CN100495261C - Half-physical emulation test system for controlling and guiding, navigating and controlling soft landing for moon - Google Patents

Abstract

A semi-physical simulation testing system of moon soft-land guidance, navigation and control consists of a three-dimensional translation analog module, a three-dimensional turning analog module, a control computer, a simulation computer and a ground survey and master-control computer. The three-dimensional translation analog module and the three-dimensional turning analog module adopt a practicality modeling. Wherein, the three-dimensional translation analog module is used to imitate an orbit motion of a lander relative to the moon surface. A sand table of the moon surface is used to imitate ground features of the moon. The three-dimensional turning analog module is used to imitate an attitude motion of the lander. And other well-developed regular sensors, executing agencies, lander dynamics and kinematics can be replaced by accurate mathematical models built up by the computer. Compared with mathematical simulation, the system can make a GNC proposal and algorithm more effectively and truly verified. Compared with full-practicality analogue system, the system has advantages of low development costs and easy operation.

Description

Soft lunar landing guidance, navigation and control semi-physical simulation system

Technical field

The present invention relates to a kind of soft lunar landing guidance, navigation and control (GNC) semi-physical simulation system, can be used for checking soft lunar landing GNC scheme and algorithm.

Background technology

Twentieth century six the seventies, the Luna of the Apollo of the U.S. and Surveyor and the Soviet Union has successively repeatedly realized the soft lunar landing detection mission.Enter 21st century, world's space industry starts the moon exploration climax once more, how to design once overlapping soft lunar landing guidance, navigation and controlling schemes and making it effectively be verified the key that becomes the soft lunar landing success or not.Along with further developing of mathematical simulation technology, at present the GNC algorithm based on the earth satellite of conventional sensor can be able to effective checking by mathematical simulation greatly.And survey for the soft landing that with the moon is center gravitation body, all there were significant differences with earth satellite for its gravity field and surrounding environment and navigation means etc.Therefore, need to consider to come GNC scheme and algorithm are verified by half material object or all-real object analogue system.Bellman and Matranga have been the Apollo planned a kind of all-real object simulation checking system that is called LLRV (Lunar Landing Research Vehicle).This system provides 5/6 self gravitation with simulation lunar gravity environment by a jet engine, serve as the soft landing retroengine by other trust engines, is used for verifying the GNC scheme of the following final landing process of hundreds of rice.But because the main task of this system is for the cosmonaut provides a ground manual operation training platform, thereby they a lot of parts that comprise weight and push system are all different with the Apollo lunar module; Obviously, this pilot system also is different from the soft lunar landing GNC plan-validation of autonomous realization.Aspect development cost, the development cost of two LLRV was just up to 3,600,000 dollars at that time, and in modified LLTV development subsequently, the development cost of three LLTV is more up to 7,500,000 dollars.And under tackling problems in key technologies and the completed prerequisite of conceptual design, the manufacturing of LLRV has also been used 14 months.Estimate in Japan " Luna " (the SELenological and Engineering Explorer of emission in 2007, SELENE) the soft landing plan has been adopted the flight validation platform of an all-real object equally (Fly Test Bed FTB) has been come GNC algorithm and hardware are verified.This pilot system has been used for reference the LLRV scheme of Apollo, and by a jet engine simulation lunar gravity environment, its checking scope also is apart from the final landing process below the lunar surface hundreds of rice.This verification experimental verification comprise range finding, test the speed and navigation sensor such as IMU but do not have the hover three-dimensional imaging and the translation of process of checking to keep away the barrier control strategy.

Summary of the invention

Technology of the present invention is dealt with problems and is: overcome the deficiencies in the prior art, a kind of soft lunar landing guidance, navigation and control semi-physical simulation system are provided, this system places the control system loop with the most key outside navigation sensor as true parts, other conventional sensor and topworks and lander dynamics and kinematics then adopt ripe mathematical model to replace, and can make the GNC scheme obtain verifying when reducing development cost more authentic and validly.

Technical solution of the present invention is: soft lunar landing guidance, navigation and control semi-physical simulation system, it is characterized in that comprising: three-dimensional translating analog module, Three dimensional rotation analog module, control computer 4, simulation computer 11 and ground test and total control computer system 5, wherein:

The three-dimensional translating analog module is used to simulate the topographical features of the moon, and the three-dimensional translating of the relative lunar surface of steering order simulation lander that sends according to ground test and total control computer system 5;

The Three dimensional rotation analog module, be used to simulate the flight attitude of lander, the last cut-offing instruction that transmits according to control computer 4 carries out work, six-freedom motion to lander is measured, obtain the line-of-sight distance and the velocity information of the relative lunar surface of lander, the three-dimensional image information of the attitude angle of lander self and attitude angular velocity information and lunar surface; The speed and the lunar surface three-dimensional image information of described line-of-sight distance, the relative lunar surface of lander are passed to control computer 4, simultaneously described attitude angle and attitude angular velocity information are passed to simulation computer 11; Carry out Three dimensional rotation according to ground test and the steering order that total control computer system 5 sends;

Control computer 4, receive the telecommand that ground test and total control computer system 5 send, control Three dimensional rotation analog module work is also obtained distance, speed and image measurement information, obtain specific force and gyro angular velocity measurement information from simulation computer 11, carry out GNC according to described metrical information and calculate, obtain the propulsion system steering order and pass to simulation computer 11; Metrical information of obtaining and the control information that calculates are passed to ground test and total control computer system 5 as telemetry;

Simulation computer 11, receive the telecommand that ground test and total control computer system 5 send, obtain attitude angle and attitude angular velocity information from the Three dimensional rotation analog module, draw gyro angular velocity measurement information through the gyro to measure Model Calculation, receive the propulsion system steering order from control computer 4, draw than force information, propulsion system parameter and position and attitude information through simulation calculation, described specific force and gyro angular velocity measurement information are passed to control computer 4, described propulsion system parameter and position and attitude information are passed to ground test and total control computer system 5;

Ground test and total control computer system 5, send telecommand to control computer 4 and simulation computer 11, receiving telemetry from control computer 4 goes forward side by side the storage of line data, shared and real-time the demonstration, receive propulsion system parameter and position and attitude information from simulation computer 11, and providing translation and rotation control instruction according to position and attitude information wherein, control three-dimensional translating analog module and Three dimensional rotation analog module move on request.

Described three-dimensional translating analog module is made up of landing lunar surface simulator and three-dimensional translating telecontrol equipment 3; Landing lunar surface simulator comprises light simulator 1 and lunar surface sand table screen 2, and light simulator 1 is used to simulate the illumination condition of lunar surface, and lunar surface sand table screen 2 is used to simulate the topographical features of lunar surface, and landing lunar surface simulator places on the three-dimensional translating telecontrol equipment 3.

Described Three dimensional rotation analog module is made up of outside navigation sensor and three shaft mechanical turntables 10, outside navigation sensor comprises ranging and range rate target simulator 6, stadimeter 7, knotmeter 8 and laser imaging sensor 9, ranging and range rate target simulator 6 is used to receive the wave beam of stadimeter 6 and knotmeter 7, and after delayed and the Doppler's translation wave beam being returned to stadimeter 6 and knotmeter 7, laser imaging sensor 9 is used for lunar surface sand table screen 2 is carried out three-dimensional imaging; Outside navigation sensor places on the three shaft mechanical turntables 10, the flight attitude of the motion simulation lander by controlling three shaft mechanical turntables 10.

Described control computer 4 receives the instruction of ground test and total control computer system 5, outside navigation sensor work in the control Three dimensional rotation analog module, by stadimeter 7, knotmeter 8 and ranging and range rate target simulator 6 are measured and are obtained true line-of-sight distance and the velocity information of lander with respect to lunar surface, carry out three-dimensional imaging by 9 pairs of lunar surface sand table screens 2 of laser imaging sensor, obtain the three-dimensional image information of lunar surface, control computer 4 obtains specific force and gyro angular velocity measurement information from simulation computer 11 simultaneously, according to described distance, speed, image, specific force and gyro angular velocity information adopt the GNC algorithm to carry out navigation calculating, obtain the propulsion system steering order and pass to simulation computer 11; Metrical information of obtaining and the control information that calculates are passed to ground test and total control computer system 5 as telemetry;

Described GNC algorithm comprises the independent navigation module, track system guide module and attitude control module; The independent navigation module receives specific force and the gyro angular velocity measurement information that simulation computer 11 transmits, carry out inertial navigation and calculate the position that the back obtains lander, speed, attitude, the attitude angular velocity six-degree-of-freedom information, and described six-degree-of-freedom information passed to track system guide module, described attitude and attitude angular velocity information are passed to the attitude control module, the independent navigation module also receives the line-of-sight distance that the outside navigation sensor in the Three dimensional rotation analog module records simultaneously, speed and lunar surface image information, described metrical information is highly resolved respectively, velocity calculated and touchdown area are passed to track system guide module with result of calculation after selecting to calculate; Track system guide module is received from leading model plane piece and selects to calculate the information of acquisition through inertial navigation calculating and height, velocity calculated and touchdown area, and selects guided mode to carry out corresponding Guidance Law according to computer instruction and calculate; The attitude and the attitude angular velocity information that calculate through inertial navigation that the attitude control module is received from that leading model plane piece transmits, and carry out attitude control law according to different attitude demands for control and calculate; Result of calculation according to track system guide module and attitude control module midcourse guidance control law can draw the propulsion system steering order that is used for track and attitude control.

The guided mode of described track system guide module comprises gravity turning nominal trajectory guided mode, becomes thrust guided mode, Bang-Bang and phase plane guided mode, and the nominal trajectory guided mode.

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

Described simulation computer 11 receives the telecommand of ground test and total control computer system 5, obtain attitude angle and attitude angular velocity information from three shaft mechanical turntables 10, after dynamics and the calculating of kinematics realistic model, obtain simulating the angular velocity measurement information of gyro; Simulation computer 11 also receives the propulsion system steering order from control computer 4, obtains the position and attitude information of propulsion system parameter and lander after dynamics and the calculating of kinematics realistic model, also can obtain the specific force metrical information simultaneously; Simulation computer 11 is passed to control computer 4 with described specific force and gyro angular velocity measurement information, and described propulsion system parameter and position and attitude information are passed to ground test and total control computer system 5.

Described dynamics comprises that with the kinematics realistic model topworks's mathematical model, dynamics contract than computation model and conventional sensor measurement model with kinematics model, track.Topworks's mathematical model receives the propulsion system steering order that control computer 4 provides, and draws the propulsion system parameter and pass to dynamics and kinematics model and ground test and total control computer system 5 after simulation calculation; Dynamics combines the propulsion system parameter with kinematics model, draw position, speed, attitude angle, the attitude angular velocity six-degree-of-freedom information of lander as calculated, and described six-degree-of-freedom information passed to ground test and total control computer system 5, simultaneously position and velocity information are passed to track and contract than computation model and conventional sensor measurement model; Track contracts and than computation model position and velocity information is contracted than after calculating, will contract than after position and velocity information pass to ground test and total control computer system 5; Position and velocity information that conventional sensor measurement model reception dynamics and kinematics model provide, after the accelerometer measures Model Calculation, obtain the specific force metrical information, simultaneously, conventional sensor measurement model also receives attitude angle and the attitude angular velocity information that three shaft mechanical turntables 10 provide, obtain gyro angular velocity measurement information through the gyro to measure Model Calculation, and described specific force and gyro angular velocity measurement information are passed to control computer 4.

The present invention compared with prior art has following advantage:

(1) compares with mathematical simulation, critical components such as outside navigation sensor in will comprising range finding, test the speed and being imaged on adopt true parts, effectively the simulation lander makes the checking of soft landing process GNC algorithm effective more comprehensively with respect to the three-dimensional translating and the Three dimensional rotation of lunar surface;

(2) compare with all-real object emulation, with the conventional sensor and the employing precise math model replacements such as topworks and lander dynamics kinematic calculation of maturation, more simple than all-real object emulation;

(3) increased the one-tenth image sensor lunar surface carried out the imaging of three-dimensional high definition rate, for the safety of lander keep away the barrier and landing lay a good foundation.

Description of drawings

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

Fig. 2 is lunar surface sand table screen dimensions of the present invention and lunar surface simulation drawing;

Fig. 3 is three-dimensional translating telecontrol equipment figure of the present invention;

Fig. 4 is a control computer GNC algorithm data process flow diagram of the present invention;

Fig. 5 is a simulation computer realistic model data flowchart of the present invention.

Embodiment

One, critical component specific design and enforcement

(1) three shaft mechanical turntable 10

Three shaft mechanical turntables 10 are used to simulate the three-axis attitude motion of lander, the attitude output of simulation gyro after the attitude angle of turntable motion and the gyro to measure Model Calculation of attitude angular velocity information in simulation computer 11.The embodiment of the invention will be utilized the existing French import three shaft mechanical turntables of our unit.

(2) outside navigation sensor

Outside navigation sensor comprises ranging and range rate target simulator 6, stadimeter 7, knotmeter 8 and laser imaging sensor 9.

Owing to be subjected to the restriction in simulation laboratory space, stadimeter 7 and knotmeter 8 can't directly be measured lunar surface sand table screen.The present invention receives by the radar beam that increases by 6 pairs of stadimeters 7 of a ranging and range rate target simulator and knotmeter 8 and send, and by sending to stadimeter 7 and knotmeter 8 again after time-delay and the Doppler's translation processing, thereby finish corresponding distance and velocity survey.Ranging and range rate target simulator 6 carries out layout according to the installation requirement of stadimeter 7 and knotmeter 8.

Laser imaging sensor 9 is mainly used in when lunar surface 100m hovers moon surface imaging.

Embodiment of the invention employing tellurometer survey is tested the speed, the outside navigation sensor assembled scheme of laser infrared radar imaging, adopts the ratio that contracts of 1:10 during imaging, promptly at the 10m place lunar surface sand table screen 2 is carried out the imaging of three-dimensional high definition rate.

(3) lunar surface sand table screen 2

The main application of lunar surface sand table screen 2 is landing lunar surface characteristic informations that simulation is provided for laser imaging sensor 9.The foundation of its size design is: make that the imaging scope is as far as possible big, be consistent after the imaging scope before the translation and the translation as far as possible and the big translation space of trying one's best is arranged.

Becoming the visual field of image sensor is 30 ° * 30 °, is 53.6m * 53.6m in the lunar surface zone that the imaging of 100m place covered.The embodiment of the invention adopts the ratio that contracts of 1:10, and the size of the lunar surface sand table screen 2 that imaging covered should be 5.36m * 5.36m.If become the image sensor optical axis to be positioned at turntable table top center,, become image sensor to form complete 3-D view in each 2.68m up and down at the distance optical axis then at the 10m height.Size is as shown in Figure 2 set, and left and right sides both direction all satisfies the requirement that is not less than 2.68m before and after translation, can become complete left and right sides image; Last direction by roll sand table screen of the mode of roller bearing, can guarantee its size constancy up and down under the prerequisite of simulation sand table screen upper and lower translation, therefore the top size also satisfies the requirement that is not less than 2.68m.Have only sand table screen below size at present owing to limited by the place to be difficult to meet the demands, but this does not influence control system obtain information from three-dimensional imaging.

Set sand table translation downwards to the right, translation distance is 1.5 meters to the right, and translation distance is 1 meter downwards, and the translation control of both direction can be carried out simultaneously.

(4) the three-dimensional translating telecontrol equipment 3

Among the present invention, the translation of shielding 2 relative three shaft mechanical turntables 10 with the lunar surface sand table is simulated in the practical flight lander with respect to the translational motion of lunar surface.This three-dimensional translating is realized by three-dimensional translating telecontrol equipment 3.

In the embodiment of the invention, three-dimensional translating telecontrol equipment 3 will adopt the form of guide rail.As shown in Figure 3, lay two of longitudinal rails about 15m,, simulate the vertical decline of the relative lunar surface of lander by cross slide way moving forward and backward on longitudinal rail at the cross slide way of laying on the longitudinal rail about 10m along the turret axis direction.Lunar surface sand table screen 2 is positioned on the cross slide way, can be along the cross slide way move left and right, simulate motion in one dimension in the real standard face with this; The translation of other direction can realize by the roller bearing that the lunar surface sand table shields 2 belows in the surface level.

(5) control computer 4 and GNC algorithm

Control computer 4 is nucleus equipments of soft lunar landing GNC semi-physical simulation system.Control computer 4 is carried out the GNC algorithm, should have guidance, navigation and control function under the cooperation of other equipment.Consider the restriction of three-axle table, control computer 4 is moved on on the ground by turntable payload weight and volume.In the embodiment of the invention, control computer 4 adopts industrial computer to replace.

The GNC algorithm is the application program that operates on the control computer 4, should have following mode of operation and function: attitude adjustment section gravity turning+nominal trajectory guided mode, landing phase nominal trajectory guided mode, the process of hovering become thrust guided mode, touchdown area selection, translation track control, Attitude Tracking and stable control etc.

The GNC algorithm data flow process of the embodiment of the invention as shown in Figure 4.Comprise the independent navigation module, track system guide module and attitude control module.The independent navigation module receives specific force and the gyro angular velocity measurement information that simulation computer 11 transmits, carry out inertial navigation and calculate the position that the back obtains lander, speed, attitude, the attitude angular velocity six-degree-of-freedom information, and described six-degree-of-freedom information passed to track system guide module, described attitude and attitude angular velocity information are passed to the attitude control module, the independent navigation module also receives the line-of-sight distance that the outside navigation sensor in the Three dimensional rotation analog module records simultaneously, speed and lunar surface image information, described metrical information is highly resolved respectively, velocity calculated and touchdown area are passed to track system guide module with result of calculation after selecting to calculate; Track system guide module is received from leading model plane piece and selects to calculate the information of acquisition through inertial navigation calculating and height, velocity calculated and touchdown area, and selects guided mode to carry out corresponding Guidance Law according to computer instruction and calculate; The attitude and the attitude angular velocity information that calculate through inertial navigation that the attitude control module is received from that leading model plane piece transmits, and carry out attitude control law according to different attitude demands for control and calculate; Result of calculation according to track system guide module and attitude control module midcourse guidance control law can draw the propulsion system steering order that is used for track and attitude control.

The inertial navigation computing formula that the embodiment of the invention adopts is as follows:

(a) attitude and attitude angular velocity calculate

$\left\{\begin{array}{c}{\widehat{\mathrm{\ω}}}_b=\widetilde{\mathrm{\ω}}\\ \stackrel{\·}{\widehat{\mathrm{\ϵ}}}=\frac{1}{2}({\widehat{\mathrm{\ϵ}}}^{\×}+\widehat{\mathrm{\η}}{E}_3)\widehat{\mathrm{\ω}},\stackrel{\·}{\widehat{\mathrm{\η}}}=-\frac{1}{2}{\widehat{\mathrm{\ϵ}}}^T\widetilde{\mathrm{\ω}}\end{array}\right.$

Wherein,

With

Be respectively the lander body attitude angular velocity vector after the calculating and the plain expression of quaternary (vector part and scalar part) of attitude angle;

Be gyro to measure angular velocity.Moon spin velocity is ignored in this calculating.

(b) position and speed calculation

$\left\{\begin{array}{c}\stackrel{\·}{\hat{r}=\hat{v}}\\ \stackrel{\·}{\hat{v}}=A(\widehat{\mathrm{\ϵ}},\widehat{\mathrm{\η}})\·\tilde{f}+g_m\end{array}\right.$

Wherein,

Be respectively the position and the velocity of the relative lunar surface of lander after the calculating;

Ratio force vector for accelerometer measures; Matrix A is for being tied to month coordinate transform battle array of flat system, its variable from body

Be respectively the vector and the scalar part of attitude quaternary element, identical with the expression in " attitude and attitude angular velocity calculate "; g _mBe moonscape acceleration of gravity vector.

The height solution formula that the embodiment of the invention adopts is as follows:

${\hat{h}}_R=\frac{3V}{S}=\frac{\left|{\left[\begin{array}{ccc}{l}_1^b& l_2^b& l_3^b\end{array}\right]}^T\right|}{|(l_2^b-l_1^b)\×(l_3^b-l_1^b)|}$

Wherein,

Be the estimated value of lander apart from the lunar surface height; V, S are respectively hexahedral volume and bottom surface (on the lunar surface) area that is formed by three wave beams of stadimeter and lunar surface;

Be respectively not three beam vector of coplane of stadimeter, wherein subscript b represents this vector representation under the lander body coordinate system, can directly record the sight line length of three wave beams by stadimeter, and the angular relationship of wave beam under the lander body series is known.

The velocity calculated formula that the embodiment of the invention adopts is as follows:

${\tilde{v}}_i=<l_i^b\&CenterDot;v^b>$

Wherein,＜expression two vectors dot product; Subscript i represents radar beam (i=1,2,3); Subscript b and vector

Meaning and " the stadimeter height resolves " in identical; v ^bThe expression lander is the unknown with respect to the velocity of lunar surface; That expression is recorded by knotmeter, resultant velocity v ^bAt the speed component of wave beam i direction, for known.Bring the measured value of radar three wave beams into following formula respectively and constitute a system of equations, separate this system of equations and can try to achieve resultant velocity.

The system of selection of touchdown area is as follows in the embodiment of the invention:

Month surface imaging sensor can provide the elevation information of lunar surface touchdown area each point in Polaroid.Shared lunar surface area is S if lander lands the back, and the tentative programme that touchdown area is selected is: be unit with S, be that starting point is carried out counterclockwise (clockwise) and searched for current landing point.To the zone of each piece S size, judge landing conditions---whether the gradient of touchdown area internal projection height and touchdown area meets the demands, until searching safe touchdown area.

Guidance control law mode of operation and computing formula that the embodiment of the invention adopts are as follows:

(a) attitude is adjusted section---the guidance of gravity turning nominal trajectory

$u_{\mathrm{Fc}}=\frac{1}{g_m\cos \mathrm{\&Psi;}}[g_m(1-\frac{\mathrm{\&tau;}{\mathrm{<figure-callout\; id="2"\; label="sin"\; filenames="C200710121319E00211.png"\; state="\{\{state\}\}">sin</figure-callout>}}^2\mathrm{\&Psi;}}{v+\mathrm{\&tau;}})+{\stackrel{\&CenterDot;\&CenterDot;}{h}}_D-[k_p(h-h_D)+k_d(\stackrel{\&CenterDot;}{h}-{\stackrel{\&CenterDot;}{h}}_D)]]$

Wherein, u _FcFor the sustainer thrust-weight ratio is guidanceed command; H,

Represent respectively practical flight apart from lunar surface height and vertical speed; h _D, Nominal height, nominal vertical speed and the vertical acceleration component of representing lander in the nominal trajectory respectively; V is the size of resultant velocity; Ψ is the angle of decline rate direction and local vertical direction; k _p, k _dBe respectively ratio and differential coefficient; τ is a given little constant; g _mBe moonscape acceleration of gravity.

(b) the translation process of hovering: vertical direction---become the thrust guidance; Horizontal direction---Bang-Bang+ phase plane guidance

On the vertical direction, control main thrust engine make its thrust all the time with lander gravity equal and opposite in direction, direction is opposite; On the horizontal direction, utilize the translational thrust device at first to adopt the positive and negative switch control strategy that quickens afterwards to slow down earlier, near the target location time, switch to phase plane control strategy again about position and speed.Under this phase plane control strategy, the translational thrust device has only " just opening ", " negative opening " and " pass " three states, so on position and speed phase plane, formed four switching lines (two burst at the seams, two close lines), phase plane is divided into four zones, control computer is according to the zones of different switch translational thrust device of lander place phase plane, thereby the control lander is near impact point.

(c) final landing section---nominal trajectory guidance

$u_{\mathrm{Fc}}=\frac{1}{g_m}[g_m+{\stackrel{\&CenterDot;\&CenterDot;}{h}}_D-(k_p(h-h_D)+k_d(\stackrel{\&CenterDot;}{h}-{\stackrel{\&CenterDot;}{h}}_D))]$

Identical among meaning of parameters in the following formula and (a).

Final landing section guided mode is the special case that attitude is adjusted the section guided mode, promptly requires Ψ=0 on the basis of attitude adjustment section Guidance Law, and this represents that promptly lander vertically descends.

(d) attitude control mode---three-axis attitude stabilization phase plane control

In the embodiment of the invention, control the phase plane control strategy that all adopts about attitude angle and attitude angular velocity apart from the attitude below the lunar surface 150m.Position-speed phase plane control strategy among this phase plane control strategy and above-mentioned (b) is similar.The thruster that is used for attitude control has only " just opening ", " negative opening " and " pass " three states, so on attitude angle and attitude angular velocity phase plane, formed four switching lines (two burst at the seams, two close lines), phase plane is divided into four zones, control computer is according to the zones of different switch appearance control thruster of lander place phase plane, thereby control lander attitude stabilization is near the nominal attitude.

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

Simulation computer 11 main track and the tasks such as attitude dynamics and kinematics emulation be responsible for specifically should possess following function: lander decline process track and attitude dynamics and kinematics emulation; Conventional sensor mathematical simulation such as IMU; Topworks's mathematical simulations such as propulsion system; Orbital motion is contracted than calculating; And the resource sharing between ground test and the total control computer.In the embodiment of the invention, simulation computer 11 adopts high performance PC.

The function of simulation computer 11 is finished by dynamics and kinematics realistic model, and dynamics and kinematics realistic model should comprise that track/attitude dynamics and kinematics model, IMU measurement model, topworks's (comprising main thrust engine, translational thrust device and attitude control thruster) mathematical model and track contract than computation model etc.

The data flow of embodiment of the invention dynamics and kinematics realistic model as shown in Figure 5.Comprise that topworks's mathematical model, dynamics contract than computation model and conventional sensor measurement model with kinematics model, track; Topworks's mathematical model receives the propulsion system steering order that control computer 4 provides, and draws the propulsion system parameter and pass to dynamics and kinematics model and ground test and total control computer system 5 after simulation calculation; Dynamics combines the propulsion system parameter with kinematics model, draw lander position, speed, attitude angle, attitude angular velocity six-degree-of-freedom information as calculated, and described six-degree-of-freedom information passed to ground test and total control computer system 5, simultaneously position and velocity information are passed to track and contract than computation model and conventional sensor measurement model; Track contracts and than computation model position and velocity information is contracted than after calculating, will contract than after position and velocity information pass to ground test and total control computer system 5; Position and velocity information that conventional sensor measurement model reception dynamics and kinematics model provide, after the accelerometer measures Model Calculation, obtain the specific force metrical information, simultaneously, conventional sensor measurement model also receives attitude angle and the attitude angular velocity information that three shaft mechanical turntables 10 provide, obtain gyro angular velocity measurement information through the gyro to measure Model Calculation, and described specific force and gyro angular velocity measurement information are passed to control computer 4.

Topworks's model that the embodiment of the invention adopts is as follows:

(a) appearance control thruster model

The opening and closing of ignoring thruster postpone, and can adopt following naive model in the emulation:

F _atti(t)＝F _0atti[I(t)-I(t-T)]

Wherein, F _Atti(t) the actual output of expression appearance control thruster; F _0attiThe nominal thrust size of expression appearance control thruster; I () represents step function, and T is the time width of jet command pulse.

(b) rail control Variable Thrust Engine model

The opening and closing of ignoring thruster postpone, and consider the thruster control accuracy that influence is bigger, can adopt following naive model during emulation:

F _obt(t)＝(1+δ)F _0obt

Wherein, F _Obt(t) the actual output of expression rail control engine; F _0obtThe nominal thrust size of expression rail control engine; δ is the Thrust Control precision of rail control engine.

Dynamics and kinematics model that the embodiment of the invention adopts are as follows:

(a) gyro to measure model

$\widetilde{\mathrm{\&omega;}}={\mathrm{\&omega;}}_r+{\mathrm{\&omega;}}_0+{\mathrm{\&omega;}}_1$

Wherein,

Be gyro to measure angular velocity; ω _rBe the actual attitude angular velocity of lander, in the l-G simulation test, this is promptly for learning the attitude angular velocity that Model Calculation obtains by attitude motion; ω ₀, ω ₁Be respectively the constant value drift and the random drift of gyro, can demarcate.

(b) accelerometer measures model

$\tilde{f}=(1+k)(f_r+\&dtri;+w)$

Wherein,

Specific force value for accelerometer measures; f _rBe the actual specific force of lander; K is the scale factor error;

For accelerometer zero partially; W is for measuring noise.

Dynamics and kinematics model that the embodiment of the invention adopts are as follows:

(a) dynamics of orbits model

$\stackrel{\&CenterDot;\&CenterDot;}{r}=F-\frac{{\mathrm{\&mu;}}_m}{{\mathrm{<figure-callout\; id="3"\; label="r"\; filenames="C200710121319E00211.png"\; state="\{\{state\}\}">r</figure-callout>}}^3}\&CenterDot;r+F_{\mathrm{\&epsiv;}}-{\mathrm{\&mu;}}_e\&CenterDot;(\frac{1}{{\mathrm{<figure-callout\; id="3"\; label="r\; e"\; filenames="C200710121319E00211.png"\; state="\{\{state\}\}">r</figure-callout>}}_e^{}}\&CenterDot;r_e+\frac{1}{{\mathrm{\&Delta;}}_{\mathrm{<figure-callout\; id="3"\; label="e"\; filenames="C200710121319E00211.png"\; state="\{\{state\}\}">e</figure-callout>}}^3}\&CenterDot;{\mathrm{\&Delta;}}_e)-{\mathrm{\&mu;}}_s\&CenterDot;(\frac{1}{{\mathrm{<figure-callout\; id="3"\; label="r\; s"\; filenames="C200710121319E00211.png"\; state="\{\{state\}\}">r</figure-callout>}}_s^{}}\&CenterDot;r_s+\frac{1}{{\mathrm{\&Delta;}}_{\mathrm{<figure-callout\; id="3"\; label="s"\; filenames="C200710121319E00211.png"\; state="\{\{state\}\}">s</figure-callout>}}^3}\&CenterDot;{\mathrm{\&Delta;}}_s)+f$

Wherein, μ _m, μ _e, μ _sBe respectively a moon heart, the earth's core and heliocentric gravitational constant; Δ _e=r-r _e, Δ s=r-r _s, r, r _eAnd r _sBe respectively the lander barycenter and arrive a month heart, month heart arrive the day heart to the earth's core and month heart radius vector.First initiative brake that F is a propulsion system in the right, second is the center gravitation of the moon, the 3rd is the non-spherical gravitation perturbation of the moon, the 4th and the 5th gravitation perturbation that is respectively the earth and the sun.These several is the main perturbation source of nearly month spacecraft flight.F is other external perturbation power except that above-mentioned three perturbations.

(b) orbital motion model

$\stackrel{\&CenterDot;}{r}=v$

(c) attitude dynamics model

$I\stackrel{\&CenterDot;}{\mathrm{\&omega;}}+\mathrm{\&omega;}\&times;\mathrm{I\&omega;}=T_c+T_d$

Wherein, I is the moment of inertia battle array of lander, and ω is the attitude angular velocity vector of lander under inertial space, T _cAnd T _dBe respectively control moment and disturbance torque.Among the present invention, T _cMainly be meant jet moment, T _dComprise jet disturbance torque, gravity gradient torque, solar radiation pressure moment etc.

This attitude dynamics model is the simplest rigid model.For the present invention,, therefore should take into full account the influence of liquid sloshing to the lander attitude because Fuel Consumption is relatively large.

(d) attitude motion is learned model

$\left\{\begin{array}{c}\stackrel{\&CenterDot;}{\mathrm{\&epsiv;}}=\frac{1}{2}({\mathrm{\&epsiv;}}^{\&times;}+\mathrm{\&eta;}{E}_3){\mathrm{\&omega;}}_b\\ \stackrel{\&CenterDot;}{\mathrm{\&eta;}}=-\frac{1}{2}{\mathrm{\&epsiv;}}^T{\mathrm{\&omega;}}_b\end{array}\right.$

Following formula is that the lander attitude motion of the plain expression of quaternary is learned equation.Wherein, ω _bBe the attitude angular velocity vector of lander under the body coordinate system, itself and 3) in the relation of ω be: ω=ω _b+ ω _o, ω here _oThe angular velocity vector of decline track under inertial system of expression lander.ε and η represent the vector and the scalar part of quaternary element respectively,

Wherein, a and Be respectively the Euler's axle and the Eulerian angle that are used for coordinate system rotation.

(7) ground test and total control computer system 5

Ground test and total control computer system 5 are used for management, data acquisition, processing and the control of pilot system, and this system forms computer network, can finish demonstration, storage and the management of data.In the embodiment of the invention, ground test and total control computer system 5 adopt high performance PC to replace.Ground test and master control computer software are installed on each nodes PC of computer network, are achieved as follows function with the configuration modular form: to the management and the control of pilot system; Collection, processing, storage and demonstration in real time to test figure; Resource sharing with simulation computer 11.

Two, workflow

Workflow of the present invention is as follows:

(1) initial time, three-dimensional translating telecontrol equipment 3 have certain initial relative position and speed, and three shaft mechanical turntables 10 have certain initial attitude angle and attitude angular velocity, and they are used for simulating relative position and the attitude information that lander has;

(2) (a) control computer 4 is opened stadimeter 7, knotmeter 8 and ranging and range rate target simulator 6 according to the telecommand of ground test and total control computer system 5, opens laser imaging sensor 9 simultaneously;

(b) record the distance and the velocity information of the relative lunar surface of lander by stadimeter 7, knotmeter 8 and ranging and range rate target simulator 6, carry out three-dimensional imaging, obtain three-dimensional image information by 9 pairs of lunar surface sand tables screens 2 of laser imaging sensor;

(c) the current attitude angle and the attitude angular velocity information that provide according to three shaft mechanical turntables 10 of simulation computer 11, after the gyro to measure Model Calculation by one of conventional sensor measurement model, the simulation gyro provides lander attitude angular velocity metrical information;

(d) utilize the accelerometer measures Model Calculation of one of conventional sensor measurement model in the simulation computer 11 to obtain the specific force metrical information of lander;

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

(3) control computer 4 is according to the distance, speed, image, specific force and the attitude angular velocity metrical information that transmit, calculate by corresponding navigation, guidance and control, obtain finally that main thrust is guidanceed command (sustainer thrust-weight ratio), touchdown area selection instruction (translation direction), the translational thrust device is guidanceed command (translational thrust device working time) and attitude stabilization steering order (appearance control thruster working time), and above-mentioned propulsion system steering order is passed to simulation computer 11;

(4) the propulsion system steering order that provides according to control computer 4, simulation computer 11 begins to carry out dynamics and the kinematic calculation and the conventional sensor navigation calculating of lander, obtain next position, speed, attitude angle and attitude angular velocity six-degree-of-freedom information of lander constantly, then described six-degree-of-freedom information is contracted than calculating, to contract and pass to ground test and total control computer system 5, and carry out data storage, share and show in real time than front-back direction attitude information;

(5) position and the attitude information that transmits according to simulation computer 11, ground test and total control computer system 5 will contract than after position and rate control instruction and angle and angular velocity steering order be transferred to three-dimensional translating telecontrol equipment 3 and three shaft mechanical turntables 10 respectively, drive two devices by instruction motion separately, ground test and total control computer system 5 also will contract and be transferred to ranging and range rate target simulator 6 than preceding position and rate control instruction simultaneously, be used for next distance and velocity survey constantly;

(6) next control cycle will repeat the step of (2)～(5), thereby constitute closed loop GNC semi-physical simulation system.

The content that is not described in detail in the instructions of the present invention belongs to this area professional and technical personnel's known prior art.

Claims (9)

1, soft lunar landing guidance, navigation and control semi-physical simulation system, it is characterized in that comprising: three-dimensional translating analog module, Three dimensional rotation analog module, control computer (4), simulation computer (11) and ground test and total control computer system (5), wherein:

The three-dimensional translating analog module is used to simulate the topographical features of the moon, and the three-dimensional translating of the relative lunar surface of steering order simulation lander that sends according to ground test and total control computer system (5);

The Three dimensional rotation analog module, be used to simulate the flight attitude of lander, the last cut-offing instruction that transmits according to control computer (4) carries out work, six-freedom motion to lander is measured, obtain the line-of-sight distance and the velocity information of the relative lunar surface of lander, the three-dimensional image information of the attitude angle of lander self and attitude angular velocity information and lunar surface; The speed and the lunar surface three-dimensional image information of described line-of-sight distance, the relative lunar surface of lander are passed to control computer (4), simultaneously described attitude angle and attitude angular velocity information are passed to simulation computer (11); Carry out Three dimensional rotation according to ground test and the steering order that total control computer system (5) sends;

Control computer (4), receive the telecommand that ground test and total control computer system (5) send, control Three dimensional rotation analog module work is also obtained distance, speed and image measurement information, obtain specific force and gyro angular velocity measurement information from simulation computer (11), carry out GNC according to described metrical information and calculate, obtain the propulsion system steering order and pass to simulation computer (11); Metrical information of obtaining and the control information that calculates are passed to ground test and total control computer system (5) as telemetry;

Simulation computer (11), receive the telecommand that ground test and total control computer system (5) send, obtain attitude angle and attitude angular velocity information from the Three dimensional rotation analog module, draw gyro angular velocity measurement information through the gyro to measure Model Calculation, receive the propulsion system steering order from control computer (4), draw through simulation calculation and to compare force information, propulsion system parameter and position and attitude information, described specific force and gyro angular velocity measurement information are passed to control computer (4), described propulsion system parameter and position and attitude information are passed to ground test and total control computer system (5);

Ground test and total control computer system (5), send telecommand to control computer (4) and simulation computer (11), receiving telemetry from control computer (4) goes forward side by side the storage of line data, shared and real-time the demonstration, receive propulsion system parameter and position and attitude information from simulation computer (11), and providing translation and rotation control instruction according to position and attitude information wherein, control three-dimensional translating analog module and Three dimensional rotation analog module move on request.

2, soft lunar landing guidance according to claim 1, navigation and control semi-physical simulation system, it is characterized in that: described three-dimensional translating analog module is made up of landing lunar surface simulator and three-dimensional translating telecontrol equipment (3); Landing lunar surface simulator comprises light simulator (1) and lunar surface sand table screen (2), light simulator (1) is used to simulate the illumination condition of lunar surface, lunar surface sand table screen (2) is used to simulate the topographical features of lunar surface, and landing lunar surface simulator places on the three-dimensional translating telecontrol equipment (3).

3, soft lunar landing guidance according to claim 1, navigation and control semi-physical simulation system, it is characterized in that: described Three dimensional rotation analog module is made up of outside navigation sensor and three shaft mechanical turntables (10), outside navigation sensor comprises ranging and range rate target simulator (6), stadimeter (7), knotmeter (8) and laser imaging sensor (9), ranging and range rate target simulator (6) is used to receive the wave beam of stadimeter (6) and knotmeter (7), and after delayed and the Doppler's translation wave beam being returned to stadimeter (6) and knotmeter (7), laser imaging sensor (9) is used for lunar surface sand table screen (2) is carried out three-dimensional imaging; Outside navigation sensor places on the three shaft mechanical turntables (10), the flight attitude of the motion simulation lander by controlling three shaft mechanical turntables (10).

4, soft lunar landing guidance according to claim 1, navigation and control semi-physical simulation system, it is characterized in that: described control computer (4) receives the instruction of ground test and total control computer system (5), outside navigation sensor work in the control Three dimensional rotation analog module, by stadimeter (7), knotmeter (8) and ranging and range rate target simulator (6) are measured and are obtained true line-of-sight distance and the velocity information of lander with respect to lunar surface, by laser imaging sensor (9) lunar surface sand table screen (2) is carried out three-dimensional imaging, obtain the three-dimensional image information of lunar surface, control computer (4) obtains specific force and gyro angular velocity measurement information from simulation computer (11) simultaneously, according to described distance, speed, image, specific force and gyro angular velocity information adopt the GNC algorithm to carry out navigation calculating, obtain the propulsion system steering order and pass to simulation computer (11); Metrical information of obtaining and the control information that calculates are passed to ground test and total control computer system (5) as telemetry;

5, soft lunar landing guidance according to claim 4, navigation and control semi-physical simulation system, it is characterized in that: described GNC algorithm comprises the independent navigation module, track system guide module and attitude control module; The independent navigation module receives specific force and the gyro angular velocity measurement information that simulation computer (11) transmits, carry out inertial navigation and calculate the position that the back obtains lander, speed, attitude, the attitude angular velocity six-degree-of-freedom information, and described six-degree-of-freedom information passed to track system guide module, described attitude and attitude angular velocity information are passed to the attitude control module, the independent navigation module also receives the line-of-sight distance that the outside navigation sensor in the Three dimensional rotation analog module records simultaneously, speed and lunar surface image information, described metrical information is highly resolved respectively, velocity calculated and touchdown area are passed to track system guide module with result of calculation after selecting to calculate; Track system guide module is received from leading model plane piece and selects to calculate the information of acquisition through inertial navigation calculating and height, velocity calculated and touchdown area, and selects guided mode to carry out corresponding Guidance Law according to computer instruction and calculate; The attitude and the attitude angular velocity information that calculate through inertial navigation that the attitude control module is received from that leading model plane piece transmits, and carry out attitude control law according to different attitude demands for control and calculate; Result of calculation according to track system guide module and attitude control module midcourse guidance control law can draw the propulsion system steering order that is used for track and attitude control.

6, soft lunar landing guidance according to claim 5, navigation and control semi-physical simulation system, it is characterized in that: the guided mode of described track system guide module comprises gravity turning nominal trajectory guided mode, becomes thrust guided mode, Bang-Bang and phase plane guided mode, and the nominal trajectory guided mode.

7, soft lunar landing guidance according to claim 5, navigation and control semi-physical simulation system, it is characterized in that: the attitude control mode of described attitude control module adopts phase plane control.

8, soft lunar landing guidance according to claim 1, navigation and control semi-physical simulation system, it is characterized in that: described simulation computer (11) receives the telecommand of ground test and total control computer system (5), obtain attitude angle and attitude angular velocity information from three shaft mechanical turntables (10), after dynamics and the calculating of kinematics realistic model, obtain simulating the angular velocity measurement information of gyro; Simulation computer (11) also receives the propulsion system steering order from control computer (4), obtains the position and attitude information of propulsion system parameter and lander after dynamics and the calculating of kinematics realistic model, also can obtain the specific force metrical information simultaneously; Simulation computer (11) is passed to control computer (4) with described specific force and gyro angular velocity measurement information, and described propulsion system parameter and position and attitude information are passed to ground test and total control computer system (5).

9, soft lunar landing guidance according to claim 8, navigation and control semi-physical simulation system, it is characterized in that: described dynamics comprises that with the kinematics realistic model topworks's mathematical model, dynamics contract than computation model and conventional sensor measurement model with kinematics model, track.Topworks's mathematical model receives the propulsion system steering order that control computer (4) provides, and draws the propulsion system parameter and pass to dynamics and kinematics model and ground test and total control computer system (5) after simulation calculation; Dynamics combines the propulsion system parameter with kinematics model, draw position, speed, attitude angle, the attitude angular velocity six-degree-of-freedom information of lander as calculated, and described six-degree-of-freedom information passed to ground test and total control computer system (5), simultaneously position and velocity information are passed to track and contract than computation model and conventional sensor measurement model; Track contracts and than computation model position and velocity information is contracted than after calculating, will contract than after position and velocity information pass to ground test and total control computer system (5); Position and velocity information that conventional sensor measurement model reception dynamics and kinematics model provide, after the accelerometer measures Model Calculation, obtain the specific force metrical information, simultaneously, conventional sensor measurement model also receives attitude angle and the attitude angular velocity information that three shaft mechanical turntables (10) provide, obtain gyro angular velocity measurement information through the gyro to measure Model Calculation, and described specific force and gyro angular velocity measurement information are passed to control computer (4).

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