CN103466100A - Lander soft landing posture control method - Google Patents
Lander soft landing posture control method Download PDFInfo
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- CN103466100A CN103466100A CN2013103728270A CN201310372827A CN103466100A CN 103466100 A CN103466100 A CN 103466100A CN 2013103728270 A CN2013103728270 A CN 2013103728270A CN 201310372827 A CN201310372827 A CN 201310372827A CN 103466100 A CN103466100 A CN 103466100A
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Abstract
The invention discloses a lander soft landing posture control method. The method includes when posture of a target changes greatly, designing PI (proportion integration) control on the basis of nominal angular speed, and when the posture quickly approaches the target, avoiding angular speed measuring saturation; when angular speed deviation decreases, directly assigning an integral term to angular speed deviation PID (proportion integration differentiation) control by disturbance moment estimated by the integral term, and controlling switching to realize smooth transition. The lander soft landing posture control method adapts to the characteristic of changing of the disturbance moment during soft landing of a lander, so that steady-state posture control accuracy is improved; on the basis of the posture control method and aiming at the characteristic of posture jumping of a guided target during soft landing, a posture out-of-tolerance judging method is designed, so that misdiagnosis caused by conventional methods is avoided. The lander soft landing posture control method is simple, reliable and high in fault positioning accuracy rate.
Description
Technical field
The present invention relates to a kind of lander soft landing attitude control method, relate in particular to the landing attitude control method of lander during a kind of soft landing, belong to the Spacecraft Attitude Control technical field.
Background technology
Attitude maneuver schematic diagram during the detector soft landing as shown in Figure 1, guidance system provides targeted attitude according to the detector position velocity information, attitude control system is determined the current attitude provided according to attitude, control detector and trend towards targeted attitude, and final vertical landing is at moonscape.
The Apollo lunar spacecraft of the U.S. and lunar spacecraft of new generation, attitude is controlled and is all adopted the phase plane control method, for example 2006, AIAA2006-6564, the disclosed a kind of attitude control method of " A Parameterized Approach to the Design of Lunar Lander Attitude Controllers " literary composition that Michael C.Johnson delivers, in the position of phase plane, calculate the attitude controlling jetting pulsewidth according to attitude angle and cireular frequency.2008, AIAA2008-6812, " Robust Digital Autopilot Design for Spacecraft Equipped with Pulse-Operated Thrusters " literary composition that Paul B.Brugarolas etc. delivers discloses the appearance control method that a kind of PD used in the Reentry section for MSL adds pulse duration modulation.The weak point of above-mentioned appearance control control algorithm is: when there is normal value disturbance torque in system, stabilization of carriage angle is near the attitude dead band, and control accuracy is poor.In June, 1998, the U.S. controls meeting paper and concentrates, " A State-Space Fault Monitor Architecture and It ' s Application to the Cassini Spacecraft " literary composition that Glenn A.Macala etc. delivers discloses the overproof decision method of a kind of attitude, when angle with cireular frequency is equidirectional and when larger, think that attitude is overproof.The weak point of the overproof decision algorithm of above-mentioned attitude is: when angle and cireular frequency reversing sense and when larger, attitude is overproof obviously, yet this algorithm can not determine the overproof fault of attitude.
Summary of the invention
Technology of the present invention is dealt with problems and is: overcome the deficiencies in the prior art, control for the attitude of lander during soft landing, propose a kind of lander soft landing attitude control method, improved the stable state attitude control accuracy, the fault localization accuracy rate is high.
Technical solution of the present invention is: a kind of lander soft landing attitude control method, and step is as follows:
(1) measure attitude angle θ and the attitude angular velocity d θ of lander in each control cycle, the attitude angular deviation θ e=θ of acquisition lander-θ r, the wherein object attitude angle of θ r for being calculated by Guidance Law in each control cycle;
(2) determine attitude control partition angle threshold θ m, θ m=Kd * d θ r/Kp, wherein: d θ r is nominal angle speed, d θ r=0.5 * ω
max, ω
maxfor the maxim of angular velocity measurement, the pid control parameter that Kd, Kp are angular deviation, Kd=2 * J * ω * ξ, Kp=J * ω
2, J is that lander inertia, ω are that controller frequency, ξ are damping coefficient;
(3) size of judgement attitude angular deviation θ e and attitude control partition angle threshold θ m, when the absolute value of attitude angular deviation θ e is less than attitude control partition angle threshold θ m, phase point (the θ e formed by attitude angular velocity d θ, attitude angular deviation θ e, d θ) be positioned at angle and cireular frequency control area, I=I+Ki * (θ-θ r) * Δ t, Y=Kp * (θ-θ r)+Kd * d θ+I, wherein I is the integration item, Δ t is control cycle, the PID control coefficient that Ki is angular deviation, Ki=0.05 * Kp, Y is control torque;
When the absolute value of attitude angular deviation θ e is more than or equal to attitude control partition angle threshold θ m, phase point (θ e, d θ) be positioned at nominal angle speed control area, I=I+Kiv * (d θ-d θ r) * Δ t, Y=Kpv * (d θ-d θ r)+I, wherein, Kpv, Kiv are respectively the PI control coefficient of cireular frequency deviation, Kpv=Kd, Kiv=0.25 * Kp;
(4) the control torque Y obtained according to step (3) calculates the jet pulsewidth PW of attitude control thruster of each control cycle, and concrete method of calculating is:
(a) in each control cycle, control torque Y is carried out to normalization method, W=|Y|/Y
max, wherein W is normalized value, Y
maxoutput torque for attitude control thruster;
(b) if W>1+H2, normalization method pulsewidth WC=1, pulse width threshold P
on=5000, wherein H2 is the pulse duration modulation parameter;
If H1<W≤1+H2, WC=(W-H1)/(1+H2-H1), P
on=Tm * (H1-H2)/(1+H2-H1), wherein Tm, H1 are the pulse duration modulation parameter;
If W≤H1, WC=0, P
on=t
min, t wherein
minminimum pulse width for attitude control thruster;
(c) the normalization method pulsewidth WC under the different condition obtained according to step (b) calculates the accumulative total pulsewidth PWI of current control cycle, PWI=PWI0 – WC * Δ t * sgn (Y) wherein, the residue pulsewidth that PWI0 is a upper control cycle, PWI0 equals the accumulative total pulsewidth of a control cycle and the difference of the jet pulsewidth of attitude control thruster, sgn () is symbolic function, parameter in bracket is greater than 0, be output as 1, parameter in bracket is less than 0, be output as-1, parameter in bracket equals 0, is output as 0;
If | PWI| > Δ t, the jet pulsewidth PW=Δ of attitude control thruster t * sgn (PWI);
If | PWI|<P
on, the jet pulsewidth PW=0 of attitude control thruster;
If P
on≤ | PWI|≤Δ t, the jet pulsewidth PW=PWI of attitude control thruster;
(5) at each control cycle, the jet pulsewidth PW of attitude control thruster that lander computer controlled attitude control thruster calculates according to step (4) carries out jet, realizes the attitude of lander is controlled.
Also comprise afterwards the step of the overproof judgement of lander attitude in step (5), specific implementation is:
(1) establish a thruster and often open fault, obtain thruster and often open fault cireular frequency d θ g and the fault angular deviation θ eg under fault, according to fault cireular frequency d θ g and fault angular deviation θ eg, calculate fault zone parameter θ
aF2=θ eg+5 °, d θ
aF2g+3 °/s of=d θ, θ
aF1=θ
aF2-d θ
aF2, d θ
aF1=-1 °/s, wherein θ
aF1, θ
aF2for the fault zone angle parameter, d θ
aF1, d θ
aF2for fault zone cireular frequency parameter;
(2) each control cycle judges that whether phase point (θ e, the d θ) region consisted of attitude angular velocity d θ, attitude angular deviation θ e is in fault zone, when | d θ |>=d θ
aF2, θ e>=θ
aF2and d θ>=d θ
aF1, θ
e≤-θ
aF2and d θ≤-d θ
aF1, θ
e+ d θ>=θ
aF1+ d θ
aF2and d θ>=0, θ e+d θ≤-θ
aF1-d θ
aF2and d θ≤0, phase point (θ e, d θ) is in fault zone;
(3) if phase point (θ e, d θ) is positioned at fault zone, and n control cycle set up continuously, judge that the attitude of lander is overproof, otherwise the lander attitude is normal, and the span of coefficient n is 5～10.
Described pulse duration modulation parameter Tm, H1 meet Tm * (H1-H2)=t simultaneously
min, H1=Kp * θ
min/ Y
max, θ
minfor attitude control accuracy.
The present invention's beneficial effect compared with prior art is:
(1) attitude control method of the present invention, when the larger variation of targeted attitude, the PI of design based on nominal angle speed controls, and when attitude is approached target fast, avoids angular velocity measurement saturated; The disturbance torque that simultaneously the integration item estimates, the integration item that indirect assignment is controlled to angular deviation PID when angular deviation reduces, control the switching smooth transition, and the present invention adapts to during the lander soft landing to become dry and disturbs the characteristics of moment, has improved stable state appearance control precision.
(2) the present invention is based on above-mentioned attitude control method, the characteristics for guidance targeted attitude saltus step during soft landing, designed the overproof decision method of this attitude, the wrong diagnosis situation of avoiding conventional approach to occur, and the method is simple and reliable, and the fault localization accuracy rate is high.
The accompanying drawing explanation
Fig. 1 is the attitude maneuver schematic diagram during detector soft landing of the present invention;
Fig. 2 is attitude control flow chart of the present invention;
Fig. 3 is that attitude subregion of the present invention is controlled schematic diagram;
Fig. 4 is the overproof judgement of attitude of the present invention fault zone schematic diagram.
The specific embodiment
As shown in Figure 2, performing step of the present invention is as follows:
(1) measure attitude angle θ and the attitude angular velocity d θ of lander in each control cycle, the attitude angular deviation θ e=θ of acquisition lander-θ r, the wherein object attitude angle of θ r for being calculated by Guidance Law in each control cycle;
(2) determine attitude control partition angle threshold θ m, θ m=Kd * d θ r/Kp, wherein: d θ r is nominal angle speed, d θ r=0.5 * ω
max, ω
maxfor the maxim of angular velocity measurement, the pid control parameter that Kd, Kp are angular deviation, Kd=2 * J * ω * ξ, Kp=J * ω
2, J is that lander inertia, ω are that controller frequency, ξ are damping coefficient;
(3) size of judgement attitude angular deviation θ e and attitude control partition angle threshold θ m, as shown in Figure 3, when the absolute value of attitude angular deviation θ e is less than attitude control partition angle threshold θ m, phase point (the θ e formed by attitude angular velocity d θ, attitude angular deviation θ e, d θ) be positioned at angle and cireular frequency control area, I=I+Ki * (θ-θ r) * Δ t, Y=Kp * (θ-θ r)+Kd * d θ+I, wherein I is the integration item, Δ t is control cycle, the PID control coefficient that Ki is angular deviation, Ki=0.05 * Kp, Y is control torque;
When the absolute value of attitude angular deviation θ e is more than or equal to attitude control partition angle threshold θ m, phase point (θ e, d θ) be positioned at nominal angle speed control area, I=I+Kiv * (d θ-d θ r) * Δ t, Y=Kpv * (d θ-d θ r)+I, wherein, Kpv, Kiv are respectively the PI control coefficient of cireular frequency deviation, Kpv=Kd, Kiv=0.25 * Kp;
(4) the control torque Y obtained according to step (3) calculates the jet pulsewidth PW of attitude control thruster of each control cycle, and concrete method of calculating is:
(a) in each control cycle, control torque Y is carried out to normalization method, W=|Y|/Y
max, wherein W is normalized value, Y
maxoutput torque for attitude control thruster;
(b) if W>1+H2, normalization method pulsewidth WC=1, pulse width threshold P
on=5000, wherein H2 is the pulse duration modulation parameter;
If H1<W≤1+H2, WC=(W-H1)/(1+H2-H1), P
on=Tm * (H1-H2)/(1+H2-H1), wherein Tm, H1 are the pulse duration modulation parameter;
If W≤H1, WC=0, P
on=t
min, t wherein
minminimum pulse width for attitude control thruster;
(c) the normalization method pulsewidth WC under the different condition obtained according to step (b) calculates the accumulative total pulsewidth PWI of current control cycle, PWI=PWI0 – WC * Δ t * sgn (Y) wherein, the residue pulsewidth that PWI0 is a upper control cycle, PWI0 equals the accumulative total pulsewidth of a control cycle and the difference of the jet pulsewidth of attitude control thruster, sgn () is symbolic function, parameter in bracket is greater than 0, be output as 1, parameter in bracket is less than 0, be output as-1, parameter in bracket equals 0, is output as 0;
If | PWI| > Δ t, the jet pulsewidth PW=Δ of attitude control thruster t * sgn (PWI);
If | PWI|<P
on, the jet pulsewidth PW=0 of attitude control thruster;
If P
on≤ | PWI|≤Δ t, the jet pulsewidth PW=PWI of attitude control thruster;
(5) at each control cycle, the jet pulsewidth PW of attitude control thruster that lander computer controlled attitude control thruster calculates according to step (4) carries out jet, realizes the attitude of lander is controlled.
Step (5) also comprises the step of the overproof judgement of lander attitude afterwards, and specific implementation is:
(1) establish a thruster and often open fault, obtain thruster and often open fault cireular frequency d θ g and the fault angular deviation θ eg under fault, according to fault cireular frequency d θ g and fault angular deviation θ eg, calculate fault zone parameter θ
aF2=θ eg+5 °, d θ
aF2g+3 °/s of=d θ, θ
aF1=θ
aF2-d θ
aF2, d θ
aF1=-1 °/s, wherein θ
aF1, θ
aF2for the fault zone angle parameter, d θ
aF1, d θ
aF2for fault zone cireular frequency parameter;
(2) as shown in Figure 4, each control cycle judges that whether phase point (θ e, the d θ) region consisted of attitude angular velocity d θ, attitude angular deviation θ e is in fault zone, when | d θ |>=d θ
aF2, θ e>=θ
aF2and d θ>=d θ
aF1, θ e≤-θ
aF2and d θ≤-d θ
aF1, θ e+d θ>=θ
aF1+ d θ
aF2and d θ>=0, θ e+d θ≤-θ
aF1-d θ
aF2and d θ≤0, phase point (θ e, d θ) is in fault zone;
(3) if phase point (θ e, d θ) is positioned at fault zone, and n control cycle set up continuously, judge that the attitude of lander is overproof, otherwise the lander attitude is normal.The span of coefficient n is 5～10.
Pulse duration modulation parameter Tm, H1 meet Tm * (H1-H2)=t simultaneously
min, H1=Kp * θ
min/ Y
max, θ
minfor attitude control accuracy.
Embodiment
(1) lander inertia J=2000kgm
2, ω=1rad/s, ξ=0.7: Kp=2000, Kd=2800, Ki=100, Kpv=2800, Kiv=500.
(2) appearance control thruster moment Y
max=240Nm, minimum pulse width t
min=0.04s, appearance control precision θ
min=0.2 °: design Tm=2, H1=0.03, H2=0.01.
(3) the maximum capacity ω of angular velocity measurement
max=6 °/s: r=3 °/s of d θ, θ m=4.2 °.
(4) according to above-mentioned control parameter, obtain a thruster and often open the g=3 °/s of d θ under fault, θ eg=4 °: design θ
aF1=3 °, θ
aF2=9 °, d θ
aF1=-1 °/s, d θ
aF2=6 °/s.
The content be not described in detail in specification sheets of the present invention belongs to those skilled in the art's known technology.
Claims (4)
1. a lander soft landing attitude control method is characterized in that step is as follows:
(1) measure attitude angle θ and the attitude angular velocity d θ of lander in each control cycle, the attitude angular deviation θ e=θ of acquisition lander-θ r, the wherein object attitude angle of θ r for being calculated by Guidance Law in each control cycle;
(2) determine attitude control partition angle threshold θ m, θ m=Kd * d θ r/Kp, wherein: d θ r is nominal angle speed, d θ r=0.5 * ω
max, ω
maxfor the maxim of angular velocity measurement, the pid control parameter that Kd, Kp are angular deviation, Kd=2 * J * ω * ξ, Kp=J * ω
2, J is that lander inertia, ω are that controller frequency, ξ are damping coefficient;
(3) size of judgement attitude angular deviation θ e and attitude control partition angle threshold θ m, when the absolute value of attitude angular deviation θ e is less than attitude control partition angle threshold θ m, phase point (the θ e formed by attitude angular velocity d θ, attitude angular deviation θ e, d θ) be positioned at angle and cireular frequency control area, I=I+Ki * (θ-θ r) * Δ t, Y=Kp * (θ-θ r)+Kd * d θ+I, wherein I is the integration item, Δ t is control cycle, the PID control coefficient that Ki is angular deviation, Ki=0.05 * Kp, Y is control torque;
When the absolute value of attitude angular deviation θ e is more than or equal to attitude control partition angle threshold θ m, phase point (θ e, d θ) be positioned at nominal angle speed control area, I=I+Kiv * (d θ-d θ r) * Δ t, Y=Kpv * (d θ-d θ r)+I, wherein, Kpv, Kiv are respectively the PI control coefficient of cireular frequency deviation, Kpv=Kd, Kiv=0.25 * Kp;
(4) the control torque Y obtained according to step (3) calculates the jet pulsewidth PW of attitude control thruster of each control cycle, and concrete method of calculating is:
(a) in each control cycle, control torque Y is carried out to normalization method, W=|Y|/Y
max, wherein W is normalized value, Y
maxoutput torque for attitude control thruster;
(b) if W>1+H2, normalization method pulsewidth WC=1, pulse width threshold P
on=5000, wherein H2 is the pulse duration modulation parameter;
If H1<W≤1+H2, WC=(W-H1)/(1+H2-H1), P
on=Tm * (H1-H2)/(1+H2-H1), wherein Tm, H1 are the pulse duration modulation parameter;
If W≤H1, WC=0, P
on=t
min, t wherein
minminimum pulse width for attitude control thruster;
(c) the normalization method pulsewidth WC under the different condition obtained according to step (b) calculates the accumulative total pulsewidth PWI of current control cycle, PWI=PWI0 – WC * Δ t * sgn (Y) wherein, the residue pulsewidth that PWI0 is a upper control cycle, PWI0 equals the accumulative total pulsewidth of a control cycle and the difference of the jet pulsewidth of attitude control thruster, sgn () is symbolic function, parameter in bracket is greater than 0, be output as 1, parameter in bracket is less than 0, be output as-1, parameter in bracket equals 0, is output as 0;
If | PWI| > Δ t, the jet pulsewidth PW=Δ of attitude control thruster t * sgn (PWI);
If | PWI|<P
on, the jet pulsewidth PW=0 of attitude control thruster;
If P
on≤ | PWI|≤Δ t, the jet pulsewidth PW=PWI of attitude control thruster;
(5) at each control cycle, the jet pulsewidth PW of attitude control thruster that lander computer controlled attitude control thruster calculates according to step (4) carries out jet, realizes the attitude of lander is controlled.
2. a kind of lander soft landing attitude control method according to claim 1, it is characterized in that: also comprise afterwards the step of the overproof judgement of lander attitude in step (5), specific implementation is:
(1) establish a thruster and often open fault, obtain thruster and often open fault cireular frequency d θ g and the fault angular deviation θ eg under fault, according to fault cireular frequency d θ g and fault angular deviation θ eg, calculate fault zone parameter θ
aF2=θ eg+5 °, d θ
aF2g+3 °/s of=d θ, θ
aF1=θ
aF2-d θ
aF2, d θ
aF1=-1 °/s, wherein θ
aF1, θ
aF2for the fault zone angle parameter, d θ
aF1, d θ
aF2for fault zone cireular frequency parameter;
(2) each control cycle judges that whether phase point (θ e, the d θ) region consisted of attitude angular velocity d θ, attitude angular deviation θ e is in fault zone, when | d θ |>=d θ
aF2, θ e>=θ
aF2and d θ>=d θ
aF1, θ e≤-θ
aF2and d θ≤-d θ
aF1, θ e+d θ>=θ
aF1+ d θ
aF2and d θ>=0, θ e+d θ≤-θ
aF1-d θ
aF2and d θ≤0, phase point (θ e, d θ) is in fault zone;
(3) if phase point (θ e, d θ) is positioned at fault zone, and n control cycle set up continuously, judge that the attitude of lander is overproof, otherwise the lander attitude is normal.
3. a kind of lander soft landing attitude control method according to claim 2 is characterized in that: in described step (3), the span of coefficient n is 5～10.
4. a kind of lander soft landing attitude control method according to claim 1, it is characterized in that: described pulse duration modulation parameter Tm, H1 meet Tm * (H1-H2)=t simultaneously
min, H1=Kp * θ
min/ Y
max, θ
minfor attitude control accuracy.
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| CN111458150B (en) * | 2020-03-31 | 2021-11-16 | 上海航天控制技术研究所 | High-reliability rail-controlled thruster fault discrimination method based on adding table |
| CN111458150A (en) * | 2020-03-31 | 2020-07-28 | 上海航天控制技术研究所 | High-reliability rail-controlled thruster fault discrimination method based on adding table |
| CN112319860A (en) * | 2021-01-05 | 2021-02-05 | 北京航空航天大学 | Self-adaptive compensation PWPF modulation method and device for RCS of aircraft |
| CN114019793A (en) * | 2021-10-08 | 2022-02-08 | 北京控制工程研究所 | Mars EDL process robust attitude control method |
| CN114030654A (en) * | 2021-10-08 | 2022-02-11 | 北京控制工程研究所 | Atmosphere entering attitude control method based on pulse width modulation |
| CN114030654B (en) * | 2021-10-08 | 2023-06-06 | 北京控制工程研究所 | Atmospheric entry attitude control method based on pulse width modulation |
| CN114019793B (en) * | 2021-10-08 | 2024-02-09 | 北京控制工程研究所 | Mars EDL process robust attitude control method |
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