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CN110542423B - Moon soft landing vertical approach obstacle avoidance guidance method - Google Patents

Moon soft landing vertical approach obstacle avoidance guidance method Download PDF

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CN110542423B
CN110542423B CN201910668412.5A CN201910668412A CN110542423B CN 110542423 B CN110542423 B CN 110542423B CN 201910668412 A CN201910668412 A CN 201910668412A CN 110542423 B CN110542423 B CN 110542423B
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CN110542423A (en
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李骥
张洪华
关轶峰
程铭
张晓文
于萍
杨巍
于洁
王志文
王华强
王泽国
陈尧
赵宇
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Beijing Institute of Control Engineering
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Abstract

本发明一种月球软着陆垂直接近避障制导方法，步骤如下：1)设探测器制导指令计算周期为T，每N个制导指令计算周期进行一次制导参数更新；假设外部导航系统建立在惯性坐标系下，当前周期由导航系统提供的目标着陆点位置矢量为

Figure DDA0002140867150000011
探测器自身在惯性系的位置矢量为ri，速度矢量为vi；设计数器k是一个非负整数，初值为0；所述惯性坐标系用i表示，原点在月球中心，三个坐标轴在惯性空间指向固定方向；N≥1；2)以目标着陆点为中心，在空间中沿固定的方向建立制导坐标系，获得由惯性系向制导坐标系的旋转矩阵；3)解算得到制导参数；4)计算得到制导指令，并交给外部姿态控制系统和发动机执行。

Figure 201910668412

The present invention is a method for guiding the vertical approach of lunar soft landing and obstacle avoidance. Under the system, the position vector of the target landing point provided by the navigation system in the current cycle is

Figure DDA0002140867150000011
The position vector of the probe itself in the inertial system is ri, and the velocity vector is vi; the counter k is a non-negative integer, and the initial value is 0; the inertial coordinate system is represented by i , the origin is at the center of the moon, and the three coordinates are The axis points to a fixed direction in inertial space; N≥1; 2) Take the target landing point as the center, establish a guidance coordinate system along a fixed direction in space, and obtain the rotation matrix from the inertial system to the guidance coordinate system; 3) Solve to get Guidance parameters; 4) The guidance command is obtained by calculation and handed over to the external attitude control system and the engine for execution.

Figure 201910668412

Description

Moon soft landing vertical approach obstacle avoidance guidance method
Technical Field
The invention relates to a moon soft landing vertical approach obstacle avoidance guidance method, and belongs to the field of spacecraft guidance control.
Background
For soft landing of the moon, the terrain is an important factor endangering the landing safety. Therefore, in the descending and flying process, the distribution situation of the obstacle on the surface of the moon is observed, a safe landing point is searched, and the flight track is changed to implement obstacle avoidance. Existing landing probe barriers generally use an inclined descent trajectory, for example, apollo uses a descent trajectory having an angle of 16 ° to 24 ° with the horizontal plane, and Chang' e # uses a descent trajectory having an angle of 45 ° with the horizontal plane. This approach requires a relatively large flat area, which is advantageous for probes landing in the moon's area. However, this trajectory of descent is very unfavorable for landing missions that extend over the meteorite crater, to the south of the moon, to the back, etc. Firstly, for a detector navigation system depending on distance measurement relative measurement, the bumpy flight path can be aggravated by the fluctuant ground; secondly, there is a risk of accidental impact during descent in terrain with severe changes.
Therefore, for such rough terrain landing tasks, it is preferable to use a vertical approach descent trajectory. The advantages are that: firstly, the vertical projection position of the detector on the lunar surface is basically fixed when the detector vertically descends, so that the influence of terrain change is eliminated, and the method is favorable for the stability of a ranging correction navigation system; and secondly, when the landing platform descends vertically, the detector can observe the same landing area continuously and stably, and obstacle avoidance is facilitated. However, after the descending track is changed to be vertical, the original approach guidance method is not suitable any more, and the main problems include: firstly, a guidance coordinate system based on the direction of a target landing point relative to a detector can have the problem of rapid angle rotation when the guidance coordinate system vertically descends, and secondly, the guidance parameter resolving period is the same as the guidance instruction resolving period, so that guidance and attitude control self-oscillation easily occurs; and thirdly, after the landing points are updated, the guidance parameters cannot be updated in time, and the guidance response is slow.
Disclosure of Invention
The technical problem solved by the invention is as follows: the defects of the prior art are overcome, the lunar soft landing vertical approach obstacle avoidance guidance method is provided, and the safety landing requirement under the rugged terrain environment on the back of the moon or in the south pole area is met.
The technical scheme of the invention is as follows:
a moon soft landing vertical approach obstacle avoidance guidance method comprises the following steps:
1) setting a calculation period of the detector guidance instruction as T, and updating guidance parameters once in every N calculation periods of the guidance instruction; assuming that the external navigation system is established under an inertial coordinate system, the position vector of the target landing point provided by the navigation system in the current period is
Figure BDA0002140867130000021
The position vector of the detector in the inertial system is riVelocity vector is vi(ii) a Designing a counter k to be a non-negative integer, wherein the initial value of the counter k is 0; the inertial coordinate system is represented by i, the origin is at the center of the moon, and the three coordinate axes point to a fixed direction in the inertial space; n is more than or equal to 1;
2) establishing a guidance coordinate system in a fixed direction in space by taking a target landing point as a center, and obtaining a rotation matrix from an inertial system to the guidance coordinate system;
3) resolving to obtain a guidance parameter;
4) and calculating to obtain a guidance instruction, and handing the guidance instruction to an external attitude control system and an engine for execution.
The process of obtaining the rotation matrix from the inertial system to the guidance coordinate system in the step 2) is as follows:
according to the image processing of the navigation camera, finding a flat landing zone, and taking the central point of the landing zone as a new safe landing point, otherwise, keeping the original value of the safe landing point; if the updated safe landing point is obtained in the period, changing the current period into a guidance parameter resolving period; then, establishing a guidance coordinate system by taking the safe landing point as an origin and taking the local fixed direction as a reference, and further obtaining a rotation matrix of the inertial system to the guidance coordinate system;
the specific process of the step 2) is as follows:
setting the current target landing point position as
Figure BDA0002140867130000022
If the target landing point is a new safe landing point obtained again by the navigation and obstacle avoidance sensor, making k equal to 0;
establishing a guidance coordinate system by taking the safe landing point as a center, wherein the x-axis direction points to the safe landing point from the moon center and represents the local vertical direction; the two axes of y and z are in the local horizontal plane; with a predetermined reference direction p in spaceiRequiring the establishment of a z-axis and a vector p of the guidance coordinate systemiAnd if the included angle is minimum, the representation of three axes of the guidance coordinate system in the inertial space is as follows:
Figure BDA0002140867130000031
Figure BDA0002140867130000032
z＝x×y
rotation matrix from inertial system to guidance coordinate system
Figure BDA0002140867130000033
The calculation is as follows:
Figure BDA0002140867130000034
when the guidance parameters are obtained through resolving in the step 3), if the current period is a parameter resolving period, the position and speed parameters of the detector given by the navigation system are converted into a guidance coordinate system; calculating guidance time by taking the position, the speed and the acceleration of the vertical motion terminal as constraints and taking the change rate of the vertical acceleration of the terminal equal to 0 as a design target; calculating a guidance parameter according to the guidance time;
the specific process of the step 3) is as follows:
setting lead time tgoIndicates that the initial value is a number greater than 10;
(3.1) if t is satisfiedgo>And if k is equal to 0, updating the guidance parameters, specifically as follows:
firstly, converting the position and the speed of a detector into a guidance coordinate system, and acquiring the position r relative to a target landing point in the guidance coordinate systemgAnd velocity vg；
Figure BDA0002140867130000035
Figure BDA0002140867130000036
Wherein, ω ismIs the angular velocity of the moon rotating relative to the inertial space,
Figure BDA0002140867130000037
the method is a representation of the velocity direction vector of the self-rotation angle of the moon in an inertial system, and the velocity direction vector of the self-rotation angle of the moon is known quantity;
calculating to obtain the remaining guidance time:
setting the target acceleration vector of the lead terminal as
Figure BDA0002140867130000038
Target speed is
Figure BDA0002140867130000039
The target position vector is rt g(ii) a The three quantities mentioned above are the design values,
Figure BDA00021408671300000310
x component of (i.e. 1)
Figure BDA00021408671300000311
The value is larger than 0 and smaller than the difference between the acceleration generated by the maximum thrust of the engine and the gravity acceleration of the moon,
Figure BDA00021408671300000312
both the y and z components of (a) are 0;
Figure BDA00021408671300000313
x component of (i.e. 1)
Figure BDA00021408671300000314
Is a number not greater than 0,
Figure BDA00021408671300000315
both the y and z components of (a) are 0; r ist gX component of (1)
Figure BDA0002140867130000041
Is the terminal height in the vertical approach process, the value is a number greater than 0, rt gBoth the y and z components of (a) are 0;
setting the target of the vertical acceleration change rate of the terminal to be zero, and enabling
Figure BDA0002140867130000042
Figure BDA0002140867130000043
Wherein
Figure BDA0002140867130000044
Is vgThe x-component of (a) is,
Figure BDA0002140867130000045
is rgX component of (1), then the guidance time tgoThe calculation is as follows:
Figure BDA0002140867130000046
calculating updated guidance parameters c1，c2，c3：
Figure BDA0002140867130000047
Figure BDA0002140867130000048
Figure BDA0002140867130000049
(3.2) if t is not satisfiedgo>10 and k is 0, then:
tgo←tgo-T
the symbol "←" represents an assignment; guidance parameter c1，c2，c3And not updated.
The specific process of the step 4) is as follows:
let T be k.T, then command acceleration under guidance system
Figure BDA00021408671300000410
Is calculated as follows
Figure BDA00021408671300000411
Wherein, ggIs the gravity acceleration vector under the guidance system;
converting the command acceleration under the guidance system into the command acceleration under the inertia system to obtain
Figure BDA00021408671300000412
And output to an external attitude control system and an engine for execution, so that the longitudinal axis of the detector, namely the thrust direction of the engine and the thrust direction of the engine
Figure BDA00021408671300000413
Coincidence, acceleration due to engine output thrust, andtarget
Figure BDA00021408671300000414
Equal in size;
updating k ← k +1 by a counter k, and judging that k is 0 if k is larger than or equal to N;
judging the ending condition: if tgo<0, finishing the vertical approaching obstacle avoidance guidance, and returning to the step 1 in the next period).
Compared with the prior art, the invention has the beneficial effects that:
firstly, the establishing mode of the guidance coordinate system is modified, the guidance coordinate system is established in a fixed space direction by taking the target landing point as the center, and the large-scale rotation of the coordinate axis direction of the guidance coordinate system caused by the small-scale change of the detector relative to the direction of the target landing point in the vertical descending track is avoided.
Secondly, a guidance parameter updating period and a guidance instruction updating period are separated, so that guidance stability is improved;
thirdly, after the landing points are obtained again, the guidance parameters are immediately recalculated, and the response speed of the guidance law is improved.
Drawings
Fig. 1 is a structural diagram of a moon soft landing vertical approach obstacle avoidance guidance method.
Fig. 2 is a schematic diagram of guidance instruction output under a guidance system in a vertical approach obstacle avoidance process.
Fig. 3 is a schematic diagram of a motion trajectory in a vertical approaching obstacle avoidance process.
Detailed Description
As shown in fig. 1, the detailed process of the present invention is as follows:
1) obtaining external navigation data
Setting a calculation period of the detector guidance instruction as T, and updating guidance parameters once in every N calculation periods of the guidance instruction; assuming that the external navigation system is established under an inertial coordinate system, the position vector of the target landing point provided by the navigation system in the current period is
Figure BDA0002140867130000051
The position vector of the detector in the inertial system isriVelocity vector is vi(ii) a Designing a counter k to be a non-negative integer, wherein the initial value of the counter k is 0; the inertial coordinate system is represented by i, the origin is at the center of the moon, and the three coordinate axes point to a fixed direction in the inertial space; n is more than or equal to 1.
2) Establishing a guidance coordinate system
Setting the current target landing point position as
Figure BDA0002140867130000052
If the target landing point is a new safe landing point obtained again by the navigation and obstacle avoidance sensor, making k equal to 0;
establishing a guidance coordinate system by taking the safe landing point as a center, wherein the x-axis direction points to the safe landing point from the center of the moon, and the x-axis direction is the local vertical direction; the y and z axes are in the local horizontal plane, and the specific direction can be set according to the requirement: with a predetermined reference direction p in spaceiRequiring the establishment of a z-axis and a vector p of the guidance coordinate systemiAnd if the included angle is minimum, the representation of the three axes of the guidance coordinate system in the inertial space can be calculated as follows:
Figure BDA0002140867130000061
Figure BDA0002140867130000062
z＝x×y(3)
rotation matrix from inertial system to guidance coordinate system
Figure BDA0002140867130000063
Can be calculated as follows
Figure BDA0002140867130000064
3) Guidance parameter solution
T for guidance timegoThe initial value is a number greater than 10.
a) If t isgo>10 and k is 0, then
Firstly, converting the position and the speed of a detector into a guidance coordinate system, and acquiring the position r relative to a target landing point in the guidance coordinate systemgAnd velocity vg；
Figure BDA0002140867130000065
Figure BDA0002140867130000066
Wherein, ω ismIs the angular velocity of the moon rotating relative to the inertial space,
Figure BDA0002140867130000067
is the representation of the velocity direction vector of the self-rotation angle of the moon in an inertial system, and the velocity direction vector of the self-rotation angle of the moon and the inertial system are known quantities.
The remaining guidance time is then calculated.
Setting the target acceleration vector of the lead terminal as
Figure BDA0002140867130000068
Target speed is
Figure BDA0002140867130000069
The target position vector is rt g(ii) a The three quantities mentioned above are the design values,
Figure BDA00021408671300000610
x component of (i.e. 1)
Figure BDA00021408671300000611
The value is larger than 0 and smaller than the difference between the acceleration generated by the maximum thrust of the engine and the gravity acceleration of the moon,
Figure BDA00021408671300000612
both the y and z components of (a) are 0;
Figure BDA00021408671300000613
x component of (i.e. 1)
Figure BDA00021408671300000614
Is a number not greater than 0,
Figure BDA00021408671300000615
both the y and z components of (a) are 0; r ist gX component of (1)
Figure BDA00021408671300000616
Is the terminal height in the vertical approach process, the value is a number greater than 0, rt gBoth the y and z components of (a) are 0;
setting the target of the vertical acceleration change rate of the terminal to be zero, and enabling
Figure BDA00021408671300000617
Figure BDA00021408671300000618
Wherein
Figure BDA00021408671300000619
Is vgThe x-component of (a) is,
Figure BDA00021408671300000620
is rgX component of (1), then the guidance time tgoThe calculation is as follows:
Figure BDA0002140867130000071
calculating updated guidance parameters c1，c2，c3
Figure BDA0002140867130000072
Figure BDA0002140867130000073
Figure BDA0002140867130000074
b) If t is not satisfiedgo>10 and k is 0, then
tgo←tgo-T
The symbol "←" represents an assignment; guidance parameter c1，c2，c3And not updated.
4) Commanded acceleration calculation
Let T be k.T, then command acceleration under guidance system
Figure BDA0002140867130000075
Can be calculated as follows
Figure BDA0002140867130000076
Wherein, ggIs the gravity acceleration vector under the guidance system, and is known. Then, it is converted into inertia system to obtain
Figure BDA0002140867130000077
Figure BDA0002140867130000078
Then will be
Figure BDA0002140867130000079
The output is executed by an external attitude control system and an engine so that the longitudinal axis of the detector, namely the thrust direction of the engine and the thrust direction of the engine
Figure BDA00021408671300000710
Coincidence, acceleration due to engine output thrust, and
Figure BDA00021408671300000711
are equal in size.
Followed by an update of the counter k
k ← k +1, and when k is judged to be equal to or greater than N, k ═ 0
And finally, judging the ending condition: if tgo<0, finishing the vertical approaching obstacle avoidance guidance, and returning to the step 1 in the next period).
Simulation analysis
Assuming that a certain detector enters an approaching obstacle avoidance process at 3000m height at a vertical speed of-30 m/s and an upward speed direction as positive and a horizontal speed of 0m, an initial target landing point is right below the detector, and the value of a guidance terminal parameter is
Figure BDA00021408671300000712
rt g＝[3,0,0]T. The guidance instruction calculation period T is 0.1s, and the guidance parameter update period is 10 times the guidance instruction calculation period, that is, N is 10. When the probe descends to the height of 1500m, the target safe landing point is determined to be 180m away from the initial target landing point. The target acceleration vector approaching the descending process under the guidance system is shown in fig. 2, and after the safe landing point is updated, the guidance acceleration has sudden change with a certain amplitude, so that the original descending flight trend is changed; the corresponding flight trajectory is shown in fig. 3, the detector first descends in a vertical manner, and after the obstacle avoidance starts, the detector descends and translates to the position above the target safe landing point. Simulation results show that the moon soft landing vertical approach obstacle avoidance guidance method provided by the invention is effective.
Those skilled in the art will appreciate that the invention may be practiced without these specific details.

Claims (4)

1.一种月球软着陆垂直接近避障制导方法，其特征在于步骤如下：1. a lunar soft landing vertical approach obstacle avoidance guidance method is characterized in that the steps are as follows: 1)设探测器制导指令计算周期为T，每N个制导指令计算周期进行一次制导参数更新；假设外部导航系统建立在惯性坐标系下，当前周期由导航系统提供的目标着陆点位置矢量为
Figure FDA0002817263730000011
探测器自身在惯性系的位置矢量为ri，速度矢量为vi；设计数器k是一个非负整数，初值为0；所述惯性坐标系用i表示，原点在月球中心，三个坐标轴在惯性空间指向固定方向；N≥1；1) Set the detector guidance command calculation cycle to be T, and update the guidance parameters every N guidance command calculation cycles; assuming that the external navigation system is established in the inertial coordinate system, the target landing point position vector provided by the navigation system in the current cycle is
Figure FDA0002817263730000011
The position vector of the probe itself in the inertial system is ri, and the velocity vector is vi; the counter k is a non-negative integer, and the initial value is 0; the inertial coordinate system is represented by i , the origin is at the center of the moon, and the three coordinates are The axis points to a fixed direction in inertial space; N≥1; 2)以目标着陆点为中心，在空间中沿固定的方向建立制导坐标系，获得由惯性系向制导坐标系的旋转矩阵；2) Taking the target landing point as the center, establish the guidance coordinate system along a fixed direction in space, and obtain the rotation matrix from the inertial system to the guidance coordinate system; 3)解算得到制导参数；3) Solve to get the guidance parameters; 4)计算得到制导指令，并交给外部姿态控制系统和发动机执行；4) Calculate the guidance command and give it to the external attitude control system and engine for execution; 所述步骤2)获得由惯性系向制导坐标系的旋转矩阵的过程为：Said step 2) the process of obtaining the rotation matrix from the inertial system to the guidance coordinate system is: 根据导航相机图像处理，寻找到一块平坦的着陆区，将着陆区中心点作为新的安全着陆点，否则安全着陆点保持原值；如果本周期获得了更新的安全着陆点，则将当前周期改为制导参数解算周期；之后，以安全着陆点为原点，以当地固定方向为参考，建立制导坐标系，进而得到惯性系向制导坐标系的旋转矩阵；According to the image processing of the navigation camera, a flat landing area is found, and the center of the landing area is taken as the new safe landing point, otherwise the safe landing point remains the original value; if an updated safe landing point is obtained in this cycle, the current cycle will be changed to is the guidance parameter calculation period; after that, take the safe landing point as the origin and the local fixed direction as the reference to establish the guidance coordinate system, and then obtain the rotation matrix of the inertial system to the guidance coordinate system; 所述步骤2)的具体过程为：The concrete process of described step 2) is: 设当前的目标着陆点位置为
Figure FDA0002817263730000012
若该目标着陆点是由导航和避障敏感器重新获取的新的安全着陆点，则令k＝0；Let the current target landing point position be
Figure FDA0002817263730000012
If the target landing point is a new safe landing point re-acquired by the navigation and obstacle avoidance sensor, then let k=0; 以安全着陆点为中心建立制导坐标系，其中x轴方向由月心指向安全着陆点，表示当地的垂直方向；y、z两轴在当地水平面内；设空间中有一个预先设定的参考方向pi，要求建立的制导坐标系z轴与矢量pi夹角最小，则制导坐标系三个轴在惯性空间的表示如下：The guidance coordinate system is established with the safe landing point as the center, in which the x-axis direction points from the center of the moon to the safe landing point, indicating the local vertical direction; the y and z axes are in the local horizontal plane; there is a preset reference direction in the space p i , the minimum angle between the z-axis of the established guidance coordinate system and the vector pi is required, then the three axes of the guidance coordinate system are represented in inertial space as follows:
Figure FDA0002817263730000013
Figure FDA0002817263730000013
Figure FDA0002817263730000021
Figure FDA0002817263730000021
 z＝x×yz=x×y 由惯性系向制导坐标系的旋转矩阵
Figure FDA0002817263730000022
计算如下：Rotation matrix from inertial frame to guidance frame
Figure FDA0002817263730000022
The calculation is as follows:
Figure FDA0002817263730000023
Figure FDA0002817263730000023
 2.根据权利要求1所述的一种月球软着陆垂直接近避障制导方法，其特征在于：所述步骤3)解算得到制导参数时，若当前周期为参数解算周期，则将导航系统给出的探测器位置、速度参数转换到制导坐标系下；以垂向运动终端位置、速度和加速度为约束，并以终端垂向加速度的变化率等于0为设计目标，计算制导时间；根据制导时间计算制导参数。2. a kind of lunar soft landing vertical approach obstacle avoidance guidance method according to claim 1, is characterized in that: when described step 3) solution obtains guidance parameter, if current cycle is parameter solution cycle, then navigation system The given detector position and speed parameters are converted into the guidance coordinate system; the vertical motion terminal position, speed and acceleration are used as constraints, and the change rate of the terminal vertical acceleration is equal to 0 as the design goal, and the guidance time is calculated; according to the guidance Time to calculate guidance parameters. 3.根据权利要求2所述的一种月球软着陆垂直接近避障制导方法，其特征在于：所述步骤3)的具体过程为：3. a kind of lunar soft landing vertical approach obstacle avoidance guidance method according to claim 2, is characterized in that: the concrete process of described step 3) is: 设制导时间用tgo表示，初值为大于10的数；Let the guidance time be represented by t go , and the initial value is a number greater than 10; (3.1)如果满足tgo>10且k＝0，则进行制导参数更新，具体如下：(3.1) If t go >10 and k=0 are satisfied, update the guidance parameters, as follows: 首先，将探测器位置和速度转换到制导坐标系下，获得在制导坐标系下相对目标着陆点的位置rg和速度vg；First, convert the detector position and velocity into the guidance coordinate system, and obtain the position r g and velocity v g of the target landing point relative to the guidance coordinate system;
Figure FDA0002817263730000024
Figure FDA0002817263730000024
Figure FDA0002817263730000025
Figure FDA0002817263730000025
 其中，ωm是月球相对惯性空间旋转的角速度大小，
Figure FDA0002817263730000026
是月球自转角速度方向矢量在惯性系的表示，它们均为已知量；Among them, ω m is the angular velocity of the moon relative to the inertial space rotation,
Figure FDA0002817263730000026
is the representation of the direction vector of the angular velocity of the moon's rotation in the inertial frame, and they are all known quantities; 计算得到剩余制导时间：Calculate the remaining guidance time: 设制导终端的目标加速度矢量为
Figure FDA0002817263730000027
目标速度为
Figure FDA0002817263730000028
目标位置矢量为rt g；上述三个量是设计值，
Figure FDA0002817263730000029
的x分量，即
Figure FDA00028172637300000210
取值应大于0且小于发动机最大推力产生的加速度与月球重力加速度大小的差，
Figure FDA00028172637300000211
的y和z分量均为0；
Figure FDA00028172637300000212
的x分量，即
Figure FDA00028172637300000213
的取值为不大于0的数，
Figure FDA00028172637300000214
的y和z分量均为0；rt g中的x分量
Figure FDA0002817263730000031
是垂直接近过程的终端高度，取值是大于0的数，rt g的y和z分量均为0；Set the target acceleration vector of the guidance terminal as
Figure FDA0002817263730000027
The target speed is
Figure FDA0002817263730000028
The target position vector is r t g ; the above three quantities are design values,
Figure FDA0002817263730000029
the x component of , i.e.
Figure FDA00028172637300000210
The value should be greater than 0 and less than the difference between the acceleration generated by the maximum thrust of the engine and the gravitational acceleration of the moon,
Figure FDA00028172637300000211
The y and z components of are both 0;
Figure FDA00028172637300000212
the x component of , i.e.
Figure FDA00028172637300000213
The value of is a number not greater than 0,
Figure FDA00028172637300000214
The y and z components of are 0; the x component in r t g
Figure FDA0002817263730000031
is the terminal height of the vertical approach process, the value is a number greater than 0, and the y and z components of r t g are both 0; 设终端垂向加速度变化率的目标为零，令
Figure FDA0002817263730000032
Figure FDA0002817263730000033
其中
Figure FDA0002817263730000034
是vg的x分量，
Figure FDA0002817263730000035
是rg的x分量，则制导时间tgo计算如下：Set the target of the terminal vertical acceleration rate to zero, let
Figure FDA0002817263730000032
Figure FDA0002817263730000033
in
Figure FDA0002817263730000034
is the x component of v g ,
Figure FDA0002817263730000035
is the x component of r g , then the guidance time t go is calculated as follows:
Figure FDA0002817263730000036
Figure FDA0002817263730000036
 计算更新后的制导参数c1，c2，c3：Calculate the updated guidance parameters c 1 , c 2 , c 3 :
Figure FDA0002817263730000037
Figure FDA0002817263730000037
Figure FDA0002817263730000038
Figure FDA0002817263730000038
Figure FDA0002817263730000039
Figure FDA0002817263730000039
 (3.2)如果不满足tgo>10且k＝0，则：(3.2) If t go >10 and k=0 are not satisfied, then: tgo←tgo-Tt go ←t go -T 符号“←”表示赋值；制导参数c1，c2，c3不更新。The symbol "←" indicates assignment; the guidance parameters c 1 , c 2 , and c 3 are not updated. 4.根据权利要求3所述的一种月球软着陆垂直接近避障制导方法，其特征在于：所述步骤4)的具体过程为：4. a kind of lunar soft landing vertical approach obstacle avoidance guidance method according to claim 3, is characterized in that: the concrete process of described step 4) is: 令t＝k·T，则制导系下的指令加速度
Figure FDA00028172637300000310
计算如下Let t=k·T, then the commanded acceleration under the guidance system
Figure FDA00028172637300000310
Calculated as follows
Figure FDA00028172637300000311
Figure FDA00028172637300000311
 其中，gg是制导系下的重力加速度矢量；Among them, g g is the gravitational acceleration vector under the guidance system; 将制导系下的指令加速度转换到惯性系下得到
Figure FDA00028172637300000312
并输出给外部姿态控制系统和发动机执行，使得探测器纵轴，即发动机推力方向与
Figure FDA00028172637300000313
重合，发动机输出推力产生的加速度与目标
Figure FDA00028172637300000314
大小相等；Convert the command acceleration in the guidance system to the inertial system to get
Figure FDA00028172637300000312
And output to the external attitude control system and the engine for execution, so that the longitudinal axis of the detector, that is, the thrust direction of the engine is the same as the
Figure FDA00028172637300000313
Coincidence, the acceleration generated by the engine output thrust and the target
Figure FDA00028172637300000314
equal in size; 计数器k更新k←k+1，判断若k≥N，则k＝0；The counter k is updated with k←k+1, and it is judged that if k≥N, then k=0; 结束条件判断：若tgo<0，则垂直接近避障制导结束，否则在下一周期重新回到步骤1)。End condition judgment: if t go < 0, the vertical approach to obstacle avoidance guidance ends, otherwise it will return to step 1) in the next cycle.
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