JP2000046977A - Vibration-suppressing device - Google Patents

Abstract

PROBLEM TO BE SOLVED: To suppress the structural resonance between the natural frequency in the rotating direction of a drive shaft and an aerodynamic vibration frequency applied to optical equipment by fully separating them by changing an inertial coefficient in the rotary direction of the drive shaft of the optical equipment. SOLUTION: The level of the natural frequency in the rotating direction of a drive shaft 4 is changed by adjusting the gain of a servo mechanism 7. However, during normal flight, the gain of the servo mechanism 7 is set to 0, and a controller 8 gives a gain change command to the servo mechanism 7 via a signal line 7 when the aerodynamic vibration direction being applied to optical equipment 2 approaches the frequency in the rotating direction of the drive shaft 4 at the gain 0 of the servo mechanism 7, thus fully separating the natural frequency in the rotating direction of the drive shaft 4 from the frequency of aerodynamic vibration being applied to the equipment 2 and hence avoiding the structural resonance between them, eliminating the peak value of a gain at the resonance frequency in the frequency characteristics of the control system of the drive motor 4 of the optical equipment 2, and preventing the oscillation of a signal near the resonance frequency.

Description

DETAILED DESCRIPTION OF THE INVENTION

[0001]

BACKGROUND OF THE INVENTION 1. Field of the Invention The present invention relates to a resonance between a structural natural frequency in a rotation direction of a drive shaft for changing an optical field of view and an aerodynamic vibration frequency applied to an optical device in an optical device mounted on an aircraft. More specifically, when the natural frequency in the rotation direction of the drive shaft of the optical device approaches the frequency of the aerodynamic vibration applied to the optical device, the natural frequency in the rotation direction of the drive shaft is reduced. Alternatively, a device that changes the damping ratio and suppresses resonance between the two is proposed.

[0002]

2. Description of the Related Art FIG. 21 is an external view of a conventional optical device mounted on an aircraft. 1 is a part of an aircraft, 2 is an optical device mounted on the aircraft, 3 is an optical sensor provided inside the optical device 2, 4 is a drive shaft for changing the optical field of view of the optical sensor 3, 5 is a drive shaft Reference numeral 4 denotes a drive motor for controlling the rotational movement, wherein A is an air flow around the optical device 2 and A is a Karman vortex pair formed downstream of the separation point of the air flow A. FIG. 22 is a graph showing a frequency characteristic of a control system of the drive motor 5 of the conventional optical device 2 mounted on an aircraft, wherein the horizontal axis represents the frequency and the vertical axis represents the gain. C is the frequency characteristic of the control system of the drive motor 5 when resonance does not occur between the natural frequency of the rotation direction of the drive shaft 4 and the frequency of the aerodynamic vibration applied to the optical device 2, and d is the rotation of the drive shaft 4. Frequency characteristics of the control system of the drive motor 5 when resonance occurs between the natural frequency in the direction and the frequency of the aerodynamic vibration applied to the optical device 2; The resonance frequency with respect to the frequency of the aerodynamic vibration according to 2 and f is the peak value of the gain at the resonance frequency e. In FIG. 21, the equation of motion of the drive shaft 4 in the rotational direction is represented as in Equation 1.

[0003]

(Equation 1)

When the equation of motion in the rotational direction of the drive shaft 4 is expressed as in Equation 1, the natural frequency in the rotational direction of the drive shaft 4 is expressed by Equation 2.
Can be expressed as

[0005]

(Equation 2)

Further, the number of occurrences of the Karman vortex pair a per unit time is n
When, acts alternating drag number 3 of such frequencies f _A is the drag direction of the optical device 2, the mounting of it and the drive shaft 4 which can not be produced on the cut surface is a perfect circle cross section for air flow of the optical apparatus 2 Since the position cannot be set on the complete central axis of the optical device 2 or the like, the alternating drag is applied not only in the drag direction but also in the rotational direction of the drive shaft 4 as aerodynamic disturbance vibration.

[0007]

(Equation 3)

The optical device 2 mounted on an aircraft rotates the drive shaft 4 by a drive motor 5 when the target is captured and tracked, thereby changing the optical field of view of the optical sensor 3 to change the position, angle or distance from the target. The spatial stability of the drive shaft 4 that changes the optical field of view of the optical sensor 3 is an important technical issue in order to measure the natural frequency. However, the natural frequency in the rotational direction of the drive shaft 4 shown in Expression 2 is shown in Expression 3. When the frequency is close to the frequency of the aerodynamic vibration applied to the optical device 2, there is a problem that a structural resonance phenomenon occurs between the two and the drive shaft 4 vibrates.

[0009]

In the optical equipment 2 mounted on an aircraft, the natural frequency of the rotation direction of the drive shaft 4 of the optical equipment 2 is close to the frequency of the aerodynamic vibration applied to the optical equipment 2 and between them. When a structural resonance phenomenon occurs in the control system of the rotation direction of the drive shaft 4 by the drive motor 5,
As shown in FIG. 22, a gain peaks around a specific resonance frequency E, and a signal in a frequency range around the resonance frequency O that should be attenuated oscillates. When the drive shaft 4 oscillates in the rotation direction, the drive motor 5 oscillates following the oscillation of the drive shaft 4 in the rotation direction, so that problems such as wasteful power consumption in the drive motor 5 and heat generation due to excessive power consumption occur. other,
Oscillation of the drive shaft 4 in the rotational direction makes it impossible to accurately measure the position, angle, or distance to the target by the optical sensor 3, and thus has a problem such as adversely affecting the target supplementary tracking control performance of the optical device 2. The present invention relates to an optical device 2 mounted on an airplane, and a natural frequency of the drive shaft 4 in the rotational direction and the optical device 2
It is an object of the present invention to suppress structural resonance with the frequency of the aerodynamic vibration.

[0010]

According to a first aspect of the present invention, there is provided a vibration suppression device provided in an optical device mounted on an aircraft, provided inside the optical device, and provided on a drive shaft for changing an optical field of view and on the drive shaft. A drive motor that controls the rotational movement of the drive shaft, a balancer whose distance from the drive shaft changes, and a change in the inertia coefficient in the rotational direction of the drive shaft by changing the distance of the balancer from the drive shaft. A servo mechanism that changes the natural frequency in the rotation direction, and the aerodynamic vibration applied to the optical device that is provided inside the optical device and whose natural frequency in the rotation direction of the drive shaft is estimated and calculated based on speed information from the aircraft And a controller for adjusting the gain of the servo mechanism so as not to approach the frequency.

Further, a vibration suppressing device according to a second aspect of the present invention includes:
In an optical device mounted on an aircraft, a drive shaft provided inside the optical device and changing an optical field of view, a drive motor provided on the drive shaft and controlling a rotational motion of the drive shaft, and a rotation direction of the drive shaft A servo mechanism that changes the elastic modulus of the drive shaft in the rotation direction by changing the expansion and contraction length of the spring and changes the natural frequency of the drive shaft in the rotation direction; A controller that adjusts the gain of the servo mechanism so that the natural frequency in the rotational direction of the drive shaft does not approach the frequency of the aerodynamic vibration applied to the optical device estimated and calculated based on the speed information from the aircraft. is there.

Further, the vibration suppressing device according to the third aspect of the present invention comprises:
In an optical device mounted on an aircraft, a drive shaft provided inside the optical device and changing an optical field of view, a drive motor provided on the drive shaft and controlling a rotational motion of the drive shaft, and a distance from the drive shaft A servo mechanism that changes the inertia coefficient in the rotation direction of the drive shaft by changing the distance of the balancer from the drive shaft, and changes the natural frequency around the drive shaft, and torque in the rotation direction of the drive shaft. A spring to be loaded, a servo mechanism that changes the elastic modulus of the drive shaft in the rotation direction by changing the expansion and contraction length of the spring, and changes the natural frequency of the drive shaft in the rotation direction, and a servo mechanism provided inside the optical device, A controller that adjusts the gain of each servo mechanism so that the natural frequency in the rotational direction does not approach the frequency of the aerodynamic vibration applied to the optical equipment estimated and calculated based on the speed information from the aircraft Those were example.

Further, a vibration suppressing device according to a fourth aspect of the present invention includes:
In an optical device mounted on an aircraft, a drive shaft provided inside the optical device and changing an optical field of view, a drive motor provided on the drive shaft and controlling a rotational motion of the drive shaft, and a rotation direction of the drive shaft A damper that applies damping force to the
A servo mechanism that adjusts the damping force of the damper to change the damping coefficient in the rotation direction of the drive shaft to change the damping ratio in the rotation direction of the drive shaft; A controller that adjusts the gain of the servo mechanism and increases the damping ratio in the rotational direction of the drive shaft when the frequency approaches the frequency of the aerodynamic vibration applied to the optical device estimated and calculated based on the speed information from the aircraft Things.

Further, a vibration suppressing device according to a fifth aspect of the present invention includes:
In an optical device mounted on an aircraft, a drive shaft provided inside the optical device and changing an optical field of view, a drive motor provided on the drive shaft and controlling a rotational motion of the drive shaft, and a distance from the drive shaft A servo mechanism that changes the inertia coefficient in the rotation direction of the drive shaft by changing the distance from the drive shaft of the balancer to change the natural frequency in the rotation direction of the drive shaft. A spring that applies torque, a servo mechanism that changes the elastic modulus of the drive shaft in the rotation direction by changing the expansion and contraction length of the spring, and a damping force in the rotation direction of the drive shaft that changes the natural frequency of the drive shaft in the rotation direction A servo mechanism that changes the damping force of the drive shaft in the rotation direction by adjusting the damping force of the damper to change the damping ratio of the drive shaft in the rotation direction, When the natural frequency of the rolling direction approaches the frequency of the aerodynamic vibration applied to the optical device estimated and calculated based on the speed information from the aircraft, the gain of each servo mechanism is adjusted and the damping ratio of the rotation direction of the drive shaft is adjusted. And a controller for increasing.

Further, a vibration suppressing device according to a sixth aspect of the present invention includes:
In an optical device mounted on an aircraft, a drive shaft provided inside the optical device and changing an optical field of view, a drive motor provided on the drive shaft and controlling a rotational motion of the drive shaft, and a distance from the drive shaft And a servo mechanism that changes the inertia coefficient in the rotation direction of the drive shaft by changing the distance from the drive shaft of the balancer to change the natural frequency in the rotation direction of the drive shaft, and the outer wall of the optical device. A vibration sensor that detects the frequency of aerodynamic vibration applied to the optical device; and a vibration sensor that is provided inside the optical device and detects a natural frequency in the rotation direction of the drive shaft detected by the vibration sensor. And a controller for adjusting the gain of the servo mechanism so as not to approach the frequency.

Further, a vibration suppressing device according to a seventh aspect of the present invention includes:
In an optical device mounted on an aircraft, a drive shaft provided inside the optical device and changing an optical field of view, a drive motor provided on the drive shaft and controlling a rotational motion of the drive shaft, and a rotation direction of the drive shaft A servo mechanism that changes the elastic modulus of the drive shaft in the rotation direction by changing the expansion and contraction length of the spring, and changes the natural frequency of the drive shaft in the rotation direction. A vibration sensor that detects the frequency of aerodynamic vibration applied to the optical device; and a frequency of aerodynamic vibration applied to the optical device that is provided inside the optical device and whose natural frequency in the rotation direction of the drive shaft is detected by the vibration sensor. And a controller that adjusts the gain of the servo mechanism so as not to approach the controller.

Further, the vibration suppressing device according to the eighth aspect of the present invention comprises:
In an optical device mounted on an aircraft, a drive shaft provided inside the optical device and changing an optical field of view, a drive motor provided on the drive shaft and controlling a rotational motion of the drive shaft, and a distance from the drive shaft A servo mechanism that changes the inertia coefficient in the rotation direction of the drive shaft by changing the distance of the balancer from the drive shaft, and changes the natural frequency around the drive shaft, and torque in the rotation direction of the drive shaft. A spring to be loaded, a servo mechanism that changes the elastic modulus in the rotation direction of the drive shaft by changing the expansion and contraction length of the spring and changes the natural frequency in the rotation direction of the drive shaft, and is provided on the outer wall surface of the optical device,
A vibration sensor that detects the frequency of aerodynamic vibration applied to the optical device, and a natural frequency in the rotation direction of the drive shaft that is provided inside the optical device and is close to the frequency of the aerodynamic vibration applied to the optical device detected by the vibration sensor And a controller that adjusts the gain of each servo mechanism so as not to cause the problem.

Further, a vibration suppressing device according to a ninth aspect of the present invention comprises:
In an optical device mounted on an aircraft, a drive shaft provided inside the optical device and changing an optical field of view, a drive motor provided on the drive shaft and controlling a rotational motion of the drive shaft, and a rotation direction of the drive shaft A damper that applies damping force to the
A servo mechanism that changes the damping force of the drive shaft in the rotational direction by adjusting the damping force of the damper to change the damping ratio of the drive shaft in the rotational direction, and aerodynamic vibration that is provided on the outer wall of the optical device and acts on the optical device And a gain of the servo mechanism provided inside the optical device when the natural frequency in the rotation direction of the drive shaft is close to the frequency of the aerodynamic vibration applied to the optical device detected by the vibration sensor. A controller for performing the adjustment and increasing the damping ratio in the rotational direction of the drive shaft.

A vibration suppression device according to a tenth aspect of the present invention is an optical device mounted on an aircraft, wherein the drive shaft is provided inside the optical device and changes an optical field of view.
A drive motor provided on the drive shaft to control the rotational movement of the drive shaft, a balancer whose distance from the drive shaft changes, and a change in the inertia coefficient of the drive shaft in the rotational direction by changing the distance of the balancer from the drive shaft A servo mechanism that changes the natural frequency of the drive shaft in the rotation direction, a spring that applies torque in the rotation direction of the drive shaft, and a spring that changes the elasticity of the drive shaft in the rotation direction by changing the expansion and contraction length of the spring. A servo mechanism that changes the natural frequency in the rotation direction of the shaft, a damper that applies a damping force in the rotation direction of the drive shaft, and a drive shaft that changes the damping coefficient in the rotation direction of the drive shaft by adjusting the damping force of the damper A servo mechanism for changing the damping ratio in the rotation direction of the optical device, a vibration sensor provided on the outer wall surface of the optical device for detecting the frequency of aerodynamic vibration applied to the optical device, and a rotation direction of the drive shaft provided inside the optical device. of A controller for adjusting the gain of each servo mechanism and increasing the damping ratio in the rotational direction of the drive shaft when the frequency of the vibration approaches the frequency of the aerodynamic vibration applied to the optical device detected by the vibration sensor. is there.

[0020]

DESCRIPTION OF THE PREFERRED EMBODIMENTS Embodiment 1 FIG. 1 is a configuration diagram showing Embodiment 1 of the present invention, in which 1 is a part of an aircraft, 2 is an optical device mounted on the aircraft, 3 is an optical sensor provided inside the optical device 2, 4 is a drive shaft for changing the optical field of view of the optical sensor 3, 5 is a drive motor for controlling the rotational movement of the drive shaft 4, 6 is a balancer, 7 is a servo mechanism for changing the distance of the balancer 6 from the drive shaft 4, Reference numeral 8 denotes a controller for adjusting the gain of the servo mechanism 7; A is an airflow around the optical device 2; A is a Karman vortex pair formed downstream of the separation point of the airflow A; G is a signal line of speed information from an aircraft; Is the servo mechanism 7 from the controller 8
To the command signal line. FIG. 2 shows Embodiment 1 of the present invention.
1 is a part of the same aircraft as in FIG. 1, 2 is the same optical device as in FIG. 1, 3 is the same optical sensor as in FIG. 1, 4 is the same drive shaft as in FIG. 1, 5 is 1 is the same drive motor as in FIG. 1, 6 is the same balancer as in FIG. 1, 7 is the same servo mechanism as in FIG. 1, 8 is the same controller as in FIG. 1, and key is a signal line for speed information from the same aircraft as in FIG. Is a command signal line from the controller 8 to the servo mechanism 7 as in FIG.

In FIG. 1, the distance of the balancer 6 provided on the drive shaft 4 from the drive shaft 4 is one of the parameters for determining the inertia coefficient of the drive shaft 4 in the rotation direction. By changing the distance from the drive shaft 4 as described above, in the equation of motion of the optical device 2 in the rotational direction of the drive shaft 4, the inertia term changes as shown in Expression 4.

[0022]

(Equation 4)

For this reason, the inertia coefficient of the drive shaft 4 in the rotational direction changes as shown in Expression 5.

[0024]

(Equation 5)

Therefore, the natural frequency f _{S1 in} the rotation direction of the drive shaft 4 changes as shown in Equation 6.

[0026]

(Equation 6)

The magnitude of the natural frequency f _{S1 in} the rotation direction of the drive shaft 4 is changed by adjusting the gain of the servo mechanism 7, but during normal flight, the gain of the servo mechanism 7 is set to 0 and applied to the optical device 2. The frequency f _A of the aerodynamic vibration is the frequency f in the rotational direction of the drive shaft 4 at a gain of 0 of the servo mechanism 7 shown in Expression 2.
When approaching _S , the controller 8 gives a command to change the gain to the servo mechanism 7 via the signal line C, and the drive shaft shown in Expression 12 with respect to the frequency f _{A of the} aerodynamic vibration applied to the optical device 2. The natural frequency f _S1 in the rotational direction of No. 4 is sufficiently separated. The frequency f _{A of the} aerodynamic vibration applied to the optical device 2
Is estimated and calculated by the controller 8 using the speed information of the aircraft obtained from the signal line G and the arithmetic expression of Expression 7.

[0028]

(Equation 7)

The frequency f of the aerodynamic vibration applied to the optical device 2
_{By making} the natural frequency f _S1 in the rotational direction of the drive shaft 4 far from A, structural resonance between the two can be avoided, and in the frequency characteristic of the control system of the drive motor 4 of the optical device 2, The peak value of the gain at the resonance frequency e shown in FIG. 22 disappears, and the signal oscillation around the resonance frequency e disappears.

Embodiment 2 FIG. FIG. 3 is a configuration diagram showing Embodiment 2 of the present invention, in which 1 is a part of an aircraft, 2 is an optical device mounted on the aircraft, 3 is an optical sensor provided inside the optical device 2, 4 is a drive shaft for changing the optical field of view of the optical sensor 3, 5 is a drive motor for controlling the rotational movement of the drive shaft 4, 9 is a spring, 7 is a servo mechanism for changing the expansion and contraction length of the spring 9, 8 is a servo mechanism. 7 is a controller for adjusting the gain of 7; A is an airflow around the optical device 2; A is a Karman vortex pair formed downstream of the separation point of the airflow A; G is a signal line of speed information from the aircraft; This is a command signal line to the servo mechanism 7. FIG. 4 is a partial configuration diagram showing Embodiment 2 of the present invention. In FIG. 4, 1 is a part of the same aircraft as FIG. 3, 2 is the same optical device as FIG. 3, 3 is the same optical sensor as FIG. 3 is the same drive shaft as in FIG. 3, 5 is the same drive motor as in FIG. 3, 9 is the same spring as in FIG. 3, 7 is the same servo mechanism as in FIG. 3, 8 is the same controller as in FIG. A signal line for speed information from the aircraft is a command signal line from the controller 8 to the servo mechanism 7 as in FIG.

In FIG. 3, the expansion and contraction length of the spring 9 provided on the drive shaft 4 is one of the parameters for determining the elastic modulus of the drive shaft 4 in the rotation direction. Is changed, in the equation of motion of the rotation direction of the drive shaft 4 of the optical device 2, the elastic term changes as shown in Expression 8.

[0032]

(Equation 8)

Therefore, the elastic modulus of the drive shaft 4 in the rotational direction changes as shown in the following equation (9).

[0034]

(Equation 9)

Therefore, the natural frequency f _{S2 in} the rotational direction of the drive shaft 4 changes as shown in Expression 10.

[0036]

(Equation 10)

The magnitude of the natural frequency f _{S2 in} the rotation direction of the drive shaft 4 is changed by adjusting the gain of the servo mechanism 7, but during normal flight, the gain of the servo mechanism 7 is set to 0 and applied to the optical device 2. The frequency f _A of the aerodynamic vibration is the frequency f in the rotational direction of the drive shaft 4 at a gain of 0 of the servo mechanism 7 shown in Expression 2.
When approaching _S , the controller 8 gives a command to change the gain to the servo mechanism 7 via the signal line C, and the drive shaft shown in Expression 12 with respect to the frequency f _{A of the} aerodynamic vibration applied to the optical device 2. The natural frequency f _S2 in the rotational direction of No. 4 is sufficiently separated. The frequency f _{A of the} aerodynamic vibration applied to the optical device 2
Is estimated and calculated by the controller 8 using the speed information of the aircraft obtained from the signal line G and the arithmetic expression of Expression 7. By causing the natural frequency f _S2 in the rotational direction of the drive shaft 4 to be sufficiently separated from the frequency f _A of the aerodynamic vibration applied to the optical device 2, structural resonance between the two is avoided, and the driving of the optical device 2 is performed. In the frequency characteristic of the control system of the motor 4, the peak value of the gain at the resonance frequency e shown in FIG. 22 disappears, and the signal oscillation near the resonance frequency e disappears.

Embodiment 3 FIG. 5 is a configuration diagram showing Embodiment 3 of the present invention, in which 1 is a part of an aircraft, 2 is an optical device mounted on the aircraft, 3 is an optical sensor provided inside the optical device 2, 4 is a drive shaft for changing the optical field of view of the optical sensor 3, 5 is a drive motor for controlling the rotational movement of the drive shaft 4, 6 is a balancer, 9 is a spring, 7
Is a servo mechanism for changing the distance of the balancer 6 from the drive shaft 4 and the length of expansion and contraction of the spring 9; 8 is a controller for adjusting the gain of the servo mechanism 7; Karman vortex pairs that can be created downstream of the separation point,
G is a signal line of speed information from the aircraft, and K is a command signal line from the controller 8 to the servo mechanism 7. FIG. 6 is a partial configuration diagram showing Embodiment 3 of the present invention, in which 1 is a part of the same aircraft as that of FIG. 5, 2 is the same optical device as that of FIG. 5, 3 is the same optical sensor as that of FIG. 5 is the same drive shaft as in FIG. 5, 5 is the same drive motor as in FIG. 5, 6 is the same balancer as in FIG. 5, 9 is the same spring as in FIG. 5, 7 is the same servo mechanism as in FIG.
Reference numeral 8 denotes the same controller as that of FIG. 5, and reference character K denotes a signal line of speed information from the same aircraft as that of FIG. 5, and reference character K denotes a command signal line from the controller 8 to the servo mechanism 7 same as that of FIG.

In FIG. 5, the distance of the balancer 6 provided on the drive shaft 4 from the drive shaft 4 is one of the parameters for determining the inertia coefficient of the drive shaft 4 in the rotational direction. Since the length of expansion and contraction of the spring 9 is one of the parameters that determine the elastic modulus of the drive shaft 4 in the rotational direction, the distance between the balancer 6 and the drive shaft 4 is changed as shown in FIG. By changing the expansion and contraction length as shown in 6, in the equation of motion of the rotation direction of the drive shaft 4 of the optical device 2, the inertia term and the elasticity term are changed as shown in Expression 11.

[0040]

[Equation 11]

Therefore, the inertia coefficient in the rotation direction of the drive shaft 4 changes as shown in Equation 5, and the elastic coefficient changes as shown in Equation 9. Therefore, the natural frequency f _{S3 in} the rotation direction of the drive shaft 6 changes as shown in Expression 12.

[0042]

(Equation 12)

The magnitude of the natural frequency f _{S3 in} the rotation direction of the drive shaft 4 is changed by adjusting the gain of the servo mechanism 7, but during normal flight, the gain of the servo mechanism 7 is set to 0 and applied to the optical device 2. The frequency f _S of the rotational direction of the drive shaft 4 at the gain of the servo mechanism 7 where the frequency f _A of the aerodynamic vibration is expressed by the equation (2)
The controller 8 gives a command to change the gain to the servo mechanism 7 via the signal line C when approaching to the drive shaft 4 so that the frequency f _A of the aerodynamic vibration applied to the optical device 2 is equal to The natural frequency f _S3 in the rotation direction is sufficiently separated. The frequency f _{A of the} aerodynamic vibration applied to the optical device 2
Is estimated and calculated by the controller 8 using the speed information of the aircraft obtained from the signal line G and the arithmetic expression of Expression 7. By causing the natural frequency f _S3 in the rotational direction of the drive shaft 4 to be sufficiently separated from the frequency f _A of the aerodynamic vibration applied to the optical device 2, structural resonance between the two is avoided, and the drive of the optical device 2 is performed. In the frequency characteristic of the control system of the motor 5, the peak value of the gain at the resonance frequency e shown in FIG. 22 disappears, and the oscillation of the signal near the resonance frequency e disappears.

Embodiment 4 FIG. FIG. 7 is a configuration diagram showing a fourth embodiment of the present invention, in which 1 is a part of an aircraft, 2 is an optical device mounted on the aircraft, 3 is an optical sensor provided inside the optical device 2, Reference numeral 4 denotes a drive shaft for changing the optical field of view of the optical sensor 3, 5 a drive motor for controlling the rotational movement of the drive shaft 4, 10 a damper, 7 a servo mechanism for changing the damping force of the damper 10, and 8 a servo mechanism. 7 is a controller for adjusting the gain of 7; A is an air flow around the optical device 2; A is a Karman vortex pair formed downstream of the separation point of the air flow A; K is a signal line of speed information from the aircraft;
ク is a command signal line from the controller 8 to the servo mechanism 7. FIG. 8 is a partial configuration diagram showing Embodiment 4 of the present invention, in which 1 is a part of the same aircraft as in FIG. 7, 2 is the same optical device as in FIG. 7, 2 is the same optical sensor as in FIG. 7 is the same rotary mechanism as in FIG. 7, 4 is the same drive shaft as in FIG. 7, and 5 is FIG.
7 is the same damper as in FIG. 7, 7 is the same servo mechanism as in FIG. 7, 8 is the same controller as in FIG. 7, G is the signal line of speed information from the same aircraft as in FIG. 7 is a command signal line from the controller 8 to the servo mechanism 7 which is the same as 7.

In FIG. 7, the damping force of the damper 10 provided on the drive shaft 4 is one of the parameters for determining the damping coefficient of the drive shaft 4 in the rotational direction.
By changing the damping force of the optical device 2 as shown in FIG. 8, in the equation of motion in the rotational direction of the drive shaft 4 of the optical device 2, the damping term changes as shown in Expression 13.

[0046]

(Equation 13)

For this reason, the damping coefficient of the drive shaft 4 in the rotation direction changes as shown in Expression 14.

[0048]

[Equation 14]

Therefore, the damping ratio in the rotational direction of the drive shaft 4 駆 動_S4
Changes as shown in Expression 15.

[0050]

(Equation 15)

The magnitude of the rotational direction of the damping ratio zeta _S4 of the drive shaft 4 is varied by the gain adjustment of the servo mechanism 7, during normal flight leave the gain of the servo mechanism 7 to 0, the optical apparatus 2
When the frequency f _A of the aerodynamic vibration according to the formula (2) approaches the frequency f _S in the rotational direction of the drive shaft 4 at a gain of 0 of the servo mechanism 7 shown in Equation 2, the controller 8 performs servo control via the signal line c. sufficiently large damping ratio zeta _S4 provides an instruction to increase the gain in the mechanism 7. Note that the frequency f _A of the aerodynamic vibration applied to the optical device 2 is estimated and calculated by the controller 8 using the aircraft speed information obtained from the signal line G and the arithmetic expression of Expression 7. By sufficiently increasing the rotational direction of the damping ratio zeta _S4 of the drive shaft 4, structural resonance occurring between the frequency of the aerodynamic vibrations according to the rotational direction of the natural frequency and the optics 2 of the drive shaft 4 is sufficiently In the frequency characteristic of the control system of the drive motor 5 of the optical device 2, the peak value of the gain at the resonance frequency e shown in FIG. 22 is sufficiently relaxed, and the oscillation of the signal near the resonance frequency e is reduced. Disappears.

Embodiment 5 FIG. FIG. 9 is a configuration diagram showing Embodiment 5 of the present invention, in which 1 is a part of an aircraft, 2 is an optical device mounted on the aircraft, 3 is an optical sensor provided inside the optical device 2, 4 is a drive shaft for changing the optical field of view of the optical sensor 3, 5 is a drive motor for controlling the rotational movement of the drive shaft 4, 6 is a balancer, 9 is a spring, 1
0 is a damper, 7 is a servo mechanism for changing the distance of the balancer 6 from the drive shaft 4, the length of expansion and contraction of the spring 9, and the damping force of the damper 10. 8 is a controller for adjusting the gain of the servo mechanism 7. The air flow around the device 2, A is a Karman vortex pair formed downstream of the separation point of the air flow A, G is a signal line of speed information from the aircraft, and G is a command signal line from the controller 8 to the servo mechanism 7. FIG. 10 is a partial configuration diagram showing a fifth embodiment of the present invention.
9, 2 is the same optical device as in FIG. 9, 3 is the same optical sensor as in FIG. 9, 3 is the same rotating mechanism as in FIG. 9, 4 is the same drive shaft as in FIG. 9, and 5 is the same drive as in FIG. Motor, 6 is FIG.
9 is the same spring as in FIG. 9, 10 is the same damper as in FIG. 9, 7 is the same servo mechanism as in FIG. 9, 8 is the same controller as in FIG. 9, and K is the speed information from the same aircraft as in FIG. Are signal signal lines from the controller 8 to the servo mechanism 7 as in FIG.

In FIG. 9, the distance of the balancer 6 provided on the drive shaft 4 from the drive shaft 4 is one of the parameters for determining the inertia coefficient of the drive shaft 4 in the rotation direction. The length of expansion and contraction of the spring 9 is one of the parameters for determining the elastic modulus of the drive shaft 4 in the rotational direction.
Since the damping force of the damper 10 provided in the above is one of the parameters for determining the damping coefficient of the drive shaft 4 in the rotation direction,
By changing the distance of the balancer 6 from the drive shaft 4 as shown in FIG. 10, changing the length of expansion and contraction of the spring 9 as shown in FIG. 10, and changing the damping force of the damper 10 as shown in FIG. In the equation of motion of the two drive shafts 4 in the rotation direction, the inertia term, the elasticity term, and the damping term change as shown in Expression 16.

[0054]

(Equation 16)

Therefore, the inertia coefficient of the drive shaft 4 in the rotational direction changes as shown in Equation 17, the elastic coefficient changes as shown in Equation 18, and the damping coefficient changes as shown in Equation 14.

[0056]

[Equation 17]

[0057]

(Equation 18)

Therefore, the damping ratio of the drive shaft 4 in the rotational direction 駆 動_S5
Changes as shown in Expression 19.

[0059]

[Equation 19]

[0060] magnitude of the rotational direction of the damping ratio zeta _S5 of the drive shaft 4 is varied by the gain adjustment of the servo mechanism 7, during normal flight leave the gain of the servo mechanism 7 to 0, the optical apparatus 2
When the frequency f _A of the aerodynamic vibration according to the formula (2) approaches the frequency f _S in the rotational direction of the drive shaft 4 at a gain of 0 of the servo mechanism 7 shown in Equation 2, the controller 8 performs servo control via the signal line c. sufficiently large damping ratio zeta _S5 provides an instruction to increase the gain in the mechanism 7. Note that the frequency f _A of the aerodynamic vibration applied to the optical device 2 is estimated and calculated by the controller 8 using the aircraft speed information obtained from the signal line G and the arithmetic expression of Expression 7. By sufficiently increasing the rotational direction of the damping ratio zeta _S5 of the drive shaft 4, structural resonance occurring between the frequency of the aerodynamic vibrations according to the rotational direction of the natural frequency and the optics 2 of the drive shaft 4 is sufficiently In the frequency characteristic of the control system of the drive motor 5 of the optical device 2, the peak value of the gain at the resonance frequency e shown in FIG. 22 is sufficiently relaxed, and the oscillation of the signal near the resonance frequency e is reduced. Disappears.

Embodiment 6 FIG. FIG. 11 is a block diagram showing a sixth embodiment of the present invention, in which 1 is a part of an aircraft, 2 is an optical device mounted on the aircraft, 3 is an optical sensor provided inside the optical device 2, 4 is a drive shaft for changing the optical field of view of the optical sensor 3, 5 is a drive motor for controlling the rotational movement of the drive shaft 4, 6 is a balancer, 7 is a servo mechanism for changing the distance of the balancer 6 from the drive shaft 4, 8 is a controller for adjusting the gain of the servo mechanism 7;
Is a vibration sensor provided on the outer wall surface of the optical device 2,
Optical device A is an airflow around the optical device 2, A is a Karman vortex pair formed downstream of the separation point of the airflow A, K is a command signal line from the controller 8 to the servo mechanism 7, and K is vibration information from the vibration sensor 11. This is a signal line. FIG. 12 is a partial configuration diagram showing Embodiment 6 of the present invention, in which 1 is the same as FIG.
Part of the same aircraft as 1, 2 is the same optical equipment as in FIG.
Is the same optical sensor as in FIG. 11, 4 is the same drive shaft as in FIG. 11,
5 is the same drive motor as in FIG. 11, 6 is the same balancer as in FIG. 11, 7 is the same servo mechanism as in FIG. 11, 8 is the same controller as in FIG. A command signal line and a signal line are signal lines for vibration information from the vibration sensor 11 as in FIG.

In FIG. 11, the distance of the balancer 6 provided on the drive shaft 4 from the drive shaft 4 is one of the parameters for determining the inertia coefficient of the drive shaft 4 in the rotation direction. By changing the distance from the drive shaft 4 as described above, in the equation of motion of the optical device 2 in the rotational direction of the drive shaft 4, the inertia term changes as shown in Expression 4. Therefore, the inertia coefficient of the drive shaft 4 in the rotation direction changes as shown in Expression 5. Therefore, the natural frequency f _{S1 in} the rotational direction of the drive shaft 4 changes as shown in Expression 6. The magnitude of the natural frequency f _{S1 in} the rotation direction of the drive shaft 4 is changed by adjusting the gain of the servo mechanism 7, but during normal flight, the gain of the servo mechanism 7 is set to 0, and the aerodynamic vibration applied to the optical device 2 is set. When the frequency f _A approaches the frequency f _S in the rotational direction of the drive shaft 4 at the gain 0 of the servo mechanism 7 shown in Equation 2, the controller 8 gives the gain to the servo mechanism 7 via a signal line. A change command is given, and the natural frequency f _S1 in the rotational direction of the drive shaft 4 shown in Expression 12 is sufficiently deviated from the frequency f _A of the aerodynamic vibration applied to the optical device 2. As for the frequency f _A of the aerodynamic vibration applied to the optical device 2, a detection value obtained by actually measuring the frequency of the aerodynamic vibration applied to the optical device 2 by the vibration sensor 11 and passing through the signal line is used. By making the natural frequency f _S1 in the rotational direction of the drive shaft 4 sufficiently far from the frequency f _A of the aerodynamic vibration applied to the optical device 2, structural resonance between the two can be avoided, and the driving of the optical device 2 can be prevented. In the frequency characteristic of the control system of the motor 4, the peak value of the gain at the resonance frequency e shown in FIG. 22 disappears, and the signal oscillation near the resonance frequency e disappears.

Embodiment 7 FIG. FIG. 13 is a configuration diagram showing a seventh embodiment of the present invention, in which 1 is a part of an aircraft, 2 is an optical device mounted on the aircraft, 3 is an optical sensor provided inside the optical device 2, 4 is a drive shaft for changing the optical field of view of the optical sensor 3, 5 is a drive motor for controlling the rotational movement of the drive shaft 4, 9 is a spring, 7 is a servo mechanism for changing the expansion and contraction length of the spring 9, 8 is a servo mechanism. 7 is a controller for adjusting the gain, 11 is a vibration sensor provided on the outer wall surface of the optical device 2, a is an airflow around the optical device 2, a is a Karman vortex pair formed downstream of the separation point of the airflow a, Is a command signal line from the controller 8 to the servo mechanism 7, and is a signal line for vibration information from the vibration sensor 11. FIG. 14 is a partial configuration diagram showing Embodiment 7 of the present invention, in which 1 is a part of the same aircraft as FIG.
13 is the same optical device as in FIG. 13, 3 is the same optical sensor as in FIG. 13, 4 is the same drive shaft as in FIG. 13, 5 is the same drive motor as in FIG. 13, 9 is the same spring as in FIG. 13, and 7 is the same servo as in FIG. The mechanism 8 is the same controller as in FIG.
Reference numeral 3 denotes a command signal line from the controller 8 to the servo mechanism 7, and reference numeral 3 denotes a vibration information signal line from the vibration sensor 11 as in FIG.

In FIG. 13, the expansion and contraction length of the spring 9 provided on the drive shaft 4 is one of the parameters for determining the elastic modulus of the drive shaft 4 in the rotation direction. Is changed, in the equation of motion of the rotation direction of the drive shaft 4 of the optical device 2, the elastic term changes as shown in Expression 8. Therefore, the elastic modulus of the drive shaft 4 in the rotation direction changes as shown in Expression 9. therefore,
The natural frequency f _{S2 in} the rotation direction of the drive shaft 4 changes as shown in Expression 10. The magnitude of the natural frequency f _{S2 in} the rotation direction of the drive shaft 4 is changed by adjusting the gain of the servo mechanism 7, but during normal flight, the gain of the servo mechanism 7 is set to 0, and the aerodynamic vibration applied to the optical device 2 is set. frequency f _a frequency of the rotating direction of the drive shaft 4 of the gain 0 in the servo mechanism 7 shown in Formula 2 f _S
The controller 8 gives a command to change the gain to the servo mechanism 7 via the signal line C when approaching to the drive shaft 4 so that the frequency f _A of the aerodynamic vibration applied to the optical device 2 is equal to The natural frequency f _S2 in the rotation direction is sufficiently separated. As for the frequency f _A of the aerodynamic vibration applied to the optical device 2, a detection value obtained by actually measuring the frequency of the aerodynamic vibration applied to the optical device 2 by the vibration sensor 11 and passing through the signal line is used. The natural frequency f _{S2 in} the rotation direction of the drive shaft 4 is calculated from the frequency f _A of the aerodynamic vibration applied to the optical device 2.
, The structural resonance between them is avoided, and the frequency characteristic of the control system of the drive motor 4 of the optical device 2 shows the gain peak at the resonance frequency e shown in FIG. There is no value, and signal oscillation around the resonance frequency e is eliminated.

Embodiment 8 FIG. FIG. 15 is a configuration diagram showing Embodiment 8 of the present invention, in which 1 is a part of an aircraft, 2 is an optical device mounted on the aircraft, 3 is an optical sensor provided inside the optical device 2, 4 is a drive shaft for changing the optical field of view of the optical sensor 3, 5 is a drive motor for controlling the rotational movement of the drive shaft 4, 6 is a balancer, 9 is a spring, 7
Is a servo mechanism for changing the distance of the balancer 6 from the drive shaft 4 and the length of expansion and contraction of the spring 9, 8 is a controller for adjusting the gain of the servo mechanism 7, and 11 is a vibration sensor provided on the outer wall surface of the optical device 2. A is an airflow around the optical device 2, A is a Karman vortex pair formed downstream of the separation point of the airflow A, K is a signal line of speed information from an aircraft, and K is a controller 8.
Command signal line from the servo mechanism 7 to the vibration sensor 11
This is a signal line for the vibration information from. FIG. 16 is a partial configuration diagram showing Embodiment 8 of the present invention, in which 1 is a part of the same aircraft as in FIG. 15, 2 is the same optical device as in FIG.
3 is the same optical sensor as in FIG. 15, 4 is the same drive shaft as in FIG. 15, 5 is the same drive motor as in FIG. 15, 6 is the same balancer as in FIG. 15, 9 is the same spring as in FIG. 15, and 7 is the same servo as in FIG. 15 is the same controller as in FIG. 15, and ク is a command signal line from the controller 8 to the servo mechanism 7 as in FIG. 15, and ケ is a signal line for vibration information from the vibration sensor 11 as in FIG.

In FIG. 15, the distance of the balancer 6 provided on the drive shaft 4 from the drive shaft 4 is one of the parameters for determining the inertia coefficient of the drive shaft 4 in the rotational direction. Since the length of expansion and contraction of the spring 9 is one of the parameters for determining the elastic modulus of the drive shaft 4 in the rotation direction, the distance between the balancer 6 and the drive shaft 4 is changed as shown in FIG. By changing the expansion and contraction length as shown in 6, in the equation of motion of the rotation direction of the drive shaft 4 of the optical device 2, the inertia term and the elasticity term are changed as shown in Expression 11. Therefore, the inertia coefficient in the rotation direction of the drive shaft 4 changes as shown in Equation 5, and the elastic coefficient changes as shown in Equation 9. Therefore, the natural frequency f in the rotation direction of the drive shaft 6
_S3 changes as shown in Expression 12. The magnitude of the natural frequency f _{S3 in} the rotation direction of the drive shaft 4 is changed by adjusting the gain of the servo mechanism 7, but during normal flight, the gain of the servo mechanism 7 is set to 0, and the aerodynamic vibration applied to the optical device 2 is set. Frequency f _A
Is approaching the frequency f _S in the rotational direction of the drive shaft 4 at a gain of 0 of the servo mechanism 7 shown in Equation 2, the controller 8 sends a command to the servo mechanism 7 to change the gain via the signal line C. The frequency f _{A of the} aerodynamic vibration applied to the optical device 2
In contrast, the natural frequency f _S3 in the rotational direction of the drive shaft 4 shown in Expression 12 is sufficiently separated. As for the frequency f _A of the aerodynamic vibration applied to the optical device 2, a detection value obtained by actually measuring the frequency of the aerodynamic vibration applied to the optical device 2 by the vibration sensor 11 and passing through the signal line is used. Optical equipment 2
The natural frequency f _S3 in the rotational direction of the drive shaft 4 is sufficiently deviated from the frequency f _A of the aerodynamic vibration according to the above, whereby structural resonance between the two is avoided, and the drive motor 5 of the optical device 2 In the frequency characteristic of the control system, the peak value of the gain at the resonance frequency e shown in FIG.
Signal oscillation around the resonance frequency e is eliminated.

Embodiment 9 FIG. 17 is a configuration diagram showing a ninth embodiment of the present invention, in which 1 is a part of an aircraft, 2 is an optical device mounted on the aircraft, 3 is an optical sensor provided inside the optical device 2, Reference numeral 4 denotes a drive shaft for changing the optical field of view of the optical sensor 3, 5 a drive motor for controlling the rotational movement of the drive shaft 4, 10 a damper, 7 a servo mechanism for changing the damping force of the damper 10, and 8 a servo mechanism. 7 is a controller for adjusting the gain, 11 is a vibration sensor provided on the outer wall surface of the optical device 2, a is an airflow around the optical device 2, a is a Karman vortex pair formed downstream of the separation point of the airflow a, Is a command signal line from the controller 8 to the servo mechanism 7, and is a signal line for vibration information from the vibration sensor 11. FIG. 18 is a partial configuration diagram showing a ninth embodiment of the present invention, in which 1 is a part of the same aircraft as in FIG. 17, 2 is the same optical device as in FIG. 17, 2 is the same optical sensor as in FIG. 17 is the same rotation mechanism as in FIG. 17, 4 is the same drive shaft as in FIG. 17, 5 is the same drive motor as in FIG. 17, and 10 is FIG.
17 is the same servo mechanism as in FIG. 17, 8 is the same controller as in FIG. 17, １７ is a command signal line from the controller 8 to the servo mechanism 7 as in FIG.
7 is a signal line of vibration information from the same vibration sensor 11 as FIG.

In FIG. 17, the damping force of the damper 10 provided on the drive shaft 4 is one of the parameters for determining the damping coefficient of the drive shaft 4 in the rotation direction.
By changing the damping force of 0 as shown in FIG. 18, in the equation of motion of the drive shaft 4 of the optical device 2 in the rotation direction, the damping term changes as shown in Expression 13. Therefore, the damping coefficient of the drive shaft 4 in the rotation direction changes as shown in Expression 14. Thus, the rotation direction of the damping ratio zeta _S4 of the drive shaft 4 is changed as the number 15 of. The magnitude of the damping ratio ζ _{S4 in} the rotational direction of the drive shaft 4 is changed by adjusting the gain of the servo mechanism 7, but during normal flight, the gain of the servo mechanism 7 is set to 0 to reduce the aerodynamic vibration applied to the optical device 2. Frequency f _A is Equation 2
When approaching the frequency f _S in the rotational direction of the drive shaft 4 at a gain of 0 of the servo mechanism 7 as shown in the following, the controller 8 gives a command to increase the gain to the servo mechanism 7 via a signal line 減 衰 and the damping ratio ζ _{Make S4} large enough. As for the frequency f _A of the aerodynamic vibration applied to the optical device 2, a detection value obtained by actually measuring the frequency of the aerodynamic vibration applied to the optical device 2 by the vibration sensor 11 and passing through the signal line is used. By sufficiently increasing the rotational direction of the damping ratio zeta _S4 of the drive shaft 4, structural resonance occurring between the frequency of the aerodynamic vibrations according to the rotational direction of the natural frequency and the optics 2 of the drive shaft 4 is sufficiently In the frequency characteristic of the control system of the drive motor 5 of the optical device 2, the peak value of the gain at the resonance frequency e shown in FIG. 22 is sufficiently relaxed, and the oscillation of the signal near the resonance frequency e is reduced. Disappears.

Embodiment 10 FIG. FIG. 19 is a configuration diagram showing Embodiment 10 of the present invention, in which 1 is a part of an aircraft, 2 is an optical device mounted on the aircraft, 3 is an optical sensor provided inside the optical device 2, 4 is a drive shaft for changing the optical field of view of the optical sensor 3, 5 is a drive motor for controlling the rotational movement of the drive shaft 4, 6 is a balancer, 9 is a spring, 10 is a damper, and 7 is a drive shaft 4 of the balancer 6. Is a servo mechanism for changing the distance of the spring 9, the expansion and contraction length of the spring 9, and the damping force of the damper 10, 8 is a controller for adjusting the gain of the servo mechanism 7, and 11 is a vibration sensor provided on the outer wall surface of the optical device 2. Is a pair of Karman vortices that can be generated downstream of the separation point of the air flow a, A is a command signal line from the controller 8 to the servo mechanism 7, and D is a signal line of vibration information from the vibration sensor 11. is there. 20 is a partial configuration diagram showing Embodiment 10 of the present invention, in which 1 is a part of the same aircraft as in FIG. 19, 2 is the same optical device as in FIG. 19, 3 is the same optical sensor as in FIG. Is FIG.
19 is the same drive shaft as in FIG. 19, and 5 is the same drive shaft as in FIG.
19, the same balancer as FIG. 19, 9 the same spring as FIG. 19, 10 the same damper as FIG. 19, 7 the same servo mechanism as FIG. 19, and 8 the same controller as FIG. Command signal lines from the controller 8 to the servo mechanism 7 are the same as those in FIG.
This is a signal line for the vibration information from.

In FIG. 19, the distance of the balancer 6 provided on the drive shaft 4 from the drive shaft 4 is one of the parameters for determining the inertia coefficient of the drive shaft 4 in the rotation direction. The expansion and contraction length of the spring 9 is one of the parameters for determining the elastic modulus of the drive shaft 4 in the rotation direction, and the damping force of the damper 10 provided on the drive shaft 4 determines the damping coefficient of the drive shaft 4 in the rotation direction. 20, the distance between the balancer 6 and the drive shaft 4 is changed as shown in FIG. 20, the length of the spring 9 is changed as shown in FIG. 20, and the damping force of the damper 10 is reduced as shown in FIG. Thus, in the equation of motion of the optical device 2 in the rotational direction of the drive shaft 4, the inertia term, the elasticity term, and the damping term change as shown in Expression 16. Therefore, the inertia coefficient of the drive shaft 4 in the rotational direction changes as shown in Equation 17, the elastic coefficient changes as shown in Equation 18, and the damping coefficient changes as shown in Equation 14. Thus, the rotation direction of the damping ratio zeta _S5 of the drive shaft 6 is changed as in Equation 19. Damping ratio in the rotation direction of the drive shaft 4 ζ
The size of the _S4, vary the gain adjustment of the servo mechanism 7, but during normal flight leave the gain of the servo mechanism 7 to 0, the frequency f _A of the aerodynamic vibrations according to the optical device 2 Number 2
When approaching the frequency f _S in the rotational direction of the drive shaft 4 at a gain of 0 of the servo mechanism 7 as shown in the following, the controller 8 gives a command to increase the gain to the servo mechanism 7 via a signal line 減 衰 and the damping ratio ζ _{Make S4} large enough. As for the frequency f _A of the aerodynamic vibration applied to the optical device 2, a detection value obtained by actually measuring the frequency of the aerodynamic vibration applied to the optical device 2 by the vibration sensor 11 and passing through the signal line is used. By sufficiently increasing the rotational direction of the damping ratio zeta _S5 of the drive shaft 4, structural resonance occurring between the frequency of the aerodynamic vibrations according to the rotational direction of the natural frequency and the optics 2 of the drive shaft 4 is sufficiently In the frequency characteristic of the control system of the drive motor 5 of the optical device 2, the peak value of the gain at the resonance frequency e shown in FIG. 22 is sufficiently relaxed, and the oscillation of the signal near the resonance frequency e is reduced. Disappears.

[0071]

According to the first aspect of the invention, in the optical device 2 mounted on an aircraft, the natural vibration of the drive shaft 4 in the rotational direction is changed by changing the inertia coefficient of the drive shaft 4 of the optical device 2 in the rotational direction. The number can be sufficiently deviated from the frequency of the aerodynamic vibration applied to the optical device 2, and the structural resonance between the two can be suppressed.

According to the second aspect, in the optical device 2 mounted on an aircraft, the natural vibration of the drive shaft 4 in the rotational direction is changed by changing the elastic coefficient of the drive shaft 4 of the optical device 2 in the rotational direction. The number can be sufficiently deviated from the frequency of the aerodynamic vibration applied to the optical device 2, and the structural resonance between the two can be suppressed.

According to the third aspect of the present invention, in the optical device 2 mounted on an aircraft, the inertia coefficient and the elastic coefficient in the rotation direction of the drive shaft 4 of the optical device 2 are changed.
The natural frequency in the rotation direction of the drive shaft 4 can be sufficiently deviated from the frequency of the aerodynamic vibration applied to the optical device 2, and the structural resonance between the two can be suppressed.

According to the fourth aspect of the present invention, in the optical device 2 mounted on an aircraft, the damping coefficient in the rotation direction of the drive shaft 4 of the optical device 2 is changed to thereby provide the attenuation ratio of the drive shaft 4 in the rotation direction. And the structural resonance occurring between the natural frequency of the drive shaft 4 in the rotation direction and the frequency of the aerodynamic vibration applied to the optical device 2 can be sufficiently attenuated.

According to the fifth aspect, in the optical device 2 mounted on an aircraft, the inertia coefficient, the elastic coefficient, and the damping coefficient of the optical device 2 in the rotation direction of the drive shaft 4 are changed, so that the drive shaft 4 , The structural resonance occurring between the natural frequency of the drive shaft 4 in the rotational direction and the frequency of the aerodynamic vibration applied to the optical device 2 can be sufficiently attenuated.

According to the sixth aspect of the present invention, in the optical device 2 mounted on an aircraft, the natural vibration of the drive shaft 4 in the rotational direction is changed by changing the inertia coefficient in the rotational direction of the drive shaft 4 of the optical device 2. The number can be sufficiently deviated from the frequency of the aerodynamic vibration applied to the optical device 2, and the structural resonance between the two can be suppressed.

According to the seventh aspect, in the optical device 2 mounted on an aircraft, the natural vibration of the drive shaft 4 in the rotational direction is changed by changing the elastic coefficient of the drive shaft 4 of the optical device 2 in the rotational direction. The number can be sufficiently deviated from the frequency of the aerodynamic vibration applied to the optical device 2, and the structural resonance between the two can be suppressed.

According to the eighth aspect of the present invention, in the optical device 2 mounted on an aircraft, the inertia coefficient and the elastic coefficient in the rotation direction of the drive shaft 4 of the optical device 2 are changed.
The natural frequency in the rotation direction of the drive shaft 4 can be sufficiently deviated from the frequency of the aerodynamic vibration applied to the optical device 2, and the structural resonance between the two can be suppressed.

According to the ninth aspect of the present invention, in the optical device 2 mounted on an aircraft, the damping coefficient in the rotational direction of the drive shaft 4 of the optical device 2 is changed by changing the attenuation coefficient in the rotational direction of the drive shaft 4. And the structural resonance occurring between the natural frequency of the drive shaft 4 in the rotation direction and the frequency of the aerodynamic vibration applied to the optical device 2 can be sufficiently attenuated.

According to the tenth aspect, in the optical device 2 mounted on an aircraft, the drive shaft 4 of the optical device 2
By changing the inertia coefficient, the elastic coefficient and the damping coefficient in the rotational direction of the drive shaft 4, the damping ratio in the rotational direction of the drive shaft 4 is increased,
It is possible to sufficiently attenuate the structural resonance occurring between the natural frequency in the rotation direction of the drive shaft 4 and the frequency of the aerodynamic vibration applied to the optical device 2.

[Brief description of the drawings]

FIG. 1 is a configuration diagram showing a first embodiment of the present invention.

FIG. 2 is a partial configuration diagram showing the first embodiment of the present invention.

FIG. 3 is a configuration diagram showing a second embodiment of the present invention.

FIG. 4 is a partial configuration diagram showing a second embodiment of the present invention.

FIG. 5 is a configuration diagram showing a third embodiment of the present invention.

FIG. 6 is a partial configuration diagram showing a third embodiment of the present invention.

FIG. 7 is a configuration diagram showing a fourth embodiment of the present invention.

FIG. 8 is a partial configuration diagram showing a fourth embodiment of the present invention.

FIG. 9 is a configuration diagram showing a fifth embodiment of the present invention.

FIG. 10 is a partial configuration diagram showing a fifth embodiment of the present invention.

FIG. 11 is a configuration diagram showing a sixth embodiment of the present invention.

FIG. 12 is a partial configuration diagram showing a sixth embodiment of the present invention.

FIG. 13 is a configuration diagram showing a seventh embodiment of the present invention.

FIG. 14 is a partial configuration diagram showing Embodiment 7 of the present invention.

FIG. 15 is a configuration diagram showing an eighth embodiment of the present invention.

FIG. 16 is a partial configuration diagram showing Embodiment 8 of the present invention.

FIG. 17 is a configuration diagram showing a ninth embodiment of the present invention.

FIG. 18 is a partial configuration diagram showing a ninth embodiment of the present invention.

FIG. 19 is a configuration diagram showing a tenth embodiment of the present invention.

FIG. 20 is a partial configuration diagram showing Embodiment 10 of the present invention.

FIG. 21 is an external view of a conventional optical device mounted on an aircraft.

FIG. 22 is a graph showing frequency characteristics of a control system of a drive motor 5 of a conventional optical device 2 mounted on an aircraft.

[Explanation of symbols]

1 part of aircraft, 2 optical devices, 3 optical sensors, 4
Drive shaft, 5 drive motor, 6 balancer, 7 servo mechanism, 8 controller, 9 spring, 10 damper,
11 Vibration sensor, a. Air flow around optical device 2, b.
Karman vortex pair, (c) frequency characteristics of the drive motor 5 when it does not resonate, (d) frequency characteristics of the drive motor 5 when it resonates, (e) resonance frequency, (f) peak value of the gain at the resonance frequency (e), H Command signal line to servo mechanism 7, signal line for vibration information.

Claims (10)

[Claims] 1. An optical device mounted on an aircraft, wherein a drive shaft provided inside the optical device and changing an optical field of view, a drive motor provided on the drive shaft and controlling a rotational movement of the drive shaft, A balancer whose distance from the drive shaft changes, a servo mechanism that changes the inertia coefficient of the drive shaft in the rotation direction by changing the distance of the balancer from the drive shaft, and an optical mechanism that changes the natural frequency of the drive shaft in the rotation direction. A controller that is provided inside the device and that adjusts the gain of the servo mechanism so that the natural frequency in the rotational direction of the drive shaft does not approach the frequency of the aerodynamic vibration applied to the optical device estimated and calculated based on speed information from the aircraft. And a vibration suppressing device. 2. An optical device mounted on an aircraft, wherein a drive shaft provided inside the optical device and changing an optical field of view, a drive motor provided on the drive shaft and controlling a rotational movement of the drive shaft, A spring that applies torque in the rotation direction of the drive shaft, a servo mechanism that changes the elastic modulus of the drive shaft in the rotation direction by changing the expansion and contraction length of the spring, and changes the natural frequency of the drive shaft in the rotation direction; A controller for adjusting the gain of the servo mechanism so that the natural frequency in the rotation direction of the drive shaft is not close to the frequency of the aerodynamic vibration applied to the optical device estimated and calculated based on the speed information from the aircraft. A vibration suppressing device comprising: 3. An optical device mounted on an aircraft, wherein a drive shaft provided inside the optical device and changing an optical field of view, a drive motor provided on the drive shaft and controlling a rotational movement of the drive shaft, A balancer whose distance from the drive shaft changes, a servo mechanism that changes the inertia coefficient in the rotational direction of the drive shaft by changing the distance of the balancer from the drive shaft and changes the natural frequency around the drive shaft, A spring that applies torque in the rotation direction, a servo mechanism that changes the elastic modulus in the rotation direction of the drive shaft by changing the expansion and contraction length of the spring, and changes the natural frequency in the rotation direction of the drive shaft, and is provided inside the optical device The gain of each servo mechanism is adjusted so that the natural frequency in the rotational direction of the drive shaft does not approach the frequency of the aerodynamic vibration applied to the optical device estimated and calculated based on the speed information from the aircraft. A vibration suppressing device comprising a roller. 4. An optical device mounted on an aircraft, wherein: a drive shaft provided inside the optical device for changing an optical field of view; a drive motor provided on the drive shaft for controlling a rotational movement of the drive shaft; A damper for applying a damping force in the rotation direction of the drive shaft, a servo mechanism for adjusting the damping force of the damper to change the damping coefficient in the rotation direction of the drive shaft to change the damping ratio in the rotation direction of the drive shaft, and an optical device Provided inside,
When the natural frequency in the rotational direction of the drive shaft is close to the frequency of the aerodynamic vibration on the optical device estimated and calculated based on the speed information from the aircraft, the gain of the servo mechanism is adjusted to attenuate the rotational direction of the drive shaft. A vibration suppression device comprising a controller for increasing a ratio. 5. An optical device mounted on an aircraft, wherein: a drive shaft provided inside the optical device for changing an optical field of view; a drive motor provided on the drive shaft for controlling a rotational movement of the drive shaft; A balancer whose distance from the drive shaft changes, a servo mechanism that changes the inertia coefficient in the rotation direction of the drive shaft by changing the distance of the balancer from the drive shaft and changes the natural frequency in the rotation direction of the drive shaft, A spring that applies torque in the rotation direction of the shaft, a servo mechanism that changes the elastic modulus in the rotation direction of the drive shaft by changing the expansion and contraction length of the spring, and changes the natural frequency in the rotation direction of the drive shaft; A damper that applies damping force in the rotation direction, a servo mechanism that adjusts the damping force of the damper to change the damping coefficient in the rotation direction of the drive shaft and changes the damping ratio in the rotation direction of the drive shaft, and installed inside the optical device La When the natural frequency in the rotational direction of the drive shaft approaches the frequency of the aerodynamic vibration applied to the optical device estimated and calculated based on the speed information from the aircraft, the gain of each servo mechanism is adjusted to rotate the drive shaft. A controller for increasing a damping ratio in a direction. 6. An optical device mounted on an aircraft, comprising: a drive shaft provided inside the optical device for changing an optical field of view; a drive motor provided on the drive shaft for controlling rotational movement of the drive shaft; A balancer whose distance from the drive shaft changes, a servo mechanism that changes the inertia coefficient of the drive shaft in the rotation direction by changing the distance of the balancer from the drive shaft, and an optical mechanism that changes the natural frequency of the drive shaft in the rotation direction. A vibration sensor provided on the outer wall surface of the device and detecting the frequency of aerodynamic vibration applied to the optical device, and an optical device provided inside the optical device and the natural frequency of the rotation direction of the drive shaft is detected by the vibration sensor A vibration suppression device comprising: a controller for adjusting a gain of a servo mechanism so as not to approach the frequency of the aerodynamic vibration. 7. An optical device mounted on an aircraft, wherein: a drive shaft provided inside the optical device for changing an optical field of view; a drive motor provided on the drive shaft for controlling a rotational motion of the drive shaft; A spring that applies torque in the rotation direction of the drive shaft, a servo mechanism that changes the elastic modulus of the drive shaft in the rotation direction by changing the expansion and contraction length of the spring, and changes the natural frequency of the drive shaft in the rotation direction; A vibration sensor that is provided on the outer wall surface of the optical device and detects the frequency of aerodynamic vibration applied to the optical device; and an optical device that is provided inside the optical device and has a natural frequency in the rotation direction of the drive shaft detected by the vibration sensor. A vibration suppression device, comprising: a controller for adjusting a gain of a servo mechanism so as not to approach a frequency of aerodynamic vibration. 8. An optical device mounted on an aircraft, wherein a drive shaft provided inside the optical device for changing an optical field of view, a drive motor provided on the drive shaft and controlling a rotational movement of the drive shaft, A balancer whose distance from the drive shaft changes, a servo mechanism that changes the inertia coefficient in the rotational direction of the drive shaft by changing the distance of the balancer from the drive shaft and changes the natural frequency around the drive shaft, A spring that applies torque in the rotation direction, a servo mechanism that changes the elastic modulus of the drive shaft in the rotation direction by changing the expansion and contraction length of the spring, and changes the natural frequency of the drive shaft in the rotation direction, and the outer wall surface of the optical device And a vibration sensor that detects the frequency of aerodynamic vibration applied to the optical device, and an aerodynamic force that is provided inside the optical device and whose natural frequency in the rotation direction of the drive shaft is detected by the vibration sensor. And a controller for adjusting the gain of each servo mechanism so as not to approach the frequency of dynamic vibration. 9. An optical device mounted on an aircraft, wherein: a drive shaft provided inside the optical device for changing an optical field of view; a drive motor provided on the drive shaft for controlling a rotational movement of the drive shaft; A damper for applying a damping force in the rotation direction of the drive shaft, a servo mechanism for adjusting the damping force of the damper to change the damping coefficient in the rotation direction of the drive shaft to change the damping ratio in the rotation direction of the drive shaft, and an optical device A vibration sensor that is provided on the outer wall surface of the optical device and detects the frequency of aerodynamic vibration applied to the optical device; and an optical device that is provided inside the optical device and has a natural frequency in the rotation direction of the drive shaft detected by the vibration sensor. A vibration suppression device, comprising: a controller that adjusts a gain of a servo mechanism when the frequency approaches an aerodynamic vibration and increases a damping ratio in a rotation direction of a drive shaft. 10. An optical device mounted on an aircraft,
A drive shaft that is provided inside the optical device and changes the optical field of view, a drive motor that is provided on the drive shaft and controls the rotational movement of the drive shaft, a balancer that changes the distance from the drive shaft, and a balancer A servo mechanism that changes the inertia coefficient in the rotation direction of the drive shaft by changing the distance from the drive shaft to change the natural frequency in the rotation direction of the drive shaft, a spring that applies torque in the rotation direction of the drive shaft, and a spring A servo mechanism that changes the elastic modulus of the drive shaft in the rotation direction by changing the expansion and contraction length of the drive shaft and changes the natural frequency of the drive shaft in the rotation direction; a damper that applies a damping force in the rotation direction of the drive shaft; A servo mechanism for adjusting a damping force to change a damping coefficient in a rotation direction of the drive shaft to change a damping ratio in a rotation direction of the drive shaft;
A vibration sensor provided on the outer wall surface of the optical device for detecting a frequency of aerodynamic vibration applied to the optical device, and an optical device provided inside the optical device and having a natural frequency in a rotation direction of the drive shaft detected by the vibration sensor A controller that adjusts the gain of each servo mechanism when the frequency is close to the frequency of the aerodynamic vibration to increase the damping ratio in the rotational direction of the drive shaft.