CN109189086A - A kind of spacecraft attitude based on magnetic suspension control sensitivity gyro and vibration integrated control method - Google Patents
A kind of spacecraft attitude based on magnetic suspension control sensitivity gyro and vibration integrated control method Download PDFInfo
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Abstract
The present invention relates to a kind of spacecraft attitudes based on magnetic suspension control sensitivity gyro (MSCSG) and vibration integrated control method.Spacecraft During Attitude Maneuver is controlled using the gyroscopic couple of the gyro gimbal rotation output in MSCSG configuration, while inhibiting spacecraft micro-vibration to interfere using micro- frame torque of magnetic suspension rotor deflection output；Efficiently separating for attitude control ring and vibration suppression ring is realized by designing reasonable bandpass filter；Attitude control/vibration suppression integrated controller of spacecraft is recycled to ensure that the stabilization of attitude control ring/vibration suppression ring double closed-loop control system；It is integrated with vibration suppression from the system-level attitude control for realizing spacecraft.The invention belongs to magnetic suspension wipping top control technology field, the spacecraft Vibration Absorbing System Design for using MSCSG as attitude control executing agency can be applied to.
Description
Technical field
The present invention relates to a kind of controls integrated with vibration of spacecraft attitude based on magnetic suspension control sensitivity gyro (MSCSG)
Method processed, suitable for inhibiting design as the micro-vibration of the spacecraft of spacecraft attitude control executing agency using MSCSG.
Technical background
Such as momenttum wheel, the rotation of mechanical control moment gyro, the shaking of refrigerant, solar battery windsurfing iso-deflection on star
Property attachment disengaging shade when the thermal shock that induces of alternating hot and cold quiver can all make spacecraft generate slightly, the micro-vibration of wideband.Micro-vibration
It has become and restricts the bottleneck that spacecraft payload plays a role.
Micro-vibration suppression technology includes component-level and system-level vibration isolation at present, and component-level vibration isolation is mainly for vibration source or has
Load expansion is imitated, passive vibration isolation, Active vibration suppression, semi-active vibration-isolating, active-passive integratedization vibration suppression are roughly divided into.Passively
Vibration isolation is mainly used to inhibit high-frequency vibration, and since the vibration isolation element spring of use is there are resonant frequency, this method exists high
Frequency vibration isolation conflicts with what resonance inhibited, causes vibration isolation precision not high.Active vibration suppression is mainly good to low-frequency vibration inhibitory effect
It is good, but external energy, controller are needed, higher cost.Semi-active vibration-isolating compensates for the defect of passive vibration isolation, but its weight
It is larger, it is poorly suitable for the spacecraft of strict control weight indicator.Active-passive integratedization vibration control vibration isolation frequency band is wider, still
Structure is more complicated.In addition, component-level vibration isolation leads to whole star due to having ignored the coupling between individual vibration source and Rigid Base
Inhibition of vibration is limited, not can guarantee the good dynamic property of spacecraft.System-level vibration suppression usually scrambles vibration as outside one kind
Dynamic, the method by increasing robust control in attitude control law is inhibited.But due to being limited to existing posture control system band
Width, along with the constraint of manipulation rule and attitude control rule, when convergence needed for such control method will meet certain attitude stability
Between it is longer, the precision of gesture stability and the speed of vibration suppression are difficult to reach ideal effect simultaneously.There are also scholars to utilize appearance
It controls executing agency and carries out Disturbance Rejection, Wang Ping et al. is in " a kind of decoupling of rotor system of magnetically suspended control moment gyroscope and disturbance suppression
Method processed " it proposes to carry out small magnitude Disturbance Rejection using magnetic suspension rotor, due to rotor deflection range very little, this method can not be same
Step realizes the large angle maneuver of spacecraft.
MSCSG is a kind of novel inertia actuator, has both had the power output of magnetic suspension control torque gyroscope (MSCMG)
Square function has the function of the angular rate measurement of magnetic suspension sensitivity gyro again.Mainly watched by magnetic suspension rotor system (MSR) and frame
Dress system composition.MSR has the micro- frame torque output capability of 2DOF and 2DOF angular speed sensitive function.Since rotor is inclined
It turn to be generated by Lorentz force magnetic bearing, therefore, micro- frame torque output accuracy ratio MSCMG is higher, simultaneously as MSCSG
Special construction, rotor deflection angle ratio MSCMG is bigger, therefore, micro- frame torque of MSCSG have inhibit spacecraft micro-vibration
Advantage.
Summary of the invention
Technology of the invention solves the problems, such as: inhibiting problem for the micro-vibration of spacecraft, proposes that one kind is not ignoring boat
Under conditions of coupling between its device Rigid Base and vibration source and not increasing additional hardware resources, attitude control executing agency MSCSG sheet is utilized
Body inhibits the micro-vibration of spacecraft, it realizes attitude control/vibration suppression integration of spacecraft from system-level angle, improve space flight
The convergence rate of device attitude control accuracy and vibration suppression.Not only the space application range of MSCSG had been expanded, but also has been the micro- vibration of spacecraft
The completely new technological approaches of the one kind in dynamic inhibition field.
Technical solution of the invention:
Spacecraft During Attitude Maneuver is controlled using the gyroscopic couple of the gyro gimbal rotation output in MSCSG configuration, is utilized
Micro- frame torque of magnetic suspension rotor deflection output in MSCSG configuration inhibits spacecraft micro-vibration, specifically includes the following steps:
(1) consider the in-orbit single MSCSG Dynamic Modeling under the conditions of rotor deflection
In-orbit single MSCSG angular momentumAnd its change rateIt may be expressed as:
Wherein, h represents the angular momentum of gyro, and δ represents frame around gyro coordinate system OXCSGYCSGZCSGMiddle OXCSGRotation angle,
α represents magnetic suspension rotor around stator coordinate OXfYfZfMiddle OXfDeflection angle；In order to avoid gyro gimbal and rotor rotation are mutual
Coupling, selects the deflection direction of rotor consistent with frame direction of rotation；
(2) consider the in-orbit MSCSG pyramid configuration Dynamic Modeling under the conditions of rotor deflection
By taking pyramid configuration as an example, the angular momentum of in-orbit pyramid configuration MSCSGAnd its change rateIt can indicate
For；
In formula, h represents the angular momentum of single gyro, and σ represents pyramid apex angle, δ1, δ2, δ3, δ4Represent pyramid configuration
The deflection angle of 4 gyro gimbals,The yaw rate of 4 gyro gimbals of pyramid configuration is represented,
α1, α2, α3, α4The rotor deflection angle of 4 gyros of pyramid configuration is represented,Represent pyramid configuration
The rotor deflection angular speed of 4 gyros；
(3) spacecraft attitude control/vibration suppression integrated controller
In the micro-vibration T being subject to using adaptive notch filter to spacecraftdOn the basis of being recognized, it is based on Reverse Step Control
Thought design integration controller:
Wherein, a, b, k1, k2It is normal number, G is the estimation parameter matrix of spacecraft rotary inertia parameter matrix J, qev=
[qev1 qev2 qev3]TRepresent the vector of rear three compositions of spacecraft attitude quaternary number error, ωib=[ωib1 ωib2
ωib3]TRepresent spacecraft attitude angular speed, ωh=[ωh1 ωh2 ωh3]TThe high fdrequency component of spacecraft attitude angular speed is represented,
ωl=[ωl1 ωl2 ωl3]TRepresent the low frequency component of spacecraft attitude angular speed；Y1, Y2It respectively indicates are as follows:
(4) Design of Bandpass
Since spacecraft attitude signal bandwidth is lower, usually in 0.1Hz hereinafter, micro-vibration signal bandwidth is higher, usually exist
20Hz or more, in order to realize attitude control signal and vibration suppression signal efficiently separate and frame corners and rotor deflection angle command signal
Separation, the cutoff frequency of bandpass filter is set as 0.1Hz, and low-pass filter selects inertial element: 1/ (10s+1), high pass
Filter selects differentiation element combination inertial element: s/ (s+0.1).Oscillation element, it is ensured that low frequency signals decay is larger, but
It is that can not retain original signal, so high-pass filter does not select oscillation element since the gain in high frequency treatment is excessive.
Inventive principle of the invention is: MSCSG magnetic suspension rotor can do small size, high frequency radial inclined in magnetic bearing
Turn, export micro- frame torque of high bandwidth, this just provides possibility for the micro-vibration inhibition of spacecraft.Using in MSCSG configuration
Gyro gimbal rotation generate gyroscopic couple carry out spacecraft attitude control, meanwhile, utilize magnetic suspension rotor deflection generate micro- frame
Booster square carries out micro-vibration inhibition.
Specific control system functional block diagram is as shown in Figure 1.The spacecraft attitude control of outer ring/vibration suppression integrated controller output
It instructs torque to manipulate to pseudoinverse to restrain, manipulation rule calculates the frame angular speed and rotor deflection of each gyro according to input instruction
The sum of angular speed, since frame corners rate frequency is lower, rotor deflection angular speed frequency is higher, is had by bandpass filter
Effect separation, exports respectively to frame and rotor as reference signal, the actual frame of MSCSG pyramid 4 gyros of Structure Configuration Synthesis
Angle and rotor deflection angle, output gyroscopic couple act on spacecraft, and the attitude angular rate of spacecraft passes through bandpass filter point
It separates out low-and high-frequency signal and feeds back to attitude control/vibration suppression integrated controller, realize two close cycles synchronously control.
Firstly, to determine single in-orbit MSCSG kinetic model: OXcsgYcsgZcsg、OXgYgZg、OXfYfZf、OXrYrZrPoint
It is not gyro coordinate system, frame coordinates system, the stator coordinate, rotor coordinate of MSCSG, rotor is around OZfIt is done with angular velocity vector Ω
High speed rotation, meanwhile, it can be around OXf、OYfRadial deflection is done, yaw rate isIn order to avoid frame rotates and turns
It intercouples between son deflection, definition rotor deflection direction is identical as frame direction of rotation, i.e., rotor is only around OXfIt does radial inclined
Turn.Its rotor deflection schematic diagram is as shown in Figure 2.
According to euler dynamical equations, bonding force square suffered by rotor is indicated are as follows:
In formula,The angular speed in rotor relative inertness space is represented,The angular momentum in rotor relative inertness space is represented,
Expression formula are as follows:
In formula, Ir、IzRespectively represent the radial and axial rotary inertia of magnetic suspension rotor.
In formula,The projection of the angular speed in gyro coordinate system in frame relative inertness space is represented,Represent rotor phase
To the angular speed of stator stator coordinate projection,Respectively represent stator coordinate to rotor coordinate,
Frame coordinates system is to stator coordinate, the coordinate conversion matrix of gyro coordinate system to frame coordinates system, their expression formula difference
Are as follows:
According to IzΩ > > Ir, bonding force square expression formula suffered by rotor:
In formula, h=IzΩ。
Gyrodynamics equation meets following relationship:
Obtain the angular momentum of the single MSCSG under the conditions of rotor deflectionAnd its change rate
Wherein, h represents the angular momentum of gyro, and δ represents frame around gyro coordinate system OXCSGYCSGZCSGMiddle OXCSGRotation angle,
α represents magnetic suspension rotor around stator coordinate OXfYfZfMiddle OXfDeflection angle；In order to avoid gyro gimbal and rotor rotation are mutual
Coupling, selects the deflection direction of rotor consistent with frame direction of rotation；
The present invention solves the angular momentum and its change rate of gyro group by taking pyramid configuration as an example:
In formula,Represent the coordinate conversion matrix that n-th of gyro (n=1,2,3,4) arrives spacecraft this system:
According to formula (11), (12), (13), the in-orbit MSCSG pyramid configuration angular momentum under the conditions of rotor deflection
And its change rateIt indicates are as follows:
In formula, h represents the angular momentum of single gyro, and σ represents pyramid apex angle, δ1, δ2, δ3, δ4Represent pyramid configuration
The deflection angle of 4 gyro gimbals,The yaw rate of 4 gyro gimbals of pyramid configuration is represented,
α1, α2, α3, α4The rotor deflection angle of 4 gyros of pyramid configuration is represented,Represent pyramid configuration
The rotor deflection angular speed of 4 gyros；
After obtaining the kinetic model of in-orbit MSCSG pyramid configuration, in next step, outer ring integrated controller is solved.
Spacecraft dynamics equation based on MSCSG:
In formula, TdThe micro-vibration interference that spacecraft is subject to is represented, can be recognized by adaptive notch filter；J represents space flight
Device moment of inertia matrix can not be accurately obtained:
According to angular momentum conservation law:
For spacecraft when by micro-vibration disturbed condition, angular speed includes to be less than 0.1Hz low frequency posture ωlBe greater than
The disturbance ω of 0.1HzhTwo parts are further derived according to above formula (15), (16):
Wherein, ulMSCSG pyramid configuration is represented by the low frequency torque less than 0.1Hz of frame rotation output, for navigating
Its device attitude control；uhMicro- frame torque greater than 0.1Hz that MSCSG pyramid configuration is exported by rotor deflection is represented, space flight is used for
Device vibration suppression.Above formula can be analyzed to following two parts:
Controller design target is to guarantee spacecraft attitude quaternary number error qev=[0；0；0], Lyapunov function is designed:
Wherein, qev=[qev1 qev2 qev3]TThe vector of rear three compositions in spacecraft attitude quaternary number error is represented,
qe0Represent the first item of spacecraft attitude quaternary number error.
Its derivation is obtained:
In order to makeDesign linear virtual controlling rule:
In formula, a, b are normal numbers.
Design system-wide Lyapunov function:
In formula, η=[J11 J22 J33 J12 J13 J23]T, since it cannot be accurately obtained, takeEstimation parameter as η
Vector, estimation parameter matrix of the G as J, in order to allowIntegrated controller design are as follows:
Wherein, a, b, k1, k2It is normal number, G is the estimation parameter matrix of spacecraft rotary inertia parameter matrix J, qev=
[qev1 qev2 qev3]TRepresent the vector of rear three compositions of spacecraft attitude quaternary number error, ωib=[ωib1 ωib2
ωib3]TRepresent spacecraft attitude angular speed, ωh=[ωh1 ωh2 ωh3]TThe high fdrequency component of spacecraft attitude angular speed is represented,
ωl=[ωl1 ωl2 ωl3]TRepresent the low frequency component of spacecraft attitude angular speed；Y1, Y2It respectively indicates are as follows:
Since spacecraft attitude signal bandwidth is lower, usually in 0.1Hz hereinafter, micro-vibration signal bandwidth is higher, usually exist
20Hz or more, in order to realize attitude control signal and vibration suppression signal efficiently separate and frame corners and rotor deflection angle command signal
Separation, the cutoff frequency of bandpass filter is set as 0.1Hz, and low-pass filter selects inertial element: 1/ (10s+1), high pass
Filter selects differentiation element combination inertial element: s/ (s+0.1).
Other than pyramid configuration, common pentagonal pyramid configuration can also be realized using similar method based on magnetcisuspension
The floating spacecraft attitude control for controlling sensitive gyro is integrated with vibration suppression.
So far, the spacecraft attitude control based on MSCSG/vibration suppression two close cycles synchronous control system is realized.
The solution of the present invention and existing scheme ratio, major advantage are: having expanded the space application range of MSCSG, not only
It is confined to the gesture stability of spacecraft, micro-vibration inhibition can also be carried out.One kind is provided for spacecraft micro-vibration suppression technology
Completely new means save hardware resource, while vibration suppression frequency band is wider, effectively realizes compared with existing micro-vibration suppression technology
Spacecraft high-precision attitude control is integrated with vibration suppression.
Detailed description of the invention
Spacecraft attitude control of the Fig. 1 based on MSCSG/vibration suppression integral control system functional block diagram.
Fig. 2 magnetic suspension rotor deflects schematic diagram.
Fig. 3 specific embodiment figure.
Spacecraft attitude angle and angular speed analogous diagram under Fig. 4 integrated control method.
Specific embodiment
Objective for implementation of the invention is the spacecraft based on MSCSG as attitude control executing agency, and magnetic suspension rotor is radially inclined
The micro- frame torque of high-precision, the high bandwidth transferred out makes it possible that the micro-vibration of spacecraft inhibits.
Specific embodiments of the present invention are as shown in figure 3, specific implementation step is as follows:
(1) consider the in-orbit single MSCSG Dynamic Modeling under the conditions of rotor deflection
In-orbit single MSCSG angular momentum and its change rate may be expressed as:
Wherein, h represents the angular momentum of gyro, and δ represents frame around gyro coordinate system OXCSGYCSGZCSGMiddle OXCSGRotation angle,
α represents magnetic suspension rotor around stator coordinate OXfYfZfMiddle OXfDeflection angle；In order to avoid gyro gimbal and rotor rotation are mutual
Coupling, selects the deflection direction of rotor consistent with frame direction of rotation；
(2) consider the in-orbit MSCSG pyramid configuration Dynamic Modeling under the conditions of rotor deflection
Pyramid configuration and pentagonal pyramid configuration, this patent is widely used using the spacecraft of MSCSG as attitude control executing agency
Mainly pyramid configuration MSCSG is studied, which can be extended to the boat based on MSCSG arbitrary configuration
In its device.
In formula, h represents the angular momentum of single gyro, and σ represents pyramid apex angle, δ1, δ2, δ3, δ4Represent pyramid configuration
The deflection angle of 4 gyro gimbals,The yaw rate of 4 gyro gimbals of pyramid configuration is represented,
α1, α2, α3, α4The rotor deflection angle of 4 gyros of pyramid configuration is represented,Represent pyramid configuration
The rotor deflection angular speed of 4 gyros；
(3) spacecraft attitude control/vibration suppression integrated controller
In the micro-vibration T being subject to using adaptive notch filter to spacecraftdOn the basis of being recognized, it is based on
Backstepping Reverse Step Control thought design integration controller:
Wherein, a, b, k1, k2It is normal number, G is the estimation parameter matrix of spacecraft rotary inertia parameter matrix J, qev=
[qev1 qev2 qev3]TRepresent the vector of rear three compositions of spacecraft attitude quaternary number error, ωib=[ωib1 ωib2
ωib3]TRepresent spacecraft attitude angular speed, ωh=[ωh1 ωh2 ωh3]TThe high fdrequency component of spacecraft attitude angular speed is represented,
ωl=[ωl1 ωl2 ωl3]TRepresent the low frequency component of spacecraft attitude angular speed；Y1, Y2It respectively indicates are as follows:
(4) Design of Bandpass
Since spacecraft attitude signal bandwidth is lower, usually in 0.1Hz hereinafter, micro-vibration signal bandwidth is higher, usually exist
20Hz or more, in order to realize attitude control signal and vibration suppression signal efficiently separate and frame corners and rotor deflection angle command signal
Separation, the cutoff frequency of bandpass filter is set as 0.1Hz, and low-pass filter selects inertial element: 1/ (10s+1), high pass
Filter selects differentiation element combination inertial element: s/ (s+0.1).
Bandpass filter is mainly used for two places: one be manipulation rate output.Manipulation rate is calculated according to input instruction
The sum of the frame angular speed of each gyro and rotor angular speed out are efficiently separated out by bandpass filter less than 0.1Hz's
The frame corners rate instruction signal of low frequency and the rotor angle rate instruction signal for being greater than 0.1Hz.The other is spacecraft attitude angle
At speed measurement, bandpass filter isolates low-and high-frequency signal and feeds back to outer ring attitude control/vibration suppression integrated controller.
For the effect for verifying the measurement method, simulation analysis, examination are carried out using overall-in-one control schema method proposed by the invention
Test result such as Fig. 4 (a) -4 (f).Abscissa indicates time, unit s in Fig. 4；4 (a) -4 (c) ordinates indicate angle, single
Position for °, 4 (d) -4 (f) ordinates indicate angular speed, unit be °/s.
It should be noted that current error, displacement error, tachometric survey that this emulation considers magnetic suspension rotor system miss
The many factors such as the installation error in difference and gyro configuration.
Wherein, blue line represents the spacecraft obtained using conventional method (rotate using gyro gimbal and carry out spacecraft attitude control)
Attitude angle and angular speed curve, red line represent the spacecraft attitude angle obtained using integral method and angular speed curve.
It can be seen that attitude control proposed by the present invention/vibration suppression integrated control method by above-mentioned the simulation experiment result, protecting
In the case where card and the consistent attitude control convergence rate of conventional method, vibration suppression precision and system stability greatly improved.Through
It calculates, attitude stability is by 10-5It is increased to 10-6.This, which illustrates method proposed by the present invention well, realizes the appearance of spacecraft
Control/vibration suppression integration, and calculating realization is simpler, engineering is strong.
The content being not described in detail in present specification belongs to the prior art well known to professional and technical personnel in the field.
Claims (2)
1. a kind of spacecraft attitude based on magnetic suspension control sensitivity gyro and vibration integrated control method, it is characterised in that:
Spacecraft During Attitude Maneuver is controlled using the gyroscopic couple of the gyro gimbal rotation output in magnetic suspension control sensitivity gyro configuration, together
When, micro- frame torque using the magnetic suspension rotor deflection output in magnetic suspension control sensitivity gyro configuration inhibits the micro- vibration of spacecraft
It is dynamic, specifically includes the following steps:
(1) consider that the in-orbit single magnetic suspension control sensitivity gyrodynamics under the conditions of rotor deflection models in-orbit single magnetic suspension
Control sensitive gyro angular momentumAnd its change rateIt may be expressed as:
Wherein, h represents the angular momentum of gyro, and δ represents frame around gyro coordinate system OXCSGYCSGZCSGMiddle OXCSGRotation angle, α generation
Table magnetic suspension rotor is around stator coordinate OXfYfZfMiddle OXfDeflection angle；In order to avoid gyro gimbal and rotor rotate phase mutual coupling
It closes, selects the deflection direction of rotor consistent with frame direction of rotation；
(2) consider the in-orbit magnetic suspension control sensitivity gyro pyramid configuration Dynamic Modeling under the conditions of rotor deflection
By taking pyramid configuration as an example, the angular momentum of in-orbit pyramid configuration magnetic suspension control sensitivity gyroAnd its change rateIt is represented by；
In formula, h represents the angular momentum of single gyro, and σ represents pyramid apex angle, δ1, δ2, δ3, δ4Represent 4 tops of pyramid configuration
The deflection angle of spiral shell frame,Represent the yaw rate of 4 gyro gimbals of pyramid configuration, α1, α2,
α3, α4The rotor deflection angle of 4 gyros of pyramid configuration is represented,Represent 4 tops of pyramid configuration
The rotor deflection angular speed of spiral shell；
(3) spacecraft attitude control/vibration suppression integrated controller
In the micro-vibration T being subject to using adaptive notch filter to spacecraftdOn the basis of being recognized, it is based on Reverse Step Control thought
Design integration controller:
Wherein, a, b, k1, k2It is normal number, G is the estimation parameter matrix of spacecraft rotary inertia parameter matrix J, qev=[qev1
qev2 qev3]TRepresent the vector of rear three compositions of spacecraft attitude quaternary number error, ωib=[ωib1 ωib2 ωib3]TGeneration
Table spacecraft attitude angular speed, ωh=[ωh1 ωh2 ωh3]TRepresent the high fdrequency component of spacecraft attitude angular speed, ωl=
[ωl1 ωl2 ωl3]TRepresent the low frequency component of spacecraft attitude angular speed；Y1, Y2It respectively indicates are as follows:
(4) Design of Bandpass
Since spacecraft attitude signal bandwidth is lower, usually in 0.1Hz hereinafter, micro-vibration signal bandwidth is higher, usually in 20Hz
More than, in order to realize attitude control signal and vibration suppression signal efficiently separate and point of frame corners and rotor deflection angle command signal
From the cutoff frequency of bandpass filter is set as 0.1Hz, and low-pass filter selects inertial element: 1/ (10s+1), high-pass filtering
Device selects differentiation element combination inertial element: s/ (s+0.1).
2. a kind of spacecraft attitude based on magnetic suspension control sensitivity gyro according to claim 1 control integrated with vibration
Method processed, it is characterised in that: this method can not only be applied to pyramid configuration, can also be applied to pentagonal pyramid configuration.
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| CN110502024A (en) * | 2019-07-23 | 2019-11-26 | 北京控制工程研究所 | A kind of universal posture executing agency of standard based on space parallel mechanism |
| CN111775142A (en) * | 2020-08-12 | 2020-10-16 | 电子科技大学 | Model identification and self-adaptive control method for hydraulic mechanical arm |
| CN112231861A (en) * | 2020-10-16 | 2021-01-15 | 哈尔滨工业大学 | Cluster control moment gyroscope vibration isolation method for inhibiting attitude-adjusting resonance |
| CN112947614A (en) * | 2021-01-28 | 2021-06-11 | 哈尔滨工业大学 | Active vibration control method of variable speed tilting momentum wheel |
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| CN110502024B (en) * | 2019-07-23 | 2020-10-20 | 北京控制工程研究所 | Quasi-universal attitude executing mechanism based on space parallel mechanism |
| CN111775142A (en) * | 2020-08-12 | 2020-10-16 | 电子科技大学 | Model identification and self-adaptive control method for hydraulic mechanical arm |
| CN112231861A (en) * | 2020-10-16 | 2021-01-15 | 哈尔滨工业大学 | Cluster control moment gyroscope vibration isolation method for inhibiting attitude-adjusting resonance |
| CN112947614A (en) * | 2021-01-28 | 2021-06-11 | 哈尔滨工业大学 | Active vibration control method of variable speed tilting momentum wheel |
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