Patent Description:
The present application relates to the technical field of space inertial attitude actuators, and in particular to a low-disturbance dual-gimbal flywheel having a spatial parallel mechanism.

A flywheel and a control moment gyroscope have been widely used in a spacecraft such as a remote sensing satellite, a communication satellite and a space telescope. They are the key inertial attitude actuators for a long-life spacecraft to achieve attitude maneuver and stability. With the development of aerospace science and technology, users have put forward higher requirements for an ultra-quiet and ultra-agile satellite platform; with the rapid development of a small and medium-sized satellite, the weight control for the product on the satellite has become more and more stringent. Therefore, an ultra-light, ultra-quiet, and ultra-agile satellite platform is required, and, the attitude actuator on the satellite platform is required to have the property of light weight, small vibration, and large output moment.

The flywheel can only output a moment in a single direction, and the amplitude of the moment is small; although the existing gimbal flywheel has a certain yaw ability on the angular momentum vector in a radial direction, the yaw range is generally relatively small; while a single gimbal control moment gyroscope can output a large moment in <NUM>°, and it has a large moment magnification and a high efficiency; a double gimbal control moment gyroscope can output a large moment in a whole space, and the moment magnification thereof is limited by the locking moment of a motor, the structure is relatively complicated, and the reliability is low, so it is rarely used in orbit. It can be seen that in order to achieve a three-axis attitude control for the spacecraft, at least three flywheel products or four single gimbal control moment gyroscope products or one double gimbal control moment gyroscope product is required; in order to achieve the zero-momentum three-axis attitude control for the spacecraft, more flywheels and control moment gyroscopes are required, making it difficult to reduce the weight of the attitude control actuator system. In addition, the gyroscopic moment generated by the coupling of the speed of the spacecraft during rapid maneuver and the angular momentum of the rotor may act on the frame motor, in this situation, the frame motor needs to be locked, but limited locking moment limits the maneuvering speed of the spacecraft such as a satellite.

The flywheel and the rotor in the control moment gyroscope may produce broadband micro-vibration during high-speed rotation, and the systemic and effective vibration control is not performed on the product. If they are directly mounted on the spacecraft, the broadband micro-vibration generated by the flywheel or the control moment gyroscope may become one of the main vibration sources of the spacecraft, which may affect the attitude stability and ultra-quietness of the spacecraft platform, and to a certain extent, affect the realization of the performance index of the load, and even the attitude stability of the satellite platform.

In view of the small output moment of the flywheel, it is impossible to achieve the rapid attitude maneuvering of the spacecraft such as a satellite. The flywheel and the single gimbal control moment gyroscope require multiple products to work together to achieve the three-axis attitude control for the spacecraft; while the output moment of the double gimbal control moment gyroscope is limited by the locking moment of the frame motor, which causes the moment amplification effect to be not fully exerted. In addition, the gyroscopic moment generated by the coupling of the maneuvering speed of the spacecraft and the angular momentum of the rotor may act on the frame along the gimbal axis of the control moment gyroscope, which needs to rely on the frame motor to lock. Therefore, the amplitude of the locking moment of the frame motor also limits the maneuvering speed of the spacecraft such as a satellite. On the other hand, the flywheel and the rotor in the control moment gyroscope may produce broadband micro-amplitude vibration during high-speed rotation, which becomes one of the main vibration sources of the spacecraft, thereby affecting the attitude stability and ultra-quietness of the spacecraft platform. In order to keep up with the rapid development of small and medium-sized spacecraft, it is necessary to develop a lighter, quieter, and more agile attitude control platform, and develop an attitude control actuator with better comprehensive performance such as weight, micro-vibration, anti-mechanical performance, attitude control moment and locking ability to match with the attitude control platform. <CIT> discloses a momentum-control system. The momentum-control system comprises a plurality of momentum actuators and a platform, upon which the plurality of momentum actuators are mounted. The momentum control system further comprises a plurality of active struts mounted on the bottom side of the platform. The active struts are configured to produce a force to steer the plurality of momentum actuators and the platform to produce forces and moments for spacecraft attitude control and disturbance suppression.

The technical problem to be solved by the present invention is to overcome the disadvantages of the prior art and a low-disturbance dual-gimbal flywheel is provided according to the appended claims. The spatial parallel mechanism is used for achieving arbitrary continuous angular maneuvering of the rotor angular momentum in a local area, thereby outputting a larger gyroscopic moment in a direction orthogonal to an angular momentum and a frame angular velocity, and a speed change of a rotor can output a small reaction moment in a direction of the angular momentum vector, which may all be used as the attitude control moment for the spacecraft, so as to achieve the omnidirectional attitude control for the spacecraft. The self-locking function of the nut and the lead screw may be used for achieving the locking of the non-maneuvering branch chain, which overcomes the limitation of the frame motor locking moment on the output moment of the double gimbal control moment gyroscope and the maneuvering speed of the spacecraft. A rod on the vibration transmission path may be designed as a spring damper to form a passive vibration isolator, so as to isolate the micro-vibration generated by the rotor assembly. The present invention is beneficial to achieve the functions of an ultra-light, ultra-quiet, ultra-agile, omnidirectional control, etc., and is not be limited by the locking moment.

A low-disturbance dual-gimbal flywheel is defined in claim <NUM>. Embodiments thereof are defined in the dependent claims.

The advantages of the present invention compared with the prior art are as follows.

The present application will be further described below in conjunction with the drawings.

The high-speed rotation of a rotor produces a certain angular momentum. The speed change of the motor changes the magnitude of the angular momentum; while the rotation of the frame changes the direction of the angular momentum; and changes in the magnitude and direction of the angular momentum may cause the exchange of the angular momentum between an actuator and the spacecraft, which can control the attitude maneuver and stability of the spacecraft. The change of the magnitude of the angular momentum may output a smaller reaction moment, while the change of the direction of the angular momentum may output a larger gyroscopic moment; and they are collectively used as the attitude control moment of the spacecraft such as a satellite. Therefore, a low-disturbance dual-gimbal flywheel having a spatial parallel mechanism is provided according to the present application. The spatial parallel mechanism is mainly used as the frame platform to drag the rotor to change the direction of the angular momentum, thereby outputting a larger gyroscopic moment; and outputting the smaller reaction moment through the speed change of the rotor. They work together to create a three-dimensional attitude control moment space, and one set of this product can output attitude control moment in any direction, thereby achieving three-axis attitude control for the spacecraft.

As shown in <FIG>, a low-disturbance dual-gimbal flywheel having a spatial parallel mechanism according to the present application includes a spatial parallel mechanism <NUM>, one wheel body <NUM>, one high-speed motor <NUM>, a bearing <NUM>, and a housing <NUM>; the wheel body <NUM> is a circular frame structure, and a circle is supported by spokes circularly radiating outward from the circular frame. A certain acute angle is present between the spokes and the circular plane, which improves the mechanical resistance of the wheel body <NUM> in the axial direction. The bearing <NUM> is mounted at the center of the wheel body <NUM>, and the rotor of the high-speed motor <NUM> is mounted on the bearing <NUM>; the bearing <NUM> and the stator of the high-speed motor <NUM> are mounted on the movable platform <NUM> of the spatial parallel mechanism <NUM> together.

The wheel body <NUM>, the stator and rotor of the high-speed motor <NUM>, and the bearing <NUM> are coaxial mounted. The spatial parallel mechanism <NUM>, the wheel body <NUM>, the high-speed motor <NUM> and the bearing <NUM> form a whole, which is mounted inside the housing <NUM> through a port of the static base <NUM> of the spatial parallel mechanism <NUM>; and a lug outside the housing <NUM> is connectable to the outer spacecraft deck. The housing <NUM> must enclose the maximum moving range of the rotor, and includes an upper housing and a lower housing, which are connected to each other by screw and sealed by soldering. The housing <NUM> forms a closed cavity after being pumped by the air extraction nozzle assembly, which provides a clean, airtight, low-pressure working environment for the platform, rotor, motor and other components therein. It not only prevents the rapid volatilization of the lubricating oil, but also prevents the entry of extras, and also reduces the wind resistance of the product during the ground test.

The rotating parts in the wheel body <NUM>, the high-speed motor <NUM> and the bearing <NUM> together form the rotor assembly; the wheel body <NUM> provides a certain moment of inertia; the bearing <NUM> stably supports the rotor assembly for a long time and provides a rotating pair for the rotor assembly; the high-speed motor <NUM> drives the rotor assembly to rotate at a high speed, and the speed of the high-speed motor <NUM> generally ranges from tens of revolutions per minute to tens of thousands of revolutions per minute, which combines with the moment of inertia to generate a certain angular momentum; and the angular momentum of the rotor assembly is changed by the speed change of the high-speed motor <NUM>, so as to output the reaction moment required for the attitude control of the spacecraft.

In this embodiment, the spatial parallel mechanism <NUM> is a <NUM>-PUU parallel mechanism including three drive branch chains, and the schematic view of the structure is shown in <FIG>, including a low-speed motor <NUM>, a static base <NUM>, a lead screw <NUM>, and a nut <NUM>, a movable platform <NUM>, a first hinge <NUM>, a spring damper <NUM>, a second hinge <NUM>, and a guide rod <NUM>; the movable platform <NUM> is connected to the spring damper <NUM> through the first hinge <NUM>, and the spring damper <NUM> is connected to the nut <NUM> through the second hinge <NUM>; and the nut <NUM> is mounted on the lead screw <NUM>, the lead screw <NUM> is mounted on the leg of the static base <NUM> through the bearing, the static base <NUM> includes a platform and three legs, the platform is support by three legs evenly distributed along the circumferential direction at a bottom of the platform; the rotor of the low-speed motor <NUM> is fixed to the end of the lead screw <NUM> by a nut, the stator of the low-speed motor <NUM> is mounted on the end of the leg of the static base <NUM> by a screw, and the stator and the rotor of the low-speed motor <NUM> are coaxially mounted; the guide rod <NUM> is mounted along the leg of the static base <NUM>, the nut <NUM> is movable along the guide rod <NUM>, and the guide rod <NUM> is configured to guide the nut. An angle between the lead screw <NUM> and a plane of an external mounting interface of the housing <NUM> is an angle α, and the angle α is optimized according to kinematics and dynamics characteristics of the low-disturbance dual-gimbal flywheel having the spatial parallel mechanism.

The low-speed motor <NUM> drives the lead screw <NUM> to rotate, and the rotational angular velocity of the lead screw is relatively low, generally not exceeding <NUM> rad/s. The lead screw <NUM> drives the nut <NUM> to move back and forth along the lead screw <NUM>; and then, the rotation of the low-speed motor <NUM> is transmitted to the movement of the relevant branch chain of the movable platform <NUM> through the second hinge <NUM>, the spring damper <NUM> and the first hinge <NUM>; finally, the movement of three driving branch chains enables the movable platform <NUM> to have at least two degrees of freedom of rotation in mutually perpendicular directions of a plane, which can achieve the yaw of arbitrarily continuous angle in a large range of a certain local area. The rotor assembly is mounted on the movable platform <NUM> of the frame assembly. If the direction of the angular momentum vector of the rotor assembly is not parallel to the direction of the rotation axis, the direction of the angular momentum vector can be changed by the rotation of the movable platform <NUM>, so as to output the gyroscopic moment required for the attitude control of the spacecraft. In this embodiment, the direction of the angular momentum vector is perpendicular to the plane of the movable platform <NUM>, so that the angular momentum vector can be yaw at arbitrarily continuous angle in a certain local larger spherical area to output a large gyroscopic moment, thereby achieving the super agility of the satellite platform. During the rapid maneuver of the spacecraft, the gyroscopic moment generated by the coupling between the maneuvering angular velocity of the spacecraft and the angular momentum vector of the rotor is transmitted to the spacecraft through the lead screw, and through the self-locking formed by the lead screw <NUM> and the nut <NUM>, no motor locking is needed, and the limitation of the motor locking moment on the maneuvering speed is overcome.

According to the dual-gimbal flywheel of the present application, a displacement of the nut <NUM> is calculated by measuring a rotation angle of the lead screw <NUM> in the spatial parallel mechanism <NUM>, or a real-time displacement of the nut <NUM> is directly measured, a real-time attitude of the movable platform <NUM> in the spatial parallel mechanism <NUM> is calculated by using the displacement of each nut <NUM>, and a real-time deflection angular velocity of the movable platform <NUM> is determined through relative change of the attitude of the movable platform <NUM>. The magnitude and direction of the gyroscopic moment required for attitude control can be obtained by the cross product of the deflection angular momentum vector and the angular velocity vector.

According to the dual-gimbal flywheel of the present application, any continuous yaw of an angular momentum vector in the corresponding area is achieved through the spatial parallel mechanism <NUM> arbitrarily continuous rotation in a local area to change a direction of the angular momentum vector, thereby outputting gyroscopic moment; through speed change of the high-speed motor <NUM>, magnitude of the angular momentum vector is changed, and reaction moment is output; and the gyroscopic moment and the reaction moment are used alone or together as attitude control moment of the spacecraft such as satellites.

The connecting rod in the spatial parallel mechanism <NUM> is a spring damper <NUM>, which forms a passive vibration isolator. The vibration isolation frequency is determined according to the optimization of structural dynamics and rotor dynamics, etc., so as to attenuate the high-frequency micro-vibration generated by the rotor. The present application has different modes at different positions and different attitudes, so the vibration isolation frequency is determined according to the optimization of structural dynamics and rotor dynamics, etc., to attenuate the high-frequency micro-vibration generated by the rotor, thereby improving the ultra-quietness performance of the dual-gimbal flywheel to a certain extent.

The forces and moments such as attitude control moment and micro-vibration are finally transmitted to the spacecraft deck through the platform <NUM>, the first hinge <NUM>, the spring damper <NUM>, the second hinge <NUM>, the nut <NUM>, the lead screw <NUM>, the static base <NUM> and the housing <NUM>, etc., avoiding the transmission of large forces and moments to the low-speed motor <NUM>, which protects the low-speed motor <NUM> and overcomes the locking problem of low-speed motor <NUM>.

Since the spatial parallel mechanism <NUM> of the dual-gimbal flywheel drives the rotor assembly to rotate around its radial direction to output a larger gyroscopic moment, and the axial direction of the rotor assembly outputs a smaller moment through the speed change, by mounting the axis of this product to the ground, the three-axis attitude control of the satellite (offset angular momentum) can be achieved, and the rapid maneuvering of the attitude of the roll axis and pitch axis can be achieved; and while the yaw that requires a small maneuvering speed is to achieve attitude control through the speed change of the rotor. It can be seen that the number of the products is reduced; and the use of two products of this type can achieve the satellite zero momentum three-axis attitude control.

The dual-gimbal flywheel has the advantages of minimum required number of attitude control, light weight, combination of attitude control and micro-vibration control, and no motor locking moment limit, so it has the characteristics of ultra-light, ultra-quiet and ultra-agile. It provides a new possibility for the rapid attitude maneuvering and stability control of the spacecraft, especially for the small and medium-sized spacecraft.

Claim 1:
A low-disturbance dual-gimbal flywheel for spacecraft attitude control, comprising
a spatial parallel mechanism (<NUM>), one single wheel body (<NUM>), one single high-speed motor (<NUM>), a bearing (<NUM>),
and a housing (<NUM>), wherein the wheel body, the high-speed motor and the bearing together form a rotor assembly; and wherein
the spatial parallel mechanism (<NUM>) comprises n drive branch chains, a static base (<NUM>) and a movable platform (<NUM>) having at least two rotational degrees of freedom;
wherein the bearing (<NUM>) is mounted at a center of the wheel body (<NUM>), a rotor of the high-speed motor (<NUM>) is mounted on the bearing (<NUM>); the bearing (<NUM>) and a stator of the high-speed motor (<NUM>) are together mounted on the movable platform (<NUM>), and the wheel body (<NUM>), the stator and rotor of the high-speed motor (<NUM>), and the bearing (<NUM>) are coaxially mounted; the spatial parallel mechanism (<NUM>), the wheel body (<NUM>), the high-speed motor (<NUM>) and bearing (<NUM>) are mounted inside the housing (<NUM>) by means of a port of the static base (<NUM>); a lug outside the housing (<NUM>) is connectable to an outer spacecraft deck; the housing (<NUM>) encloses a maximum moving range of the wheel body (<NUM>); wherein
the spatial parallel mechanism (<NUM>) controls movement of the movable platform (<NUM>) through movement of the n drive branch chains, wherein n is a positive integer equal to or larger than <NUM>;
wherein the static base (<NUM>) comprises a platform and n legs, and the platform is supported by the n legs evenly distributed in a circumferential direction of the platform at a bottom of the platform;
wherein the low-disturbance dual-gimbal flywheel is configured to achieve continuous angular maneuvering of the rotor assembly angular momentum vector in a local area through continuous rotation of the movable platform (<NUM>) in the local area, wherein a direction of the rotor assembly angular momentum vector is changed and a gyroscopic moment is output,
and to change the magnitude of the angular momentum vector through a speed change of the high-speed motor, wherein a reaction moment is output; wherein the gyroscopic moment and the reaction moment can be used separately or together as an attitude control moment of the spacecraft.