Mechanical arm cabin section assembly process posture prediction and on-orbit test method and electronic device

Abstract

The application relates to a mechanical arm cabin section assembly process posture prediction and on-orbit test method and electronic equipment, which comprises the following steps: establishing a floating base mechanical arm dynamics model; establishing a cabin section assembly process simulation model; developing mechanical arm cabin section assembly process dynamics pre-simulation; developing uncontrolled flight posture test of the flight posture of a combination body; correcting the simulation model and dynamics parameters by using simulation results and telemetry data; and developing simulation prediction again according to the cabin section assembly task state by using the corrected simulation model and dynamics parameters. The application adopts a simulation-on-orbit test-model correction-re-simulation mode, corrects dynamics parameters by combining on-orbit test data on the basis of a floating base mechanical arm dynamics basic equation, overcomes the difficulty that the ground cannot systematically verify the dynamics in the transfer process, improves simulation prediction accuracy, improves task measurement and control coverage, effectively guarantees reliable implementation of the task, and accumulates experience and technical foundation for space station construction and application development.

Description

TECHNICAL FIELD

[0001] The application relates to a mechanical arm cabin section assembly process posture prediction and on-orbit test method and electronic equipment. BACKGROUND

[0002] The Chinese space station is composed of a core cabin, an experiment cabin I and an experiment cabin II, which are assembled on-orbit. The experiment cabin I and the experiment cabin II are axially docked with the core cabin to form an assembly, and then the mechanical arm is used to complete the final configuration of the permanent parking.

[0003] During the mechanical arm transfer process, the space station assembly exhibits flexible multi-body dynamics and control coupling, and the assembly configuration size, mechanical arm motion path, out-of-cabin accessory dynamics load, energy balance and ground and space measurement and control link multi-constraints. In order to meet the design constraints in all aspects and decouple dynamics and control, the cabin section assembly process adopts an attitude uncontrolled design scheme. The accurate prediction of the assembly attitude is particularly important.

[0004] The space mechanical arm plays an important role in the assembly and construction of the foreign "Peace" space station and the international space station, but China lacks technical foundation and experience in the attitude prediction of the space station assembly and construction, especially in the cabin section transfer process.

[0005] (1) The "Peace" space station is equipped with a 2-DOF transfer mechanical arm, and the transfer process is automatically completed in about 1 hour. During the transfer, the assembly stops attitude control, and after the cabin section transfer is completed, the assembly is automatically controlled according to the program, and the transfer process does not depend on ground measurement and control. Since the transfer mechanical arm has only 2 degrees of freedom, the ground established a physical verification system before the first transfer of the Peace space station, and used the degree of freedom decomposition method to carry out full physical verification for on-orbit tasks.

[0006] (2) At the beginning of the construction of the international space station, the docking task of the Russian cabin section and the American cabin section is also completed with the help of the Canadian 1 arm of the space shuttle. During this period, the attitude of the international space station + space shuttle assembly is stopped, and then the attitude control engine of the space shuttle is used to complete the capture connection of the APAS type docking mechanism between the Russian cabin section and the American cabin section. The simulation model based on the space shuttle mechanical arm has been verified in a large number of on-orbit tasks, and the simulation analysis is mainly carried out on the ground before the task is executed.

[0007] (3) In 2003, the Columbia space shuttle disintegrated on reentry due to the loss of thermal protection tiles. This incident led to an increase in the functional requirements of the space shuttle's robotic arm, and NASA developed a space shuttle maintenance maneuver scheme: in the "time-sharing operation mode", the robotic arm moves freely in the International Space Station, and when the attitude drift approaches the allowable boundary (determined by the dynamic load, TT&C and energy balance, etc.), the robotic arm joints are locked, the attitude is controlled by the Russian cabin segment attitude control engine jet, and then the attitude is stopped and the robotic arm operation is continued. Eventually, this scheme was not implemented on-orbit, and a more stable and reliable system scheme was adopted - the robotic arm gripped the upper extension rod to complete the observation task of the space shuttle heat dissipation structure.

[0008] China lacks technical foundation and experience in space station assembly and construction, mechanical arm application engineering implementation, especially in the attitude prediction of cabin segment transfer process. SUMMARY

[0009] In view of the above technical problems, the present application provides a mechanical arm cabin assembly process attitude prediction and on-orbit test method, which is used to develop a mechanical arm cabin transfer scheme system design and analysis, improve the attitude simulation prediction accuracy of the uncontrolled process of cabin transfer, and provide a technical basis for accurately developing a TT&C scheme design during the mission flight implementation and effectively ensuring the reliable implementation of the mission.

[0010] The technical solution for achieving the object of the present application is: a mechanical arm cabin assembly process attitude prediction and on-orbit test method, comprising the following steps:

[0011] Step S1, establishing a floating base mechanical arm dynamics model;

[0012] Step S2, establishing a cabin assembly process simulation model;

[0013] Step S3, developing a mechanical arm cabin assembly process dynamics pre-simulation;

[0014] Step S4, developing an uncontrolled flight attitude test of the flight attitude of the combination;

[0015] Step S5, using the simulation results obtained in step S3 and the telemetry data obtained in step S4 to correct the simulation model and the dynamics parameters;

[0016] Step S6, using the corrected simulation model and dynamics parameters to develop simulation prediction again according to the cabin assembly task state.

[0017] According to one aspect of the present application, in the step S1, specifically comprising:

[0018] Step S101, based on the floating base manipulator system and the basic equation of dynamics, considering the effect of environmental torque on the floating base manipulator system, a system dynamics mathematical model of the floating base manipulator is established, and the specific formula is as follows:

[0019]

[0020] Wherein, H m is the inertia matrix of the manipulator system itself, H b is the inertia tensor of the core cabin, H bm is the coupling inertia between the core cabin and the manipulator, Θ is the joint angle vector, x0 is the core cabin pose vector, c b , c m are nonlinear forces related to the core cabin motion and the manipulator motion respectively, including centripetal force and Coriolis force; F b is the control force and control torque acting on the base, τ m is the driving torque of the manipulator joint, F e is the external force and external torque received by the manipulator end and the environment, J b is the Jacobian matrix of the core cabin as the floating base of the manipulator, J m is the Jacobian matrix of the manipulator;

[0021] Step S102, a joint control model of the manipulator is established, and the manipulator joint controller is simplified as a PD controller, and the driving torque and control rate of each joint of the manipulator are as follows:

[0022]

[0023] Wherein, i=1, 2, …, 7. k pi and k di are control parameters of each joint controller, q di and q i are the expected angle and actual angle of each joint respectively;

[0024] Step S103, considering the effect of environmental torque on the floating base manipulator system, an environmental disturbance model is established.

[0025] According to one aspect of the present application, in step S103, the environmental torque includes gravity gradient torque and aerodynamic torque, and the calculation of the gravity gradient torque specifically includes:

[0026] According to the rotation matrix corresponding to the attitude angle Wherein ψ represents the yaw attitude angle, is the roll attitude angle, and θ is the pitch attitude angle:

[0027]

[0028] The gravity gradient moment M is calculated using a theoretical model formula g , where ω0 is the orbit angular velocity, is the inertia matrix of the spacecraft,

[0029]

[0030] According to one aspect of the present application, in step S103, calculating the aerodynamic moment specifically includes:

[0031] The space station cabin is simplified as a combination model of "cylinder segments and sailboard planes", and the aerodynamic moment of the space station cabin and sailboard is calculated in real time,

[0032] wherein the cabin diameters of the core cabin, the forward cargo spaceship, the rear cargo spaceship and the radial manned spaceship are D0, D2, D3 and D 13 , respectively; A ha , B ha , C ha , D ha , A hb , B hb , C hb , D hb are the left and right sailboard corner point coordinates of the core cabin, A 3a , B 3a , C 3a , D 3a , A 3b , B 3b , C 3b , D 3b are the left and right sailboard corner point coordinates of the rear cargo spaceship, A 2a , B 2a , C 2a , D 2a , A 2b , B 2b , C 2b , D 2b are the left and right sailboard corner point coordinates of the forward cargo spaceship, and A 13a , B 13a , C 13a , D 13a , A 13b , B 13b , C 13b , D 13b are the left and right sailboard corner point coordinates of the radial manned spaceship.

[0033] According to one aspect of the present application, in step S103, calculating the aerodynamic moment of the space station cabin and sailboard specifically includes:

[0034] In step S1031, the aerodynamic drag force F of the combination cabin and sailboard in the orbit system is calculated, and the formula is:

[0035] F = [F ox F oy F oz ] T , F oy = 0, F oz = 0,

[0036] wherein, the atmospheric resistance coefficient C d is taken as 2.5, the atmospheric density p is taken as 4.47E-12 kgm -3 , for the space station operating orbit spacecraft, the wind speed component v x of the orbit system x axis is -7.310 km / s, the wind speed in the y axis and z axis directions is ignored;

[0037] Step S1032, the projection transformation of the coordinate system is carried out through the flight attitude angle, and the aerodynamic resistance F d of the space station body system is calculated, and the formula is:

[0038]

[0039] Step S1033, the aerodynamic moment M a of the space station body system is calculated, and the formula is:

[0040] M a = r ca x F d ,

[0041] wherein, M ax , M ay , M az are components of the aerodynamic moment in the space station body control coordinate system,

[0042] r ca is the vector of the center of mass pressure center moment vector in the body control coordinate system, and the formula is:

[0043]

[0044] r ca × is the corresponding antisymmetric matrix of r ca , and the formula is:

[0045]

[0046] According to one aspect of the present application, in step S2, specifically comprising:

[0047] Step S201, a three-dimensional model of the cabin and the mechanical arm is established and imported into ADAMS software, a dynamic model is set according to a flight task simulation, and a Simulink module of the dynamic model is obtained

[0048] The flight task at least includes parameters of a coordinate system and mass characteristics, an initial flight attitude angle and an initial angular velocity of the assembly, joint motion of the mechanical arm, a space environment and an engine thrust of the cargo ship;

[0049] Step S202, a core cabin / cargo ship gravity gradient torque calculation module and a core cabin / cargo ship aerodynamic torque calculation module are established in Simulink.

[0050] Step S203, a repositioning attitude test simulation model is generated in a state that the mechanical arm keeps the initial state and does not move.

[0051] Step S204, a cabin section assembly simulation model is generated in a state that the mechanical arm drags the cargo ship to move.

[0052] According to one aspect of the present application, in step S3, pre-simulations of two states are respectively carried out according to the initial dynamic parameter conditions in step S1 and the simulation model established in step S2:

[0053] Step S301, system dynamics pre-simulation of the assembly attitude without control in a state that the mechanical arm keeps the initial state and does not move.

[0054] Step S302, system dynamics pre-simulation of the assembly attitude without control in a state that the mechanical arm drags the cargo ship to move in the cabin section assembly process.

[0055] According to one aspect of the present application, in step S4, specifically includes:

[0056] Step S401, according to a cabin section assembly scheme design state and a nominal stop control initial flight attitude angle design state, a full system is organized and arranged to carry out a no-control flight attitude test of the assembly repositioning flight attitude.

[0057] Step S402, initial attitude angles and initial attitude angular velocities of the assembly at a stop control moment, and attitude angle telemetry sequences of the assembly during the no-control period are collected.

[0058] According to one aspect of the present application, in step S5, specifically includes:

[0059] Step S501, matching consistency and correctness of the scheme design and the simulation result are confirmed, and deviations between flight attitude test telemetry data obtained in step S401 during the no-control attitude test of the assembly and simulation results obtained in step S301 are compared and analyzed.

[0060] Step S502, parameters are adjusted for the simulation result with the attitude prediction deviation greater than the preset deviation;

[0061] Step S503, the telemetry parameters of the attitude test are taken as the reference for the simulation comparison, and the simulation analysis of the attitude test is carried out again;

[0062] Step S504, steps S502 and S503 are repeated until the deviation between the simulation result and the telemetry parameters of the attitude test is less than the preset threshold accepted by the tracking and control capability.

[0063] According to one aspect of the present application, it comprises one or more processors, one or more memories, and one or more computer programs; wherein the processor is connected with the memory, and the one or more computer programs stored in the memory are executed by the processor when the electronic device is running, so that the electronic device performs a mechanical arm cabin section assembly process attitude prediction and on-orbit test method according to any one of claims 1-9.

[0064] According to the concept of the present application, a mechanical arm cabin section assembly process attitude prediction and on-orbit test method and an electronic device are proposed, comprising the following steps: Step S1, establishing a floating base mechanical arm dynamics model; Step S2, establishing a cabin section assembly process simulation model; Step S3, carrying out a mechanical arm cabin section assembly process dynamics pre-simulation; Step S4, carrying out an uncontrolled flight attitude test of the flight attitude of the combination; Step S5, using the simulation result obtained in step S3 and the telemetry data obtained in step S4 to correct the simulation model and the dynamics parameters; Step S6, using the corrected simulation model and the dynamics parameters to carry out simulation prediction again according to the cabin section assembly task state. The present application adopts the simulation-on-orbit test-model correction-re-simulation method, uses the floating base mechanical arm dynamics basic equation as the basis, corrects the dynamics parameters combined with the on-orbit test data, overcomes the difficulty that the dynamics of the transfer process cannot be verified on the ground, improves the accuracy of the attitude simulation prediction during uncontrolled flight, improves the task tracking and control coverage, effectively ensures the reliable implementation of the task, and accumulates experience and technical foundation for the construction and application development of space stations. BRIEF DESCRIPTION OF DRAWINGS

[0065] Figure 1 A flowchart schematically showing a mechanical arm cabin section assembly process attitude prediction and on-orbit test method according to an embodiment of the present application;

[0066] Figure 2 A schematic diagram showing the overall process of the cabin section assembly process attitude prediction according to an embodiment of the present application;

[0067] Figure 3 A schematic diagram showing a floating base mechanical arm dynamics model according to an embodiment of the present application;

[0068] Figure 4 The schematic representation illustrates a simplified geometric model of an assembly for calculating aerodynamic torque according to an embodiment of the present invention.

[0069] Figure 5 This schematic representation illustrates an ADAMS dynamic model of a compartment assembly process according to one embodiment of the present invention.

[0070] Figure 6 This illustration illustrates a simulation modeling process for the module assembly process according to one embodiment of the present invention.

[0071] Figure 7 This schematic representation illustrates a Simulink simulation model of a module assembly according to one embodiment of the present invention.

[0072] Figure 8 This schematic representation illustrates a Simulink simulation model of an on-orbit test of a transposition attitude according to an embodiment of the present invention.

[0073] Figure 9 The diagram illustrates a comparison of the pre-simulated posture curves of a robotic arm dragging a cargo ship for rotation and the robotic arm not moving, according to one embodiment of the present invention. Detailed Implementation

[0074] To more clearly illustrate the technical solutions in the embodiments of the present invention or the prior art, the accompanying drawings used in the embodiments will be briefly described below. Obviously, the drawings described below are merely some embodiments of the present invention, and those skilled in the art can obtain other drawings based on these drawings without any creative effort.

[0075] The present invention will now be described in detail with reference to the accompanying drawings and specific embodiments. The embodiments cannot be described in detail here, but the embodiments of the present invention are not limited to the following embodiments.

[0076] like Figures 1 to 9 As shown, the present invention provides a method for attitude prediction and on-orbit testing during the assembly process of a robotic arm module, comprising the following steps:

[0077] Step S1: Establish the dynamic model of the floating base robotic arm;

[0078] Step S2: Establish a simulation model of the module assembly process;

[0079] Step S3: Conduct dynamic pre-simulation of the robotic arm module assembly process;

[0080] Step S4: Conduct uncontrolled flight attitude tests of the combined vehicle's shifting flight attitude;

[0081] Step S5, the simulation model and the dynamic parameters are corrected by using the simulation result obtained in step S3 and the telemetry data obtained in step S4;

[0082] Step S6, the simulation prediction is carried out again according to the cabin segment assembly task state by using the corrected simulation model and the dynamic parameters.

[0083] In this embodiment, the simulation-in-orbit test-model correction-re-simulation mode is adopted, the dynamic parameters are corrected based on the floating base manipulator dynamics basic equation and the in-orbit test data, the difficulty that the dynamics of the transfer process cannot be verified on the ground is overcome, the attitude simulation prediction accuracy during the uncontrolled period is improved as a whole, the task measurement and control coverage is improved, the reliable implementation of the task is effectively ensured, and the experience and technical foundation are accumulated for the space station construction and application development.

[0084] As shown in the figure, Figure 3 in one embodiment of the present application, preferably, in the step S1, specifically includes:

[0085] Step S101, based on the floating base manipulator system and the dynamics basic equation, the action of the environmental moment on the floating base manipulator system is considered, and the system dynamics mathematical model of the floating base manipulator is established, and the specific formula is as follows:

[0086]

[0087] Wherein, H m is the inertia matrix of the manipulator system itself, H b is the inertia tensor of the core cabin, H bm is the coupling inertia between the core cabin and the manipulator, Θ is the joint angle vector, x0 is the core cabin pose vector, c b , c m are nonlinear forces related to core cabin motion and manipulator motion respectively, including centripetal force and Coriolis force; F b is the control force and control moment acting on the base, τ m is the driving torque of the manipulator joint, F e is the external force and external moment received by the manipulator end and the environment, J b is the Jacobian matrix of the core cabin as the floating base of the manipulator, J m is the Jacobian matrix of the manipulator;

[0088] Step S102, the joint control model of the manipulator is established, and the manipulator joint controller is simplified as a PD controller, and the manipulator joint driving torque and control rate are as follows:

[0089]

[0090] Wherein, i=1, 2, …, 7. kpi and k di are control parameters of each joint controller, respectively, q di and q i are the desired angle and the actual angle of each joint, respectively;

[0091] In step S103, the space station flight orbit is mainly affected by the gravity gradient torque and the aerodynamic torque generated by the space rare gas. The gravity gradient torque can be calculated according to a theoretical model formula. In order to calculate the aerodynamic torque of the space station cabin and the sail plate in real time, the space station cabin is simplified to a “cylinder segment + sail plate plane” model for calculation, and on this basis, an environmental disturbance model is established.

[0092] In an embodiment of the present application, preferably, in step S103, the environmental torque includes the gravity gradient torque and the aerodynamic torque, and the calculation of the gravity gradient torque specifically includes:

[0093] The rotation matrix corresponding to the attitude angle is where ψ represents the yaw attitude angle, the roll attitude angle, and θ represents the pitch attitude angle:

[0094]

[0095] The gravity gradient torque M g is calculated by using a theoretical model formula.

[0096] The formula is as follows, where ω0 is the orbit angular velocity, is the inertia matrix of the spacecraft,

[0097]

[0098] As shown in FIG. Figure 4 In an embodiment of the present application, preferably, in step S103, the calculation of the aerodynamic torque specifically includes:

[0099] The space station cabin is simplified to a combined model of “cylinder segment and sail plate plane”, and the aerodynamic torque of the space station cabin and the sail plate is calculated in real time,

[0100] where the cabin diameters of the core cabin, the forward cargo spaceship, the rear cargo spaceship, and the radial manned spaceship are D0, D2, D3, and D 13 , respectively; A ha , B ha , C ha , D ha , A hb , B hb , C hb , and D hb are the left and right sail plate corner point coordinates of the core cabin.3a , B 3a , C 3a , D 3a , A 3b , B 3b , C 3b , D 3b are left and right sail panel corner point coordinates of the aft cargo spaceship, in A 2a , B 2a , C 2a , D 2a , A 2b , B 2b , C 2b , D 2b are left and right sail panel corner point coordinates of the forward cargo spaceship, in A 13a , B 13a , C 13a , D 13a , A 13b , B 13b , C 13b , D 13b are left and right sail panel corner point coordinates of the radial manned spaceship, the attitude angle and rotation matrix of the space station assembly are defined in accordance with the foregoing gravity gradient torque calculation.

[0101] In an embodiment of the present application, preferably, in step S103, the aerodynamic torque of the space station cabin and sail panel is calculated, specifically comprising:

[0102] Step S1031, the aerodynamic drag F of the assembly cabin and sail panel in the orbit system is calculated, the formula is:

[0103] F = [F ox F oy F oz ] T , F oy = 0, F oz = 0,

[0104] Wherein, the atmospheric drag coefficient C d Take 2.5, the atmospheric density p takes 4.47E-12 kgm -3 , the wind speed component v x of the orbit system x axis is -7.310 km / s, the wind speed in the y axis and z axis direction is ignored, and S is the projection area and of the cabin and sail panel.

[0105] As Figure 4 shown, the projection area of the cabin column section structure along the -x direction of the orbit coordinate system is calculated as follows:

[0106]

[0107] When the sailboards of each core cabin, forward and rear cargo spacecraft and manned spacecraft are all in the horizontal zero position, the three corner points A ha , B ha , C ha of the left sailboard of the core cabin can be expressed in the OYZ plane of the orbit coordinate system as A oha , B oha , C oha , and the coordinates can be expressed as:

[0108]

[0109]

[0110]

[0111] The projected area (upstream area) of the left sailboard of the core cabin in the OYZ plane of the orbit coordinate system is:

[0112] S ha = A oha (2)B oha (3)+B oha (2)C oha (3)+C oha (2)A oha (3)-A oha (2)C oha (3)-C oha (2)B oha (3)-B oha (2)A oha (3)

[0113] The projected area (upstream area) of the right sailboard of the core cabin in the OYZ plane of the orbit coordinate system is the same as that of the sailboard in the II quadrant:

[0114] S hb =S ha

[0115] Since the sailboards of the forward and rear cargo spacecraft are coplanar with the sailboards of the core cabin, only the size is different, so the upstream area of each sailboard of the cargo spacecraft is:

[0116] S 2a =S 2b =S 3a =S 3b =kS ha , where k is the area proportionality coefficient, k = 4.5

[0117] The three corner points A 13a , B 13a , C 13a of the right sailboard of the manned spacecraft can be expressed in the OYZ plane of the orbit coordinate system as A o13a , Bo13a , C o13a , whose coordinates can be expressed as:

[0118]

[0119]

[0120]

[0121] The projected area (upstream area) of the right sailboard of the manned spacecraft in the OYZ plane of the orbit coordinate system is:

[0122] S 13a = A o13a (2)B o13a (3)+B o13a (2)C o13a (3)+C o13a (2)A o13a (3)-A o13a (2)C o13a (3)-C o13a (2)B o13a (3)-B o13a (2)A o13a (3)

[0123] The projected area (upstream area) of the left sailboard of the manned spacecraft in the OYZ plane of the orbit coordinate system is the same as that of the right sailboard:

[0124] S 13b =S 13a

[0125] According to the formula , the aerodynamic drag F of the combination body cabin column segment structure and the eight sailboards can be calculated.

[0126] Step S1032, the aerodynamic drag F of the space station in the body coordinate system is calculated by projection transformation of the coordinate system through the flight attitude angle d , and the formula is:

[0127]

[0128] Step S1033, the aerodynamic moment M of the space station in the body coordinate system is calculated a , and the formula is:

[0129] M a =r ca ×F d ,

[0130] wherein M ax , M ay , M azr is the component of aerodynamic moment in the body control coordinate system, r ca is the vector of center of mass pressure center moment in the body control coordinate system, and the formula is:

[0131]

[0132] r ca is the component of aerodynamic moment in the body control coordinate system, r ca is the corresponding antisymmetric matrix, and the formula is:

[0133]

[0134] As Figure 5 and Figure 6 shown, in an embodiment of the present application, preferably, in order to have a better human-computer interaction interface, preferably, ADAMS is used to carry out full-system dynamic simulation modeling of the cabin segment assembly process, and in step S2, specifically includes:

[0135] Step S201, a three-dimensional model of the cabin and the mechanical arm is established and imported into the ADAMS software, a dynamic model is set according to the flight mission simulation, a Simulink module of the dynamic model is obtained, and the flight mission at least includes parameters of coordinate system and mass characteristics, initial flight attitude angle and initial angular velocity of the combined body, joint motion of the mechanical arm, space environment and cargo ship engine thrust.

[0136] The coordinate system and mass characteristics at least include: establishing an inertial coordinate system, an orbital coordinate system, a geometric coordinate system of the cabin, a control coordinate system and mass characteristics thereof in ADAMS, a joint coordinate system of the mechanical arm, a center of mass coordinate system and mass characteristics thereof.

[0137] The initial flight attitude angle and initial angular velocity of the combined body at least include: setting the initial attitude angular velocity of the orbital system relative to the inertial space according to the flight mission, setting the initial attitude and initial angular velocity of the combined body according to the simulation condition, and measuring the output as the attitude and angular velocity of the combined body control coordinate system relative to the orbital coordinate system for output to the Matlab joint simulation.

[0138] The joint motion of the mechanical arm at least includes: the mechanical arm is driven to move or kept in the initial state of locking by inputting the planned angle of the mechanical arm in the form of a spline curve into the model. The joint controls the PD controller of the mechanical arm joint model in step S102 above in ADAMS in a closed loop.

[0139] The space environment at least includes: setting the gravity acceleration g to be zero; setting a moment input interface at the center of mass position of the core cabin / cargo ship for receiving the Matlab calculation results for joint simulation.

[0140] The cargo ship engine thrust at least includes: setting the start time, start time length and thrust size according to the flight mission.

[0141] As shown in Figure 6 Step S202, a core cabin / cargo ship gravity gradient torque calculation module and a core cabin / cargo ship aerodynamic torque calculation module are established in Simulink.

[0142] Step S203, as shown in Figure 7 A repositioning attitude test simulation model is generated in a state that the mechanical arm keeps the initial state without movement.

[0143] Step S204, as shown in Figure 8 A cabin section assembly simulation model is generated in a state that the mechanical arm drags the cargo ship to move.

[0144] Figure 7 In the simulation, the initial flight attitude angle and initial angular velocity of the combination body, the mass characteristics of the combination body, the mass characteristics of the core cabin and the cargo ship, the dynamics parameters and movement trajectory of the mechanical arm, the engine starting parameters, etc. are taken as the simulation conditions, and the gravity gradient torque and the aerodynamic torque are taken as the inputs of the cabin section assembly simulation model, and the flight attitude angle and the angular velocity of the combination body are outputted over time.

[0145] Figure 8 In the simulation, the initial flight attitude angle and initial angular velocity of the combination body, the mass characteristics of the core cabin and the cargo ship, the dynamics parameters and movement trajectory of the mechanical arm, the engine starting parameters, etc. are taken as the simulation conditions, and the gravity gradient torque and the aerodynamic torque of the core cabin and the gravity gradient torque and the aerodynamic torque of the cargo ship are taken as the inputs of the repositioning attitude test simulation model, and the flight attitude angle of the core cabin and the flight attitude angle of the cargo ship are outputted.

[0146] In an embodiment of the present application, preferably, in step S3, the cabin section assembly process dynamics pre-simulation in two states that the mechanical arm drags the cargo ship to move and the mechanical arm keeps the initial state without movement is carried out respectively according to the initial dynamics parameter conditions in step S1 and the simulation model established in step S2, so as to establish the initial conditions and the basis for model correction, as shown in Figure 2 Link M1 and link M2.

[0147] Under the simulation initial conditions that the initial attitude and the initial attitude angular velocity of the combination body are the same, it can be found through comparative analysis that the mechanical arm movement and the mechanical arm without movement mainly reflect an upper convex or lower concave of angular momentum exchange in the combination body attitude curve, and the overall movement trend is consistent, as shown in Figure 9

[0148] In addition, in this embodiment, the cargo ship trajectory control engine action is added after 2000s in the mechanical arm movement simulation, therefore, there is a large difference between the two groups of curves after 2000s.

[0149] In an embodiment of the present application, preferably, in step S4, it specifically includes:

[0150] ​Step S401, before the real cabin segment assembly flight task, according to the cabin segment assembly scheme design state, the nominal stop control initial flight attitude angle design state is arranged, the whole system is organized and arranged to carry out the uncontrolled flight attitude test of the combined body position, such as Figure 2 Middle link M3;

[0151] Wherein, the on-orbit attitude test is basically the same as the cabin segment assembly task flight state, the main difference is that the mechanical arm keeps power off and does not move.

[0152] Step S402, collect the initial attitude angle and initial attitude angular velocity of the combined body at the stop control moment, and the attitude angle telemetry sequence of the combined body during uncontrolled period.

[0153] In an embodiment of the present application, preferably, in step S5, specifically includes:

[0154] Step S501, confirm the matching consistency and correctness of the scheme design and simulation results, and compare and analyze the deviation between the flight attitude test telemetry data of the combined body during the uncontrolled attitude test in step S401 and the simulation results obtained in step S301, such as Figure 2 Middle link M4;

[0155] Step S502, adjust the parameters of the simulation results with attitude prediction deviation greater than the preset deviation;

[0156] Wherein, the main adjustment parameters are the mass and inertia matrix of the mechanical arm base end spacecraft (core cabin) and the mechanical arm load end spacecraft. The mass and inertia matrix can be obtained by calculating the initial information of the structure dry weight, the remaining amount of propellant, the on-orbit material management software of astronauts, and the calculation results of the three-dimensional model layout. The residual deviation of the mass characteristics (mainly caused by astronaut video, water, urine / feces, waste, etc.) can be adjusted by modifying the inertia parameters in the model.

[0157] Step S503, taking the telemetry parameters of the attitude test as the reference for simulation comparison, redeveloping the simulation analysis of the attitude test; if the deviation between the simulation results and the telemetry parameters of the attitude test is less than the preset threshold value accepted by the measurement and control tracking ability, the model correction is ended; otherwise, return to step S502 to re-adjust the parameters;

[0158] Step S504, repeat step S502 and step S503 until the deviation between the simulation results and the telemetry parameters of the attitude test is less than the preset threshold value accepted by the measurement and control tracking ability.

[0159] According to an aspect of the present application, there is provided an electronic device comprising one or more processors, one or more memories, and one or more computer programs; wherein the processor is connected with the memory, and the one or more computer programs are stored in the memory, and when the electronic device is running, the processor executes the one or more computer programs stored in the memory, so that the electronic device performs a mechanical arm cabin section assembly process posture prediction and on-orbit test method according to any one of the above technical solutions.

[0160] According to an aspect of the present application, there is provided a computer readable storage medium for storing computer instructions, which, when executed by a processor, implement a mechanical arm cabin section assembly process posture prediction and on-orbit test method according to any one of the above technical solutions.

[0161] To sum up, the present application provides a mechanical arm cabin section assembly process posture prediction and on-orbit test method and an electronic device, comprising the following steps: step S1, establishing a floating base mechanical arm dynamics model; step S2, establishing a cabin section assembly process simulation model; step S3, carrying out mechanical arm cabin section assembly process dynamics pre-simulation; step S4, carrying out uncontrolled flight attitude test of the flight attitude of the combination; step S5, using the simulation results obtained in step S3 and the telemetry data obtained in step S4 to correct the simulation model and the dynamics parameters; and step S6, using the corrected simulation model and the dynamics parameters to carry out simulation prediction again according to the cabin section assembly task state. The present application adopts the simulation-on-orbit test-model correction-re-simulation mode, and uses the floating base mechanical arm dynamics basic equation as the basis, and corrects the dynamics parameters in combination with the on-orbit test data, overcomes the difficulty that the dynamics of the transfer process cannot be verified on the ground, improves the accuracy of the uncontrolled period attitude simulation prediction, improves the task measurement and control coverage, effectively ensures the reliable implementation of the task, and accumulates experience and technical foundation for the construction and application development of the space station.

[0162] In addition, it should be noted that the present application can be provided as a method, device or computer program product. Therefore, the embodiments of the present application can adopt a completely hardware embodiment, a completely software embodiment or an embodiment combining software and hardware aspects. Moreover, the embodiments of the present application can adopt the form of a computer program product implemented on one or more computer usable storage media containing computer usable program codes.

[0163] It is also to be noted that, as used in this specification and the appended claims, the singular forms "a," "an," and "the" include plural referents unless the context clearly dictates otherwise. Thus, for example, reference to "a component" or "the component" can include a plurality of such components unless the context clearly dictates otherwise. Similarly, the word "or" as used in a phrase such as "A or B" does not exclude the presence of both A and B, unless the context clearly dictates otherwise.

[0164] Finally, it is to be understood that the above description is intended to be illustrative, and not restrictive. Many other embodiments will be apparent to those of skill in the art upon reviewing the above description. The scope of the application should, therefore, be determined not with reference to the above description, but instead with reference to the appended claims, along with their full scope of equivalents.

Claims

1. A method for attitude prediction and on-orbit testing during the assembly process of a robotic arm module, comprising the following steps: Step S1: Establish the dynamic model of the floating base robotic arm; Step S2: Establish a simulation model of the module assembly process; Step S3: Conduct dynamic pre-simulation of the robotic arm module assembly process; Step S4: Conduct uncontrolled flight attitude tests of the combined vehicle's shifting flight attitude; Step S5: Use the simulation results obtained in step S3 and the telemetry data obtained in step S4 to correct the simulation model and dynamic parameters; Step S6: Using the corrected simulation model and dynamic parameters, conduct simulation prediction again according to the module assembly mission status; In step S3, pre-simulations for two states are conducted according to the initial dynamic parameter conditions in step S1 and the simulation model established in step S2: Step S301: Pre-simulation of the system dynamics of the uncontrolled assembly posture while the robotic arm remains in its initial state and does not move; Step S302: Pre-simulation of the system dynamics of the assembly of the cargo ship without attitude control during the assembly process of the robotic arm dragging the moving compartment. Step S4 specifically includes: Step S401: Based on the design state of the module assembly scheme and the design state of the nominal initial flight attitude angle of the stop control, organize and arrange the uncontrolled flight attitude test of the combined body's shift flight attitude for the entire system. Step S402: Collect the initial attitude angle and initial attitude angular velocity of the combined body at the moment of stopping control, and the attitude angle telemetry sequence of the combined body during the uncontrolled period; Step S5 specifically includes: Step S501: Confirm the consistency and correctness of the matching between the scheme design and the simulation results, and compare and analyze the deviation between the flight attitude test telemetry data obtained in step S401 during the uncontrolled attitude test of the combined body and the simulation results obtained in step S301. Step S502: Adjust the parameters of the simulation results where the attitude prediction deviation is greater than the preset deviation; Step S503: Using the telemetry parameters of the attitude test as the benchmark for simulation comparison, conduct the simulation analysis of the attitude test again. Step S504: Repeat steps S502 and S503 until the deviation between the simulation result and the attitude test telemetry parameters is less than the preset threshold acceptable for the measurement and control tracking capability.

2. The method according to claim 1, characterized in that, Step S1 specifically includes: Step S101: Based on the floating-base manipulator system and its fundamental dynamic equations, and considering the effect of environmental torque on the floating-base manipulator system, establish a system dynamic mathematical model for the floating-base manipulator. The specific formulas are as follows: Among them, H m H is the inertia matrix of the robotic arm's own system. b For the inertial tensor of the core module, H bm Let Θ be the coupling inertia between the core module and the robotic arm, Θ be the joint angle vector, x0 be the core module pose vector, and c be the joint angle vector. b c m These are the nonlinear forces related to the core module's motion and the robotic arm's motion, including centripetal force and Coriolis force; F b τ represents the control force and control torque acting on the base. m F is the driving torque of the robotic arm joint. e J is the external force and torque that the end effector of the robotic arm comes into contact with the environment. b For the Jacobian matrix of the core module serving as the floating base of the robotic arm, J m Let be the Jacobian matrix of the robotic arm; Step S102: Establish the joint control model of the robotic arm, simplifying the robotic arm joint controller into a PD controller. The driving torque and control law of each joint of the robotic arm are as follows: Where i = 1, 2, ..., 7, k pi and k di These are the control parameters for each joint controller, q di and q i These are the expected angles and actual angles of each joint, respectively. Step S103: Consider the effect of environmental torque on the floating base robotic arm system and establish an environmental disturbance model.

3. The method according to claim 2, characterized in that, In step S103, the environmental torque includes the gravity gradient torque and the aerodynamic torque. Calculating the gravity gradient torque specifically includes: The rotation matrix corresponding to the attitude angle is as follows: Where ψ represents the yaw attitude angle, θ is the roll attitude angle, and θ is the pitch attitude angle. Calculate the gravitational gradient torque M using theoretical model formulas g The formula is as follows, where ω0 is the angular velocity of the orbit. The inertia matrix of the aircraft.

4. The method according to claim 3, characterized in that, In step S103, calculating the aerodynamic torque specifically includes: The space station module is simplified into a combined model of "cylinder segments and solar panels", and the aerodynamic moments of the space station module and solar panels are calculated in real time. The core module, forward cargo spacecraft, backward cargo spacecraft, and radial manned spacecraft have diameters of D0, D2, D3, and D, respectively. 13 ; with A ha B ha C ha D ha A hb B hb C hb D hb Let A be the coordinates of the left and right corner points of the core module's solar panels. 3a B 3a C 3a D 3a A 3b B 3b C 3b D 3b Let A be the coordinates of the left and right corner points of the rearward cargo spacecraft's solar panels. 2a B 2a C 2a D 2a A 2b B 2b C 2b D 2b Let A be the coordinates of the left and right corner points of the forward cargo spacecraft's solar panels. 13a B 13a C 13a D 13a A 13b B 13b C 13b D 13b These are the coordinates of the left and right corner points of the radial manned spacecraft's sails.

5. The method according to claim 4, characterized in that, In step S103, the aerodynamic moments of the space station module and solar panels are calculated, specifically including: Step S1031: Calculate the aerodynamic drag F of the combined cabin and solar panels in the orbital system. The formula is: F=[F ox F oy F oz ] T , F oy =0,F oz =0, Among them, the atmospheric drag coefficient C d Take 2.5, and take the atmospheric density ρ as 4.47E-12kgm. -3 For spacecraft orbiting the space station, the wind speed component v along the x-axis of the orbital system. x The wind speed is -7.310 km / s, and the wind speeds in the y and z axes are negligible. S is the sum of the projected areas of the cabin and the sail. Step S1032: Perform coordinate system projection transformation using flight attitude angles to calculate the aerodynamic drag F under the space station's intrinsic system. d The formula is: Step S1033: Calculate the aerodynamic moment M under the space station's own system. a The formula is: M a =r ca ×F d , Among them, M ax M ay M az This represents the component of the aerodynamic torque in the space station's main control coordinate system. r ca Let the center of mass compression moment vector be the vector in the body control coordinate system, and the formula is: r ca × For r ca The corresponding antisymmetric matrix is ​​represented by the formula:

6. The method according to claim 1, characterized in that, Step S2 specifically includes: Step S201: Create a 3D model of the cabin and robotic arm, import it into ADAMS software, set up the dynamic model according to the flight mission simulation, and obtain the Simulink module of the dynamic model. The flight mission includes at least the parameters of the coordinate system and mass characteristics, the initial flight attitude angle and initial angular velocity of the combined body, the joint motion of the robotic arm, the space environment, and the thrust of the cargo ship's engine. Step S202: Create a core compartment / cargo ship gravity gradient moment calculation module and a core compartment / cargo ship aerodynamic moment calculation module in Simulink; Step S203: Generate a simulation model for the rotation posture test while keeping the robotic arm in its initial, non-moving state; Step S204: Generate a simulation model of the compartment assembly in the state of the cargo ship being dragged by a robotic arm.

7. An electronic device, characterized in that, include: One or more processors, one or more memories, and one or more computer programs; wherein the processor is connected to the memory, and the one or more computer programs are stored in the memory, and when the electronic device is running, the processor executes the one or more computer programs stored in the memory to cause the electronic device to perform a method for attitude prediction and on-orbit testing of a robotic arm module assembly process as described in any one of claims 1-6.

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