CN102566578A - Singular value decomposition-based coordination control method of single gimbal control moment gyros (SGCMGs) - Google Patents

Abstract

The invention relates to a singular value decomposition-based coordination control method of single gimbal control moment gyros (SGCMGs). The singular value decomposition-based coordination control method comprises the following steps of: distributing instruction moment required for controlling an entire spacecraft to two sets of SGCMGs according to a certain proportion; decomposing the instruction moment distributed to the set A of SGCMGs again by using a singular value decomposition method; distributing an instruction moment component in the singular direction of the set A of SGCMGs to the set B of SGCMGs; still distributing an instruction moment component perpendicular to the singular direction of the set A of SGCMGs to the set A of SGCMGs; after distribution, solving instruction gimbal angular velocities of the two sets of SGCMGs respectively; operating according to the respective instruction gimbal angular velocity; and acting the sum of output moment on the spacecraft to finish accurate attitude control. According to the singular value decomposition-based coordination control method, the gyros can still accurately and effectively output control moment to control the attitude of the spacecraft while a part of gyros fails and the gyros are singular and the probability of premature saturation of a single set of SGCMGs is also avoided to the greatest extent without configuring an extra execution mechanism.

Description

Single frame control-moment gyro crowd control method for coordinating based on svd

Technical field

The present invention relates to a kind of attitude control method of spacecraft, be specifically related to a kind of single frame control-moment gyro crowd's control method.

Background technology

Along with the development of aerospace industry, modern spacecraft to precision, life-span and the reliability of attitude control system require increasingly high.Spacecraft mainly is to realize through topworks's output control moment in the control of rail attitude.

The attitude control actuator of spacecraft employing at present mainly contains jet thrust device, angular momentum switch, magnetic torquer etc.Wherein the angular momentum switch has can provide continuous attitude control moment, non-consume fuel, do not pollute optical device and flight environment of vehicle, the advantages such as vibration of not easy excitated spacecraft flexible appendage, thereby as the main actuating mechanism of spacecraft attitude control system and be widely used in high precision, long-life spacecraft.

The principle of work of angular momentum switch is based upon on the basis of the conservation of angular momentum; When the big or small or direction of its angular momentum changes according to certain rules; Act on the spacecraft body producing the continuous moment of reaction, thereby reach the purpose of controlling spacecraft attitude.In all kinds of angular momentum switches; Single frame control-moment gyro crowd (Single Gimbal Control Moment Gyros; SGCMGs) can not only export big amplitude control moment; Also has simple in structure, advantage such as reliability is high, faster system response, control are more accurate; Become the first-selected attitude control actuator of the actual medium-and-large-sized long-life spacecraft of engineering, all adopted SGCMGs to control main actuating mechanism as attitude like the large-scale space telescope (LST) of the U.S. and the Mir space station (MIR) of USSR's emission.China starts late about the research of CMGs, and Beijing Control Engineering Inst. began to develop mechanical bearing SGCMGs in 1999, and is successfully applied to target aircraft of Heavenly Palace of in September, 2011 emission first.

When utilization SGCMGs carries out attitude control to spacecraft, need at first design the manipulation rule of SGCMGs, confirm gyro gimbal angular velocity by instruction control moment, make the gyro output torque consistent with the instruction moment of spacecraft attitude control system requirement.Yet the intrinsic configuration singular problem of SGCMGs is but given to handle to restrain to design and has been brought very big difficulty.The configuration of SGCMGs is unusual to be meant when being in some frame corners combination; The output torque vector coplane of each gyro; And make that in the direction perpendicular to this plane be the moment that requirement can't be provided on the unusual direction; Particularly when having the part gyro to lose efficacy among the SGCMGs, the quantity that makes up corresponding to unusual frame corners can sharply increase, and makes singular problem more serious.Though many scholars have carried out big quantity research to this, still there are some problems in the manipulation that is designed rule, handle rule like zero motion and can't avoid showing singular point, and at the SGCMGs configuration when unusual, the frame corners velocity solution is excessive even do not have and separate; Robust pseudoinverse and broad sense robust pseudoinverse are handled rule all can introduce the moment error, and attitude control accuracy is descended.

On the other hand, the fit spacecraft of at present existing in the world large-scale groups mostly adopts the structure of many cabins section, and its attitude control actuator includes two cover pentagonal pyramid configuration SGCMGs at least, is installed on one of application cabin of core cabin and butt joint respectively.In traditional controlling schemes, core cabin SGCMGs is generally used for the attitude control of the whole assembly in independent core cabin and butt joint back, and uses the attitude control of using the cabin before cabin SGCMGs only is used to dock.The maximum problem of this scheme is, when only utilizing core cabin SGCMGs to carry out assembly control, if the part gyro breaks down, the then existing rule of handling can't guarantee that SGCMGs can realize the unusual accurate output of avoiding fully with moment simultaneously.

The present invention to this difficult point problem, proposes a kind of control method for coordinating based on svd that is applied to SGCMGs just, and being intended to provides technical support for the domestic Large Spacecraft attitude control task with future now.

Summary of the invention

The objective of the invention is to spacecraft with two cover pentagonal pyramid configuration SGCMGs controls; A kind of SGCMGs control method for coordinating is proposed; Guarantee when the part gyro lost efficacy and gyro when unusual, still can make gyro export the attitude of control moment accurately and efficiently with the control spacecraft.

The invention provides a kind of single frame control-moment gyro crowd control method for coordinating based on svd; Have two cover pentagonal pyramid configuration SGCMGs at spacecraft; And wherein the part gyro (comprising 1,2 or 3) of A cover lost efficacy; Under the situation of B cover operate as normal, can be suitable for method of the present invention.

Method of the present invention may further comprise the steps:

Step 1, will control the required instruction moment of whole spacecraft distribute to by a certain percentage two the cover SGCMGs;

Step 2, the method for utilizing svd are decomposed the instruction moment of distributing to A cover SGCMGs once more; With wherein distributing to B cover SGCMGs, overlap SGCMGs and still distribute to A perpendicular to the instruction moment components of the unusual direction of A cover SGCMGs along the instruction moment components of the unusual direction of A cover SGCMGs;

Step 3, assigned after, A cover SGCMGs utilizes pseudoinverse to handle rule and solves its instruction frame corners speed, B cover SGCMGs utilizes pseudoinverse to add zero motion and handles rule and solve its instruction frame corners speed;

Step 4, two cover SGCMGs press instruction frame corners speed running separately respectively, and the output torque sum acts on spacecraft, accomplish accurate attitude control.

Beneficial effect

Need not to dispose under the situation of extra topworks; The inventive method makes full use of the control ability of two cover SGCMGs; Through the coordination control of two cover SGCMGs, well solved insurmountable problem when utilizing a cover SGCMGs to carry out spacecraft attitude control separately, guarantee when the part gyro lost efficacy and gyro when unusual; Still can make gyro export the attitude of control moment accurately and efficiently, also avoid single cover SGCMGs saturated possibility too early to the full extent with the control spacecraft.

Description of drawings

Fig. 1 is the structural representation of single frame control-moment gyro (SGCMG).

Fig. 2 is the configuration synoptic diagram of two cover SGCMGs.

Fig. 3 is two cover SGCMGs control method for coordinating schematic diagrams based on svd.

Fig. 4 is the assembly spacecraft attitude control system based on two cover SGCMGs.

Fig. 5 is the unusual tolerance of A cover SGCMGs figure as a result.

Fig. 6 is the actual frame angular velocity of A cover SGCMGs figure as a result.

Fig. 7 is the error result figure of the actual output torque and instruction control moment of two cover SGCMGs.

Embodiment

Below in conjunction with accompanying drawing, specify preferred implementation of the present invention.

Be the clearer present embodiment of introducing, at first the principle of simple declaration SGCMG output torque combines the enforcement of SGCMGs explanation this method of two cover pentagonal pyramid configurations again.It is emphasized that this method only needs the SGCMGs of two cover pentagonal pyramid configurations, and and do not rely on concrete mounting means.

Referring to Fig. 1; SGCMG is made up of the rotor of a constant speed rotation and the framework of support rotor;

is rotor spin axis direction;

is the gimbal axis rotary speed direction,

in the opposite direction with the output control moment.Rotor spin axis and gimbal axis quadrature are installed, and are driven by rotor electric machine and frame motor respectively.Rotor electric machine drives rotor around the rotation of spin axis constant speed, produces a constant angle momentum.Frame motor makes framework turn over frame corners δ around the gimbal axis that is fixed on the spacecraft body with angular velocity

according to steering order.Because the rotation of gimbal axis causes rotor spin axis direction to change, and the angular momentum of rotor is changed, thereby export a gyroscopic couple.For single SGCMG,, can obtain its control moment of exporting and do according to the principle of work of above introduction

$\stackrel{\&RightArrow;}{T}=-\left(\stackrel{\&CenterDot;}{\mathrm{\&delta;}}\stackrel{\&RightArrow;}{g}\right)\&times;\left(h_0\stackrel{\&RightArrow;}{s}\right)=-{\mathrm{<figure-callout\; id="0"\; label="h"\; filenames="HDA0000130508730000032.png"\; state="\{\{state\}\}">h</figure-callout>}}_0\stackrel{\&CenterDot;}{\mathrm{\&delta;}}\stackrel{\&RightArrow;}{t}---\left(1\right)$

Wherein,

The moment vector of representing single SGCMG output, h ₀Nominal angular momentum for gyrorotor.

The gyroscopic couple that single SGCMG produces and in order spacecraft to be carried out three controls, generally need be no less than 3 SGCMG and form gyro groups only on the plane vertical with its gimbal axis, through the direction and the size of different frames angle array mode adjustment output torque.For the SGCMGs system that is made up of a plurality of gyros, simple for making steering logic, the single gyro in the gyro group is in quality, and equal values is got in parameter aspects such as rotor speed and moment of inertia, therefore single the angular momentum amplitude h that gyro provides ₀All identical.If gyro group is made up of n gyro, the gimbal axis direction unit vector of i gyro for

rotor angular momentum direction unit vector for

gyro output torque in the other direction unit vector for

then the total angular momentum of gyro group can be expressed as

${\stackrel{\&RightArrow;}{h}}_c={\mathrm{<figure-callout\; id="0"\; label="h"\; filenames="HDA0000130508730000032.png"\; state="\{\{state\}\}">h</figure-callout>}}_0\underset{i=1}{\overset{n}{\mathrm{\&Sigma;}}}{\stackrel{\&RightArrow;}{s}}_i---\left(2\right)$

Wherein,

Be the total angular momentum vector of system, it is write as at spacecraft body coordinate system f _b(o _bx _by _bz _b) under component array form do

h _c＝A _sh ₀ (3)

Wherein, A _s=[s ₁s ₂S _n], s _iAngular momentum direction unit vector for i SGCMG of correspondence

At f _bIn the component array.

In like manner, according to formula (1), the resultant couple vector that can obtain gyro group output does

${\stackrel{\&RightArrow;}{T}}_c={-\mathrm{<figure-callout\; id="0"\; label="h"\; filenames="HDA0000130508730000032.png"\; state="\{\{state\}\}">h</figure-callout>}}_0\underset{i=1}{\overset{n}{\mathrm{\&Sigma;}}}{\stackrel{\&CenterDot;}{\mathrm{\&delta;}}}_i\stackrel{\&RightArrow;}{t_i}---\left(4\right)$

Wherein,

is the resultant moment vector that n gyro produces, and is the frame corners speed of i gyro.It is write as at f _b(o _bx _by _bz _b) under component array form do

$T_c={-h}_0{A}_t\stackrel{\&CenterDot;}{\mathrm{\&delta;}}---\left(5\right)$

Wherein, A _t=[t ₁t ₂T _n], t _iBe the output torque of i SGCMG of correspondence unit vector in the other direction

At f _bIn the component array, $\stackrel{\&CenterDot;}{\mathrm{\&delta;}}={\left[\begin{array}{cccc}{\stackrel{\&CenterDot;}{\mathrm{\&delta;}}}_1& {\stackrel{\&CenterDot;}{\mathrm{\&delta;}}}_2& \&CenterDot;\&CenterDot;\&CenterDot;& {\stackrel{\&CenterDot;}{\mathrm{\&delta;}}}_n\end{array}\right]}^T$ Be frame corners speed column vector.

In formula (3) and formula (5), A _sAnd A _tBe variable, change, can be written as with gyro gimbal angle δ

A _s＝A _s0d[cosδ]+A _t0d[sinδ] (6)

A _t＝A _t0d[cosδ]-A _s0d[sinδ] (7)

In the formula, A _S0And A _T0Be respectively A _sAnd A _tInitial value, cos δ=[cos δ ₁Cos δ ₂... cos δ _n] ^T, sin δ=[sin δ ₁Sin δ ₂... sin δ _n] ^TTo any n-dimensional vector x=[x ₁x ₂X _n] ^T, operator d [x] is defined as following diagonal matrix

d[x]＝diag(x ₁?x ₂?…?x _n) (8)

Angular momentum equation (3) and the momental equation (5) of the SGCMGs of n gyro composition have below been obtained respectively.

Referring to Fig. 2, the pentagonal pyramid configuration is made up of 6 SGCMG, and they are mounted respectively on 6 adjacent sides of regular dodecahedron, and the angle of two adjacent surfaces is 116.51 ° arbitrarily, and the gimbal axis of each gyro is symmetrically distributed, respectively perpendicular to its side, place.Consider to have the SGCMGs of A and B two cover pentagonal pyramid configurations to be installed in the spacecraft; Supposing has the part gyro to lose efficacy among the A cover SGCMGs; The number of the gyro of residue operate as normal is n (3≤n≤6), the then complete operate as normal of B cover SGCMGs, and the spacecraft body coordinate system still is designated as f _b(o _bx _by _bz _b).Promptly be practical implementation step of the present invention below.

The first step: will control the required instruction moment of whole spacecraft and distribute to two cover SGCMGs by a certain percentage.Preferred scheme is to carry out moment according to the minimum envelop angular momentum size of two cover SGCMGs and distribute.Suppose that the total instruction control moment that obtains from the attitude controller that is designed is T _c, from avoiding two angles that gyro is saturated, can at first distribute total instruction control moment according to the minimum envelop angular momentum size of two cover SGCMGs, wherein, the instruction control moment of distributing to A cover inefficacy SGCMGs is T _Ca, be expressed as

$T_{\mathrm{ca}}=\frac{h_{\mathrm{an}}}{h_{\mathrm{an}}+4.4719}{T}_c---\left(9\right)$

The instruction control moment of distributing to B cover SGCMGs is T _Cb, be expressed as

$T_{\mathrm{cb}}=\frac{4.4719}{h_{\mathrm{an}}+4.4719}{T}_c---\left(10\right)$

Wherein, h _AnEqual the smallest angles momentum of envelope of the individual SGCMGs of n (3≤n≤6) of A cover operate as normal; Can utilize numerical method from formula (3), to calculate (reference: Zhang Renwei; " satellite orbit and attitude dynamics and control ", publishing house of BJ University of Aeronautics & Astronautics, 281-285).

Second step is based on the theoretical moment reallocation of svd.The theory of utilizing svd is to T _CaCarry out sub-distribution again, obtain moment components T respectively perpendicular to the unusual direction of A cover SGCMGs _Ca1With moment components T along the unusual direction of A cover SGCMGs _Ca2For specifying the implementation process of present embodiment, will specify how to obtain T below _Ca1And T _Ca2Expression formula.

Referring to Fig. 3; is the unit vector of the unusual direction of A cover SGCMGs; Make

(i=1; 2; ..., n) the output torque opposite direction unit vector of remaining n normal gyro among the unusual moment A cover of (3≤n≤6) expression SGCMGs.For to

sub-distribution again, confirm that unusual direction is crucial.Output torque matrix of coefficients C to inefficacy SGCMGs _aCarry out svd, obtain

C _a＝h ₀A _at＝USV ^T (11)

In the formula, A _At=A _At0D [cos δ _a]-A _As0D [sin δ _a], U ∈ R ^{3 * 3}, V ∈ R ^{N * n}, be unitary matrix.S ∈ R ^{3 * n}Can be written as following form

S＝[S ₁?0 _3×(n-3)] (12)

In the formula

S ₁＝diag(σ ₁?σ ₂?σ ₃) (13)

σ wherein ₁, σ ₂And σ ₃Be C _aSingular value, and satisfy σ ₁>=σ ₂>=σ ₃V also can be written as

V＝[V ₁?V ₂] (14)

V in the formula ₁∈ R ^{N * 3}, V ₂∈ R ^{N * (n-3)}U and V ₁Can be expressed as U=[U by its column vector ₁U ₂U ₃], V ₁=[V ₁₁V ₁₂V ₁₃], then the output torque equation to inefficacy SGCMGs can be written as

$C_a{\stackrel{\&CenterDot;}{\mathrm{\&delta;}}}_a={\mathrm{US}}_1{V}_1^T{\stackrel{\&CenterDot;}{\mathrm{\&delta;}}}_a=\underset{i=1}{\overset{3}{\mathrm{\&Sigma;}}}{\mathrm{\&sigma;}}_i{U}_i{V}_{1i}^T{\stackrel{\&CenterDot;}{\mathrm{\&delta;}}}_a=-T_{\mathrm{ca}1}---\left(15\right)$

T in the formula _Ca1Be the instruction moment of distributing to inefficacy SGCMGs perpendicular to unusual direction.When SGCMGs is absorbed in when unusual σ fully ₃=0, can know U this moment ₃The direction output torque is zero, and U also is described ₃Be the unusual direction of SGCMGs.And when SGCMGs when unusual, σ ₃Also approach zero, at this moment U ₃Also, can be described as accurate unusual direction just near unusual direction.In case confirmed unusual direction U ₃, just can be with SGCMGs near unusual time instruction moment at U ₃On part component and SGCMGs complete when unusual instruction moment at U ₃On whole components distribute to B cover SGCMGs, by B cover SGCMGs output, with the A cover SGCMGs that avoids losing efficacy σ when unusual ₃→ 0 causes the phenomenon that the frame corners velocity solution is excessive or nothing is separated to take place.

Order is assigned to the U of B cover SGCMGs ₃Instruction moment on the direction does

${\mathrm{<figure-callout\; id="2"\; label="T\; ca"\; filenames="HDA0000130508730000022.png"\; state="\{\{state\}\}">T</figure-callout>}}_{\mathrm{ca}}=\frac{\mathrm{\&alpha;}}{{\mathrm{\&sigma;}}_3+\mathrm{\&alpha;}}{U}_3{U}_3^T{T}_{\mathrm{ca}}---\left(16\right)$

In the formula

$\mathrm{\&alpha;}=\left\{\begin{array}{cc}0& D_a\&GreaterEqual;\mathrm{\&epsiv;}\\ k{(D_a-\mathrm{\&epsiv;})}^2,& \mathrm{\&epsiv;}>D_a\&GreaterEqual;0\end{array}\right.---\left(17\right)$

D wherein _aUnusual tolerance for inefficacy SGCMGs is expressed as

D _a＝det(A _atA _at ^T) (18)

ε is positive threshold value, and k is positive scalar parameter, and both can be selected according to actual conditions.The instruction moment of SGCMGs of then losing efficacy does

T _ca1＝T _ca-T _ca2＝US _aU ^TT _ca (19)

Wherein

$S_a=\mathrm{<figure-callout\; id="1"\; label="diag"\; filenames="HDA0000130508730000012.png,HDA0000130508730000022.png"\; state="\{\{state\}\}">diag</figure-callout>}\left(\begin{array}{ccc}1& 1& \frac{{\mathrm{\&sigma;}}_3}{{\mathrm{\&sigma;}}_3+\mathrm{\&alpha;}}\end{array}\right)---\left(20\right)$

Can know by formula (16) and (19), work as D _a>=ε can think A cover SGCMGs away from unusual, α=0, σ ₃＞0, S then _a=E ₃, T _Ca1=T _Ca, T _Ca2=0, three instruction moment T _CaDistribute to A cover SGCMGs fully.And work as D _a＜ε, A cover SGCMGs move closer to when unusual, and α increases, σ ₃Reduce, then distribute to B cover SGCMGs at unusual direction U ₃On instruction moment also constantly increase, distribute to simultaneously A cover SGCMGs at unusual direction U ₃On instruction moment then constantly reduce.Finally, when SGCMGs is unusual, σ ₃=0,

U ₃On instruction moment distribute to B cover SGCMGs fully, guarantee that topworks still can accurately export three instruction moments when unusual.

So far, obtained T fully _Ca1And T _Ca2Expression formula suc as formula shown in (19) and (16).

The 3rd goes on foot, and confirms to be finally allocated to the moment of two cover SGCMGs.After twice moment distribution, last T _cIn distribute to the instruction control moment components T ' of A cover SGCMGs _CaBe expressed as

$T_{\mathrm{ca}}^{\&prime;}={\mathrm{<figure-callout\; id="1"\; label="T\; ca"\; filenames="HDA0000130508730000012.png,HDA0000130508730000022.png"\; state="\{\{state\}\}">T</figure-callout>}}_{\mathrm{ca}}={\mathrm{US}}_a{U}^T{T}_{\mathrm{ca}}=\frac{h_{\mathrm{an}}}{h_{\mathrm{an}}+4.4719}{\mathrm{US}}_a{U}^T{T}_c---\left(21\right)$

Distribute to the instruction control moment components T ' of B cover SGCMGs _CbBe expressed as

$T_{\mathrm{cb}}^{\&prime;}=T_{\mathrm{cb}}+{\mathrm{<figure-callout\; id="2"\; label="T\; ca"\; filenames="HDA0000130508730000022.png"\; state="\{\{state\}\}">T</figure-callout>}}_{\mathrm{ca}}=\frac{4.4719}{h_{\mathrm{an}}+4.4719}{T}_c+\frac{\mathrm{\&alpha;}}{{\mathrm{\&sigma;}}_3+\mathrm{\&alpha;}}\&CenterDot;\frac{h_{\mathrm{an}}}{h_{\mathrm{an}}+4.4179}{U}_3{U}_3^T{T}_c---\left(22\right)$

The 4th step, the manipulation rule design of two cover SGCMGs.For A cover SGCMGs, when it is absorbed in when unusual, only need the moment of output perpendicular to unusual direction, therefore can directly ask the pseudoinverse of frame corners speed to separate according to formula (15) and (21), obtain

${\stackrel{\&CenterDot;}{\mathrm{\&delta;}}}_{\mathrm{ca}}=-\frac{h_{\mathrm{an}}}{h_{\mathrm{an}}+4.4719}{V}_1{S}_1^{-1}{S}_a{U}^T{T}_c---\left(23\right)$

For B cover SGCMGs,, just can effectively use zero motion to handle and restrain because its each gyro operate as normal all can be set to be the spheroid of radius, in this angular momentum body, latent singular point only to be arranged like this with envelope smallest angles momentum by its controlled angular momentum body.It is unusual to be that convenient A cover SGCMGs is absorbed in, and utilize zero motion to handle rule and also can handle B cover SGCMGs fully and hide singular point, and output is along the moment of the unusual direction of A cover SGCMGs.Zero motion of B cover SGCMGs is handled rule and is designed to

${\stackrel{\&CenterDot;}{\mathrm{\&delta;}}}_{\mathrm{cb}}=-{A_{\mathrm{bt}}}^T{\left(A_{\mathrm{bt}}{A_{\mathrm{bt}}}^T\right)}^{-1}\frac{T_{\mathrm{cb}}^{\&prime;}}{{\mathrm{<figure-callout\; id="0"\; label="h"\; filenames="HDA0000130508730000032.png"\; state="\{\{state\}\}">h</figure-callout>}}_0}+\mathrm{\&beta;}(I_{6\&times;6}-{A_{\mathrm{bt}}}^T{\left(A_{\mathrm{bt}}{A_{\mathrm{bt}}}^T\right)}^{-1}{A}_{\mathrm{bt}})\frac{{\&PartialD;D}_b\left({\mathrm{\&delta;}}_b\right)}{{\&PartialD;\mathrm{\&delta;}}_b}---\left(24\right)$

In the formula, A _Bt=A _Bt0D [cos δ _b]-A _Bs0D [sin δ _b], scalar parameter β chooses as follows,

$\left\{\begin{array}{cc}\mathrm{\&beta;}=0,& D_b>1\\ \mathrm{\&beta;}=5,& D_b\&le;0.5\\ \mathrm{\&beta;}=20{(D_b-1)}^2,& 0.5<D_b\&le;1\end{array}\right.---\left(25\right)$

Wherein, D _b=det (A _BtA _Bt ^T) be the unusual tolerance of B cover SGCMGs.The concrete theory of this part can with reference to pertinent literature (reference: Zhang Renwei, " satellite orbit and attitude dynamics and control ", publishing house of BJ University of Aeronautics & Astronautics, 291-293).

So far; Having obtained A cover and separately skeleton instruction angular velocity

of B cover SGCMGs fully only need rotate with the gimbal axis that these two instruction angular speed drive A cover and B cover SGCMGs respectively with and just can guarantee when the part gyro lost efficacy and gyro when unusual; Control moment is able to export accurately and efficiently, and then the attitude of control spacecraft.

Referring to Fig. 4, scheme of the present invention residing position in whole Spacecraft Control loop is frame of broken lines part among the figure.Among the figure, (1) is actual attitude angle and angular velocity information; (2) for estimating attitude angle and angular velocity information; (3) be expectation attitude angle and angular velocity information; (4) be instruction control moment; (5) for distributing to the instruction control moment of inefficacy gyro group; (6) for distributing to the instruction control moment of operate as normal gyro group; (7) be perpendicular to the moment components of the unusual direction of inefficacy gyro group in (6); (8) be (5) moment components along the unusual direction of inefficacy gyro group; (9) for the instruction frame corners speed of inefficacy gyro group; (10) be the instruction frame corners speed of operate as normal gyro group; (11) for the actual frame angular velocity of inefficacy gyro group; (12) be the actual frame angular velocity of operate as normal gyro group; (13) for the actual output torque of inefficacy gyro group; (14) be the actual output torque of operate as normal gyro group; (15) be the total output torques of two cover gyro groups; (16) be outer disturbance torque.Spacecraft attitude control system based on two cover SGCMGs constitutes close loop control circuit together by attitude sensor, attitude controller, topworks (two cover SGCMGs) and spacecraft body.Attitude sensor is measured and is confirmed spacecraft with respect to some known reference target direction of space, confirms to confirm spacecraft attitude after algorithm is further handled the information that records through attitude again.Confirm the attitude information that link obtains according to the attitude information and the attitude of spacecraft expectation then, select appropriate control algorithm design attitude controller, thus the required instruction control moment of controlled spacecraft.Next be exactly to utilize the two cover SGCMGs that proposed to coordinate controlling schemes, thereby the framework of handling two each gyros of cover SGCMGs move and guarantee that two cover SGCMGs can produce required control moment by steering order according to certain rules.At last, the resultant couple of two cover SGCMGs outputs acts on the spacecraft body, can obtain the attitude response of spacecraft according to the spacecraft attitude dynamics equation of being set up.Because two cover SGCMGs can accurately export control moment, so spacecraft will rotate according to the attitude angle and the angular velocity characteristics of motion of expectation.

Do bright specifically below in conjunction with some assembly spacecraft attitude control simulation results to this programme.

Referring to Fig. 2, suppose that the core cabin in the assembly of space station all respectively is equipped with a cover pentagonal pyramid configuration SGCMGs with one of application cabin.Among the figure, o _bx _by _bz _bBe core cabin body coordinate system, initial point o _bBe taken at the barycenter of core cabin module, x _b, y _bAnd z _bBe fixed on the core cabin,

With

Be respectively the gimbal axis direction unit vector of the 1-6 gyro of the A cover SGCMGs that is installed on the core cabin,

With

Be respectively the armature spindle direction unit vector of the 1-6 gyro of A cover SGCMGs.

and

are respectively the gimbal axis direction unit vector of the 1-6 gyro that is installed on the B cover SGCMGs that uses the cabin, and

and

is respectively the armature spindle direction unit vector of the 1-6 gyro of B cover SGCMGs.Two cover SGCMGs are with respect to installation position such as Fig. 2 of core cabin body coordinate system; It is consistent with

direction that the gimbal axis

of first gyro is formed the projection on plane among the A cover SGCMGs with

at

along

, among the B cover SGCMGs gimbal axis

of first gyro along

at

Yu

projection on overlap plane Yu direction is Yi Zhi.Select the reason of this installation position to be operate as normal or when having the part gyro to lose efficacy, this installation position can both guarantee that gyro group has envelope and unusual performance index preferably all as two cover SGCMGs.Each the gyro initial time rotor angular momentum direction unit vector of two cover SGCMGs and output torque the component array of unit vector under the body series of core cabin in the other direction do

s _a10＝[1?0?0] ^T，s _a20＝[-sin18°?0?-cos18°] ^T，s _a30＝[-1?0?0] ^T，s _a40＝[-sin18°?0?cos18°] ^T，

s _a50＝[cos36°?0?sin36°] ^T，s _a60＝[cos36°?0?-sin36°] ^T，t _a10＝[0?0?-1] ^T，

t _a20＝[-sin26.57°cos18°?-cos26.57°?sin26.57°sin18°] ^T，

t _a30＝[0?-cos26.57°?sin26.57°] ^T，

t _a40＝[sin26.57°cos18°?-cos26.57°?sin26.57°sin18°] ^T，

t _a50＝[sin26.57°sin36°?-cos26.57°?-sin26.57°cos36°] ^T，

t _a60＝[-sin26.57°sin36°?-cos26.57°?-sin26.57°cos36°] ^T。

s _b10＝[0?0?1] ^T，s _b20＝[0-cos18°?-sin18°] ^T，s _b30＝[0?0?-1] ^T，s _b40＝[0cos18°?-sin18°] ^T，

s _b50＝[0sin36°cos36°] ^T，s _b60＝[0?-sin36°?cos36°] ^T，t _b10＝[0?-1?0] ^T，

t _b20＝[-cos26.57°sin26.57°sin18°-sin26.57°cos18°] ^T，

t _b30＝[-cos26.57°sin26.57°0] ^T，

t _b40＝[-cos26.57°sin26.57°sin18°sin26.57°cos18°] ^T，

t _b50＝[-cos26.57°-sin26.57°cos36°sin26.57°sin36°] ^T，

t _b60＝[-cos26.57°-sin26.57°cos36°-sin26.57°sin36°] ^T。

The nominal angular momentum of supposing each gyro is 180Nms, and the 5th and the 6th gyro of hypothesis A cover SGCMGs lost efficacy the initial frame corners δ of A cover SGCMGs _A0=[pi/2 00 0] ^T, the initial frame corners δ of B cover SGCMGs _B0=[pi/2 0000 0] ^TIn order to carry out numerical simulation, also need synthesize A cover and the final output torque of B cover SGCMGs.Specific practice is: the framework servo-drive system of supposing each gyro can realize accurately control; Can think that then the actual frame angular velocity and instruction frame corners speed of each gyro equates; After promptly

arrived the instruction frame corners speed of A cover SGCMGs and each gyro of B cover SGCMGs in the 4th step in implementation step; So again according to gyro output torque equation (5), the actual output torque that can obtain A cover and B cover SGCMGs is respectively

$T_{\mathrm{ra}}={-h}_0{A}_{\mathrm{at}}{\stackrel{\&CenterDot;}{\mathrm{\&delta;}}}_{\mathrm{ra}}---\left(26\right)$

$T_{\mathrm{rb}}={-h}_0{A}_{\mathrm{bt}}{\stackrel{\&CenterDot;}{\mathrm{\&delta;}}}_{\mathrm{rb}}---\left(27\right)$

The final moment of passing through is synthesized, and can obtain the total output torque of two cover SGCMGs to do

T _r＝T _ra+T _rb (28)

Adopt the PID control law to carry out the large angle maneuver control of assembly spacecraft.Visible like Fig. 5 and Fig. 6, when the 2800s left and right sides was arrived in simulation run, A cover SGCMGs was very near unusual, but this moment, its frame corners speed was not undergone mutation, and A cover SGCMGs keeps controlled in whole process.Simultaneously, as shown in Figure 7, the total output torque and instruction of two cover SGCMGs control moment error all maintains in the 10-14Nm in the whole process, and the moment output accuracy is very high.

In sum, the present invention has provided a kind of single frame control-moment gyro crowd control method for coordinating based on svd.When a cover pentagonal pyramid configuration SGCMGs who installs on the assembly of space station has 3 gyros to break down at the most, can unite mounted another set of SGCMGs and coordinate control.Utilize the method for svd to make fault SGCMGs only need export instruction moment, and will distribute to normal SGCMGs along the instruction moment of unusual direction perpendicular to unusual direction.Adopt this coordination controlling schemes, can guarantee the complete controllability of fault SGCMGs when experience is unusual, can guarantee that again total output torque and instruction moment of two cover SGCMGs conforms to fully, thereby improved the precision of attitude control.The present invention can be applied in Large Spacecraft tasks such as space station.

The above only is a preferred implementation of the present invention; Should be understood that; For those skilled in the art, under the prerequisite that does not break away from the principle of the invention, can also make some improvement; Perhaps part technical characterictic wherein is equal to replacement, these improvement and replacement also should be regarded as protection scope of the present invention.

Claims (2)

1. single frame control-moment gyro crowd control method for coordinating based on svd; This method is applicable to that spacecraft has two cover pentagonal pyramid configuration SGCMGs; And wherein A is with 1,2 or 3 gyros inefficacies, and the situation of B cover operate as normal may further comprise the steps:

Step 1, will control the required instruction moment of whole spacecraft distribute to by a certain percentage two the cover SGCMGs;

Step 2, the method for utilizing svd are decomposed the instruction moment of distributing to A cover SGCMGs once more; With wherein distributing to B cover SGCMGs, overlap SGCMGs and still distribute to A perpendicular to the instruction moment components of the unusual direction of A cover SGCMGs along the instruction moment components of the unusual direction of A cover SGCMGs;

Step 3, assigned after, A cover SGCMGs utilizes pseudoinverse to handle rule and solves its instruction frame corners speed, B cover SGCMGs utilizes pseudoinverse to add zero motion and handles rule and solve its instruction frame corners speed;

Step 4, two cover SGCMGs press instruction frame corners speed running separately respectively, and the output torque sum acts on spacecraft, accomplish accurate attitude control.

2. a kind of single frame control-moment gyro crowd control method for coordinating according to claim 1 based on svd; It is characterized in that; In step 1, carry out moment according to the minimum envelop angular momentum size of two cover SGCMGs and distribute: suppose that the total instruction control moment that obtains from the attitude controller that is designed is T _c, the instruction control moment of distributing to A cover inefficacy SGCMGs is T _Ca, be expressed as

The instruction control moment of distributing to B cover SGCMGs is T _Cb, be expressed as

Wherein, h _AnBe the smallest angles momentum of the envelope of n SGCMGs of A cover operate as normal, the span of n is 3≤n≤6.

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