CN112027115B - Landing and ascending spacecraft integrated control system - Google Patents
Landing and ascending spacecraft integrated control system Download PDFInfo
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- CN112027115B CN112027115B CN202010762819.7A CN202010762819A CN112027115B CN 112027115 B CN112027115 B CN 112027115B CN 202010762819 A CN202010762819 A CN 202010762819A CN 112027115 B CN112027115 B CN 112027115B
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- B—PERFORMING OPERATIONS; TRANSPORTING
- B64—AIRCRAFT; AVIATION; COSMONAUTICS
- B64G—COSMONAUTICS; VEHICLES OR EQUIPMENT THEREFOR
- B64G1/00—Cosmonautic vehicles
- B64G1/22—Parts of, or equipment specially adapted for fitting in or to, cosmonautic vehicles
- B64G1/24—Guiding or controlling apparatus, e.g. for attitude control
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- B—PERFORMING OPERATIONS; TRANSPORTING
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- B64G—COSMONAUTICS; VEHICLES OR EQUIPMENT THEREFOR
- B64G1/00—Cosmonautic vehicles
- B64G1/22—Parts of, or equipment specially adapted for fitting in or to, cosmonautic vehicles
- B64G1/24—Guiding or controlling apparatus, e.g. for attitude control
- B64G1/36—Guiding or controlling apparatus, e.g. for attitude control using sensors, e.g. sun-sensors, horizon sensors
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- B—PERFORMING OPERATIONS; TRANSPORTING
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- B64G—COSMONAUTICS; VEHICLES OR EQUIPMENT THEREFOR
- B64G1/00—Cosmonautic vehicles
- B64G1/22—Parts of, or equipment specially adapted for fitting in or to, cosmonautic vehicles
- B64G1/24—Guiding or controlling apparatus, e.g. for attitude control
- B64G1/36—Guiding or controlling apparatus, e.g. for attitude control using sensors, e.g. sun-sensors, horizon sensors
- B64G1/361—Guiding or controlling apparatus, e.g. for attitude control using sensors, e.g. sun-sensors, horizon sensors using star sensors
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Abstract
The integrated control system comprises a master controller, a slave controller, a terrain measurement sensor, a distance measurement sensor, a speed measurement sensor and a star sensor, can realize high-reliability control in the process of spacecraft ascending and landing, can discard partial products in ascending, reduces the weight of the system, solves the problems of poor reliability and large weight of the traditional spacecraft control structure, and considers the reliability requirement and the weight constraint.
Description
Technical Field
The invention relates to an integrated control system for a landing and ascending spacecraft, and belongs to the technical field of spacecraft control.
Background
With the continuous development of unmanned deep space exploration space missions, the landing and ascending control of extraterrestrial celestial bodies needs to be completed simultaneously in one mission. For example, the flying object "landing and ascending assembly" of Chang E five task in China needs to complete the lunar landing and ascending and landing tasks.
Whether the ascent or the landing mission consumes a large amount of fuel, the landing process needs to reduce the lunar velocity of the spacecraft from about 1.7km/s to 0, and the ascent mission needs to raise the lunar velocity of the spacecraft from 0 to about 1.7 km/s. Compared with the conventional spacecraft, the landing and ascending mission spacecraft can meet the launching weight constraint of the spacecraft only by reducing the dry weight of the system. Meanwhile, as the landing or rising process has irreversibility, the reliability of the system must be ensured, and the redundant components of related functions can be switched in real time when the system fails. In summary, two strict requirements for the design of the spacecraft control system architecture for landing and ascending tasks are: high reliability and light weight. Therefore, how to effectively meet the two requirements is the problem to be solved first by the design of the control system.
For the moon landing and ascending control task, the most typical representative is the apollonian moon cabin in the united states. Under the current technical development, the landing control and ascending control tasks of the control system of the Apollodenum lunar chamber are distributed in two computing systems, namely PNGS and AGS, and the requirement of the whole-course control task can be met only if the two systems are reliable; the method is characterized in that no star sensor is provided, and an astronaut is adopted to manually obtain an inertial attitude through an optical collimation system; the PNGS uses a platform inertial navigation system, the AGS uses a strapdown inertial navigation system, and the two sets of inertial navigation systems belong to a mechanical inertial navigation system. Two sets of computer systems and end products are installed in the riser. The landing distance measurement and the speed measurement both adopt a microwave system, namely a landing radar without backup products, and the landing radar is arranged on a lander. Obviously, the design is limited to the technical level of low performance and heavy weight of the electronic product at the time, and the reliability of the whole system is poor.
The Chang' e five-number landing and ascending combined body is a spacecraft which can continuously complete the landing and ascending tasks for the first time after the Apollo task. The current technology has advanced far beyond the Apollo era. Therefore, how to design a control system meeting the task requirement under the condition of high reliability and considering the requirement of weight optimization in the prior art is the problem which is mainly solved by a detector system.
Therefore, in the design of the control system of the ChangE five-model landing and ascending combination, firstly, an integrated design is carried out, namely, the landing and ascending share an inertia measurement product and a high-reliability main controller, wherein the high-reliability main controller comprises a star sensor, an inertia measurement assembly and a top layer controller. The common products can work in landing and lifting control, so that related products do not need to be configured for landing and lifting tasks respectively, and the aim of saving weight is fulfilled. Secondly, the design makes full use of the opportunity that part of the products can be discarded when the landing device ascends, and the slave controller and the products for landing are arranged on the landing device and discarded when the landing device ascends, so that the weight of the landing device is reduced. In order to ensure the reliability of the information link, the main controller is directly communicated with each product, when a single communication link fails, the redundant product can also be normally communicated with the main controller, and the slave controller only executes low-frequency tasks such as power distribution, remote measurement, remote control and the like. Then, for the safety terrain selection function without landing, since it is only required at landing, it is also assigned to the slave controller, thereby reducing the weight of the riser. Finally, heterogeneous backup is adopted for terminal products, and the reliability of the system is improved.
Disclosure of Invention
The technical problem solved by the invention is as follows: aiming at the problems of poor technical reliability, old product technical level and the like of the existing landing and ascending task control system, and combining new requirements of unmanned landing for autonomous identification of a safety zone and the like, the landing and ascending spacecraft integrated control system is provided.
The technical scheme for solving the technical problems is as follows:
the integrated control system of the landing and ascending spacecraft comprises a main controller, a slave controller, a terrain measurement sensor, a distance measurement sensor, a speed measurement sensor and a star sensor, wherein:
a main controller: the system is arranged on an ascender part of the spacecraft, performs navigation guidance and control on the whole flight process of the spacecraft, performs power supply and distribution and remote control and remote measurement control on a terminal product arranged in the ascender, realizes direct communication between the ascender and the terminal product in the lander, sends a control instruction for the lander to the slave controller and receives landing safety point information returned by the slave controller;
from the controller: the lander part is arranged on the spacecraft, receives a control instruction which is sent by the main controller and faces the lander, performs power supply and distribution and remote control and remote measurement control on a terminal product on the lander according to the control instruction, communicates with the terrain measuring sensor and processes a landing image which is obtained by the landimage sensor and faces the terrain measuring sensor to obtain a landing safety point, and feeds information of the landing safety point back to the main controller;
a topographic measurement sensor: the lander part is arranged for automatically acquiring landing terrain in the landing process and sending the acquired landing image to the slave controller for processing according to the instruction of the slave controller;
distance measuring sensor: the device is arranged on the lander part and is used for measuring the distance from the spacecraft to the surface of the star body to be landed in the landing process of the lander in real time;
speed measuring sensor: the spacecraft speed measurement device is arranged on the lander part and is used for measuring the speed of a spacecraft relative to the surface of a landing satellite in the landing process of the lander in real time;
and (3) star sensor: the device is arranged on the riser part and is used for measuring inertial attitude parameters of the whole process of the spacecraft;
an inertia measurement assembly: and the device is arranged on the riser part and is used for measuring inertial attitude angular velocity parameters of the whole process of the spacecraft motion.
The main controller adopts a three-machine redundancy hot backup mode to ensure the reliability and uninterrupted switching of the flight full-process controller.
The slave controller adopts a dual-machine redundancy hot backup mode to ensure the short-time reliability and uninterrupted switching in the half-way after landing.
The star sensors are in redundant configuration according to task requirements, the star sensors comprise star sensors with dust covers and sensors without dust covers, the star sensors with the dust covers have the capacity of preventing moon dust from polluting lenses in the landing and takeoff processes, the star sensors without the dust covers are used for preventing the dust covers from being opened due to faults and losing the measurement function, the star sensors are arranged on the ascenders, and the inertial attitude acquisition capacity of the system is not affected due to the fact that a single product fails.
The distance measuring sensor and the speed measuring sensor are both provided with laser and microwave system measuring sensors.
The topographic survey sensor is provided with a laser and an optical system survey sensor.
The inertia measurement assembly is provided with a laser and optical fiber system inertia measurement assembly.
The communication between the main controller and the lander product is divided into direct communication and indirect communication according to the frequency and importance of information communication, the main controller is indirectly communicated with the terrain measurement sensor through the sub-controller and is directly communicated with other sensors, and the condition that the sub-controller fails to cause the incapability of acquiring the landing information can be avoided.
Compared with the prior art, the invention has the advantages that:
(1) according to the integrated control system for the landing and ascending spacecraft, the landing sensors and the slave controllers can be left on the lunar surface by utilizing the takeoff time of the spacecraft surface through the control framework of the master controller and the slave controllers, and the realization of the ascending function is not influenced. The method can reduce the weight of the system during rising, achieve the effect of reducing fuel consumption, realize the landing safety point calculation through the slave controller, because the calculation amount of the landing safety point is large, need more calculation resources, set up the slave controller can reduce the calculation ability demand of the master controller, help to reduce the weight of the master controller and guarantee the real-time of the calculation of the master controller;
(2) compared with the conventional redundancy method for increasing the number of products, the heterogeneous backup method for the sensors obviously avoids common cause actual effect and better ensures the reliability of the system, the high-reliability three-machine redundancy hot backup mode is adopted by the main controller to ensure the whole-course reliability and real-time uninterrupted flight process, and the safety point identification function of the sub controller is only used for short-term work after landing, so that the high-reliability two-machine redundancy mode is adopted to effectively ensure the reliability of the system. According to the actual working process, the redundancy forms of the master controller and the slave controller are set in a differentiated mode, and the reliability requirement and the weight constraint are both considered;
(3) the invention integrates necessary information paths on the premise of preferentially ensuring the reliability of the power process, thereby reducing the weight of the system. The information paths of the landing sensor information, the control instruction and other key information of the power process are not combed by the slave controller even if the product is positioned on the lander, and a redundant path is adopted to be directly connected with the master controller, so that the transmission nodes of the information paths are prevented from increasing and reducing the reliability. Information paths with low power process criticality such as power supply and distribution, remote control and remote measurement are combed, integrated and uniformly managed through the slave controller, the number of links is reduced, and the weight of the system is optimized.
Drawings
FIG. 1 is a schematic structural diagram of an integrated spacecraft control system provided by the invention;
Detailed Description
The utility model provides a landing and rising spacecraft integrated control system, is applicable to and carries out the unmanned landing and rising task of extraterrestrial celestial body, specifically includes main control unit, from controller, landing image sensor, distance measurement sensor, speed measurement sensor, star sensor, and landing and rising spacecraft control system adopts layered structure, and the top layer sets up main control unit, and the bottom is provided with from the controller, and main control unit sets up in the riser part, sets up in the lander part from the controller, wherein:
the spacecraft comprises an ascender and a lander part, wherein a main controller, a star sensor and an inertia measurement assembly are arranged on the ascender, and a slave controller and other sensors are arranged on the lander, and the main controller, the star sensor and other sensors are specifically as follows:
the main controller performs navigation guidance and control on the whole flight process of the spacecraft, performs power supply and distribution and remote control and remote measurement control on a terminal product arranged in the ascender, realizes communication between the ascender and the terminal product in the lander, sends a control instruction for the lander to the slave controller and receives information returned by the slave controller; the main controller adopts a three-machine redundancy hot backup mode to ensure the reliability and uninterrupted switching of the flight full-process controller;
the slave controller receives a control instruction which is sent by the master controller and faces the lander, performs power supply and distribution and remote control and remote measurement control on a terminal product on the lander according to the control instruction, communicates with the terrain measurement sensor and processes landing terrain information acquired by the terrain measurement sensor to acquire a landing safety point; the slave controller adopts a dual-computer redundancy hot backup mode to ensure the short-time reliability and uninterrupted switching in the half-way after landing;
the method comprises the steps that a terrain measurement sensor acquires landing terrain information in a landing process and sends the obtained landing terrain information to a slave controller for processing according to an instruction of the slave controller, and the terrain measurement sensor is provided with an optical system terrain sensor and a laser system terrain sensor;
the distance measuring sensor measures the distance from the spacecraft to the surface of the star body to be landed in the landing process of the lander in real time; the distance measuring sensor is provided with a laser and microwave system measuring sensor;
the speed measuring sensor measures the landing speed of the spacecraft in real time in the landing process of the lander; the speed measuring sensor is provided with a laser and microwave system measuring sensor;
the star sensor measures the inertial attitude parameters of the whole process of the spacecraft; the quantity of the star sensors is determined according to task requirements, the star sensors are provided with dust covers, the star sensors are not provided with the dust covers, the star sensors provided with the dust covers are used for preventing moon dust pollution in the measuring process, and the sensors not provided with the dust covers can avoid the fault mode that the star sensors fail due to mechanical faults of the dust covers. The sensor without the dust cover and the star sensor with the dust cover are arranged at different positions of the raiser;
inertia measurement assembly for measuring angular velocity of spacecraft in inertia space and provided with laser and optical fiber system
In the design of a communication link, remote control and remote measurement instructions and power supply and distribution instructions facing the land device are uniformly managed by the slave controller, and communication information of a terminal product is directly connected with the master controller by adopting a redundant path.
The following is further illustrated with reference to specific examples:
in this embodiment, as shown in fig. 1, the landing and ascending spacecraft control system adopts a layered structure, and a product shared by a main controller and a landing and ascending task is installed on the spacecraft and mainly includes a satellite sensor and an inertia measurement component; the main controller is responsible for power supply and distribution and remote control and remote measurement of the riser product, is responsible for direct communication with the lander product and the riser product, does not contain a terrain measuring sensor, and is responsible for instruction calculation and output of navigation guidance and control in the whole flight process;
the slave controller is responsible for responding to the control instruction of the master controller and completing power supply and distribution and remote control and remote measurement control of the lander product according to the instruction of the master controller; the main controller is responsible for processing the landing image to calculate a landing safety point and feeding the information of the landing safety point back to the main controller;
the main controller adopts a high-reliability three-machine redundancy hot backup mode, so that the reliability and uninterrupted switching of the flight full-process controller are ensured; the slave controller adopts a dual-computer redundancy hot backup mode with higher reliability, so that the short-time reliability requirement and uninterrupted switching in the half-way after landing are ensured;
the landing sensors in the control system are configured with: the distance measuring sensor adopts measuring sensors of two systems of laser and microwave, the speed measuring sensor adopts measuring sensors of two systems of laser and microwave, and the topographic survey adopts two sensors of laser and optical systems;
the control system is provided with a plurality of inertia measurement assemblies, each assembly can complete the measurement of the inertia angular velocity and the acceleration, and different measurement system products are adopted for redundancy backup. A laser system inertia assembly and an optical fiber system inertia assembly are configured;
the star sensors are configured in a plurality of numbers, and a local heterogeneous design is adopted. Some star sensors are provided with dust covers, so that the pollution of lunar dust can be prevented. Some star sensors are not provided with dust covers, so that the failure of products caused by mechanism faults can be avoided.
The elevator is provided with a three-machine redundant hot backup main controller, the controller comprises three computer systems which take TSC695 as a core processor, the correctness of a control instruction is confirmed by mutual comparison and two-out of three, the lander is provided with a slave controller, the controller comprises two controllers which take SMJ320C6701 as the core processor, the current aircraft is selected by the designation of the master controller, the control instruction of the master controller is responded, the power supply and distribution and remote control and remote measurement control of lander products are completed according to the instruction of the master controller, and the controllers are communicated with a terrain measuring sensor and process land shape information to obtain a landing safety point;
the landing sensors arranged on the lander mainly comprise a laser distance measuring sensor taking a laser LD diode as a core device, a Ka-band microwave pulse distance measuring sensor, a laser speed measuring sensor taking a fiber laser as a core component, a Ka-band microwave continuous wave speed measuring sensor, an imaging type terrain measuring sensor of an optical system and a laser system scanning type terrain measuring sensor. The lifter is provided with a laser inertia measurement assembly and an optical fiber inertia measurement assembly, and is provided with 2 star sensors with dustproof devices and 1 star sensor without dustproof devices.
The main controller receives instruction information (including power supply instructions and remote control and remote measurement instructions) which are totally sent to a control system by the detector from other systems of the ascender in a 1553B bus and remote control instruction line mode, and sends the instruction information related to the lander and the instruction information which is autonomously generated by the main controller and is sent to the slave controller through a serial port. And the main controller receives information such as landing safety points and the like fed back from the slave controller through the serial port. The main controller and each terminal product carry out one-to-one communication through a serial port, and measurement and remote measurement information of the products are directly obtained. The main controller sends power supply instructions to each terminal product of the ascender in a remote control instruction line mode, and remote measurement information of each terminal product of the ascender is obtained in an analog quantity remote measurement acquisition mode. All serial ports and command lines adopt a double-point double-line mode to carry out redundancy of transmission links.
The slave controller receives command information (including power supply commands and remote control and telemetering commands) sent by the detector from other systems of the lander in a 1553B bus and remote control command line mode, receives the command information sent by the master controller through a serial port, receives the terrain information sent by the terrain sensor through an LVDS bus, sends the remote control and telemetering information to the terrain sensor through the serial port, sends the power supply commands to terminal products of the lander in a remote control command line mode, and obtains the telemetering information of the terminal products of the lander in an analog quantity telemetering and collecting mode. All serial ports and command lines adopt a double-point double-line mode to carry out redundancy of transmission links.
Those skilled in the art will appreciate that those matters not described in detail in the present specification are well known in the art.
Claims (8)
1. The integrated control system of the landing and ascending spacecraft is characterized in that: the device comprises a master controller, a slave controller, a terrain measurement sensor, a distance measurement sensor, a speed measurement sensor and a star sensor, wherein:
a main controller: the system is arranged on an ascender part of the spacecraft, performs navigation guidance and control on the whole flight process of the spacecraft, performs power supply and distribution and remote control and remote measurement control on a terminal product arranged in the ascender, realizes direct communication between the ascender and the terminal product in the lander, sends a control instruction for the lander to the slave controller and receives landing safety point information returned by the slave controller;
from the controller: the lander part is arranged on the spacecraft, receives a control instruction which is sent by the main controller and faces the lander, performs power supply and distribution and remote control and remote measurement control on a terminal product on the lander according to the control instruction, communicates with the terrain measurement sensor and processes a landing image which is obtained by the landimage sensor and faces the terrain measurement sensor to obtain a landing safety point, and feeds the information of the landing safety point back to the main controller;
a topographic measurement sensor: the lander part is arranged for automatically acquiring landing terrain in the landing process and sending the acquired landing image to the slave controller for processing according to the instruction of the slave controller;
distance measuring sensor: the device is arranged on the lander part and is used for measuring the distance from the spacecraft to the surface of the star body to be landed in the landing process of the lander in real time;
speed measuring sensor: the spacecraft relative to the surface of the star body to be landed is arranged on the lander part and used for measuring the speed of the spacecraft relative to the surface of the star body to be landed in the landing process of the lander in real time;
and (3) star sensor: the device is arranged on the riser part and is used for measuring inertial attitude parameters of the whole process of the spacecraft;
an inertia measurement assembly: and the device is arranged on the riser part and is used for measuring inertial attitude angular velocity parameters of the whole process of the spacecraft motion.
2. The integrated landing and ascending spacecraft control system according to claim 1, wherein: the main controller adopts a three-machine redundancy hot backup mode to ensure the reliability and uninterrupted switching of the main controller in the whole flight process.
3. The integrated landing and ascending spacecraft control system according to claim 1, wherein: the slave controller adopts a dual-machine redundancy hot backup mode to ensure the short-time reliability and uninterrupted switching in the half-way after landing.
4. The integrated landing and ascending spacecraft control system according to claim 1, wherein: the star sensors are in redundant configuration according to task requirements, the star sensors comprise star sensors with dust covers and sensors without dust covers, the star sensors with the dust covers have the capacity of preventing moon dust from polluting lenses in the landing and takeoff processes, the star sensors without the dust covers are used for preventing the dust covers from being opened due to faults and losing the measurement function, the star sensors are arranged on the ascenders, and the inertial attitude acquisition capacity of the system is not affected due to the fact that a single product fails.
5. The integrated landing and ascending spacecraft control system according to claim 1, wherein: the distance measuring sensor and the speed measuring sensor are both provided with laser and microwave system measuring sensors.
6. The integrated landing and ascending spacecraft control system according to claim 1, wherein: the topographic survey sensor is provided with a laser and an optical system survey sensor.
7. The integrated landing and ascending spacecraft control system according to claim 1, wherein: the inertia measurement assembly is provided with a laser and optical fiber system inertia measurement assembly.
8. The integrated landing and ascending spacecraft control system according to claim 1, wherein: the communication between the main controller and the lander product is divided into direct communication and indirect communication according to the frequency and importance of information communication, the main controller is indirectly communicated with the terrain measurement sensor through the sub-controller and is directly communicated with other sensors, and the condition that the sub-controller fails to cause the incapability of acquiring the landing information can be avoided.
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| CN113091490A (en) * | 2021-02-26 | 2021-07-09 | 北京空间飞行器总体设计部 | Combined heat sink-based moon sampling return detector autonomous thermal management system |
| CN114019788B (en) * | 2021-10-08 | 2024-03-26 | 北京控制工程研究所 | Partition-based rapid translation obstacle avoidance method in landing process |
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