CN100495261C - Lunar soft landing guidance, navigation and control semi-physical simulation test system - Google Patents

Abstract

The lunar soft landing guidance, navigation and control semi-physical simulation test system consists of a three-dimensional translation simulation module, a three-dimensional rotation simulation module, a control computer, a simulation computer, and a ground test and master control computer system. The three-dimensional translation simulation module and the three-dimensional rotation simulation module adopt physical models, among which, the three-dimensional translation simulation module is used to simulate the orbital motion of the lander relative to the lunar surface, and the lunar sand table screen on it is used to simulate the surface characteristics of the moon; the three-dimensional The rotation simulation module is used to simulate the attitude movement of the lander. Other mature conventional sensors and actuators, as well as lander dynamics and kinematics, are replaced by precise mathematical models established by computers. Compared with mathematical simulation, this system can make the GNC scheme and algorithm more realistically and effectively verified, and compared with the whole physical simulation system, it has the advantages of low development cost, simple operation and so on.

Description

Lunar soft landing guidance, navigation and control semi-physical simulation test system

technical field

The invention relates to a lunar soft landing guidance, navigation and control (GNC) semi-physical simulation test system, which can be used to verify the lunar soft landing GNC scheme and algorithm.

Background technique

In the 1960s and 1970s, Apollo and Surveyor of the United States and Luna of the Soviet Union successively realized soft landing missions on the moon. Entering the 21st century, the world’s aerospace field once again set off a climax of lunar exploration. How to design a set of lunar soft landing guidance, navigation and control schemes and make them effective verification has become the key to the success of lunar soft landing. With the further development of mathematical simulation technology, most of the GNC algorithms of earth satellites based on conventional sensors can be effectively verified through mathematical simulation. As for the soft landing detection with the moon as the center gravitational body, its gravity field, surrounding environment and navigation means are significantly different from those of earth satellites. Therefore, it is necessary to consider verifying the GNC scheme and algorithm through a half-physical or full-physical simulation system. Bellman and Matranga designed a full physical simulation verification system called LLRV (Lunar Landing Research Vehicle) for the Apollo program. The system uses a jet engine to provide 5/6 of its own gravity to simulate the lunar gravity environment, and other thrust engines act as soft landing brake engines to verify the GNC program for the final landing process below a few hundred meters. However, since the main task of the system is to provide astronauts with a ground manual operation training platform, many of its components, including the weight and thrust system, are different from the Apollo lunar module; obviously, the test system is also different from the self-realized Verification of the GNC program for lunar soft landing. In terms of development costs, the development costs of the two LLRVs at that time were as high as 3.6 million US dollars, and in the subsequent development of the improved LLTV, the development costs of the three LLTVs were as high as 7.5 million US dollars. On the premise that key technical breakthroughs and scheme design have been completed, the manufacture of LLRV also took 14 months. The Japanese "Luna" (SELenological and Engineering Explorer, SELENE) soft landing plan, which is expected to be launched in 2007, also uses a full-body flight verification platform (Fly Test Bed, FTB) to verify the GNC algorithm and hardware. The test system draws on Apollo's LLRV program, and simulates the lunar gravity environment through a jet engine. The verification range is also the final landing process below a few hundred meters from the lunar surface. The test verified navigation sensors including distance measurement, speed measurement, and IMU, but did not verify the 3D imaging and translational obstacle avoidance control strategies during the hovering process.

Contents of the invention

The technical solution problem of the present invention is: to overcome the deficiencies in the prior art, provide a kind of lunar soft landing guidance, navigation and control semi-physical simulation test system, this system puts the most critical external navigation sensor as a real component in the control system In the loop, other conventional sensors and actuators, as well as lander dynamics and kinematics are replaced by mature mathematical models, which can make the GNC program more realistic and effective while reducing development costs.

The technical solution of the present invention is: lunar soft landing guidance, navigation and control semi-physical simulation test system, which is characterized in that it includes: three-dimensional translation simulation module, three-dimensional rotation simulation module, control computer 4, simulation computer 11 and ground test and overall Control computer system 5, wherein:

The three-dimensional translation simulation module is used to simulate the surface characteristics of the moon, and simulates the three-dimensional translation of the lander relative to the lunar surface according to the ground test and the control command issued by the master control computer system 5;

The three-dimensional rotation simulation module is used to simulate the flight attitude of the lander. It works according to the power-on and power-off instructions from the control computer 4, measures the six-degree-of-freedom movement of the lander, and obtains the line-of-sight distance and speed of the lander relative to the lunar surface. information, the attitude angle and attitude angular velocity information of the lander itself and the three-dimensional image information of the lunar surface; the line-of-sight distance, the speed of the lander relative to the lunar surface and the three-dimensional image information of the lunar surface are transmitted to the control computer 4, and the attitude Angle and attitude angular velocity information is sent to simulation computer 11; Carry out three-dimensional rotation according to the control instruction that ground test and total control computer system 5 send;

The control computer 4 receives the remote control instructions sent by the ground test and the master control computer system 5, controls the three-dimensional rotation simulation module to work and obtains distance, speed and image measurement information, and obtains specific force and gyro angular velocity measurement information from the simulation computer 11. According to the The measurement information is calculated by GNC, and the propulsion system control command is obtained and transmitted to the simulation computer 11; the obtained measurement information and the calculated control information are transmitted to the ground test and master control computer system 5 as telemetry data;

The simulation computer 11 receives the remote control command issued by the ground test and the master control computer system 5, obtains the attitude angle and attitude angular velocity information from the three-dimensional rotation simulation module, calculates the gyro angular velocity measurement information through the gyro measurement model, and receives the propulsion system information from the control computer 4. Control command, the specific force information, propulsion system parameters and position and attitude information are obtained through simulation calculation, and the specific force and gyro angular velocity measurement information are transmitted to the control computer 4, and the propulsion system parameters and position and attitude information are transmitted to the ground test And master control computer system 5;

The ground test and general control computer system 5 sends remote commands to the control computer 4 and the simulation computer 11, receives telemetry data from the control computer 4 and performs data storage, sharing and real-time display, and receives the propulsion system parameters and position attitude from the simulation computer 11 information, and give translation and rotation control commands according to the position and attitude information in it, and control the three-dimensional translation simulation module and the three-dimensional rotation simulation module to move as required.

The three-dimensional translation simulation module is composed of a landing lunar surface simulator and a three-dimensional translation motion device 3; the landing lunar surface simulator includes a light simulator 1 and a lunar sand table screen 2, and the light simulator 1 is used to simulate the movement of the lunar surface. Lighting conditions, the lunar surface sand table screen 2 is used to simulate the surface characteristics of the lunar surface, and the landing lunar surface simulator is placed on the three-dimensional translational motion device 3 .

The three-dimensional rotation simulation module is made up of an external navigation sensor and a three-axis mechanical turntable 10. The external navigation sensor includes a range-finding and speed-measuring target simulator 6, a range finder 7, a speedometer 8 and a laser imaging sensor 9. The speed measuring target simulator 6 is used to receive the beams of the rangefinder 6 and the speedometer 7, and return the beams to the rangefinder 6 and the speedometer 7 after delay and Doppler translation, and the laser imaging sensor 9 is used for the lunar The surface sand table screen 2 performs three-dimensional imaging; the external navigation sensor is placed on the three-axis mechanical turntable 10, and the flight attitude of the lander is simulated by controlling the movement of the three-axis mechanical turntable 10.

The control computer 4 receives instructions from the ground test and the master control computer system 5, and controls the operation of the external navigation sensor in the three-dimensional rotation simulation module, which is measured and obtained by the rangefinder 7, the speedometer 8 and the distance measurement speed measurement target simulator 6. The real line-of-sight distance and speed information of the lander relative to the lunar surface, the laser imaging sensor 9 performs three-dimensional imaging on the lunar surface sand table screen 2 to obtain the three-dimensional image information of the lunar surface, and at the same time the control computer 4 obtains the specific force and The gyro angular velocity measurement information, according to the distance, speed, image, specific force and gyro angular velocity information, uses the GNC algorithm to perform navigation calculations, obtains the propulsion system control instructions and transmits them to the simulation computer 11; the acquired measurement information and the calculated control The information is sent to the ground test and master control computer system as telemetry data 5;.

The GNC algorithm includes an autonomous navigation module, a trajectory guidance module and an attitude control module; the autonomous navigation module receives the specific force and gyro angular velocity measurement information from the simulation computer 11, and obtains the position, speed, and attitude of the lander after performing inertial navigation calculations , Attitude angular velocity six degrees of freedom information, and the six degrees of freedom information is passed to the trajectory guidance module, the attitude and attitude angular velocity information is passed to the attitude control module, and the autonomous navigation module also receives the external navigation in the three-dimensional rotation simulation module The line-of-sight distance, speed and lunar surface image information measured by the sensor, after performing altitude calculation, speed calculation and landing area selection calculation on the measurement information, the calculation results are transmitted to the trajectory guidance module; the trajectory guidance module receives the autonomous The information obtained by the navigation module through inertial navigation calculation, altitude, speed calculation and landing area selection calculation, and select the guidance mode according to the computer instruction to calculate the corresponding guidance law; the attitude control module receives the information from the autonomous navigation module through inertial navigation calculation. attitude and attitude angular velocity information, and calculate the attitude control law according to different attitude control requirements; according to the calculation results of the guidance control law in the trajectory guidance module and attitude control module, the propulsion system control commands for orbit and attitude control can be obtained .

The guidance modes of the trajectory guidance module include gravity turn nominal trajectory guidance mode, variable thrust guidance mode, Bang-Bang and phase plane guidance mode, and nominal trajectory guidance mode.

The attitude control mode of the attitude control module adopts phase plane control.

The simulation computer 11 receives the ground test and the remote control command of the master control computer system 5, obtains the attitude angle and attitude angular velocity information from the three-axis mechanical turntable 10, and obtains the angular velocity measurement information of the simulated gyroscope after calculation by the dynamics and kinematics simulation model The simulation computer 11 also receives the propulsion system control command from the control computer 4, and obtains the propulsion system parameters and the position and attitude information of the lander after the calculation of the dynamics and kinematics simulation model, and can also obtain the relative force measurement information at the same time; the simulation computer 11 will The specific force and gyro angular velocity measurement information is transmitted to the control computer 4, and the propulsion system parameters and position and attitude information are transmitted to the ground testing and master control computer system 5.

The dynamics and kinematics simulation model includes a mathematical model of an actuator, a dynamics and kinematics model, a track scaling calculation model and a conventional sensor measurement model. The mathematical model of the actuator receives the propulsion system control instructions given by the control computer 4, and obtains the propulsion system parameters after simulation calculation and transmits them to the dynamics and kinematics model and the ground test and master control computer system 5; the dynamics and kinematics model Combining the parameters of the propulsion system, the six degrees of freedom information of the lander's position, velocity, attitude angle, and attitude angular velocity are obtained through calculation, and the six degrees of freedom information is transmitted to the ground test and master control computer system 5, and the position and velocity The information is transmitted to the orbit scale calculation model and the conventional sensor measurement model; the orbit scale calculation model performs scale calculation on the position and velocity information, and then transmits the scaled position and velocity information to the ground test and master control computer system 5 The conventional sensor measurement model receives the position and velocity information given by the dynamics and kinematics model, and obtains the specific force measurement information after calculation by the accelerometer measurement model. At the same time, the conventional sensor measurement model also receives the three-axis mechanical turntable 10. The attitude angle and attitude angular velocity information are calculated by the gyro measurement model to obtain the gyro angular velocity measurement information, and the specific force and gyro angular velocity measurement information are transmitted to the control computer 4.

Compared with the prior art, the present invention has the following advantages:

(1) Compared with mathematical simulation, the key components such as external navigation sensors including distance measurement, speed measurement and imaging are used as real components to effectively simulate the three-dimensional translation and three-dimensional rotation of the lander relative to the lunar surface, so as to make a soft landing The verification of the process GNC algorithm is more comprehensive and effective;

(2) Compared with the full physical simulation, the mature conventional sensors and actuators and lander dynamic kinematics calculations are replaced by accurate mathematical models, which is simpler and easier than the full physical simulation;

(3) The imaging sensor is added to perform three-dimensional high-resolution imaging of the lunar surface, which lays the foundation for the safe obstacle avoidance and landing of the lander.

Description of drawings

Fig. 1 is a system composition block diagram of the present invention;

Fig. 2 is the moon surface sand table screen size and the moon surface simulation diagram of the present invention;

Fig. 3 is the figure of three-dimensional translation motion device of the present invention;

Fig. 4 is the control computer GNC algorithm data flowchart of the present invention;

Fig. 5 is a data flow chart of the simulation computer simulation model of the present invention.

Detailed ways

1. Specific design and implementation of key components

(1) Three-axis mechanical turntable 10

The three-axis mechanical turntable 10 is used to simulate the three-axis attitude movement of the lander, and the attitude angle and attitude angular velocity information of the turntable movement is calculated by the gyro measurement model in the simulation computer 11 to simulate the attitude output of the gyro. The embodiment of the present invention will utilize the existing French imported three-axis mechanical turntable of the unit.

(2) External navigation sensor

The external navigation sensor includes a distance measuring and speed measuring target simulator 6 , a range finder 7 , a speed meter 8 and a laser imaging sensor 9 .

Due to the limitation of the space of the simulation laboratory, the rangefinder 7 and the speedometer 8 cannot directly measure the lunar surface sand table screen. The present invention receives the radar beams sent by the rangefinder 7 and the speedometer 8 by adding a rangefinder and speedometer target simulator 6, and sends them to the rangefinder 7 and the speedometer 8 after delay and Doppler translation processing , so as to complete the corresponding distance and speed measurements. The distance-measuring and speed-measuring target simulator 6 is laid out according to the installation requirements of the rangefinder 7 and the speedometer 8 .

The laser imaging sensor 9 is mainly used for imaging the lunar surface when hovering 100m away from the lunar surface.

The embodiment of the present invention adopts the external navigation sensor combination scheme of microwave distance measurement and speed measurement and laser radar imaging, and a 1:10 scaling is used for imaging, that is, three-dimensional high-resolution imaging of the lunar surface sand table screen 2 at a distance of 10m.

(3) Lunar Sand Table Screen 2

The main purpose of the lunar surface sand table screen 2 is to provide the laser imaging sensor 9 with simulated landing lunar surface characteristic information. The basis for its size design is to make the imaging range as large as possible, to keep the imaging range before translation as consistent as possible after translation, and to have as large a translation space as possible.

The field of view of the imaging sensor is 30°×30°, and the area of the lunar surface covered by imaging at 100m is 53.6m×53.6m. The embodiment of the present invention adopts a scale ratio of 1:10, and the size of the lunar sand table screen 2 covered by the imaging should be 5.36m×5.36m. If the optical axis of the imaging sensor is located at the center of the turntable, at a height of 10m, the imaging sensor can form a complete three-dimensional image within 2.68m from the optical axis. With the size setting shown in Figure 2, the left and right directions both meet the requirement of not less than 2.68m before and after translation, and a complete left and right image can be formed; in the upper direction, the sand table screen is rolled by means of a roller, which can be simulated on the sand table screen. Under the premise of moving up and down, the upper and lower dimensions are guaranteed to remain unchanged, so the upper dimension also meets the requirement of not less than 2.68m. At present, only the size of the bottom of the sand table screen is difficult to meet the requirements due to the limitation of the site, but this does not affect the control system to obtain information from 3D imaging.

Set the sand table to move to the right and down, the distance to the right is 1.5 meters, and the distance to move down is 1 meter, and the translation control in both directions can be performed at the same time.

(4) Three-dimensional translational motion device 3

In the present invention, the translational movement of the lunar surface sand table screen 2 relative to the three-axis mechanical turntable 10 is used to simulate the translational movement of the lander relative to the lunar surface during actual flight. This three-dimensional translation is realized by the three-dimensional translation motion device 3 .

In the embodiment of the present invention, the three-dimensional translational movement device 3 will take the form of a guide rail. As shown in Figure 3, two longitudinal guide rails of about 15m are laid along the axis of the turntable, and horizontal guide rails of about 10m are laid on the longitudinal guide rails, and the vertical descent of the lander relative to the lunar surface is simulated by the horizontal guide rails moving back and forth on the longitudinal guide rails . The lunar sand table screen 2 is located on the horizontal guide rail, and can move left and right along the horizontal guide rail to simulate the one-dimensional movement in the actual horizontal plane; the translation in the other direction in the horizontal plane can be realized by the roller under the lunar sand table screen 2.

(5) Control computer 4 and GNC algorithm

The control computer 4 is the core equipment of the GNC semi-physical simulation test system for lunar soft landing. The control computer 4 executes the GNC algorithm, and should have guidance, navigation and control functions with the cooperation of other equipment. Considering the limitation of the payload weight and volume by the three-axis turntable, the control computer 4 is moved from the turntable to the ground. In the embodiment of the present invention, the control computer 4 is replaced by an industrial computer.

The GNC algorithm is an application program running on the control computer 4, and should have the following working modes and functions: gravity turn + nominal trajectory guidance mode in the attitude adjustment section, nominal trajectory guidance mode in the landing section, variable thrust guidance mode during hovering, Landing area selection, translation orbit control, attitude tracking and stability control, etc.

The data flow of the GNC algorithm in the embodiment of the present invention is shown in FIG. 4 . Contains autonomous navigation module, trajectory guidance module and attitude control module. The autonomous navigation module receives the specific force and gyro angular velocity measurement information from the simulation computer 11, and obtains the six degrees of freedom information of the lander's position, velocity, attitude, and attitude angular velocity after performing inertial navigation calculations, and transmits the six degrees of freedom information to The trajectory guidance module transmits the attitude and attitude angular velocity information to the attitude control module, and the autonomous navigation module also receives the line-of-sight distance, speed and lunar surface image information measured by the external navigation sensor in the three-dimensional rotation simulation module. After the measurement information is calculated by height, speed and landing area selection, the calculation results are sent to the trajectory guidance module; the trajectory guidance module receives the autonomous navigation module and obtains it through inertial navigation calculation, altitude, speed calculation and landing area selection calculation. information, and select the guidance mode according to the computer instruction to calculate the corresponding guidance law; the attitude control module receives the attitude and attitude angular velocity information calculated by the inertial navigation from the autonomous navigation module, and performs the attitude control law according to different attitude control requirements. Calculation: According to the calculation results of the guidance control law in the trajectory guidance module and the attitude control module, the propulsion system control instructions for trajectory and attitude control can be obtained.

The inertial navigation calculation formula adopted in the embodiment of the present invention is as follows:

(a) Calculation of attitude and attitude angular velocity

$\left\{\begin{array}{c}{\stackrel{<span\; class="notranslate">\; ^</span>}{\mathrm{<span\; class="notranslate">\; \&omega;</span>}}}_{\mathrm{<span\; class="notranslate">\; b</span>}}<span\; class="notranslate">\; =</span>\stackrel{<span\; class="notranslate">\; ~</span>}{\mathrm{<span\; class="notranslate">\; \&omega;</span>}}\\ \stackrel{<span\; class="notranslate">\; \&CenterDot;</span>}{\stackrel{<span\; class="notranslate">\; ^</span>}{\mathrm{<span\; class="notranslate">\; \&epsiv;</span>}}}<span\; class="notranslate">\; =</span>\frac{\mathrm{<span\; class="notranslate">\; 1</span>}}{\mathrm{<span\; class="notranslate">\; 2</span>}}<span\; class="notranslate">\; (</span>{\stackrel{<span\; class="notranslate">\; ^</span>}{\mathrm{<span\; class="notranslate">\; \&epsiv;</span>}}}^{<span\; class="notranslate">\; \&times;</span>}<span\; class="notranslate">\; +</span>\stackrel{<span\; class="notranslate">\; ^</span>}{\mathrm{<span\; class="notranslate">\; \&eta;</span>}}{\mathrm{<span\; class="notranslate">\; E.</span>}}_{\mathrm{<span\; class="notranslate">\; 3</span>}}<span\; class="notranslate">\; )</span>\stackrel{<span\; class="notranslate">\; ^</span>}{\mathrm{<span\; class="notranslate">\; \&omega;</span>}}<span\; class="notranslate">\; ,</span>\stackrel{<span\; class="notranslate">\; \&CenterDot;</span>}{\stackrel{<span\; class="notranslate">\; ^</span>}{\mathrm{<span\; class="notranslate">\; \&eta;</span>}}}<span\; class="notranslate">\; =</span><span\; class="notranslate">\; -</span>\frac{\mathrm{<span\; class="notranslate">\; 1</span>}}{\mathrm{<span\; class="notranslate">\; 2</span>}}{\stackrel{<span\; class="notranslate">\; ^</span>}{\mathrm{<span\; class="notranslate">\; \&epsiv;</span>}}}^{\mathrm{<span\; class="notranslate">\; T</span>}}\stackrel{<span\; class="notranslate">\; ~</span>}{\mathrm{<span\; class="notranslate">\; \&omega;</span>}}\end{array}\right.$

和

分别为计算后的着陆器本体姿态角速度矢量和姿态角的四元素表示(矢量部分和标量部分)；

为陀螺测量角速度。该计算忽略月球自转角速度。in,

and

are the four-element representation (vector part and scalar part) of the calculated lander body attitude angular velocity vector and attitude angle, respectively;

Measure angular velocity for a gyroscope. This calculation ignores the Moon's angular velocity.

(b) Position and velocity calculation

$\left\{\begin{array}{c}\stackrel{<span\; class="notranslate">\; \&Center\; Dot;</span>}{\stackrel{<span\; class="notranslate">\; ^</span>}{\mathrm{<span\; class="notranslate">\; r</span>}}<span\; class="notranslate">\; =</span>\stackrel{<span\; class="notranslate">\; ^</span>}{\mathrm{<span\; class="notranslate">\; v</span>}}}\\ \stackrel{<span\; class="notranslate">\; \&CenterDot;</span>}{\stackrel{<span\; class="notranslate">\; ^</span>}{\mathrm{<span\; class="notranslate">\; v</span>}}}<span\; class="notranslate">\; =</span>\mathrm{<span\; class="notranslate">\; A</span>}<span\; class="notranslate">\; (</span>\stackrel{<span\; class="notranslate">\; ^</span>}{\mathrm{<span\; class="notranslate">\; \&epsiv;</span>}}<span\; class="notranslate">\; ,</span>\stackrel{<span\; class="notranslate">\; ^</span>}{\mathrm{<span\; class="notranslate">\; \&eta;</span>}}<span\; class="notranslate">\; )</span><span\; class="notranslate">\; \&CenterDot;</span>\stackrel{<span\; class="notranslate">\; ~</span>}{\mathrm{<span\; class="notranslate">\; f</span>}}<span\; class="notranslate">\; +</span>{\mathrm{<span\; class="notranslate">\; g</span>}}_{\mathrm{<span\; class="notranslate">\; m</span>}}\end{array}\right.$

分别为计算后的着陆器相对月面的位置和速度矢量；

为加速度计测量的比力矢量；矩阵A为从本体系到月平系的坐标变换阵，其变量

分别为姿态四元素的矢量和标量部分，与“姿态和姿态角速度计算”中的表示相同；g_m为月球表面重力加速度矢量。in,

are the calculated position and velocity vectors of the lander relative to the lunar surface, respectively;

is the specific force vector measured by the accelerometer; the matrix A is the coordinate transformation matrix from the local system to the lunar level system, and its variable

are the vector and scalar parts of the four elements of attitude, respectively, which are the same as those in "Calculation of Attitude and Attitude Angular Velocity"; g _m is the gravitational acceleration vector on the lunar surface.

The height calculation formula adopted in the embodiment of the present invention is as follows:

${\stackrel{<span\; class="notranslate">\; ^</span>}{\mathrm{<span\; class="notranslate">\; h</span>}}}_{\mathrm{<span\; class="notranslate">\; R</span>}}<span\; class="notranslate">\; =</span>\frac{\mathrm{<span\; class="notranslate">\; 3</span>}\mathrm{<span\; class="notranslate">\; V</span>}}{\mathrm{<span\; class="notranslate">\; S</span>}}<span\; class="notranslate">\; =</span>\frac{<span\; class="notranslate">\; |</span>{\left[\begin{array}{ccc}{\mathrm{<span\; class="notranslate">\; l</span>}}_{\mathrm{<span\; class="notranslate">\; 1</span>}}^{\mathrm{<span\; class="notranslate">\; b</span>}}& {\mathrm{<span\; class="notranslate">\; l</span>}}_{\mathrm{<span\; class="notranslate">\; 2</span>}}^{\mathrm{<span\; class="notranslate">\; b</span>}}& {\mathrm{<span\; class="notranslate">\; l</span>}}_{\mathrm{<span\; class="notranslate">\; 3</span>}}^{\mathrm{<span\; class="notranslate">\; b</span>}}\end{array}\right]}^{\mathrm{<span\; class="notranslate">\; T</span>}}<span\; class="notranslate">\; |</span>}{<span\; class="notranslate">\; |</span><span\; class="notranslate">\; (</span>{\mathrm{<span\; class="notranslate">\; l</span>}}_{\mathrm{<span\; class="notranslate">\; 2</span>}}^{\mathrm{<span\; class="notranslate">\; b</span>}}<span\; class="notranslate">\; -</span>{\mathrm{<span\; class="notranslate">\; l</span>}}_{\mathrm{<span\; class="notranslate">\; 1</span>}}^{\mathrm{<span\; class="notranslate">\; b</span>}}<span\; class="notranslate">\; )</span><span\; class="notranslate">\; \&times;</span><span\; class="notranslate">\; (</span>{\mathrm{<span\; class="notranslate">\; l</span>}}_{\mathrm{<span\; class="notranslate">\; 3</span>}}^{\mathrm{<span\; class="notranslate">\; b</span>}}<span\; class="notranslate">\; -</span>{\mathrm{<span\; class="notranslate">\; l</span>}}_{\mathrm{<span\; class="notranslate">\; 1</span>}}^{\mathrm{<span\; class="notranslate">\; b</span>}}<span\; class="notranslate">\; )</span><span\; class="notranslate">\; |</span>}$

为着陆器距月面高度的估算值；V，S分别为由测距仪三条波束与月面形成的六面体的体积和底面(月面上)面积；

分别为测距仪不共面的三条波束矢量，其中上标b代表该矢量表示在着陆器本体坐标系下，由测距仪可直接测得三条波束的视线长度，而波束在着陆器本体系下的角度关系是已知的。in,

is the estimated value of the height of the lander from the lunar surface; V and S are respectively the volume and the area of the bottom surface (lunar surface) of the hexahedron formed by the three beams of the range finder and the lunar surface;

are the three beam vectors of the range finder that are not coplanar, where the superscript b represents that the vector indicates that in the lander body coordinate system, the range finder can directly measure the line-of-sight lengths of the three beams, and the beams in the lander body system The following angular relationship is known.

The speed solution formula adopted in the embodiment of the present invention is as follows:

${\stackrel{<span\; class="notranslate">\; ~</span>}{\mathrm{<span\; class="notranslate">\; v</span>}}}_{\mathrm{<span\; class="notranslate">\; i</span>}}<span\; class="notranslate">\; =</span><span\; class="notranslate">\; <</span>{\mathrm{<span\; class="notranslate">\; l</span>}}_{\mathrm{<span\; class="notranslate">\; i</span>}}^{\mathrm{<span\; class="notranslate">\; b</span>}}<span\; class="notranslate">\; \&CenterDot;</span>{\mathrm{<span\; class="notranslate">\; v</span>}}^{\mathrm{<span\; class="notranslate">\; b</span>}}<span\; class="notranslate">\; ></span>$

的意义与“测距仪高度解算”中相同；v^b表示着陆器相对于月面的速度矢量，为未知；表示由测速仪测得的、合速度v^b在波束i方向的速度分量，为已知。将雷达三波束的测量值分别带入上式构成一个方程组，解这个方程组即可求得合速度。Among them, <·> represents the dot product of two vectors; the subscript i represents the radar beam (i=1, 2, 3); the superscript b and the vector

The meaning of is the same as that in "range finder altitude solution"; v ^b represents the velocity vector of the lander relative to the lunar surface, which is unknown; Represents the velocity component of the resultant velocity v ^b in the direction of beam i measured by the velocimeter, which is known. Put the measured values of the three radar beams into the above formula to form a system of equations, and solve this system of equations to obtain the resultant velocity.

The selection method of landing area in the embodiment of the present invention is as follows:

The lunar surface imaging sensor can give the height information of each point in the lunar surface landing area in one imaging. Assuming that the lunar area occupied by the lander after landing is S, the initial plan for selecting the landing area is: use S as the unit and start from the current landing point to search counterclockwise (clockwise). For each area of size S, it is necessary to judge the landing conditions——whether the raised height in the landing area and the slope of the landing area meet the requirements, until a safe landing area is searched.

The working mode and calculation formula of the guidance control law adopted in the embodiment of the present invention are as follows:

(a) Attitude Adjustment Section - Gravity Turn Nominal Trajectory Guidance

${\mathrm{<span\; class="notranslate">\; u</span>}}_{\mathrm{<span\; class="notranslate">\; Fc</span>}}<span\; class="notranslate">\; =</span>\frac{\mathrm{<span\; class="notranslate">\; 1</span>}}{{\mathrm{<span\; class="notranslate">\; g</span>}}_{\mathrm{<span\; class="notranslate">\; m</span>}}\mathrm{<span\; class="notranslate">\; cos</span>}\mathrm{<span\; class="notranslate">\; \&Psi;</span>}}<span\; class="notranslate">\; [</span>{\mathrm{<span\; class="notranslate">\; g</span>}}_{\mathrm{<span\; class="notranslate">\; m</span>}}<span\; class="notranslate">\; (</span>\mathrm{<span\; class="notranslate">\; 1</span>}<span\; class="notranslate">\; -</span>\frac{\mathrm{<span\; class="notranslate">\; \&tau;</span>}{\mathrm{<span\; class="notranslate">\; <figure-callout\; id="2"\; label="sin"\; filenames="C200710121319E00211.png"\; state="\{\{state\}\}">sin</figure-callout></span>}}^{\mathrm{<span\; class="notranslate">\; 2</span>}}\mathrm{<span\; class="notranslate">\; \&Psi;</span>}}{\mathrm{<span\; class="notranslate">\; v</span>}<span\; class="notranslate">\; +</span>\mathrm{<span\; class="notranslate">\; \&tau;</span>}}<span\; class="notranslate">\; )</span><span\; class="notranslate">\; +</span>{\stackrel{<span\; class="notranslate">\; \&CenterDot;</span><span\; class="notranslate">\; \&CenterDot;</span>}{\mathrm{<span\; class="notranslate">\; h</span>}}}_{\mathrm{<span\; class="notranslate">\; D.</span>}}<span\; class="notranslate">\; -</span><span\; class="notranslate">\; [</span>{\mathrm{<span\; class="notranslate">\; k</span>}}_{\mathrm{<span\; class="notranslate">\; p</span>}}<span\; class="notranslate">\; (</span>\mathrm{<span\; class="notranslate">\; h</span>}<span\; class="notranslate">\; -</span>{\mathrm{<span\; class="notranslate">\; h</span>}}_{\mathrm{<span\; class="notranslate">\; D.</span>}}<span\; class="notranslate">\; )</span><span\; class="notranslate">\; +</span>{\mathrm{<span\; class="notranslate">\; k</span>}}_{\mathrm{<span\; class="notranslate">\; d</span>}}<span\; class="notranslate">\; (</span>\stackrel{<span\; class="notranslate">\; \&CenterDot;</span>}{\mathrm{<span\; class="notranslate">\; h</span>}}<span\; class="notranslate">\; -</span>{\stackrel{<span\; class="notranslate">\; \&Center\; Dot;</span>}{\mathrm{<span\; class="notranslate">\; h</span>}}}_{\mathrm{<span\; class="notranslate">\; D.</span>}}<span\; class="notranslate">\; )</span><span\; class="notranslate">\; ]</span><span\; class="notranslate">\; ]</span>$

分别表示实际飞行的距月面高度和垂直速度；h_D，分别表示标称轨迹中着陆器的标称高度、标称垂直速度和垂直加速度分量；v为合速度的大小；Ψ为下降速度方向与当地垂线方向的夹角；k_p，k_d分别为比例和微分系数；τ为一给定的小常数；g_m为月球表面重力加速度。Among them, u _Fc is the main engine thrust-to-weight ratio guidance command; h,

Respectively represent the height from the lunar surface and the vertical velocity of the actual flight; h _D , represent the nominal altitude, nominal vertical velocity and vertical acceleration component of the lander in the nominal trajectory; v is the magnitude of the resultant velocity; Ψ is the angle between the direction of the descent velocity and the direction of the local vertical; k _p and k _d are respectively Proportional and differential coefficients; τ is a given small constant; g _m is the gravitational acceleration on the lunar surface.

(b) Hover translation process: vertical direction——variable thrust guidance; horizontal direction——Bang-Bang+phase plane guidance

In the vertical direction, the main thrust engine is controlled so that the thrust is always equal to the gravity of the lander, and the direction is opposite; in the horizontal direction, the positive and negative switch control strategy of first accelerating and then decelerating is adopted by using the translational thruster, and then switched when approaching the target position to a phase-plane control strategy with respect to position and velocity. Under this phase plane control strategy, the translational thruster only has three states of "positive open", "negative open" and "closed", so four switch lines (two open lines, two close lines) are formed on the position and velocity phase plane. line), the phase plane is divided into four areas, and the control computer switches and translates the thrusters according to the different areas of the phase plane where the lander is located, thereby controlling the lander to approach the target point.

(c) Final landing segment—nominal trajectory guidance

${\mathrm{<span\; class="notranslate">\; u</span>}}_{\mathrm{<span\; class="notranslate">\; Fc</span>}}<span\; class="notranslate">\; =</span>\frac{\mathrm{<span\; class="notranslate">\; 1</span>}}{{\mathrm{<span\; class="notranslate">\; g</span>}}_{\mathrm{<span\; class="notranslate">\; m</span>}}}<span\; class="notranslate">\; [</span>{\mathrm{<span\; class="notranslate">\; g</span>}}_{\mathrm{<span\; class="notranslate">\; m</span>}}<span\; class="notranslate">\; +</span>{\stackrel{<span\; class="notranslate">\; \&Center\; Dot;</span><span\; class="notranslate">\; \&Center\; Dot;</span>}{\mathrm{<span\; class="notranslate">\; h</span>}}}_{\mathrm{<span\; class="notranslate">\; D.</span>}}<span\; class="notranslate">\; -</span><span\; class="notranslate">\; (</span>{\mathrm{<span\; class="notranslate">\; k</span>}}_{\mathrm{<span\; class="notranslate">\; p</span>}}<span\; class="notranslate">\; (</span>\mathrm{<span\; class="notranslate">\; h</span>}<span\; class="notranslate">\; -</span>{\mathrm{<span\; class="notranslate">\; h</span>}}_{\mathrm{<span\; class="notranslate">\; D.</span>}}<span\; class="notranslate">\; )</span><span\; class="notranslate">\; +</span>{\mathrm{<span\; class="notranslate">\; k</span>}}_{\mathrm{<span\; class="notranslate">\; d</span>}}<span\; class="notranslate">\; (</span>\stackrel{<span\; class="notranslate">\; \&CenterDot;</span>}{\mathrm{<span\; class="notranslate">\; h</span>}}<span\; class="notranslate">\; -</span>{\stackrel{<span\; class="notranslate">\; \&Center\; Dot;</span>}{\mathrm{<span\; class="notranslate">\; h</span>}}}_{\mathrm{<span\; class="notranslate">\; D.</span>}}<span\; class="notranslate">\; )</span><span\; class="notranslate">\; )</span><span\; class="notranslate">\; ]</span>$

The meanings of the parameters in the above formula are the same as those in (a).

The final landing segment guidance mode is a special case of the attitude adjustment segment guidance mode, which requires Ψ=0 on the basis of the attitude adjustment segment guidance law, which means that the lander descends in the vertical direction.

(d) Attitude control mode - three-axis attitude stable phase plane control

In the embodiment of the present invention, the attitude control below 150m from the lunar surface adopts the phase plane control strategy about attitude angle and attitude angular velocity. The phase plane control strategy is similar to the position-velocity phase plane control strategy in (b) above. The thruster used for attitude control has only three states of "positive open", "negative open" and "closed", so four switch lines (two open lines, two close lines) are formed on the phase plane of attitude angle and attitude angular velocity ), the phase plane is divided into four areas, and the control computer switches the attitude control thrusters according to the different areas of the phase plane where the lander is located, so as to control the attitude of the lander to stabilize near the nominal attitude.

(6) Simulation computer 11 and dynamics and kinematics simulation model

The simulation computer 11 is mainly responsible for tasks such as orbit and attitude dynamics and kinematics simulation, and should have the following functions: orbit and attitude dynamics and kinematics simulation during the landing process of the lander; mathematical simulation of conventional sensors such as IMU; propulsion system and other actuators Mathematical simulation; orbital movement scaling calculation; resource sharing with ground test and master control computer. In the embodiment of the present invention, the simulation computer 11 adopts a high-performance PC.

The function of simulation computer 11 is completed by dynamics and kinematics simulation model, and dynamics and kinematics simulation model should include track/attitude dynamics and kinematics model, IMU measurement model, executive mechanism (comprising main thrust engine, translation thruster and Attitude control thruster) mathematical model and orbit scaling calculation model, etc.

The data flow of the dynamics and kinematics simulation model of the embodiment of the present invention is shown in FIG. 5 . Including the mathematical model of the actuator, the dynamics and kinematics model, the calculation model of orbital scaling, and the conventional sensor measurement model; the mathematical model of the actuator receives the propulsion system control command given by the control computer 4, and obtains the propulsion system parameters after simulation calculation And pass it to the dynamics and kinematics model and the ground test and master control computer system 5; the dynamics and kinematics model combines the parameters of the propulsion system to calculate the six degrees of freedom information of the lander's position, velocity, attitude angle, and attitude angle velocity. And the information of the six degrees of freedom is transmitted to the ground test and the general control computer system 5, and the position and velocity information are transmitted to the track scale calculation model and the conventional sensor measurement model; the track scale calculation model performs position and velocity information After the scaling calculation, the scaled position and velocity information is transmitted to the ground test and the master control computer system 5; the conventional sensor measurement model receives the position and velocity information given by the dynamics and kinematics model, and is measured by the accelerometer. After the calculation, the relative force measurement information is obtained. At the same time, the conventional sensor measurement model also receives the attitude angle and attitude angular velocity information provided by the three-axis mechanical turntable 10, and the gyro angular velocity measurement information is obtained through the calculation of the gyro measurement model, and the specific force and The gyro angular velocity measurement information is transmitted to the control computer 4.

The executive mechanism model that the embodiment of the present invention adopts is as follows:

(a) Attitude control thruster model

Neglecting the opening and closing delay of the thruster, the following simple model can be used in the simulation:

F _atti (t) = F _0atti [I(t)-I(tT)]

Among them, F _atti (t) represents the actual output of the attitude control thruster; F _0atti represents the nominal thrust of the attitude control thruster; I(·) represents the step function, and T is the time width of the jet command pulse.

(b) Orbit-controlled variable-thrust engine model

Neglecting the opening and closing delay of the thruster, considering the control accuracy of the thruster which has a greater influence, the following simple model can be used in the simulation:

F _obt (t)=(1+δ)F _0obt

Among them, F _obt (t) represents the actual output of the orbital control engine; F _0obt represents the nominal thrust of the orbital control engine; δ is the thrust control accuracy of the orbital control engine.

The dynamics and kinematics model that the embodiment of the present invention adopts is as follows:

(a) Gyro measurement model

$\stackrel{<span\; class="notranslate">\; ~</span>}{\mathrm{<span\; class="notranslate">\; \&omega;</span>}}<span\; class="notranslate">\; =</span>{\mathrm{<span\; class="notranslate">\; \&omega;</span>}}_{\mathrm{<span\; class="notranslate">\; r</span>}}<span\; class="notranslate">\; +</span>{\mathrm{<span\; class="notranslate">\; \&omega;</span>}}_{\mathrm{<span\; class="notranslate">\; 0</span>}}<span\; class="notranslate">\; +</span>{\mathrm{<span\; class="notranslate">\; \&omega;</span>}}_{\mathrm{<span\; class="notranslate">\; 1</span>}}$

为陀螺测量角速度；ω_r为着陆器实际姿态角速度，仿真试验中，此即为由姿态运动学模型计算得到的姿态角速度；ω₀，ω₁分别为陀螺的常值漂移和随机漂移，可标定。in,

is the angular velocity _measured by the gyro; ω _r is the actual attitude angular velocity of the lander. In the simulation test, _this is the attitude angular velocity calculated by the attitude kinematics model; .

(b) Accelerometer measurement model

$\stackrel{<span\; class="notranslate">\; ~</span>}{\mathrm{<span\; class="notranslate">\; f</span>}}<span\; class="notranslate">\; =</span><span\; class="notranslate">\; (</span>\mathrm{<span\; class="notranslate">\; 1</span>}<span\; class="notranslate">\; +</span>\mathrm{<span\; class="notranslate">\; k</span>}<span\; class="notranslate">\; )</span><span\; class="notranslate">\; (</span>{\mathrm{<span\; class="notranslate">\; f</span>}}_{\mathrm{<span\; class="notranslate">\; r</span>}}<span\; class="notranslate">\; +</span><span\; class="notranslate">\; \&dtri;</span><span\; class="notranslate">\; +</span>\mathrm{<span\; class="notranslate">\; w</span>}<span\; class="notranslate">\; )</span>$

为加速度计测量的比力值；f_r为着陆器实际比力；k为刻度因数误差；

为加速度计的零偏；w为测量噪声。in,

is the specific force value measured by the accelerometer; f _r is the actual specific force of the lander; k is the scale factor error;

is the zero bias of the accelerometer; w is the measurement noise.

The dynamics and kinematics model that the embodiment of the present invention adopts is as follows:

(a) Orbit dynamics model

$\stackrel{<span\; class="notranslate">\; \&Center\; Dot;</span><span\; class="notranslate">\; \&Center\; Dot;</span>}{\mathrm{<span\; class="notranslate">\; r</span>}}<span\; class="notranslate">\; =</span>\mathrm{<span\; class="notranslate">\; f</span>}<span\; class="notranslate">\; -</span>\frac{{\mathrm{<span\; class="notranslate">\; \&mu;</span>}}_{\mathrm{<span\; class="notranslate">\; m</span>}}}{{\mathrm{<span\; class="notranslate">\; <figure-callout\; id="3"\; label="r"\; filenames="C200710121319E00211.png"\; state="\{\{state\}\}">r</figure-callout></span>}}^{\mathrm{<span\; class="notranslate">\; 3</span>}}}<span\; class="notranslate">\; \&Center\; Dot;</span>\mathrm{<span\; class="notranslate">\; r</span>}<span\; class="notranslate">\; +</span>{\mathrm{<span\; class="notranslate">\; f</span>}}_{\mathrm{<span\; class="notranslate">\; \&epsiv;</span>}}<span\; class="notranslate">\; -</span>{\mathrm{<span\; class="notranslate">\; \&mu;</span>}}_{\mathrm{<span\; class="notranslate">\; e</span>}}<span\; class="notranslate">\; \&Center\; Dot;</span><span\; class="notranslate">\; (</span>\frac{\mathrm{<span\; class="notranslate">\; 1</span>}}{{\mathrm{<span\; class="notranslate">\; <figure-callout\; id="3"\; label="r\; e"\; filenames="C200710121319E00211.png"\; state="\{\{state\}\}">r</figure-callout></span><figure-callout\; id="3"\; label="r\; e"\; filenames="C200710121319E00211.png"\; state="\{\{state\}\}">\; </figure-callout>}}_{\mathrm{<span\; class="notranslate">\; </span>}}^{\mathrm{<span\; class="notranslate">e</span>}}}<span\; class="notranslate">\; \&Center\; Dot;</span>{\mathrm{<span\; class="notranslate">\; r</span>}}_{\mathrm{<span\; class="notranslate">\; e</span>}}<span\; class="notranslate">\; +</span>\frac{\mathrm{<span\; class="notranslate">\; 1</span>}}{{\mathrm{<span\; class="notranslate">\; \&Delta;</span>}}_{\mathrm{<span\; class="notranslate">\; <figure-callout\; id="3"\; label="e"\; filenames="C200710121319E00211.png"\; state="\{\{state\}\}">e</figure-callout></span>}}^{\mathrm{<span\; class="notranslate">\; 3</span>}}}<span\; class="notranslate">\; \&Center\; Dot;</span>{\mathrm{<span\; class="notranslate">\; \&Delta;</span>}}_{\mathrm{<span\; class="notranslate">\; e</span>}}<span\; class="notranslate">\; )</span><span\; class="notranslate">\; -</span>{\mathrm{<span\; class="notranslate">\; \&mu;</span>}}_{\mathrm{<span\; class="notranslate">\; the\; s</span>}}<span\; class="notranslate">\; \&CenterDot;</span><span\; class="notranslate">\; (</span>\frac{\mathrm{<span\; class="notranslate">\; 1</span>}}{{\mathrm{<span\; class="notranslate">\; r</span>}}_{\mathrm{<span\; class="notranslate">\; the\; <figure-callout\; id="3"\; label="s"\; filenames="C200710121319E00211.png"\; state="\{\{state\}\}">s</figure-callout></span>}}^{\mathrm{<span\; class="notranslate">\; 3</span>}}}<span\; class="notranslate">\; \&Center\; Dot;</span>{\mathrm{<span\; class="notranslate">\; r</span>}}_{\mathrm{<span\; class="notranslate">\; the\; s</span>}}<span\; class="notranslate">\; +</span>\frac{\mathrm{<span\; class="notranslate">\; 1</span>}}{{\mathrm{<span\; class="notranslate">\; \&Delta;</span>}}_{\mathrm{<span\; class="notranslate">\; the\; <figure-callout\; id="3"\; label="s"\; filenames="C200710121319E00211.png"\; state="\{\{state\}\}">s</figure-callout></span>}}^{\mathrm{<span\; class="notranslate">\; 3</span>}}}<span\; class="notranslate">\; \&Center\; Dot;</span>{\mathrm{<span\; class="notranslate">\; \&Delta;</span>}}_{\mathrm{<span\; class="notranslate">\; the\; s</span>}}<span\; class="notranslate">\; )</span><span\; class="notranslate">\; +</span>\mathrm{<span\; class="notranslate">\; f</span>}$

_Among them, μ _m , μ _e , μ _s are gravitational constants from the center _of the moon, the center of _the earth _, and the center of the sun _; The arrow diameters from the center of the moon to the center of the earth and from the center of the moon to the center of the sun. The first item F on the right is the active braking force of the propulsion system, the second item is the central gravitational force of the moon, the third item is the non-spherical gravitational perturbation of the moon, and the fourth and fifth items are the gravitational perturbation of the earth and the sun respectively . These items are the main perturbation sources for near-moon spacecraft flights. f is other external perturbations except the above three perturbations.

(b) Orbit kinematics model

$\stackrel{<span\; class="notranslate">\; \&Center\; Dot;</span>}{\mathrm{<span\; class="notranslate">\; r</span>}}<span\; class="notranslate">\; =</span>\mathrm{<span\; class="notranslate">\; v</span>}$

(c) Attitude dynamics model

$\mathrm{<span\; class="notranslate">\; I</span>}\stackrel{<span\; class="notranslate">\; \&Center\; Dot;</span>}{\mathrm{<span\; class="notranslate">\; \&omega;</span>}}<span\; class="notranslate">\; +</span>\mathrm{<span\; class="notranslate">\; \&omega;</span>}<span\; class="notranslate">\; \&times;</span>\mathrm{<span\; class="notranslate">\; I\&omega;</span>}<span\; class="notranslate">\; =</span>{\mathrm{<span\; class="notranslate">\; T</span>}}_{\mathrm{<span\; class="notranslate">\; c</span>}}<span\; class="notranslate">\; +</span>{\mathrm{<span\; class="notranslate">\; T</span>}}_{\mathrm{<span\; class="notranslate">\; d</span>}}$

Among them, I is the moment of inertia matrix of the lander, ω is the attitude angular velocity vector of the lander in the inertial space, T _c and T _d are the control torque and disturbance torque, respectively. In the present invention, T _c mainly refers to jet torque, and T _d includes jet disturbance torque, gravity gradient torque, solar radiation pressure torque and the like.

The attitude dynamics model is the simplest rigid body model. For the present invention, since the fuel consumption is relatively large, the influence of liquid sloshing on the attitude of the lander should be fully considered.

(d) Attitude kinematics model

$\left\{\begin{array}{c}\stackrel{<span\; class="notranslate">\; \&Center\; Dot;</span>}{\mathrm{<span\; class="notranslate">\; \&epsiv;</span>}}<span\; class="notranslate">\; =</span>\frac{\mathrm{<span\; class="notranslate">\; 1</span>}}{\mathrm{<span\; class="notranslate">\; 2</span>}}<span\; class="notranslate">\; (</span>{\mathrm{<span\; class="notranslate">\; \&epsiv;</span>}}^{<span\; class="notranslate">\; \&times;</span>}<span\; class="notranslate">\; +</span>\mathrm{<span\; class="notranslate">\; \&eta;</span>}{\mathrm{<span\; class="notranslate">\; E.</span>}}_{\mathrm{<span\; class="notranslate">\; 3</span>}}<span\; class="notranslate">\; )</span>{\mathrm{<span\; class="notranslate">\; \&omega;</span>}}_{\mathrm{<span\; class="notranslate">\; b</span>}}\\ \stackrel{<span\; class="notranslate">\; \&Center\; Dot;</span>}{\mathrm{<span\; class="notranslate">\; \&eta;</span>}}<span\; class="notranslate">\; =</span><span\; class="notranslate">\; -</span>\frac{\mathrm{<span\; class="notranslate">\; 1</span>}}{\mathrm{<span\; class="notranslate">\; 2</span>}}{\mathrm{<span\; class="notranslate">\; \&epsiv;</span>}}^{\mathrm{<span\; class="notranslate">\; T</span>}}{\mathrm{<span\; class="notranslate">\; \&omega;</span>}}_{\mathrm{<span\; class="notranslate">\; b</span>}}\end{array}\right.$

The above formula is the attitude kinematics equation of the lander represented by four elements. Among them, ω _b is the attitude angular velocity vector of the lander in the body coordinate system, and its relationship with ω in 3) is: ω=ω _b +ω _o , where ω _o represents the angular velocity of the landing orbit of the lander in the inertial system vector. ε and η denote the vector and scalar parts of the four elements, respectively,

Among them, a and are the Euler axes and Euler angles used for the coordinate system rotation, respectively.

(7) Ground test and master control computer system 5

The ground test and general control computer system 5 is used for the management, data collection, processing and control of the test system. This system forms a computer network and can complete the display, storage and management of data. In the embodiment of the present invention, the ground testing and master control computer system 5 is replaced by a high-performance PC. The ground test and general control computer software is installed on each node PC of the computer network, and realizes the following functions in the form of configuration modules: management and control of the test system; collection, processing, storage and real-time display of test data; and simulation computer 11 resource sharing.

2. Workflow

Work process of the present invention is as follows:

(1) At the initial moment, the three-dimensional translational motion device 3 has a certain initial relative position and velocity, and the three-axis mechanical turntable 10 has a certain initial attitude angle and attitude angular velocity, which are used to simulate the relative position and attitude information of the lander;

(2) (a) the control computer 4 starts the rangefinder 7, the speedometer 8 and the rangefinder speed measurement target simulator 6 according to the ground test and the remote control instruction of the master control computer system 5, and turns on the laser imaging sensor 9 simultaneously;

(b) The distance and speed information of the lander relative to the lunar surface are measured by the rangefinder 7, the speedometer 8, and the distance and speed measurement target simulator 6, and the laser imaging sensor 9 performs three-dimensional imaging on the lunar surface sand table screen 2 to obtain 3D image information;

(c) The simulation computer 11 calculates the current attitude angle and attitude angular velocity information according to the three-axis mechanical turntable 10, and after the calculation of the gyro measurement model of one of the conventional sensor measurement models, the simulated gyro provides the lander attitude angular velocity measurement information;

(d) using the accelerometer measurement model of one of the conventional sensor measurement models in the simulation computer 11 to calculate and obtain the specific force measurement information of the lander;

(e) the above-mentioned distance, speed, image, specific force and attitude angular velocity measurement information are sent to the control computer 4;

(3) The control computer 4 finally obtains the main thrust guidance command (thrust-to-weight ratio of the main engine) and landing area selection through corresponding navigation, guidance and control calculations based on the transmitted distance, speed, image, specific force and attitude angular velocity measurement information command (translation direction), translation thruster guidance command (translation thruster working time) and attitude stability control command (attitude control thruster working time), and the above-mentioned propulsion system control command is passed to the simulation computer 11;

(4) According to the propulsion system control command given by the control computer 4, the simulation computer 11 starts to perform dynamic and kinematic calculations of the lander and conventional sensor navigation calculations, and obtains the position, velocity, attitude angle and position of the lander at the next moment. Attitude angular velocity six-degree-of-freedom information, and then the six-degree-of-freedom information is scaled and calculated, and the position and attitude information before and after scaling is transmitted to the ground test and the master control computer system 5 for data storage, sharing and real-time display;

(5) According to the position and attitude information transmitted from the simulation computer 11, the ground test and general control computer system 5 transmits the scaled position and speed control commands and angle and angular velocity control commands to the three-dimensional translational motion device 3 and the three-dimensional motion device respectively. The axis mechanical turntable 10 drives the two devices to move according to their respective instructions. At the same time, the ground test and master control computer system 5 also transmits the position and speed control instructions before scaling to the distance measurement and speed measurement target simulator 6 for the next moment. distance and speed measurement;

(6) Steps (2) to (5) will be repeated in the next control cycle to form a closed-loop GNC semi-physical simulation test system.

The contents not described in detail in the description of the present invention belong to the prior art known to those skilled in the art.

Claims (9)

1. The semi-physical simulation test system for lunar soft landing guidance, navigation and control is characterized in that it includes: a three-dimensional translation simulation module, a three-dimensional rotation simulation module, a control computer (4), a simulation computer (11) and a ground test and master control computer System (5), wherein: The three-dimensional translation simulation module is used to simulate the surface characteristics of the moon, and simulates the three-dimensional translation of the lander relative to the lunar surface according to the control command sent by the ground test and the master control computer system (5); The three-dimensional rotation simulation module is used to simulate the flight attitude of the lander. It works according to the power-on and power-off instructions from the control computer (4), measures the six-degree-of-freedom movement of the lander, and obtains the line-of-sight distance of the lander relative to the lunar surface. and speed information, the attitude angle and attitude angular velocity information of the lander itself and the three-dimensional image information of the lunar surface; the line-of-sight distance, the speed of the lander relative to the lunar surface and the three-dimensional image information of the lunar surface are sent to the control computer (4), and simultaneously Pass the attitude angle and the attitude angular velocity information to the simulation computer (11); carry out three-dimensional rotation according to the control instructions sent by the ground test and the master control computer system (5); The control computer (4) receives the remote control commands issued by the ground test and the master control computer system (5), controls the three-dimensional rotation simulation module to work and obtains distance, speed and image measurement information, and obtains specific force and gyro angular velocity from the simulation computer (11) Measurement information, GNC calculation is carried out according to the measurement information, the propulsion system control command is obtained and passed to the simulation computer (11); the obtained measurement information and the calculated control information are passed to the ground test and the master control computer system as telemetry data (5); The simulation computer (11) receives the remote control commands sent by the ground test and the master control computer system (5), obtains attitude angle and attitude angular velocity information from the three-dimensional rotation simulation module, calculates the gyro angular velocity measurement information through the gyro measurement model, and obtains the gyro angular velocity measurement information from the control computer. (4) Receive the propulsion system control command, obtain the specific force information, propulsion system parameters and position attitude information through simulation calculation, transmit the specific force and gyro angular velocity measurement information to the control computer (4), and transfer the propulsion system parameters and the position and attitude information are sent to the ground test and the master control computer system (5); The ground test and master control computer system (5) sends remote control instructions to the control computer (4) and the simulation computer (11), receives telemetry data from the control computer (4) and performs data storage, sharing and real-time display, and sends out remote control commands from the simulation computer (11). (11) Receive propulsion system parameters and position and attitude information, and give translation and rotation control commands according to the position and attitude information, and control the three-dimensional translation simulation module and the three-dimensional rotation simulation module to move as required. 2. The semi-physical simulation test system for guidance, navigation and control of lunar soft landing according to claim 1, characterized in that: the three-dimensional translation simulation module consists of a landing lunar surface simulator and a three-dimensional translation motion device (3) Composition; the landing moon simulator includes a light simulator (1) and a moon sand screen (2), the light simulator (1) is used to simulate the lighting conditions of the moon, and the moon sand screen (2) is used to simulate the moon The surface features of the landing lunar surface are placed on the three-dimensional translation motion device (3). 3. The lunar soft landing guidance, navigation and control semi-physical simulation test system according to claim 1, characterized in that: the three-dimensional rotation simulation module is composed of an external navigation sensor and a three-axis mechanical turntable (10), and the external The navigation sensor includes a ranging and speed measuring target simulator (6), a range finder (7), a speed measuring instrument (8) and a laser imaging sensor (9), and the ranging and speed measuring target simulator (6) is used to receive the range finder (6) and the beam of the speedometer (7), and return the beam to the rangefinder (6) and the speedometer (7) after delay and Doppler translation, and the laser imaging sensor (9) is used for the lunar surface The sand table screen (2) performs three-dimensional imaging; the external navigation sensor is placed on the three-axis mechanical turntable (10), and the flight attitude of the lander is simulated by controlling the movement of the three-axis mechanical turntable (10). 4. The lunar soft landing guidance, navigation and control semi-physical simulation test system according to claim 1, characterized in that: the control computer (4) receives instructions from the ground test and master control computer system (5), and controls The external navigation sensor in the three-dimensional rotation simulation module works, and the real line-of-sight distance and speed information of the lander relative to the lunar surface are measured by the rangefinder (7), the speedometer (8) and the distance measurement speed measurement target simulator (6) , the lunar surface sand table screen (2) is three-dimensionally imaged by the laser imaging sensor (9) to obtain the three-dimensional image information of the lunar surface, and at the same time the control computer (4) obtains the specific force and gyro angular velocity measurement information from the simulation computer (11), Adopt GNC algorithm to carry out navigation calculation according to described distance, speed, image, specific force and gyro angular velocity information, obtain propulsion system control command and send to simulation computer (11); The data is transmitted to the ground test and the master control computer system (5); 5. The lunar soft landing guidance, navigation and control semi-physical simulation test system according to claim 4, characterized in that: the GNC algorithm includes an autonomous navigation module, a trajectory guidance module and an attitude control module; the autonomous navigation module receives the simulated The specific force and gyro angular velocity measurement information sent by the computer (11) is used to obtain the six degrees of freedom information of the lander's position, velocity, attitude, and attitude angular velocity after performing inertial navigation calculations, and transmit the six degrees of freedom information to the trajectory guidance module , transmit the attitude and attitude angular velocity information to the attitude control module, and at the same time, the autonomous navigation module also receives the line-of-sight distance, velocity and lunar surface image information measured by the external navigation sensor in the three-dimensional rotation simulation module, and the measurement information is respectively After altitude calculation, speed calculation and landing area selection calculation, the calculation results are sent to the trajectory guidance module; the trajectory guidance module receives the information obtained by the autonomous navigation module through inertial navigation calculation, altitude, speed calculation and landing area selection calculation, And select the guidance mode according to the computer instruction to calculate the corresponding guidance law; the attitude control module receives the attitude and attitude angular velocity information calculated by the inertial navigation from the autonomous navigation module, and calculates the attitude control law according to different attitude control requirements; The calculation results of the guidance control laws in the trajectory guidance module and attitude control module can obtain the propulsion system control instructions for trajectory and attitude control. 6. The semi-physical simulation test system for guidance, navigation and control of lunar soft landing according to claim 5, characterized in that: the guidance modes of the trajectory guidance module include gravity turning nominal trajectory guidance mode, variable thrust guidance mode, Bang -Bang and phase plane guidance modes, and nominal trajectory guidance modes. 7. The lunar soft landing guidance, navigation and control semi-physical simulation test system according to claim 5, characterized in that: the attitude control mode of the attitude control module adopts phase plane control. 8. The lunar soft landing guidance, navigation and control semi-physical simulation test system according to claim 1, characterized in that: the simulation computer (11) receives the remote control instructions from the ground test and master control computer system (5), The attitude angle and attitude angular velocity information are obtained from the three-axis mechanical turntable (10), and the angular velocity measurement information of the simulated gyroscope is obtained after calculation by the dynamics and kinematics simulation model; the simulation computer (11) also receives the propulsion system control from the control computer (4) Instructions, the propulsion system parameters and the position and attitude information of the lander are obtained after the calculation of the dynamics and kinematics simulation model, and the relative force measurement information can also be obtained at the same time; the simulation computer (11) transmits the specific force and gyro angular velocity measurement information to The control computer (4) transmits the propulsion system parameters and position and attitude information to the ground test and master control computer system (5). 9. The lunar soft landing guidance, navigation and control semi-physical simulation test system according to claim 8, characterized in that: the dynamics and kinematics simulation model includes the mathematical model of the actuator, the dynamics and kinematics model, the track Scaled calculation models and conventional sensor measurement models. The mathematical model of the actuator receives the propulsion system control instructions given by the control computer (4), and obtains the propulsion system parameters after simulation calculation and transmits them to the dynamics and kinematics model and the ground test and master control computer system (5); dynamics Combining the propulsion system parameters with the kinematic model, the lander's position, velocity, attitude angle, and attitude angular velocity six-degree-of-freedom information are obtained through calculation, and the six-degree-of-freedom information is transmitted to the ground test and master control computer system (5) At the same time, the position and velocity information is transmitted to the orbit scale calculation model and the conventional sensor measurement model; the orbit scale calculation model performs scale calculation on the position and velocity information, and then transmits the scaled position and velocity information to the ground test and the total control computer system (5); the conventional sensor measurement model receives the position and velocity information given by the dynamics and kinematics model, and obtains the specific force measurement information after calculation by the accelerometer measurement model; at the same time, the conventional sensor measurement model also The attitude angle and attitude angular velocity information given by the three-axis mechanical turntable (10) are received, the gyro angular velocity measurement information is obtained through calculation of the gyro measurement model, and the specific force and gyro angular velocity measurement information are transmitted to the control computer (4).

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