CN104063537A - Multi-body dynamics parameter determination system based on distributive time trigger and method thereof - Google Patents

Abstract

The invention discloses a multi-body dynamics parameter determination system based on the distributive time trigger and a method thereof. The method comprises the following steps: firstly, a system architecture model based on a time-triggered data bus is established, wherein each aircraft system simulation node is connected onto the bus and comprises multiple simulation modules; secondly, a system dynamics parameter is subjected to calculation through the guiding of the dynamics simulation modules; thirdly, a single aircraft system inner closed-loop simulation process is implemented through a system inner data receiving and dispatching module; fourthly, data exchange is implemented through an outer receiving and dispatching module and other aircraft system simulation nodes; fifthly, the dynamic parameters in a next time is obtained through calculation via a differential equation or a linearization formula, and the synchronization of a distributed system multi-body dynamics parameter is achieved.

Description

Many-body dynamics parameter based on the distributed time triggers is determined system and method thereof

Technical field

The many-body dynamics parameter the present invention relates to based on the distributed time triggers is determined system and method thereof, belongs to spacecraft ground l-G simulation test field.

Background technology

According to the general plan of the manned astro-engineering, manned space station forms by assembling in-orbit to build, space station is emitted to respectively space by a plurality of independent tank sections, the final space station that forms of intersection by each other docking again, during also have part manned spaceship and cargo ship to dock and stop in space station.

In the l-G simulation test and testing authentication process of GNC (GNC is guidance navigation and control) subsystem, need to face the situation that a plurality of aircraft are worked in space simultaneously, therefore need to solve the technical matters of multi-aircraft associative simulation and checking.In the related work of manned space flight the second stage of the project, in order to complete the task of two aircraft intersection docking verification experimental verifications, when two aircraft carry out joint test, use an equipment as main control computer, this main control computer is responsible for calculating the dynamics of object machine and tracker simultaneously, and result of calculation is converted into excited data sends to platform sensor simulator and intersection docking sensor simulator.In second phase work, testing apparatus by tracker is served as main control computer, and the testing apparatus of object machine is as slave, the dynamics data calculating in target device is not introduced closed loop verification system, therefore whole process of the test is not real Combined Trials, but by tracker relevant device and the leading single cabin section verification experimental verification of system.

In the planning of space station Combined Trials, may have the task of 5 aircraft associative simulations (3 cabin section+cargo ship+manned spaceships) at the most, if re-use the second stage of main control computer pattern, can face 2 insurmountable problems:

A) computational resource is not enough: the kinetic parameter that single CPU cannot complete a plurality of aircraft within the emulation cycle calculates;

B) system-level tasks synchronization difficulty, is used traditional architectures cannot accomplish large system data high-speed synchronous and share.

Therefore need to design a kind of distributed Dynamic Co-Simulation pilot system, for many cabins, space station section Combined Trials, verify.

Summary of the invention

Technology of the present invention is dealt with problems and is: overcome the deficiencies in the prior art, the many-body dynamics parameter that proposition triggered based on the distributed time is determined system and method thereof, by the system architecture triggering based on the time, coordinate through linearizing system dynamics parameter determination method, the many-body dynamics that has solved large-scale distributed system distributes and calculates and parameter problem identificatioin; The combined calculation method of the many simulation nodes by distributed system, has improved system-level arithmetic capability, has solved the problem of single CPU computational resource deficiency; By the system level data function of exchange of time trigger data bus, solved system-level tasks synchronization problem.

Technical solution of the present invention is: the many-body dynamics parameter based on the distributed time triggers is determined system, comprise a plurality of aerocraft system simulation nodes and time-trigged system bus, a plurality of aerial vehicle simulation nodes interconnect by time-trigged system bus, and each aerial vehicle simulation node comprises aerocraft system control realistic model, internal system data transmit-receive module, dynamics simulation module, data preprocessing module, system external data transceiver module; Aerocraft system is controlled realistic model and is comprised that control bus in GNC controller, GNC system, topworks, topworks monitor module, GNC sensor, sensor excitation simulator;

Data preprocessing module comprises that by the dynamics simulation module original state of setting system operation initial time, system preliminary orbit position, system initial attitude deliver to dynamics simulation module;

The initial driving force mathematic(al) parameter of the system that dynamics simulation module is set according to data preprocessing module comprises that operation initial bat time, system preliminary orbit position, system initial attitude carry out track and attitude and calculate that the system operation time, system track position and the system attitude that obtain current bat deliver to respectively internal system data transmit-receive module and system external data transceiver module;

Internal system data transmit-receive module comprises that by the kinetic parameter of current shooting system system operation time, system track position and system attitude change into sensor measurement excited data and deliver to sensor excitation simulator according to sensor model; Sensor excitation simulator is measured excited data according to sensor GNC sensor is carried out to optical excitation or microwave excitation or these physical stimulus of thermal excitation, make GNC sensor form sensor measuring-signal, GNC sensor is delivered to GNC controller by the sensor measuring-signal of formation by control bus in GNC system;

GNC controller forms guidance control rate according to sensor measuring-signal, according to the instruction of guidance control rate formation control, steering order is delivered to topworks by control bus in GNC system, topworks changes the state of flywheel or engine jet pipe according to the steering order of receiving, steering order is delivered to topworks simultaneously and monitored module, topworks monitors the steering order that module monitors topworks sends here, the steering order of simultaneously topworks being sent here sends to internal system data transmit-receive module, internal system data transmit-receive module is converted into system control and system control moment by steering order according to the specified control parameter of topworks, system control and system control moment are sent to dynamics simulation module,

System external data transceiver module comprises that by the kinetic parameter of the current shooting system receiving from dynamics simulation module system operation time, system track position and system attitude send to other aerocraft system simulation node by time-trigged system bus, the kinetic parameter that simultaneity factor external data transceiver module receives the current shooting system of other aerocraft system simulation node comprises system operation time, system track position and system attitude, synchronous to realize in system the kinetic parameter between each aerocraft system simulation node.

Many-body dynamics parameter determination method based on distributed time triggering, it is characterized in that: the feature of having utilized distributed system, service time, trigger data bus was carried out time established data exchange in system, and finally reach that system-level kinetic parameter is determined and synchronous effect, the concrete steps of method are as follows:

(1) data preprocessing module comprises that by the dynamics simulation module original state of setting system operation initial time, system preliminary orbit position, system initial attitude deliver to dynamics simulation module;

(2) the initial driving force mathematic(al) parameter of the system that dynamics simulation module is set according to data preprocessing module comprises that operation initial bat time, system preliminary orbit position, system initial attitude carry out track and attitude and calculate that the system operation time, system track position and the system attitude that obtain current bat deliver to respectively internal system data transmit-receive module and system external data transceiver module;

(3) internal system data transmit-receive module comprises that by the kinetic parameter of current shooting system system operation time, system track position and system attitude change into sensor measurement excited data and deliver to sensor excitation simulator according to sensor model;

(4) sensor excitation simulator carries out optical excitation or microwave excitation or these physical stimulus of thermal excitation according to sensor measurement excited data to GNC sensor, make GNC sensor form sensor measuring-signal, GNC sensor is delivered to GNC controller by the sensor measuring-signal of formation by control bus in GNC system;

(5) GNC controller forms guidance control rate according to sensor measuring-signal, according to the instruction of guidance control rate formation control, steering order is delivered to topworks by control bus in GNC system;

(6) topworks changes the state of flywheel or engine jet pipe according to the steering order of receiving, steering order is delivered to topworks simultaneously and is monitored module;

(7) topworks monitors the steering order that module monitors topworks sends here, and the steering order of simultaneously topworks being sent here sends to internal system data transmit-receive module;

(8) internal system data transmit-receive module is converted into system control F by steering order according to the specified control parameter of topworks _cwith system control moment T _c, system control and system control moment are sent to dynamics simulation module;

(9) system external data transceiver module comprises that by the kinetic parameter of the current shooting system receiving from dynamics simulation module system operation time, system track position and system attitude send to time-trigged system bus, and extracts the system dynamic mathematic(al) parameter of other aerocraft system simulation node from time-trigged system bus;

(10) dynamics simulation module is according to the terrestrial gravitation F of the system operation time of current bat in step (2), system track position and the current bat of system Attitude Calculation _g, terrestrial gravitation gradient moment T _gwith unexpected perturbed force F _wwith unexpected disturbance torque T _w;

(11) dynamics simulation module is according to the system control F of current bat in step (8) _cand the terrestrial gravitation F of current bat in step (10) _gwith unexpected perturbed force F _w, then according to the dynamics of orbits differential equation of classical Newton's law, calculate the second derivative of current bat dynamics of orbits parameter according to the second derivative of current bat dynamics of orbits parameter extrapolate the kinetic parameter r of the system of next bat, the dynamics of orbits differential equation of described classical Newton's law is:

$\stackrel{..}{r}=\frac{F}{m}=\frac{F_g+F_c+F_w}{m}---\left(1\right)$

In formula, represent acceleration under inertial system (ax, ay, az) ', (r represents the position (x, y, z) of aircraft under Earth central inertial system ', represent the speed (dx, dy, dz) of aircraft under Earth central inertial system '), F represents all external force (comprising inertial force) that aircraft is subject to, wherein F _grepresent terrestrial gravitation, F _crepresent actuating mechanism controls power, F _wrepresent unexpected perturbed force, m represents this weight of aircraft;

(12) dynamics simulation module is according to the system control moment T in step (8) _c, and the terrestrial gravitation gradient moment T in step (10) _gwith unexpected disturbance torque T _wobtain bonding force square projection T under body series that aircraft is received, then according to the attitude dynamics differential equation of classical Newton's law, calculate the second derivative of current bat attitude dynamics parameter and extrapolate the attitude dynamics parameter ω of next bat, the described attitude dynamics differential equation is:

$I\stackrel{.}{\mathrm{\ω}}+\widetilde{\mathrm{\ω}}*H=T=T_C+T_g+T_w---\left(2\right)$

In formula, ω is the projection of angular velocity under body series under aircraft inertial system, be the skew matrix of ω, I is the projection under body series of system moment of inertia, and H is the projection under body series of aircraft moment of inertia, and T is bonding force square projection under body series that aircraft is received, T _cactuating mechanism controls moment, T _gterrestrial gravitation gradient moment, T _wnon-expected interference moment;

(13), according to the attitude dynamics differential equation in other aerocraft system kinetic parameter in step (9), the dynamics of orbits differential equation in step 11 and step (12), the dynamics of orbits differential equation and the attitude dynamics differential equation are carried out to linearization and obtain formula (3):

$X_{t1}=e^{A{t}_1}+{\∫}_{t_0}^{t_1}{e}^{A(t_1-\mathrm{\τ})}B(F_w+F_C)\mathrm{d\τ}---\left(3\right)$

Wherein X represents dynamics of orbits parameter or attitude dynamics parameter, t ₁represent the target simulator time (being generally a control cycle), A representative is through linearizing system state matrix, and B represents that control is to the Transfer Parameters of system state, F _crepresent actuating mechanism controls power, F _wrepresent unexpected perturbed force;

(14) according to the formula (3) in step (13), calculate other aerocraft system kinetic parameter, complete the determining of kinetic parameter of all aircraft nodes in this periodic system.

The present invention's advantage is compared with prior art:

(1) the invention solves the exchange of many-body dynamics parameter and the definite problem of parameter of distributed system, provided a kind of system scheme that triggers distributed computing technology based on the time;

(2) traditional dynamics calculation system architecture based on centralized all, all calculating all concentrates on a main control computer and completes, its advantage is to be very easy to realize, shortcoming is to solve day by day complicated large system high-precision environment dynamics calculation amount problem, and cannot utilize the calculating advantage of distributed system, therefore must find a technological approaches addressing this problem;

(3) core concept of the present invention is that single aircraft system node is responsible for the complete single vehicle dynamics of emulation, other dynamics adopts the formula of reduction through deriving to carry out recursion, and periodically carry out system-level parameters correction, guarantee that terminal does not produce cumulative errors to the recursive parameter of other aircraft;

(4) the present invention has utilized the feature of system-level Distributed Calculation, has improved system-level simulation calculation ability, has also guaranteed that the dynamics precision index of all terminals in system meets the demands.

Accompanying drawing explanation

Fig. 1 is the process flow diagram of system works method of the present invention;

Fig. 2 is system architecture schematic diagram of the present invention;

Fig. 3 is single aircraft system emulation node working timing figure of the present invention.

Embodiment

Basic ideas of the present invention are: model be take time trigger data bus as basic distributed systems architecture, each aerocraft system simulation node in system has been responsible for the closed-loop simulation process of this aircraft, simultaneously with other aerocraft system simulation node exchange kinetics parameter, by linearizing formula, carry out kinetic parameter reckoning again, the definite and synchronizing process of completion system kinetic parameter.

Below in conjunction with accompanying drawing, the present invention is described in detail, and distributed system forms as shown in Figure 2, and system is comprised of 5 aerocraft system simulation nodes, and each node represents an aerospacecraft system, possesses all features of single spacecraft; Each node can independent operating, under independent operating state, is equivalent to the closed-loop simulation process of single aircraft.Under single spacecraft state, system is mainly comprised of aerocraft system control realistic model, internal system data transmit-receive module, dynamics simulation module, data preprocessing module, system external data transceiver module.

In distributed system, all single aircraft system emulation nodes are all connected in a time trigger data bus, and the system external data transceiver module in single aerocraft system is controlled.Time-trigged system bus (Time Triggered Data Bus) is different from traditional as 1553B bus and CAN bus, in this bus, there is no clear and definite master slave relation, all node priorities are identical, and according to the flow process timesharing of appointing in advance, alternately send data in bus, the representative of this class data bus is FlexRay bus, TTP bus and TTEthernet bus.Service time, the advantage of trigger data bus was that the operation of system does not rely on any one independently node, and the reconfigurability of system is stronger, can articulate 2～10 system nodes that do not wait.

It is the GNC system model of spacecraft that aerocraft system is controlled realistic model, and this model can be GNC system actual product, can be also the simulator of simulation GNC systemic-function.This model is comprised of 5 submodules: in GNC controller, GNC system, control bus, topworks, topworks monitor module, GNC sensor, sensor excitation simulator.Wherein GNC controller is the nucleus module of this model, is responsible for gathering from GNC intra-system bus the metrical information of each sensor, and by the steering order of navigation, conductance calculating aircraft processed, then send to topworks by GNC intra-system bus, GNC sensor is responsible for the external information of GNC system and is measured, then metrical information is sent to GNC controller, sensor excitation simulator is on ground, to coordinate the uphole equipment of sensor module work, first simulator obtains kinetic parameter from dynamics simulation module, again parameter is converted into the required pumping signal of different sensors, such as first star sensor simulator receives the signal that dynamics is converted to sensor measurement data, again this signal is generated to the starry sky image that star sensor should collect under this state according to actual space constellation configuration, again this starry sky image is encouraged to star sensor, make the true outputting power of star sensor learn the measurement data signals after transforming, reach the effect of ground simulation simulation, topworks is the execution module of GNC system, and the steering order that reception GNC controller sends the action that is converted into topworks, as engine jet pipe start or flywheel acceleration rotation etc., the instruction that topworks's supervision module-specific is received in supervision topworks and current running status, then be converted to actual dynamics Controlling power F _cwith control moment T _cfeed back to dynamics, form system closed-loop simulation.

Dynamics simulation module be mainly responsible for to be calculated this aerocraft system kinetic parameter, and calculates other vehicle dynamics parameter, and kinetic parameter is determined and is divided into two main contents: dynamics of orbits parameter and attitude dynamics parameter.

(1) dynamics of orbits computing formula

The emulation of environmental dynamics can be divided into two major parts, the one, dynamics of orbits emulation, the 2nd, attitude dynamics emulation.The basis of dynamics of orbits is disome lemma law, has:

F＝grad(U)

Wherein F is the gravitation that aircraft is subject to, and considers the perturbation that aircraft is subject to, and has

$\stackrel{..}{r}=-\frac{\mathrm{Gm}}{r^3}r+\mathrm{grad}\left(R\right)$

Wherein R is perturbative force bit function, has

grad(R)＝F _ru _r+F _tu _t+F _nu _n

In spacecraft orbit dynamics, terrestrial gravitation is maximal term, the control of aircraft is 3 rank a small amount of with respect to terrestrial gravitation, and perturbative force is 6 rank a small amount of with respect to terrestrial gravitation, therefore in traditional dynamics of orbits differential equation, at short notice, can be linear (time is short) permanent (external force impact is little) differential equation by equation simplification, so this differential equation should have general solution:

$X_{t1}=e^{A{t}_1}+{\∫}_{t_0}^{t_1}{e}^{A(t_1-\mathrm{\τ})}B(F_w+F_C)\mathrm{d\τ}$

F wherein _wand F _cbe respectively perturbative force and control, terrestrial gravitation is relevant to aircraft ontological property, in matrix A, with constant, embodies.According to every 160ms, exchange a secondary data,

t ₁-t ₀＝0.16

Use the deviation of this reduced equation and true environment dynamics calculation to come from F _wand F _cuncertainty, according to actual conditions, F _cspan be (300N～+ 300N), F _wspan be (5N～+ 5N), aircraft physical characteristics is cylindrical, quality is designated as 10 tons, the deviation uncertainty within 0.16 second time of general solution of differential equation is 7mm.Below 1cm, and there is not cumulative errors in the position uncertainty of the dynamics of orbits recursive algorithm in every 160ms under inertial system, meets the requirement of spacecraft ground l-G simulation test.

(2) attitude dynamics computing formula

Attitude dynamics take attitude quaternion as basic expression-form as:

$\stackrel{.}{q}=\frac{1}{2}\mathrm{\Ω}\left(\mathrm{\ω}\right)q$

Expansion is:

$\stackrel{.}{q}=\left[\begin{array}{c}{\stackrel{.}{q}}_1\\ {\stackrel{.}{q}}_2\\ {\stackrel{.}{q}}_3\\ {\stackrel{.}{q}}_4\end{array}\right]=\frac{1}{2}\left[\begin{array}{cccc}{q}_4& -q_3& q_2& q_1\\ q_3& q_4& {-q}_1& q_2\\ {-q}_2& q_1& q_4& q_3\\ {-q}_1& {-q}_2& {-q}_3& q_4\end{array}\right]\left[\begin{array}{c}{\mathrm{\ω}}_x\\ {\mathrm{\ω}}_y\\ {\mathrm{\ω}}_z\\ 0\end{array}\right]$

Because the angular velocity of spacecraft in actual moving process is very little, || ω || <0.005, therefore obtains once at each exchange cycle re-use this coordinate single order Long Gekutafa to calculate the attitude quaternion in lower each emulation cycle, have:

$q_{i+1}=q_i+{\stackrel{.}{q}}_0\&times;\mathrm{\&Delta;t},i\&Element;(\mathrm{0,1,2}....15)$

The uncertainty that single order Long Gekutafa after use is simplified carries out attitude dynamics calculating mainly comes from T _w(perturbation moment) and T _cthe uncertainty of (control moment), T _cspan be (250Nm～+ 250Nm), T _wspan be (3Nm～+ 3Nm), the aircraft principal axis of inertia is taken as 10 ⁵kg.m.m, to calculate the uncertainty that hypercomplex number result is transformed into attitude angle under inertial system be (0.047 degree～+ 0.047 degree) to the attitude dynamics recursive algorithm in every 160ms.

Determined that thus system architecture, system information flows, single aircraft system emulation nodal function form and system dynamics parameter determination method, will describe the concrete implementation step of the many-body dynamics parameter determination method based on distributed time triggering as shown in Figure 1 below:

(1) data preprocessing module comprises that by the dynamics simulation module original state of setting system operation initial time, system preliminary orbit position, system initial attitude deliver to dynamics simulation module;

(2) the initial driving force mathematic(al) parameter of the system that dynamics simulation module is set according to data preprocessing module comprises that operation initial bat time, system preliminary orbit position, system initial attitude carry out track and attitude and calculate that the system operation time, system track position and the system attitude that obtain current bat deliver to respectively internal system data transmit-receive module and system external data transceiver module;

(3) internal system data transmit-receive module comprises that by the kinetic parameter of current shooting system system operation time, system track position and system attitude change into sensor measurement excited data and deliver to sensor excitation simulator according to sensor model;

(4) sensor excitation simulator carries out optical excitation or microwave excitation or these physical stimulus of thermal excitation according to sensor measurement excited data to GNC sensor, make GNC sensor form sensor measuring-signal, GNC sensor is delivered to GNC controller by the sensor measuring-signal of formation by control bus in GNC system;

(5) GNC controller forms guidance control rate according to sensor measuring-signal, according to the instruction of guidance control rate formation control, as in-orbit over the ground in three axle control situations, the measurement data that GNC controller obtains according to star sensor, be converted into the attitude information of current flight device, and calculate control rate according to PID controller, then according to engine jet pipe, steering order is delivered to topworks by control bus in GNC system the most at last;

(6) topworks changes the state of flywheel or engine jet pipe according to the steering order of receiving, steering order is delivered to topworks simultaneously and is monitored module;

(7) topworks monitors the steering order that module monitors topworks sends here, and the steering order of simultaneously topworks being sent here sends to internal system data transmit-receive module;

(8) internal system data transmit-receive module is converted into system control F by steering order according to the specified control parameter of topworks _cwith system control moment T _c, system control and system control moment are sent to dynamics simulation module;

(9) system external data transceiver module comprises that by the kinetic parameter of the current shooting system receiving from dynamics simulation module system operation time, system track position and system attitude send to time-trigged system bus, and extracts the system dynamic mathematic(al) parameter of other aerocraft system simulation node from time-trigged system bus; 3 the aerocraft system simulation nodes (A, B and C) of take are example, this aerocraft system kinetic parameter that node A learns emulation module by this node medium power sends in time-trigged system bus by system external data transceiver module, from time-trigged system bus, obtain the system dynamic mathematic(al) parameter of Node B and node C transmission simultaneously, like this with node A, just obtained the kinetic parameter of all nodes in the system that comprises A, B and C, in like manner Node B and node C also can acquisition system in the kinetic parameter of all nodes;

(10) dynamics simulation module is according to the terrestrial gravitation F of the system operation time of current bat in step (2), system track position and the current bat of system Attitude Calculation _g, terrestrial gravitation gradient moment T _gwith unexpected perturbed force F _wwith unexpected disturbance torque T _w;

(11) dynamics simulation module is according to the system control F of current bat in step (8) _cand the terrestrial gravitation F of current bat in step (10) _gwith unexpected perturbed force T _g, then according to the dynamics of orbits differential equation of classical Newton's law, calculate the second derivative of current bat dynamics of orbits parameter according to the second derivative of current bat dynamics of orbits parameter extrapolate the kinetic parameter r of the system of next bat, the dynamics of orbits differential equation of described classical Newton's law is:

$\stackrel{..}{r}=\frac{F}{m}=\frac{F_g+F_c+F_w}{m}---\left(1\right)$

In formula, represent acceleration under inertial system (ax, ay, az) ', (r represents the position (x, y, z) of aircraft under Earth central inertial system ', represent the speed (dx, dy, dz) of aircraft under Earth central inertial system '), F represents all external force (comprising inertial force) that aircraft is subject to, wherein F _grepresent terrestrial gravitation, F _crepresent actuating mechanism controls power, F _wrepresent unexpected perturbed force, m represents this weight of aircraft;

(12) dynamics simulation module is according to the system control moment T in step (8) _c, and the terrestrial gravitation gradient moment T in step (10) _gwith unexpected disturbance torque T _wobtain bonding force square projection T under body series that aircraft is received, then according to the attitude dynamics differential equation of classical Newton's law, calculate the second derivative of current bat attitude dynamics parameter , and extrapolate the attitude dynamics parameter ω of next bat, the described attitude dynamics differential equation is:

$I\stackrel{.}{\mathrm{\&omega;}}+\widetilde{\mathrm{\&omega;}}*H=T=T_C+T_g+T_w---\left(2\right)$

In formula, ω is the projection of angular velocity under body series under aircraft inertial system, be the skew matrix of ω, I is the projection under body series of system moment of inertia, and H is the projection under body series of aircraft moment of inertia, and T is bonding force square projection under body series that aircraft is received, T _cactuating mechanism controls moment, T _gterrestrial gravitation gradient moment, T _wnon-expected interference moment;

(13), according to the attitude dynamics differential equation in other aerocraft system kinetic parameter in step (9), the dynamics of orbits differential equation in step 11 and step (12), the dynamics of orbits differential equation and the attitude dynamics differential equation are carried out to linearization and obtain formula (3):

$X_{t1}=e^{A{t}_1}+{\&Integral;}_{t_0}^{t_1}{e}^{A(t_1-\mathrm{\&tau;})}B(F_w+F_C)\mathrm{d\&tau;}---\left(3\right)$

Wherein X represents dynamics of orbits parameter or attitude dynamics parameter, t ₁represent the target simulator time (being generally a control cycle), A representative is through linearizing system state matrix, and B represents that control is to the Transfer Parameters of system state, F _crepresent actuating mechanism controls power, F _wrepresent unexpected perturbed force;

(14) according to the formula (3) in step (13), calculate other aerocraft system kinetic parameter, complete the determining of kinetic parameter of all aircraft nodes in this periodic system;

(15) as shown in Figure 3, described under a kind of typical condition, the work schedule of the single aircraft system emulation node in system, the emulation cycle of system is 10ms, and the control cycle of system is 160ms, and time-trigged system bus data exchange cycle is 160ms; Wherein A1～A8 represents the work schedule of GNC controller, A1 represents GNC controller task time that control bus obtains GNC sensor data in GNC system, A2 represents GNC controller internal data comparison task time, the A3 representative navigation calculation task time, A4 represents the conductance calculation task time processed, A5 represents GNC controller internal data comparison task time, A6 represents that GNC controller is to the time of topworks's sending controling instruction task, A7 represents Telemetering Data Processing task time, and A8 represents the fault-tolerant decision task time.The exemplary operation sequential of D1～D10 representative dynamics simulation module in the cycle at 1 controller simulation, D1 representative and other vehicle dynamics parameter exchange process, D2 represents other vehicle dynamics parameter renewal process, in D3～D7 representative system, different single aircraft system emulation nodes are to the task time that sends this aerocraft system kinetic parameter in time-trigged system bus, D10 represents that dynamics simulation module receives the actuating mechanism controls instruction that topworks monitors that module sends, and adds the task time of system closed-loop simulation; Q1～Q5 represented within 1 emulation cycle, the work schedule of dynamics simulation module, Q1 represents steering order inquiry and upgrades task time, Q2 represents topworks's model emulation task time, Q3 representative system track and attitude dynamics model calculation task time, Q4 represents external tapping Data Update task time, and Q5 represents clock and data synchronous task time.Under this typical condition, system is moved according to work schedule as shown in Figure 3, can reach the system emulation of expection and control effect.

Environmental dynamics data filtering algorithm in the present invention is on the basis of primary track and attitude dynamics, by the analysis to spacecraft dynamics characteristic, nonlinear differential equation is reduced to permanent linear differential equation at short notice, when meeting simulation accuracy requirement, simplified calculating, and realized the method for distributed system dynamics simulation, for future many spacecrafts of large system associative simulation provide an effective technology to realize approach.

The content not being described in detail in instructions of the present invention belongs to those skilled in the art's known technology.

Claims (2)

1. the many-body dynamics parameter triggering based on the distributed time is determined system, it is characterized in that: comprise a plurality of aerocraft system simulation nodes and time-trigged system bus, a plurality of aerial vehicle simulation nodes interconnect by time-trigged system bus, and each aerial vehicle simulation node comprises aerocraft system control realistic model, internal system data transmit-receive module, dynamics simulation module, data preprocessing module, system external data transceiver module; Aerocraft system is controlled realistic model and is comprised that control bus in GNC controller, GNC system, topworks, topworks monitor module, GNC sensor, sensor excitation simulator;

Data preprocessing module comprises that by the dynamics simulation module original state of setting system operation initial time, system preliminary orbit position, system initial attitude deliver to dynamics simulation module;

The initial driving force mathematic(al) parameter of the system that dynamics simulation module is set according to data preprocessing module comprises that operation initial bat time, system preliminary orbit position, system initial attitude carry out track and attitude and calculate that the system operation time, system track position and the system attitude that obtain current bat deliver to respectively internal system data transmit-receive module and system external data transceiver module;

Internal system data transmit-receive module comprises that by the kinetic parameter of current shooting system system operation time, system track position and system attitude change into sensor measurement excited data and deliver to sensor excitation simulator according to sensor model; Sensor excitation simulator is measured excited data according to sensor GNC sensor is carried out to optical excitation or microwave excitation or these physical stimulus of thermal excitation, make GNC sensor form sensor measuring-signal, GNC sensor is delivered to GNC controller by the sensor measuring-signal of formation by control bus in GNC system;

GNC controller forms guidance control rate according to sensor measuring-signal, according to the instruction of guidance control rate formation control, steering order is delivered to topworks by control bus in GNC system, topworks changes the state of flywheel or engine jet pipe according to the steering order of receiving, steering order is delivered to topworks simultaneously and monitored module, topworks monitors the steering order that module monitors topworks sends here, the steering order of simultaneously topworks being sent here sends to internal system data transmit-receive module, internal system data transmit-receive module is converted into system control and system control moment by steering order according to the specified control parameter of topworks, system control and system control moment are sent to dynamics simulation module,

System external data transceiver module comprises that by the kinetic parameter of the current shooting system receiving from dynamics simulation module system operation time, system track position and system attitude send to other aerocraft system simulation node by time-trigged system bus, the kinetic parameter that simultaneity factor external data transceiver module receives the current shooting system of other aerocraft system simulation node comprises system operation time, system track position and system attitude, synchronous to realize in system the kinetic parameter between each aerocraft system simulation node.

2. the many-body dynamics parameter determination method triggering based on the distributed time, is characterized in that: concrete steps are as follows:

(1) data preprocessing module comprises that by the dynamics simulation module original state of setting system operation initial time, system preliminary orbit position, system initial attitude deliver to dynamics simulation module;

(2) the initial driving force mathematic(al) parameter of the system that dynamics simulation module is set according to data preprocessing module comprises that operation initial bat time, system preliminary orbit position, system initial attitude carry out track and attitude and calculate that the system operation time, system track position and the system attitude that obtain current bat deliver to respectively internal system data transmit-receive module and system external data transceiver module;

(3) internal system data transmit-receive module comprises that by the kinetic parameter of current shooting system system operation time, system track position and system attitude change into sensor measurement excited data and deliver to sensor excitation simulator according to sensor model;

(4) sensor excitation simulator carries out optical excitation or microwave excitation or these physical stimulus of thermal excitation according to sensor measurement excited data to GNC sensor, make GNC sensor form sensor measuring-signal, GNC sensor is delivered to GNC controller by the sensor measuring-signal of formation by control bus in GNC system;

(5) GNC controller forms guidance control rate according to sensor measuring-signal, according to the instruction of guidance control rate formation control, steering order is delivered to topworks by control bus in GNC system;

(6) topworks changes the state of flywheel or engine jet pipe according to the steering order of receiving, steering order is delivered to topworks simultaneously and is monitored module;

(7) topworks monitors the steering order that module monitors topworks sends here, and the steering order of simultaneously topworks being sent here sends to internal system data transmit-receive module;

(8) internal system data transmit-receive module is converted into system control F by steering order according to the specified control parameter of topworks _cwith system control moment T _c, system control and system control moment are sent to dynamics simulation module;

(9) system external data transceiver module comprises that by the kinetic parameter of the current shooting system receiving from dynamics simulation module system operation time, system track position and system attitude send to time-trigged system bus, and extracts the system dynamic mathematic(al) parameter of other aerocraft system simulation node from time-trigged system bus;

(10) dynamics simulation module is according to the terrestrial gravitation F of the system operation time of current bat in step (2), system track position and the current bat of system Attitude Calculation _g, terrestrial gravitation gradient moment T _gwith unexpected perturbed force F _wwith unexpected disturbance torque T _w;

(11) dynamics simulation module is according to the system control F of current bat in step (8) _cand the terrestrial gravitation F of current bat in step (10) _gwith unexpected perturbed force F _w, then according to the dynamics of orbits differential equation of classical Newton's law, calculate the second derivative of current bat dynamics of orbits parameter according to the second derivative of current bat dynamics of orbits parameter extrapolate the kinetic parameter r of the system of next bat, the dynamics of orbits differential equation of described classical Newton's law is:

$\stackrel{..}{r}=\frac{F}{m}=\frac{F_g+F_c+F_w}{m}---\left(1\right)$

In formula represent acceleration under inertial system (ax, ay, az) ', F represents all external force that aircraft is subject to, wherein F _grepresent terrestrial gravitation, F _crepresent actuating mechanism controls power, F _wrepresent unexpected perturbed force, m represents this weight of aircraft;

(12) dynamics simulation module is according to the system control moment T in step (8) _c, and the terrestrial gravitation gradient moment T in step (10) _gwith unexpected disturbance torque T _wobtain bonding force square projection T under body series that aircraft is received, then according to the attitude dynamics differential equation of classical Newton's law, calculate the derivative of current bat attitude dynamics parameter and extrapolate the attitude dynamics parameter ω of next bat, the described attitude dynamics differential equation is:

$I\stackrel{.}{\mathrm{\&omega;}}+\widetilde{\mathrm{\&omega;}}*H=T=T_C+T_g+T_w---\left(2\right)$

In formula, ω is the projection of angular velocity under body series under aircraft inertial system, be the skew matrix of ω, I is the projection under body series of system moment of inertia, and H is the projection under body series of aircraft moment of inertia, and T is bonding force square projection under body series that aircraft is received, T _cactuating mechanism controls moment, T _gterrestrial gravitation gradient moment, T _wnon-expected interference moment;

(13), according to the attitude dynamics differential equation in other aerocraft system kinetic parameter in step (9), the dynamics of orbits differential equation in step (11) and step (12), the dynamics of orbits differential equation and the attitude dynamics differential equation are carried out to linearization and obtain formula (3):

$X_{t1}=e^{A{t}_1}+{\&Integral;}_{t_0}^{t_1}{e}^{A(t_1-\mathrm{\&tau;})}B(F_w+F_C)\mathrm{d\&tau;}---\left(3\right)$

Wherein X represents dynamics of orbits parameter or attitude dynamics parameter, t ₁represent the target simulator time, A representative is through linearizing system state matrix, and B represents that control is to the Transfer Parameters of system state, F _crepresent actuating mechanism controls power, F _wrepresent unexpected perturbed force;

(14) according to the formula (3) in step (13), calculate other aerocraft system kinetic parameter, complete the determining of kinetic parameter of all aircraft nodes in this periodic system.