Δ x 1 and Δ x 2. Measure the corresponding orthogonal components. and Establish a system of linear equations: Solving for the given information, we get: Furthermore, the following can be separated: That is, the derivative of the time difference of the washing flow is completed. and the derivative of the rotational angular rate C mq Decoupling.

2. The device for testing the derivative of the time difference of oscillation washing flow under the excitation of the variable center of rotation of an aircraft as described in claim 1, characterized in that, The structure of the motion conversion mechanism includes: A cam, which is slidably connected to the output shaft of the drive motor; A spherical bearing, the inner ring of which is slidably connected to the cam.

3. The device for testing the derivative of the time difference of oscillation washing flow under the excitation of the variable center of rotation of an aircraft as described in claim 1, characterized in that, The measuring balance is a five-component strain balance. Strain gauges are attached to the measuring beam of the measuring balance. The strain gauges are connected by a Wheatstone bridge. The signal output terminals of the strain gauges are connected to a dynamic data acquisition system.

4. The device for testing the derivative of the time difference of oscillation washing flow under the excitation of the variable center of rotation of an aircraft as described in claim 2, characterized in that, The structure of the measuring balance includes: The model connection segment is connected in series with the middle structural segment and the eccentric shaft connection segment at its rear end; The model connection segment is provided with a model connection cone at its end; The central structural segment includes: A pin connector, with a gap reserved between one end and the connecting section of the model; The pin connector is provided with a pin shaft connection hole for inserting the pin shaft. The balance measuring element is symmetrically arranged on the upper and lower sides of the pin connector. The balance measuring element is connected to the model connecting section, and a gap is reserved between the balance measuring element and the pin connector. The eccentric shaft connecting section has a U-shaped structure, and symmetrical joint bearing grooves are provided on the inner side of the eccentric shaft connecting section. The joint bearing is slidably disposed in the joint bearing groove.

5. The device for testing the derivative of the time difference of oscillation washing flow under the excitation of the variable center of rotation of an aircraft as described in claim 2, characterized in that, The single-sided balance measuring element includes three measuring beams arranged in parallel in a "three" shape. Strain gauges are attached to the measuring beams and the strain gauges are connected by a Wheatstone bridge.

6. The device for testing the derivative of the time difference of oscillation washing flow under the excitation of the variable center of rotation of an aircraft as described in claim 1, characterized in that, It also includes an angular displacement measurement and compensation module, which includes a position recording encoder mounted on the output shaft of the drive motor.

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