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

Moon soft landing vertical approach obstacle avoidance guidance method Download PDF

Info

Publication number
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
Authority
CN
China
Prior art keywords
guidance
coordinate system
inertial
moon
target
Prior art date
Legal status (The legal status is an assumption and is not a legal conclusion. Google has not performed a legal analysis and makes no representation as to the accuracy of the status listed.)
Active
Application number
CN201910668412.5A
Other languages
Chinese (zh)
Other versions
CN110542423A (en
Inventor
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
Current Assignee (The listed assignees may be inaccurate. Google has not performed a legal analysis and makes no representation or warranty as to the accuracy of the list.)
Beijing Institute of Control Engineering
Original Assignee
Beijing Institute of Control Engineering
Priority date (The priority date is an assumption and is not a legal conclusion. Google has not performed a legal analysis and makes no representation as to the accuracy of the date listed.)
Filing date
Publication date
Application filed by Beijing Institute of Control Engineering filed Critical Beijing Institute of Control Engineering
Priority to CN201910668412.5A priority Critical patent/CN110542423B/en
Publication of CN110542423A publication Critical patent/CN110542423A/en
Application granted granted Critical
Publication of CN110542423B publication Critical patent/CN110542423B/en
Active legal-status Critical Current
Anticipated expiration legal-status Critical

Links

Images

Classifications

    • G—PHYSICS
    • G01—MEASURING; TESTING
    • G01C—MEASURING DISTANCES, LEVELS OR BEARINGS; SURVEYING; NAVIGATION; GYROSCOPIC INSTRUMENTS; PHOTOGRAMMETRY OR VIDEOGRAMMETRY
    • G01C21/00—Navigation; Navigational instruments not provided for in groups G01C1/00 - G01C19/00
    • G01C21/24—Navigation; Navigational instruments not provided for in groups G01C1/00 - G01C19/00 specially adapted for cosmonautical navigation

Landscapes

  • Engineering & Computer Science (AREA)
  • Radar, Positioning & Navigation (AREA)
  • Remote Sensing (AREA)
  • Physics & Mathematics (AREA)
  • Astronomy & Astrophysics (AREA)
  • Automation & Control Theory (AREA)
  • General Physics & Mathematics (AREA)
  • Traffic Control Systems (AREA)
  • Navigation (AREA)

Abstract

The invention discloses a moon soft landing vertical approach obstacle avoidance guidance method, which 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 DDA0002140867150000011
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.

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. A moon soft landing vertical approach obstacle avoidance guidance method is characterized by comprising 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 periodMeasured as
Figure FDA0002817263730000011
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) calculating to obtain a guidance instruction, and delivering 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 FDA0002817263730000012
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 systemiMinimum included angle, the systemThe three axes of the lead frame are represented in inertial space as follows:
Figure FDA0002817263730000013
Figure FDA0002817263730000021
z＝x×y
rotation matrix from inertial system to guidance coordinate system
Figure FDA0002817263730000022
The calculation is as follows:
Figure FDA0002817263730000023
2. the moon soft landing vertical approach obstacle avoidance guidance method according to claim 1, characterized in that: 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; and calculating the guidance parameters according to the guidance time.
3. The moon soft landing vertical approach obstacle avoidance guidance method according to claim 2, characterized in that: 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, the position and the speed of the detector are converted into a guidance coordinate system to obtainObtaining the position r relative to the target landing point under the guidance coordinate systemgAnd velocity vg；
Figure FDA0002817263730000024
Figure FDA0002817263730000025
Wherein, ω ismIs the angular velocity of the moon rotating relative to the inertial space,
Figure FDA0002817263730000026
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 FDA0002817263730000027
Target speed is
Figure FDA0002817263730000028
The target position vector is rt g(ii) a The three quantities mentioned above are the design values,
Figure FDA0002817263730000029
x component of (i.e. 1)
Figure FDA00028172637300000210
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 FDA00028172637300000211
both the y and z components of (a) are 0;
Figure FDA00028172637300000212
x component of (i.e. 1)
Figure FDA00028172637300000213
Is a number not greater than 0,
Figure FDA00028172637300000214
both the y and z components of (a) are 0; r ist gX component of (1)
Figure FDA0002817263730000031
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 FDA0002817263730000032
Figure FDA0002817263730000033
Wherein
Figure FDA0002817263730000034
Is vgThe x-component of (a) is,
Figure FDA0002817263730000035
is rgX component of (1), then the guidance time tgoThe calculation is as follows:
Figure FDA0002817263730000036
calculating updated guidance parameters c1，c2，c3：
Figure FDA0002817263730000037
Figure FDA0002817263730000038
Figure FDA0002817263730000039
(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.
4. The moon soft landing vertical approach obstacle avoidance guidance method according to claim 3, characterized in that: the specific process of the step 4) is as follows:
let T be k.T, then command acceleration under guidance system
Figure FDA00028172637300000310
Is calculated as follows
Figure FDA00028172637300000311
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 FDA00028172637300000312
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 FDA00028172637300000313
Coincidence, acceleration due to engine output thrust and target
Figure FDA00028172637300000314
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).
CN201910668412.5A 2019-07-23 2019-07-23 Moon soft landing vertical approach obstacle avoidance guidance method Active CN110542423B (en)

Priority Applications (1)

Application Number Priority Date Filing Date Title
CN201910668412.5A CN110542423B (en) 2019-07-23 2019-07-23 Moon soft landing vertical approach obstacle avoidance guidance method

Applications Claiming Priority (1)

Application Number Priority Date Filing Date Title
CN201910668412.5A CN110542423B (en) 2019-07-23 2019-07-23 Moon soft landing vertical approach obstacle avoidance guidance method

Publications (2)

Publication Number Publication Date
CN110542423A CN110542423A (en) 2019-12-06
CN110542423B true CN110542423B (en) 2021-06-11

Family

ID=68709792

Family Applications (1)

Application Number Title Priority Date Filing Date
CN201910668412.5A Active CN110542423B (en) 2019-07-23 2019-07-23 Moon soft landing vertical approach obstacle avoidance guidance method

Country Status (1)

Country Link
CN (1) CN110542423B (en)

Citations (7)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
US7967255B2 (en) * 2006-07-27 2011-06-28 Raytheon Company Autonomous space flight system and planetary lander for executing a discrete landing sequence to remove unknown navigation error, perform hazard avoidance and relocate the lander and method
CN103662096A (en) * 2013-12-13 2014-03-26 北京控制工程研究所 Self-adaptation powered explicit guidance method
CN103662091A (en) * 2013-12-13 2014-03-26 北京控制工程研究所 High-precision safe landing guiding method based on relative navigation
CN106527473A (en) * 2016-10-27 2017-03-22 上海航天控制技术研究所 Obstacle-avoidance landing method on lunar surface
CN108594802A (en) * 2018-02-28 2018-09-28 北京控制工程研究所 The determination of detector target touchdown area and avoidance method of guidance and device
CN108984907A (en) * 2018-07-18 2018-12-11 哈尔滨工业大学 A kind of interative guidance method based on yaw corner condition
CN109018440A (en) * 2018-07-23 2018-12-18 哈尔滨工业大学 VTOL carrier rocket grade landing phase precise perpendicularity soft landing quartic polynomial method of guidance

Family Cites Families (4)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
SU1760868A1 (en) * 1990-02-08 1995-09-27 Всесоюзный научно-исследовательский институт радиоаппаратуры Device to measure angular coordinate of landing system
US7216069B2 (en) * 2001-01-19 2007-05-08 Honeywell International, Inc. Simulated visual glideslope indicator on aircraft display
CN105929844B (en) * 2016-04-26 2019-01-08 哈尔滨工业大学 Barrier-avoiding method under a kind of more Obstacles Constraints environment of objects outside Earth soft landing
CN107323692B (en) * 2017-07-04 2019-10-18 北京理工大学 A kind of energy optimizing method of small feature loss soft landing avoidance

Patent Citations (7)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
US7967255B2 (en) * 2006-07-27 2011-06-28 Raytheon Company Autonomous space flight system and planetary lander for executing a discrete landing sequence to remove unknown navigation error, perform hazard avoidance and relocate the lander and method
CN103662096A (en) * 2013-12-13 2014-03-26 北京控制工程研究所 Self-adaptation powered explicit guidance method
CN103662091A (en) * 2013-12-13 2014-03-26 北京控制工程研究所 High-precision safe landing guiding method based on relative navigation
CN106527473A (en) * 2016-10-27 2017-03-22 上海航天控制技术研究所 Obstacle-avoidance landing method on lunar surface
CN108594802A (en) * 2018-02-28 2018-09-28 北京控制工程研究所 The determination of detector target touchdown area and avoidance method of guidance and device
CN108984907A (en) * 2018-07-18 2018-12-11 哈尔滨工业大学 A kind of interative guidance method based on yaw corner condition
CN109018440A (en) * 2018-07-23 2018-12-18 哈尔滨工业大学 VTOL carrier rocket grade landing phase precise perpendicularity soft landing quartic polynomial method of guidance

Non-Patent Citations (2)

* Cited by examiner, † Cited by third party
Title
Vision based autonomous guidance algorithm for terminal descent phase of soft lunar landing;Avijit Banerjee，Shyam Mohan M.，Radhakant Padhi;《2016 IEEE Annual India Conference》;20161218;全文 *
嫦娥三号自主避障软着陆控制技术;张洪华,梁俊,黄翔宇,赵宇,王立,关轶峰,程铭,李骥,王鹏基,于洁,袁利;《中国科学》;20140630;全文 *

Also Published As

Publication number Publication date
CN110542423A (en) 2019-12-06

Similar Documents

Publication Publication Date Title
CN112241125B (en) Unmanned aerial vehicle trajectory tracking method based on differential flatness characteristic
Bayard et al. Vision-based navigation for the NASA mars helicopter
CN107861517B (en) Skip reentry vehicle online trajectory planning guidance method based on linear pseudo-spectrum
CN107992074A (en) A kind of reentry trajectory design method based on flight path angle planning
CN106371312B (en) Lift formula based on fuzzy controller reenters prediction-correction method of guidance
CN105865455B (en) A method of utilizing GPS and accelerometer calculating aircraft attitude angle
CN109190158B (en) Optimal orbit design method considering non-cooperative target no-fly zone constraint
CN109765919B (en) Spacecraft close-range safe operation control method based on equal collision probability line method
CN108153330A (en) Unmanned aerial vehicle three-dimensional track self-adaptive tracking method based on feasible region constraint
CN109407688B (en) Centroid motion decoupling method for vertical take-off and landing rocket online trajectory planning
Mills et al. Vision based control for fixed wing UAVs inspecting locally linear infrastructure using skid-to-turn maneuvers
CN113867143B (en) Safety soft landing analysis obstacle avoidance guidance method for extraterrestrial celestial body
JPH03172887A (en) Method of displaying attitude of airframe
CN110297500A (en) A kind of continuous path planning method giving unmanned plane under more way points
CN110542423B (en) Moon soft landing vertical approach obstacle avoidance guidance method
CN113608543A (en) Method, device, equipment and storage medium for self-adaptive planning of flight path of aircraft
Davidi et al. Formation flight using multiple integral backstepping controllers
CN108563124B (en) Self-adaptive control method of rotor unmanned aerial vehicle based on API (application program interface) user-settable track
CN113625768B (en) Mars helicopter track planning method, system, equipment and storage medium
CN115993769A (en) Integrated guidance control method for high-dynamic aircraft
JPH11105797A (en) Landing gear
Al-Sharman Auto takeoff and precision landing using integrated GPS/INS/Optical flow solution
CN111651860B (en) Predictive correction robust guidance method for re-entry section of reusable carrier
Kovalev et al. UAV's autonomous navigation principe based on Earth remote sensing data
CN114594783B (en) Four-rotor real-time trajectory planning and landing control method based on overall process constraint

Legal Events

Date Code Title Description
PB01 Publication
PB01 Publication
SE01 Entry into force of request for substantive examination
SE01 Entry into force of request for substantive examination
GR01 Patent grant
GR01 Patent grant