CN103955223A - Posture and path coupling control method for deep space exploration soft landing process - Google Patents

Abstract

The invention relates to a posture and path coupling control method for the deep space exploration soft landing process. In the deep space celestial body soft landing process, a thrust-variable swing engine is used for conducting posture control, and therefore a landing path is disturbed. Important measures for reducing posture control and path control coupling of the thrust-variable swing engine includes that a part, used for eliminating centroid skewing, in a swing angle of the thrust-variable swing engine is estimated on line, and flight postures are amended. According to the posture and path coupling control method, the swing angle of the thrust-variable swing engine is required to be estimated on line firstly, and stable components of the swing angle are extracted; then the target flight postures are amended through the stable components of the swing angle, the amended thrust direction is coincident with the thrust direction expected in a guidance law, and influences of posture control swing of the thrust-variable swing engine on guidance are initiatively eliminated. By means of the posture and path coupling control method, disturbance to the guidance is reduced to the maximum degree while the posture control is carried out, and even though the guidance has no three-direction position control capacity, it can be guaranteed that the flight path is coincident with an ideal path as far as possible.

Description

A kind of appearance rail coupling control method of survey of deep space soft landing process

Technical field

The appearance rail coupling control method that the present invention relates to a kind of survey of deep space soft landing process, belongs to the autonomous control field of survey of deep space.

Background technology

The detection of landing is a kind of important means of survey of deep space.For the detector of large-scale lander or landing large-size celestial, in order to reduce landing speed, offset target celestial body gravitation, the thrust of landing engine certainly will be very large.If adopt large-scale landing engine, be subject to the impact of the factors such as engine manufacture, installation and decline process propellant space emission imbalance, landing motor power direction can not crossed barycenter just, will produce so very large attitude disturbance moment, and attitude is controlled and had a negative impact.The method solving has two kinds.Method is larger attitude control motor to be installed offset the disturbance torque that landing engine produces, but from Project Realization, larger attitude control motor all can run into a lot of difficulties from type selecting to installing, and control accuracy and efficiency not high yet.Another method is that landing engine is selected and waved engine.This engine has servo control mechanism, can change within the specific limits the direction of thrust.By adjusting the direction of landing motor power, made it barycenter, just can reduce the generation of disturbance torque.On the other hand, because landing motor power is very large, by less thrust deflection, just can form larger attitude control moment, thereby greatly improve the efficiency that attitude is controlled, reduce the demand to attitude control engine.But should see, wave engine and can form thrust drift angle for reducing disturbance torque, there is deviation in the thrust direction of wishing with guidance system, produces horizontal perturbed force, and this will exert an influence to guidance.While therefore, selecting to wave engine as landing engine, there will be appearance rail coupled problem.

At present, only have the Apollo of the U.S. and Project Constellation to adopt in moon landing process the engine control technology of waving.In design, they separately consider guidance and attitude control.Attitude is controlled to adopt and is waved the realization of engine deflection thrust direction, and Guidance Law supposes that the thrust direction of engine is along lander y direction all the time.Like this, actual having ignored waved the disturbing force that motor power deflection produces on perpendicular to y direction.Its consequence is that the perturbed force that attitude control produces can exert an influence to descending trajectory, and this need to rely on guidance loop to go to eliminate, and can reduce the execution efficiency of guidance; And not possessing three direction position control abilities once Guidance Law, the deviation of practical flight track and desired trajectory will strengthen.

Summary of the invention

Technology of the present invention is dealt with problems: overcome and existing survey of deep space soft landing is waved to engine control technology can produce horizontal perturbed force guidance is caused to the problem of adverse effect, a kind of appearance rail coupling control method of survey of deep space soft landing process has been proposed, the method on-line identification goes out to wave the steady propulsion drift angle that engine forms for reducing disturbance torque, and the target propulsive force direction generating with this angle compensation guidance, thereby eliminated, wave the interference that the control of engine attitude forms guidance, be conducive to realize the target of attitude and TRAJECTORY CONTROL simultaneously.

Technical solution of the present invention: a kind of appearance rail coupling control method of survey of deep space soft landing process, comprises that step is as follows:

(1) setting up three-dimensional XYZ coordinate is, by X-axis be defined as three principal axis of inertia of lander the axis of rolling, by Y-axis be defined as three principal axis of inertia of lander pitch axis, Z axis is defined as to the yaw axis of three principal axis of inertia of lander, change thrust is waved engine and is arranged at along X-direction, becomes thrust and waves engine along lander Y, Z axis swing output variable thrust.

(2) according to the aimed acceleration direction and goal rate of acceleration change of lander guidance system output and posture control system output wave engine pivot angle controlled quentity controlled variable, carry out targeted attitude and the attitude angular velocity of attitude command plane-generating lander.

(3) under the three-dimensional XYZ coordinate of step (1) foundation is, the current attitude of lander and attitude angular velocity measured value that the targeted attitude of lander that step (2) is generated and the navigational system of attitude angular velocity and lander provide compare, and are formed for attitude and the angular velocity error of rolling, pitching and the jaw channel of the control of lander three-axis attitude.

(4) the three-dimensional XYZ coordinate of setting up in step (1) according to the lander quality of the navigational system estimation of lander, calculates the centroid position of lander under being; The change thrust of being exported by lander guidance system is again waved motor power ratio and is calculated and become thrust and wave the large small instruction of engine target thrust, and the large small instruction of the target propulsive force Maximum controlling moment that can produce at pitching and yaw direction; According to the free-running frequency of the pitching of storing in the controller of Maximum controlling moment and lander and jaw channel and these closed-loop control system parameters of damping ratio, adjust in real time the pitching of lander and the pid control parameter of jaw channel and comprise controller proportional coefficient, integral item coefficient and differential term coefficient.

(5) under the three-dimensional XYZ coordinate of step (1) foundation is, the roll channel attitude of calculating according to step (3) and attitude angular velocity error generate control moment instruction by proportional-integral-differential PID control method, control moment instruction is generated to instruction pulsewidth through pulse-width modulation PWM, the instruction pulsewidth of generation is sent to the output of roll channel attitude control engine, complete the attitude of the axis of rolling is controlled.

(6) under the three-dimensional XYZ coordinate of step (1) foundation is, pitching and jaw channel attitude and the attitude angular velocity error according to step (3), calculated, the pid control parameter being generated by step (4), the pivot angle controlled quentity controlled variable of engine is waved in formation, this pivot angle controlled quentity controlled variable sends to waves engine, and this pivot angle controlled quentity controlled variable also feeds back to step (2) for correction target attitude.Wave engine and adjust engine pivot angle according to this pivot angle controlled quentity controlled variable, what according to step (4), generate waves the large small instruction output engine of engine target thrust thrust simultaneously, the attitude control moment of formation to pitching and jaw channel, completes the attitude of pitching and yaw axis is controlled.

The aimed acceleration direction of waving the output of engine pivot angle controlled quentity controlled variable correction guidance system of using in described step (2) posture control system of lander to generate, has eliminated and has waved the thrust direction that the control of engine appearance produces and change the impact on guidance.

In described step (4), the quality providing according to the navigational system of the thrust ratio of lander guidance system output and lander is estimated, calculate in real time and wave the control moment that engine produces, and the pid control parameter of pitching and jaw channel is carried out to online dynamic adjustment.

The present invention compared with prior art tool has the following advantages:

(1) the present invention has proposed a kind of On-line Estimation and has waved engine drift angle in step (3), and the method for compensation guidance targeted attitude calculating, has reduced the impact of landing mission appearance control on guidance, is conducive to improve the consistance of flight path and ideal trajectory.

(2) the present invention has used and a kind ofly in line computation, has waved engine appearance control-torque size in step (4), and dynamically adjust the method for PID controller parameter, make control system change and still there is stable control characteristic huge in the situation that in landing mission mass property.

Accompanying drawing explanation

Fig. 1 is deep space soft landing process appearance rail coupling control method structured flowchart of the present invention;

Fig. 2 is that the present invention waves engine pivot angle deflection schematic diagram;

Fig. 3 is the flight path under deep space soft landing process distinct methods of the present invention;

Fig. 4 is the control flow chart of control method of the present invention.

Embodiment

Basic ideas of the present invention are: according to change thrust, wave engine and swung and made thrust cross lander barycenter by engine nozzle, to reduce disturbance torque, realize the feature that attitude is controlled, adopt filtering technique to estimate the stable component in this pivot angle, pick out the angle that the thrust direction causing because of centroid motion departs from lander X-axis, and according to this drift angle valuation adjustment aim flight attitude, making to become thrust, to wave the target propulsive force direction that the thrust direction of the final output of engine exports with Guidance Law the same, meet to greatest extent when attitude is controlled target and guidance target and realize.

Without loss of generality, definition lander body coordinate system XYZ, three coordinate axis are parallel to respectively three principal axis of inertia directions of lander, and pitch axis, the Z axis that the axis of rolling, the Y-axis that wherein X-axis is defined as three principal axis of inertia of lander is defined as three principal axis of inertia of lander is defined as the yaw axis of three principal axis of inertia of lander.Suppose that lander is provided with a change thrust along X-axis and waves engine, the Guidance Navigation and Control System block diagram of landing mission as shown in Figure 1.

Concrete computation process is as follows as shown in Figure 4:

1) navigation calculation

The navigational system of landing mission be take inertial navigation as main, assists range finding, tests the speed or image measurement.Concrete navigation algorithm does not belong to content of the present invention, supposes that it has exported the inertia attitude of lander and (be accustomed to the use of sexual stance hypercomplex number q here _sENSexpression), attitude angular velocity (is used vector represent, subscript b represents that this vector representation is in lander body coordinate system), inertial position is (with the inertial position vector r at relative celestial body center _sENSrepresent) and quality estimation (use m _prepresent).

2) guidance is resolved

Landing mission different phase adopts different method of guidances, and specifically which kind of guidance algorithm does not belong to content of the present invention.But the position that in general, it all can provide according to navigational system, velocity estimation generate the instruction of target propulsive force acceleration direction vector and (use a _cMDexpression), the instruction of the variation of target propulsive force acceleration direction rule (is used represent) and engine thrust output proportional command, thrust output (is used f with the ratio of full thrust _trepresent).

3) attitude command planning (corresponding step 2)

A) the target propulsive force acceleration direction calculating targeted attitude providing according to guidance, with targeted attitude hypercomplex number q _cMDrepresent, make the lander axis of rolling (X-axis) and a _cMDoverlap.Method is as follows:

Calculate the target directing x of lander X-axis _cMD

$x_{\mathrm{CMD}}=\frac{a_{\mathrm{CMD}}}{\left|\right|a_{\mathrm{CMD}}\left|\right|}---\left(1\right)$

The sensing of two other axle of lander retrains according to task, such as meeting sensing demand of specific sensor etc.Here suppose the sensing target celestial body center of lander yaw axis (Z axis).The target directing z of Z axis so _cMDfor

$z_{\mathrm{CMD}}=\frac{r_{\mathrm{SENS}}}{\left|\right|r_{\mathrm{SENS}}\left|\right|}---\left(2\right)$

And the target directing y of lander pitch axis (Y-axis) _cMDaccording to right-handed helix theorem, be calculated as follows

$y_{\mathrm{CMD}}=\frac{z_{\mathrm{CMD}}\×x_{\mathrm{CMD}}}{\left|\right|z_{\mathrm{CMD}}\×x_{\mathrm{CMD}}\left|\right|}---\left(3\right)$

For guaranteeing that orthogonality need recalculate the target directing of Z axis

z _CMD=x _CMD×y _CMD (4)

Can calculate targeted attitude Matrix C like this _cMDfor

C _CMD=[x _CMD y _CMD z _CMD] ^T (5)

C _cMDwith targeted attitude hypercomplex number q _cMDthere is following relation

C _CMD=Aq(q _CMD) (6)

Wherein Aq () represents hypercomplex number to be converted to the function of attitude matrix matrix.Have

$\mathrm{Aq}\left(q\right)=\left[\begin{array}{ccc}{q}_1^2-q_2^2-q_3^2+q_4^2& 2(q_1{q}_2+q_3{q}_4)& 2(q_1{q}_3-q_2{q}_4)\\ 2(q_1{q}_2-q_3{q}_4)& -q_1^2+q_2^2-q_3^2+q_4^2& 2(q_2{q}_3+q_1{q}_4)\\ 2(q_1{q}_3+q_2{q}_4)& 2(q_2{q}_3-q_1{q}_4)& -q_1^2-q_2^2+q_3^2+q_4^2\end{array}\right]---\left(7\right)$

Q ₁～q ₄respectively q _cMDfour units, i.e. q _cMD=[q ₁, q ₂, q ₃, q ₄] ^t.From formula (7), can calculate lander targeted attitude hypercomplex number q _cMD, concrete grammar is this area conventional method, no longer describes in detail here.

Targeted attitude angular velocity according to ask for,

${\mathrm{\ω}}_{\mathrm{CMD}}^b=C_{\mathrm{CMD}}\·{\stackrel{\·}{a}}_{\mathrm{CMD}}---\left(8\right)$

B) afterwards, according to a upper cycle by 7) become thrust and wave the PID of engine and control the pitch axis pivot angle instruction δ that the change thrust of calculating is waved engine _ywith yaw axis pivot angle instruction δ _z(as shown in Figure 2) engine pivot angle when estimation thrust is crossed barycenter.It can pass through δ _yand δ _zsending into low-pass filter carries out.Method is:

Targeted attitude angular velocity when formula (8) calculating when constant,

$\begin{array}{c}{\widehat{\mathrm{\δ}}}_y=(1-K_y){\widehat{\mathrm{\δ}}}_y+K_y{\mathrm{\δ}}_y\\ {\widehat{\mathrm{\δ}}}_z=(1-K_z){\widehat{\mathrm{\δ}}}_z+K_z{\mathrm{\δ}}_z\end{array}---\left(9\right)$

Subscript " ^ " represents estimated value.K _yand K _zbe to be all less than 0 real number that is greater than 1, they need to design for concrete application characteristic.

C) lander targeted attitude is revised

The lander axis of rolling (X-axis) vector is at the x that is expressed as of lander body coordinate system ^b=[1 0 0] ^t, and estimate cross barycenter time change thrust wave motor power direction a ^bfor

$a^b=\left[\begin{array}{c}\sqrt{1-{\tan }^2\mathrm{\δ}\hat{y}-{\tan }^2\mathrm{\δ}\hat{z}}\\ -\tan \mathrm{\δ}\hat{z}\\ \tan \mathrm{\δ}\hat{y}\end{array}\right]---\left(10\right)$

X ^band a ^bbetween rotating vector e be

$e=\frac{a^b\×x^b}{\left|\right|a^b\×x^b\left|\right|}---\left(11\right)$

Rotation angle Φ is

Φ=arccos(<a ^b·x ^b>) (12)

Can obtain thus rotation hypercomplex number q _rfor

$q_r=\left[\begin{array}{c}e\sin \left(\frac{\mathrm{\&Phi;}}{2}\right)\\ \cos \left(\frac{\mathrm{\&Phi;}}{2}\right)\end{array}\right]---\left(13\right)$

Utilize this hypercomplex number can calculate revised targeted attitude hypercomplex number q _cMDand angular velocity

$q_{\mathrm{CMD}}\&LeftArrow;q_{\mathrm{CMD}}\&CircleTimes;q_r---\left(14\right)$

${\mathrm{\&omega;}}_{\mathrm{CMD}}^b\&LeftArrow;\mathrm{Aq}\left(q_r\right){\mathrm{\&omega;}}_{\mathrm{CMD}}^b---\left(15\right)$

Wherein, represent hypercomplex number multiplication, ← expression assignment.

4) attitude and attitude angular velocity error are calculated (corresponding step 3)

Relatively attitude quaternion, the attitude angular velocity of navigation output and targeted attitude hypercomplex number and the angular velocity of cooking up, can be formed for attitude and angular velocity error that lander rolling, pitching and three passages of driftage are controlled.

Error attitude quaternion Δ q is

$\mathrm{\&Delta;q}=q_{\mathrm{CMD}}^{-1}\&CircleTimes;q_{\mathrm{SENS}}---\left(16\right)$

Under the effect of attitude control system, attitude error maintains a less scope.Therefore,, under low-angle condition, can be similar to and obtain attitude error θ _eRRfor

${\mathrm{\&theta;}}_{\mathrm{ERR}}\&ap;2\&times;\left[\begin{array}{c}\mathrm{\&Delta;}{q}_1\\ \mathrm{\&Delta;}{q}_2\\ \mathrm{\&Delta;}{q}_3\end{array}\right]---\left(17\right)$

Δ q wherein ₁～Δ q ₃1st～3 components of error quaternion Δ q.

Angular velocity error for

${\mathrm{\&omega;}}_{\mathrm{ERR}}^b={\mathrm{\&omega;}}_{\mathrm{SENS}}^b-\mathrm{Aq}\left(\mathrm{\&Delta;q}\right){\mathrm{\&omega;}}_{\mathrm{CMD}}^b---\left(18\right)$

θ _eRRwith three component [θ _{eRR, x}, θ _{eRR, y}, θ _{eRR, z}] and respectively attitude error and the angular velocity error of rolling, pitching and jaw channel.

5) (corresponding step 4) is calculated in gain

The lander quality m providing according to navigation _p, estimate inertia, the height of center of mass of lander, and then can calculate the fore-and-aft distance of the relative engine jet pipe of barycenter throat (thrust point).

The method of being estimated lander inertia and height of center of mass by quality is a lot, the most accurately method be according to the mass property of the dry device of lander (containing propellant) and residual propellant number resolve.In the present invention, use Function Fitting method.

If ground has obtained lander and has filled the quality m after propellant ₀, inertia J ₀with centroid position vector P _gO0, the quality m after the emptying propellant of lander in addition _f, inertia is J _fwith centroid position vector P _gOfif precision allows, so current inertia J and centroid position P _gOcan estimate by the method for linear fit

$J=(J_f-J_0)\frac{m_p-m_0}{m_f-m_0}+J_0---\left(19\right)$

$P_{\mathrm{GO}}=(P_{\mathrm{GOf}}-P_{\mathrm{GO}0})\frac{m_p-m_0}{m_f-m_0}+P_{\mathrm{GO}0}---\left(20\right)$

Owing to becoming thrust and wave the location aware of engine throat, be designated as P _engine, can calculate change thrust and wave throat to the fore-and-aft distance x of lander barycenter _t.

x _T=P _GO,x-P _engine,x (21)

P _{gO, x}for P _gOthe first component, P _{engine, x}for P _enginethe first component.

According to the change thrust of guidance output, wave motor power ratio f _t, can calculate the large small instruction F of target propulsive force _tfor

F _T=f _TT _ds (22)

T _dsthe full thrust of engine.The Maximum controlling moment T that this thrust size can form at pitching and jaw channel _maxfor

T _max=f _TT _dsx _T (23)

Gain available packages on definition pitching and jaw channel is containing the vectorial K of two units (corresponding Y-axis and Z axis respectively) _trepresent

$K_T=\frac{1}{T_{\max }}\left[\begin{array}{c}{J}_{\mathrm{yy}}\\ J_{\mathrm{zz}}\end{array}\right]---\left(24\right)$

Wherein, J _yythe second row secondary series unit of J, J _zzit is the third line the 3rd row unit of J.So just can be according to predetermining and be stored in pitching in control system and jaw channel closed-loop control system naturally frequently

Rate ω _dESand dampingratioζ _dES, calculate the pid control parameter that pitching and jaw channel wave engine, proportional coefficient vector K _pwith integral item coefficient vector K _dfor

$K_P=\left[\begin{array}{c}{K}_{\mathrm{Py}}\\ K_{\mathrm{Pz}}\end{array}\right]={\mathrm{\&omega;}}_{\mathrm{DES}}^2{K}_T---\left(25\right)$

$K_D=\left[\begin{array}{c}{K}_{\mathrm{Dy}}\\ K_{\mathrm{Dz}}\end{array}\right]=2{\mathrm{\&zeta;}}_{\mathrm{DES}}{\mathrm{\&omega;}}_{\mathrm{DES}}{K}_T---\left(26\right)$

Wherein, K _pyk _pthe first component, represent the proportional coefficient of pitch channel controller; K _pzk _psecond component, represent the proportional coefficient of jaw channel controller; K _dyk _dthe first component, represent the differential term coefficient of pitch channel controller; K _dzk _dsecond component, represent the differential term coefficient of jaw channel controller.

Integral item coefficient vector K _i=[K _iy, K _iz] ^tcan be according to K _psuitably choose, it should be less than K _p1～2 magnitude.K wherein _iyk _ithe first component, represent the integral item coefficient of pitch channel controller; K _izk _isecond component, represent the integral item coefficient of jaw channel controller.

6) roll channel PID+PWM controls (corresponding step 5)

According to the attitude error of roll channel and angular velocity error, control the attitude control that roll channel attitude control engine is carried out the axis of rolling.Because roll channel attitude control engine is pulse mode work, therefore need to modulate this direction control moment.The method that can use PID+PWM to control, by PID control law, according to attitude error and angular velocity error, generate the control moment needing, then with pulse modulation technology, control moment is modulated, become the switching signal that drives the work of roll channel attitude control engine.The jet control method of PID+PWM belongs to the mature technology in Spacecraft Attitude Control field, does not belong to content of the present invention.

7) become PID control (the corresponding step 6) that thrust is waved engine

The control that change thrust is waved engine comprises that the PID of servo control mechanism pivot angle controls and the control of motor power size.

Become thrust and wave the large small instruction of engine target thrust and in formula (22), calculate, directly it is exported to and become thrust and wave the thrust variable valve of engine and control.

The control that change thrust is waved engine servo control mechanism pivot angle utilizes 5) the middle pid parameter and 4 calculating) the middle pitching obtaining and jaw channel attitude and the calculating of angular velocity error.

$\left[\begin{array}{c}{\mathrm{\&delta;}}_y\\ {\mathrm{\&delta;}}_z\end{array}\right]=\left[\begin{array}{c}{K}_{\mathrm{Py}}{\mathrm{\&theta;}}_{\mathrm{ERR},y}+K_{\mathrm{Iy}}\&Integral;{\mathrm{\&theta;}}_{\mathrm{ERR},y}\mathrm{dt}+K_{\mathrm{Dy}}{\mathrm{\&omega;}}_{\mathrm{ERR},y}\\ K_{\mathrm{Pz}}{\mathrm{\&theta;}}_{\mathrm{ERR},z}+K_{\mathrm{Iz}}\&Integral;{\mathrm{\&theta;}}_{\mathrm{ERR},z}\mathrm{dt}+K_{\mathrm{Dz}}{\mathrm{\&omega;}}_{\mathrm{ERR},z}\end{array}\right]---\left(27\right)$

The pitch axis pivot angle instruction δ generating _ywith yaw axis pivot angle instruction δ _zexport to and wave the execution of engine servo control mechanism.Simultaneously it also exports to 3) in formula (9) for next cycle, estimate that the change thrust of offsetting disturbance torque waves engine pivot angle.

The guidance, navigation and the control that have so just completed one-period are calculated.Constantly repeat 1 afterwards)～7), just formed whole flow processs of landing mission appearance rail coupling control algorithm.

Simulation analysis

The soft lunar landing detector of take is verified as example proposes appearance rail coupling control method to the present invention.Choose the power decline process of lander and study, take that to slow down be fundamental purpose this stage, do not carry out voyage control, the attitude angle of guidance system output is approximate at the uniform velocity to be changed, and becomes thrust and waves engine operation under full thrust rating.25.5 tons of lander initial weights, become thrust and wave 80,000 Ns of the full thrusts of engine, 20 ° of maximum pendulum angles.The barycenter of supposing lander in emulation departs from 0.1 meter of X-axis in Z direction.Adopting two kinds of methods to carry out soft landing guidance controls.Method one is that guidance, the control that current Apollo and Project Constellation adopt separates the method designing, and engine pivot angle is not compensated; Method two is that coupling design method is controlled in the guidance that the present invention proposes, and engine pivot angle is compensated.Simulation result as shown in Figure 3.Wherein dot-and-dash line is the flight path of reference, without the standard flight path under centroid motion (disturbance torque); Dotted line is the flight path that has method one under centroid motion; Solid line is the flight path that has method two under centroid motion.Can see, the flight path of method two is more close to reference to flight path.This shows that the method that the present invention proposes is effective.

Non-elaborated part of the present invention belongs to techniques well known.

Claims (3)

1. an appearance rail coupling control method for survey of deep space soft landing process, is characterized in that: comprise that step is as follows:

(1) setting up three-dimensional XYZ coordinate is, by X-axis be defined as three principal axis of inertia of lander the axis of rolling, by Y-axis be defined as three principal axis of inertia of lander pitch axis, Z axis is defined as to the yaw axis of three principal axis of inertia of lander, change thrust is waved engine and is arranged at along X-direction, becomes thrust and waves engine along lander Y, Z axis swing output variable thrust.

(2) according to the aimed acceleration direction and goal rate of acceleration change of lander guidance system output and posture control system output wave engine pivot angle controlled quentity controlled variable, carry out targeted attitude and the attitude angular velocity of attitude command plane-generating lander.

(3) under the three-dimensional system of coordinate of setting up in step (1), the current attitude of lander and attitude angular velocity measured value that the targeted attitude of lander that step (2) is generated and the navigational system of attitude angular velocity and lander provide compare, and are formed for attitude and angular velocity error, the attitude of pitch channel and the attitude of angular velocity error and jaw channel and the angular velocity error of the roll channel of lander three-axis attitude control.

(4), under the three-dimensional system of coordinate of setting up in step (1), according to the lander quality of the navigational system estimation of lander, calculate the centroid position of lander; The change thrust of being exported by lander guidance system is again waved motor power ratio and is calculated and become thrust and wave the large small instruction of engine target thrust, and the large small instruction of the target propulsive force Maximum controlling moment that can produce at pitching and yaw direction; According to the free-running frequency of the pitching of storing in the controller of Maximum controlling moment and lander and jaw channel and these closed-loop control system parameters of damping ratio, adjust in real time the pitching of lander and the pid control parameter of jaw channel and comprise controller proportional coefficient, integral item coefficient and differential term coefficient.

(5) under the three-dimensional system of coordinate of setting up in step (1), the roll channel attitude of calculating according to step (3) and attitude angular velocity error generate control moment instruction by proportional-integral-differential PID control method, control moment instruction is generated to instruction pulsewidth through pulse-width modulation PWM, the instruction pulsewidth of generation is sent to the output of roll channel attitude control engine, complete the attitude of the axis of rolling is controlled.

(6) under the three-dimensional system of coordinate of setting up in step (1), pitching and jaw channel attitude and the attitude angular velocity error according to step (3), calculated, the pid control parameter being generated by step (4), the pivot angle controlled quentity controlled variable of engine is waved in formation, this pivot angle controlled quentity controlled variable sends to waves engine, and this pivot angle controlled quentity controlled variable also feeds back to step (2) for correction target attitude.Wave engine and adjust engine pivot angle according to this pivot angle controlled quentity controlled variable, what according to step (4), generate waves the large small instruction output engine of engine target thrust thrust simultaneously, the attitude control moment of formation to pitching and jaw channel, completes the attitude of pitching and yaw axis is controlled.

2. the appearance rail coupling control method of a kind of survey of deep space soft landing process according to claim 1, it is characterized in that: the aimed acceleration direction of waving the output of engine pivot angle controlled quentity controlled variable correction guidance system of using in described step (2) posture control system of lander to generate, eliminated and waved the thrust direction that the control of engine appearance produces and change the impact on guidance.

3. the appearance rail coupling control method of a kind of survey of deep space soft landing process according to claim 1, it is characterized in that: in described step (4), the quality providing according to the navigational system of the thrust ratio of lander guidance system output and lander is estimated, calculate in real time and wave the control moment that engine produces, and the pid control parameter of pitching and jaw channel is carried out to online dynamic adjustment.