CN102331785A - Method for controlling spacecraft attitude directing constraint attitude maneuver - Google Patents

Abstract

The invention relates to an independent method for controlling spacecraft attitude directing constraint attitude maneuver, belonging to the technical field of spacecraft attitude control. The method comprises the following steps of: constructing a navigation function V related to the motion of the tail end point of a current pointing vector r of a sensor on a unit spherical surface S by taking the tail end point of a target pointing vector rd of the sensor as a target point position, taking the tail end point of the current pointing vector r as a current position and taking a spherical doom formed by a pointing constraint as a barrier region; designing a control torque expression according to the navigation function, and regulating the amplitude of a control torque by changing control torque parameters to drive the spacecraft to make the sensor point to the target vector rd; and driving the spacecraft to rotate in the vector direction of the sensor by an angle theta after the sensor points to the target vector rd, so that a complete attitude maneuver process of the spacecraft is realized. According to the method, pointing avoidance of the barrier region can be processed definitely, a local minimum value can be avoided for a plurality of barrier constraints simultaneously, safe maneuver of the spacecraft to a target attitude is ensured, the requirement of boundedness on an executing mechanism is met, and independent control over spacecraft directing constraint attitude maneuver is realized.

Description

A kind of spacecraft points to constraint attitude maneuver control method

Technical field

The invention belongs to spacecraft attitude control technology field, relate to a kind of method of being pointed to the autonomous attitude maneuver control of constraint.

Background technology

Along with the development of spacecraft technology and the variation of space mission, low light level Sensitive Apparatus (like infrared telescope, star sensor etc.) need be installed on the spacecraft sometimes accomplish specific detection mission.Particularly the light of the sun is very sensitive to the high light celestial body for these instruments, if not taking precautions against light through measures necessary enters in its sensitive components, just is easy to cause the damage of sensitive components.In the attitude maneuver process, common way is to close these devices, but in some cases, still needs these instruments in running order, and this just requires to invent the sensing bypassing method of attitude maneuver.

In handling the method for this problem, generally speaking there are two types, the one, cook up the attitude maneuver sequence in advance, then through carrying out progressively tracking sequence point, thereby accomplish whole attitude maneuver process; Another kind is to be dissolved into planning and execution in the controller simultaneously; Promptly through making up energy function; Barrier zone in the whole work space is expressed as the high potential energy district, and the target place is expressed as the low-potential energy district, incorporates attitude dynamics and kinematics then; Draw the appropriate control input, rotate to targeted attitude thereby drive spacecraft attitude.Last class methods generally need higher calculation cost, on star the limited spacecraft of computational resource and inadvisable.Then one type of mode is less to the computational resource requirement, more can better satisfy the requirement of real-time on the star by contrast.

To back one class methods; Wisniewski; R.and P.Kulczycki considers that star sensor points to the situation of constraint, utilizes the energy forming method to accomplish the attitude maneuver process in " Slew maneuver control for spacecraft equipped with star camera and reaction wheels. " literary composition; But do not consider the problem of actuating mechanism controls bounded, therefore can not guarantee the smooth execution of actual engineering.

Further; Radice; G.and I.Ali is in " Autonomous attitude using potential function method under control input saturation " literary composition; Considering that a sensor points to the situation of constraint, consider the boundedness of topworks equally, obtained the control input of bounded.But also untreated areas is pointed to constraint, just makes in the process of implementation to point to axle away from the obstacle sensing as far as possible, is difficult to explanation and points to the barrier zone that whether has entered into the center that is oriented to obstacle.When retraining for a plurality of sensings in addition, there is the situation of local solution in this method.

Generally speaking; Existing to being pointed in the attitude maneuver control method of constraint; The processing of evading to the barrier zone problem does not also have clear and definite method, and local minimizing rejecting considers simultaneously that also the method for topworks's bounded also occurs when in addition many obstacles being retrained.

Summary of the invention

In order to evade the sensing of barrier zone clearly; Local minimizing problem when rejecting the constraint of many obstacles simultaneously; Under the situation of considering topworks's bounded, the present invention proposes the method that a kind of spacecraft points to the control of constraint attitude maneuver, and its concrete process that realizes is following:

Step 1, according to the structure mount message of spacecraft itself, the sensor pointing vector that obtains need to consider pointing to constraint is expressed as r under body series _b, under inertial system, be expressed as r;

According to attitude sensor information and ephemeris information, obtain the spacecraft barycenter and be respectively r in the expression under the inertial system to the vector of n relevant celestial body _Oj, j=1 ..., n; According to attitude sensor information, obtain spacecraft under inertial system with respect to the attitude matrix C under body series _Ib, subscript representes that this attitude matrix is the conversion from body series b to inertial system I;

According to the expectation attitude matrix

The target directing that obtains the sensor pointing vector is expressed as r under inertial system _d

Step 2, be the center, set up unit sphere S with the spacecraft barycenter; Visual field vertex angle theta according to sensor ₀And the view angle theta of n relevant celestial body _Oj, obtain sensor direction vector r and n day voxel vector r _OjBetween restriction relation be r ^Tr _Oj＜cos θ _j, θ wherein _j=θ ₀+ θ _Oj, promptly sensor direction vector r can not enter in the attitude maneuver process by j day voxel vector r _OjBe symcenter, the summit is at the spacecraft barycenter, and cone apex angle is θ _jSpace awl in, the sensing of n celestial body will form n space awl, these spaces bore and unit sphere S crossing after, unit sphere S is cut out n spherical crown surface, wherein, the distance at the center of j spherical crown surface and spherical crown edge is ρ _j, ${\mathrm{\ρ}}_j=\sqrt{2-2\mathrm{Cos}{\mathrm{\θ}}_j};$

Step 3, on unit sphere S, with sensor target directing vector r _dDistal point be the impact point position, the distal point of the pointing vector r that sensor is current is a current location, pointing to the formed spherical crown surface of constraint is barrier zone, makes up the navigation function V that the distal point about r moves on sphere; Navigation function V is:

$V={\left(\frac{{\mathrm{\γ}}_d^k}{{\mathrm{\γ}}_d^k+\mathrm{\β}}\right)}^{1/k}---\left(1\right)$

Wherein, k is a constant, k>=2, γ _d=|| r-r _d|| ², β and β _jBe intermediate variable, ${\mathrm{\β}}_j={\left|\right|r-r_{\mathrm{Oj}}\left|\right|}^2-{\mathrm{\ρ}}_j^2;$

Step 4, the function that will navigate are fused in the design process of control law, in conjunction with attitude dynamics and kinematics, and utilize backstepping method design control moment u expression formula, spacecraft rotation under control moment drives, and control moment u is:

Wherein, μ, η, s is for regulating parameter, and J is the inertia matrix of spacecraft, and ω is the attitude angular velocity of spacecraft, [ω ^*] be the multiplication cross matrix of spacecraft attitude angular velocity, expectation attitude angular velocity ω _sDerivative with the expectation attitude angular velocity

Provide by (3)-(7) and (8)-(14) respectively;

${\mathrm{\ω}}_s=-\frac{1}{\mathrm{\μ}}{(\mathrm{F\Ψ}-\mathrm{G\Ξ})}^T---\left(3\right)$

Wherein, F, Ψ, G, Ξ are intermediate variable, and it is following to embody formula:

$F={({\mathrm{\γ}}_d^k+\mathrm{\β})}^{-1/k}-{({\mathrm{\γ}}_d^k+\mathrm{\β})}^{-1-1/k}{\mathrm{\γ}}_d^k---\left(4\right)$

$G={\mathrm{\γ}}_d\frac{{({\mathrm{\γ}}_d^k+\mathrm{\β})}^{-1/k-1}}{k}---\left(5\right)$

Ψ＝-2(r-r _d) ^T(C _Ib[r _b×]) (6)

Wherein, r _b=[r _B1, r _B2, r _B3] ^T, [r _b*] be the multiplication cross matrix, its form does $\left[\begin{array}{ccc}0& -r_{b3}& r_{b2}\\ r_{b2}& 0& -r_{b1}\\ -r_{b2}& {\mathrm{<figure-callout\; id="1"\; label="r\; b"\; filenames="HDA0000076247450000021.png"\; state="\{\{state\}\}">r</figure-callout>}}_b& 0\end{array}\right],$

$\mathrm{\&Xi;}=\underset{j=1}{\overset{n}{\mathrm{\&Sigma;}}}\left(\left(\underset{i=1}{\overset{n}{\underset{i\&NotEqual;j}{\mathrm{\&Pi;}}}}{\mathrm{\&beta;}}_i\right)(-2{\left({r-r}_{\mathrm{oj}}\right)}^T\left(C_{\mathrm{Ib}}\right[r_b\&times;\left]\right))\right)---\left(7\right)$

Wherein,

is intermediate variable, and it is following to embody formula:

Wherein, B _j,

Be intermediate variable, it is following to embody formula:

B _j＝-2(r-r _oj) ^TC _Ib[r _b×] (13)

Because the navigation function can guarantee that sensor, all will be motor-driven to dbjective state from original state safely except that limited several saddle points of navigation function, thereby can avoid the appearance of local minimum.

Step 5, according to the control moment expression formula of confirming in the step 4, through changing controlled variable μ, η and s adjust the amplitude of control moment, thereby satisfy the output requirement of spacecraft topworks;

Wherein,

is normal value; || e|| is an intermediate variable, and it is following to embody formula:

Be constant by formula (16) and (17) visible

; Formula (18) and (19) visible || e|| is along with parameter μ; The variation of η and s and changing; Therefore; Through adjustment controlled variable μ; η and s; And then regulate

and || the size of e||, can adjust the boundary that control is imported;

Step 6, drive spacecraft, thereby make sensor definite object vector r according to step 4 and the determined control moment of step 5 _d

Step 7, spacecraft are accomplished sensor definite object vector r _dAfter, need again around sensor direction vector anglec of rotation θ, thereby realize the complete attitude maneuver process of spacecraft; Wherein, definite formula of angle θ is following:

$\mathrm{\&theta;}=2\mathrm{ar}\cos \left(Q_{e_0}\right)---\left(20\right)$

Wherein,

Be deviation attitude quaternion Q _eMark portion, Q _eBy the deviation attitude matrix

It is converted,

Be the spacecraft attitude matrix,

Be the final objective attitude matrix.

Beneficial effect

(1) the inventive method through the navigation function that the distal point that makes up about r moves on sphere, can be evaded the sensing of barrier zone clearly and handling, rather than sensing spool is pointed to away from obstacle.

Can avoid local minimum when (2) retraining for a plurality of obstacles simultaneously, guarantee that spacecraft safety is motor-driven to targeted attitude.

(3) through the adjustment of control moment and confirming of control input border, satisfied the requirement of topworks's boundedness, realized that autonomous spacecraft points to the control of constraint attitude maneuver.

Description of drawings

Fig. 1 is the sensing constraint attitude figure of embodiment of the present invention;

Fig. 2 is attitude constraint mapping graph;

Fig. 3 is the motion path of sensor direction vector end on unit sphere in the practical implementation case;

Fig. 4 is the time history of control moment input in the practical implementation case.

Embodiment

Elaborate below in conjunction with the embodiment of accompanying drawing to the inventive method.

A kind of spacecraft points to constraint attitude maneuver control method, and its concrete steps comprise:

Step 1, as shown in Figure 1, establishing needs to consider that the infrared senstive device that points to constraint points to the r that is expressed as under body series _b=[0.12,0.24,0.963] ^TThe spacecraft barycenter is to the component r of direction vector under inertial system of the sun _O1=[1,0,0] ^T, the spacecraft barycenter is to the component r of direction vector under inertial system of Saturn _O2=[0.5 ,-0.866,0] ^T, the spacecraft barycenter is to the component r of direction vector under inertial system of Jupiter _O3=[0.5,0.866,0] ^TCan obtain the attitude matrix C of inertial system according to attitude sensor information in real time with respect to the spacecraft body series _IbInitial time spacecraft attitude matrix does $C_{\mathrm{Ib}}=\left[\begin{array}{ccc}1& 0& 0\\ 0& 1& 0\\ 0& 0& 1\end{array}\right],$ Angular velocity under the body series is ω=[0,0,0] ^TThe component of sensor pointing vector under inertial system is r (r=C _Ibr _b).According to the expression r of sensor under body series _bAnd expectation attitude matrix $C_{\mathrm{Ib}}^d=\left[\begin{array}{ccc}0.4602& -0.6354& 0.6201\\ -0.8786& -0.4261& 0.2156\\ 0.1272& -0.6440& -0.7543\end{array}\right],$ The target directing that obtains the sensor pointing vector is expressed as r under inertial system _d=[0.5,0 ,-0.866].

Step 2, the visual field vertex angle theta of establishing sensor ₀The visual angle of=14 ° and relevant celestial body is respectively θ _O1=3 °, θ _O2=1 °, θ _O3=2 °, θ then ₁=θ ₀+ θ _O1=17 °, θ ₂=θ ₀+ θ _O2=15 °, θ ₃=θ ₀+ θ _O3=16 °.As shown in Figure 2, the center of the 3rd spherical crown surface that the center of the center of 1st spherical crown surface relevant with the sun and the distance at spherical crown edge are with Saturn the 2nd relevant spherical crown surface and the distance at spherical crown edge are correlated with Jupiter for

and spherical crown edge distance is

Step 3, according to the navigation functional form, with the r that obtains in step 1 and the step 2 _dBe brought into γ _dIn, with r _Oj(j=1,2,3) and ρ _j(j=1,2,3) are brought into β _jIn (j=1,2,3).In addition, k=2, N=3.

Step 4, utilize the controlled moment of backstepping method input expression formula

Wherein, the inertia matrix of spacecraft $J=\left[\begin{array}{ccc}100& & \\ & 100& \\ & & 100\end{array}\right]\mathrm{Kg}\&CenterDot;m^2.$ ω is that spacecraft is at the control component of angular velocity under body series constantly, ω _sBe control expectation angular velocity constantly,

Constantly expect the derivative of angular velocity for control.μ, η and s are the adjusting parameter, and all are arithmetic number.

Step 5, reach formula for list-directed input list, design parameter μ=12, η=30, s=0.5, the peak value that makes control moment is less than 1Nm.

Step 6, drive spacecraft, thereby make sensor definite object vector r according to step 4 and the determined control moment of step 5 _d

Step 7, at γ _dAmplitude reach 10 ^-9After, control is switched to the large angle maneuver around the sensor direction vector. $C_{\mathrm{Ib}}^s=\left[\begin{array}{ccc}-0.0632& -0.7080& 0.7034\\ -0.9825& 0.1678& 0.0806\\ -0.1751& -0.6860& -0.7062\end{array}\right]$ For the attitude matrix of spacecraft, according to the targeted attitude matrix at switching instant $C_{\mathrm{Ib}}^d=\left[\begin{array}{ccc}0.4602& -0.6354& 0.6201\\ -0.8786& -0.4261& 0.2156\\ 0.1272& -0.6440& -0.7543\end{array}\right],$ Can try to achieve the deviation attitude matrix does $C_{\mathrm{Ib}}^e=C_{\mathrm{Ib}}^d{\left(C_{\mathrm{Ib}}^s\right)}^T=\left[\begin{array}{ccc}0.8569& -0.5088& -0.0827\\ 0.5088& 0.8092& 0.2939\\ -0.0827& -0.2939& 0.9523\end{array}\right],$ And then try to achieve the deviation hypercomplex number and do ${\mathrm{<figure-callout\; id="0"\; label="Q\; e"\; filenames="HDA0000076247450000011.png,HDA0000076247450000013.png"\; state="\{\{state\}\}">Q</figure-callout>}}_{e_{}}={[\mathrm{0.9511,0.1545,0},-0.2675]}^T,$ Obtain at last around the sensor direction vector anglec of rotation

Set forth through top method, this method is simulated specifically as shown in Figure 3.With Spherical Surface S ²Be launched into the plane by longitude and latitude,, and the end orbit of sensor direction vector be plotted on this plane according to potential energy curves such as navigation function draftings.Can find out that by Fig. 3 the spherical crown surface barrier zone that is formed by three celestial bodies is walked around in the path of the distal point of sensor direction vector on sphere safely, and reach the target directing distal point.The time history curve plotting of three components of control moment u in Fig. 4, can be seen three component u _x, u _y, u _zMaximum amplitude all be no more than the 1Nm of expectation.

Though in conjunction with accompanying drawing embodiment of the present invention has been described, to those skilled in the art, under the prerequisite that does not break away from the principle of the invention, can also make some improvement, these also should be regarded as belonging to protection scope of the present invention.

Claims (1)

1. a spacecraft points to the method that the constraint attitude maneuver is controlled, and it is characterized in that: the detailed process that this method realizes is following:

Step 1, according to the structure mount message of spacecraft itself, the sensor pointing vector that obtains need to consider pointing to constraint is expressed as r under body series _b, under inertial system, be expressed as r;

According to attitude sensor information and ephemeris information, obtain the spacecraft barycenter and be respectively r in the expression under the inertial system to the vector of n relevant celestial body _Oj, j=1 ..., n; According to attitude sensor information, obtain spacecraft under inertial system with respect to the attitude matrix C under body series _Ib, subscript representes that this attitude matrix is the conversion from body series b to inertial system I;

According to the expectation attitude matrix

The target directing that obtains the sensor pointing vector is expressed as r under inertial system _d

Step 2, be the center, set up unit sphere S with the spacecraft barycenter; Visual field vertex angle theta according to sensor ₀And the view angle theta of n relevant celestial body _Oj, obtain sensor direction vector r and n day voxel vector r _OjBetween restriction relation be r ^Tr _Oj＜cos θ _j, θ wherein _j=θ ₀+ θ _Oj, promptly sensor direction vector r can not enter in the attitude maneuver process by j day voxel vector r _OjBe symcenter, the summit is at the spacecraft barycenter, and cone apex angle is θ _jSpace awl in, the sensing of n celestial body will form n space awl, these spaces bore and unit sphere S crossing after, unit sphere S is cut out n spherical crown surface, wherein, the distance at the center of j spherical crown surface and spherical crown edge is ρ _j, ${\mathrm{\&rho;}}_j=\sqrt{2-2\mathrm{Cos}{\mathrm{\&theta;}}_j};$

Step 3, on unit sphere S, with sensor target directing vector r _dDistal point be the impact point position, the distal point of the pointing vector r that sensor is current is a current location, pointing to the formed spherical crown surface of constraint is barrier zone, makes up the navigation function V that the distal point about r moves on sphere; Navigation function V is:

$V={\left(\frac{{\mathrm{\&gamma;}}_d^k}{{\mathrm{\&gamma;}}_d^k+\mathrm{\&beta;}}\right)}^{1/k}---\left(1\right)$

Wherein, k is a constant, k>=2, γ _d=|| r-r _d|| ², β and β _jBe intermediate variable,

${\mathrm{\&beta;}}_j={\left|\right|r-r_{\mathrm{Oj}}\left|\right|}^2-{\mathrm{\&rho;}}_j^2;$

Step 4, the function that will navigate are fused in the design process of control law, in conjunction with attitude dynamics and kinematics, and utilize backstepping method design control moment u expression formula, spacecraft rotation under control moment drives; Control moment u is:

Wherein, μ, η, s is for regulating parameter, and J is the inertia matrix of spacecraft, and ω is the attitude angular velocity of spacecraft, ω _sFor the expectation attitude angular velocity with

Be the derivative of expectation attitude angular velocity, [ω ^*] be the multiplication cross matrix of spacecraft attitude angular velocity;

Step 5, according to the control moment expression formula of confirming in the step 4, through changing controlled variable μ, η and s adjust the amplitude of control moment, thereby satisfy the output requirement of spacecraft topworks;

Step 6, drive spacecraft, thereby make sensor definite object vector r according to step 4 and the determined control moment of step 5 _d

Step 7, spacecraft are accomplished sensor definite object vector r _dAfter do not reach complete targeted attitude, need to rotate to an angle around the sensor direction vector again, thereby realize the complete attitude maneuver process of spacecraft.

Images