CN101033973A  Attitude determination method of miniaircraft inertial integrated navigation system  Google Patents
Attitude determination method of miniaircraft inertial integrated navigation system Download PDFInfo
 Publication number
 CN101033973A CN101033973A CN 200710021401 CN200710021401A CN101033973A CN 101033973 A CN101033973 A CN 101033973A CN 200710021401 CN200710021401 CN 200710021401 CN 200710021401 A CN200710021401 A CN 200710021401A CN 101033973 A CN101033973 A CN 101033973A
 Authority
 CN
 China
 Prior art keywords
 attitude
 amp
 step
 air vehicle
 micro air
 Prior art date
Links
 238000000034 methods Methods 0.000 claims abstract description 11
 238000001914 filtration Methods 0.000 claims description 10
 239000011295 pitch Substances 0.000 claims description 5
 230000001419 dependent Effects 0.000 claims description 4
 238000004364 calculation methods Methods 0.000 claims description 2
 230000000694 effects Effects 0.000 description 2
 239000000203 mixtures Substances 0.000 description 2
 230000001276 controlling effects Effects 0.000 description 1
 230000002401 inhibitory effects Effects 0.000 description 1
 XUIMIQQOPSSXEZUHFFFAOYSAN silicon Chemical compound data:image/svg+xml;base64,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 data:image/svg+xml;base64,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 [Si] XUIMIQQOPSSXEZUHFFFAOYSAN 0.000 description 1
Abstract
Description
One, technical field
The invention belongs to the movement parameter measurement systems technology field of micro air vehicle.
Two, background technology
Micro air vehicle is in closely military surveillance, target search, and there are special use value and wide prospect in fields such as disaster monitoring.The Core Feature of the closed loop that the Navigation, Guide and Controlling system of micro air vehicle forms is that the state of minute vehicle is measured and controlled, and makes it can reach the purpose of autonomous flight.The micro air vehicle attitude and heading reference system is the subsystem of miniature navigation system, is used for the attitude of realtime measurement minute vehicle, and the attitude information of measuring is fed back to the controller of minute vehicle.The micro air vehicle attitude and heading reference system is one of gordian technique in the micro air vehicle for attitude stabilization and flight control algorithm provide the attitude information of micro air vehicle, is the essential condition of micro air vehicle automated spacecraft.
Micro air vehicle inertia assembled gesture determines that system adopts the miniature strapdown inertial navigation system of MEMS (Micro ElectroMechanicalSystem) inertial sensor (little gyro, microacceleration gauge).Adopting the silicon microgyroscope and the accelerometer volume of MEMS technology little, in light weight, is present unique inertial sensor that satisfies minute vehicle weight and volume requirement.Strapdown inertial navigation system (Strapdown Inertial Navigation System) is a kind of system that realizes navigation feature based on Newton's laws of motion in the mode of calculating, its core sensor is the measuring unit that is made of gyroscope (angular motion sensor) and accelerometer two class inertial sensors such as (line motion sensors), and the measuring unit that is made of abovementioned two class sensors directly connects firmly on motion carrier.Processor in the strapdown inertial navigation system is by A/D conversion circuit or directly adopt digital interface, read the data of interior angular motion of abovementioned measuring unit and line motion sensor, press the principle process of inertial navigation algorithm, primary measuring data is processed, calculate the parameters such as attitude, speed and position of motion carrier.
MEMS inertial sensor measuring error is big, can disperse fast based on the attitude error of the microinertial navigation system of MEMS inertial sensor, is difficult to satisfy the demand of micro air vehicle autonomous flight.Therefore in the miniature inertial navigation system at micro air vehicle, need to introduce the attitude information that error does not increase in time, come restraining inertial navigation system attitude error to disperse.The inhibition that the error of miniature inertial navigation system is dispersed is minitype inertial integrated navigation system key in application in micro air vehicle, the method of proposition in this regard mainly contains and utilizes accelerometer to adopt the mode of weighting as obliquity sensor both at home and abroad, merge with the attitude of little inertial navigation system as: in August, 2004, the master thesis of university is raised Birmingham by the U.S.: Design Of An Autopilot For Small Unmanned Aerial Vehicles, 7578; The inventor herein proposes to utilize modern optimal State Estimation method in 2006 the 6th phases applied science journal paper realization of the MEMSINS minute vehicle attitude and heading reference system " research "and Kalman filtering realizes that microinertial navigation system and accelerometer obtain the method that horizontal attitude merges etc.
The accelerometer obliquity sensor is influenced by carrier dynamically, the conventional thinking of existing document is that the use to the accelerometer obliquity sensor retrains, information according to Inertial Measurement Unit IMU and boat appearance system is judged the dynamic process of carrier, after the carrier dynamic process surpasses certain limit, no longer utilize the accelerometer obliquity sensor to carry out attitude measurement, because the micro air vehicle moment of inertia is little, maneuverability, be subjected to the influence of airflow serious, attitude range and rate of change are much larger than general unmanned plane, and this mode is not suitable for micro air vehicle.
Three, summary of the invention
Fundamental purpose of the present invention is, improve in the existing document deficiency to accelerometer obliquity sensor error processing method under the dynamic condition, under Navigation, Guidance and Control (GNC) closed loop condition, from the GNC loop, extract key feature information, exploration is adapted to the new way of micro air vehicle characteristics, suppresses the attitude error that is dynamically caused by the accelerometer obliquity sensor inflight of micro air vehicle.
Content of the present invention is that the little inertia combined navigation system of the MEMS of micro air vehicle utilizes guidance information in the micro aircraft GNC closed loop at its place to realize that attitude determines method, and its characteristics realize grinding by following steps:
(1) guidance algorithm of micro aircraft GNC closed loop resolves step: the position and the position of expecting way point current according to micro air vehicle, calculate the heading of expectation, the heading and the current flight direction of expectation are subtracted each other, the difference of gained multiply by a coefficient k, the micro air vehicle roll angle that obtains expecting, wherein the desirable scope of coefficient k is 0.2 to 0.5;
(2) MEMS Inertial Measurement Unit signals collecting step: gather the output signal of MEMS Inertial Measurement Unit, obtain the angular velocity and the specific force of micro air vehicle;
(3) the inertial navigation algorithm resolves step: the angular velocity that step (2) is collected and than force signal by the flow process of strap inertial navigation algorithm, calculates the navigation information of attitude, speed and the position of micro air vehicle;
(4) obtain the metrical information step of position and attitude: the position, speed and the course information that read GPS, utilize the ratio force signal that constantly collects recently simultaneously, the measuring principle of pressing the accelerometer obliquity sensor is calculated the roll angle and the angle of pitch of micro air vehicle;
(5) the selfadaptation setup procedure of Kalman filter observing matrix: the roll angle of the expectation that obtains with step (1) is an independent variable, by a piecewise function, the variance of horizontal attitude in the observation noise matrix of realtime computer card Thalmann filter, keep multiple relation in real time in the observation noise matrix of the Kalman filter of calculating between the variance of horizontal attitude and its initial value, this multiple is the dependent variable of piecewise function, and the concrete funtcional relationship of this piecewise function is determined by the flight test of reality;
(6) combined filter and correction step: utilize step (5) to adjust the Kalman filter of observing matrix parameter in real time, the metrical information that step (4) is obtained is handled, the error of attitude, speed and position that estimating step (2) inertial navigation calculates, in attitude, speed and position that step (2) obtains, deduct the error that this step estimates, and attitude, speed and the positional information of replacement step (2);
(7) feedback step of attitude, positional information: the attitude information that step (6) obtains is input to the control module of micro air vehicle, and control module changes by the attitude of the rudder face control micro air vehicle of micro air vehicle; Position and course information that step (6) obtains are input to the system guide module, are used for completing steps (1).
The present invention starts with in the angle of control closedloop system from navigating, guiding, based on Kalman filtering the average essence of reinforcement is arranged most, in conjunction with differentiation to the state of flight of micro air vehicle, the method of the observation noise matrix by the little inertia combined navigation system Kalman filter of dynamic adjustment, improve the precision of little inertia combined navigation system attitude, weaken the dynamic obliquity sensor inflight of micro air vehicle to the correction of system's attitude, weaken the dynamically attitude correction in Navigation, Guidance and Control loop equivalence interference inflight, improve the flight quality of micro air vehicle.
Four, description of drawings
Fig. 1 is minute vehicle Navigation, Guidance and Control system principle diagram.
Fig. 2 is a miniature Navigation, Guidance and Control of the present invention system closed loop block diagram.
Fig. 3 is the little inertia combined navigation system composition frame chart of MEMS among Fig. 2.
The roll angle curve synoptic diagram of micro air vehicle autonomous flight in Fig. 4 prior art.
Fig. 5 is the roll angle curve synoptic diagram of micro air vehicle autonomous flight of the present invention.
Five, embodiment
Principle of the present invention is:
In traditional GNC closed loop, navigational system provides navigational parameter to feed back to guidance and flight control algorithm, shown in the theory diagram that each module and solid line among Fig. 1 constitute.The present invention is then on the basis in traditional GNC loop, the feedforward control of increase from the system guide module to navigation subsystem, shown in the dotted line of Fig. 1, utilize guidance information, control combination Kalman filter noise matrix, realize of the selfadaptation adjustment of Kalman filter observation noise battle array with state of flight, by parameter adjustment, improve the adaptivity of little inertia combined navigation system to state of flight, realize that the attitude of existing miniaircraft inertial integrated navigation system under dynamic flying condition determine, improve the attitude accuracy under the dynamic flying condition and the stability of flight.
Determine for the attitude that realizes miniaircraft inertial integrated navigation system, adapt to dynamic flying condition, need finish the work:
(1) guidance algorithm of micro aircraft GNC closed loop resolves
Micro air vehicle calculates the course angle ψ of expectation according to the position of current position and expectation way point _{e}, the course angle ψ that the current actual measurement of expecting of course angle and micro navigation system is obtained subtracts each other, and the difference of gained multiply by a coefficient k, and the micro air vehicle roll angle that obtains expecting, k are than row coefficient, but span is 0.2 to 0.5.
With the micro air vehicle takeoff point is initial point, with the north orientation is x axle positive dirction, is y axle positive dirction with the east orientation, sets up local the earth horizontal coordinates, under this coordinate system, course angle turns to then course angle increase of right side to be 0 degree along x axle positive dirction, and the scope of course angle is (0,360) degree, the real time position coordinate of micro air vehicle is that (X, Y), the way point coordinate representation that current expectation is flown to is (P _{x}, P _{y}), then Qi Wang heading is calculated as follows:
As Px＞X, Py 〉=Y,
Work as Px=X, Py＞Y, ψ _{e}=pi/2;
As Px＜X,
Work as Px=X, Py＜Y, ψ _{e}=3 pi/2s;
As Px＞X, Py＜Y
Guidance algorithm is pressed the roll angle γ that following formula generates expectation _{e}: γ _{e}=k (ψ _{e}ψ),
Wherein ψ is the course angle that little inertia combined navigation system actual measurement obtains.
(2) MEMS Inertial Measurement Unit signals collecting step: gather the output signal of MEMS Inertial Measurement Unit, obtain the angular velocity and the specific force of micro air vehicle;
(3) the inertial navigation algorithm resolves
The angular velocity that utilization collects and than force signal by the flow process of strap inertial navigation algorithm, calculates the navigation informations such as attitude, speed and position of micro air vehicle.The initial attitude that inertial navigation resolves, speed and position are imported by the outside.
Body axis system is followed successively by angular velocity omega around the roll axle with respect to the component at three axles of body system of the angular velocity of local geographic coordinate system _{x}, around the angular velocity omega of pitch axis _{y}, around the angular velocity omega of azimuth axis _{z}, the computation period of inertial navigation is Δ t, three components of the angle delta θ that body turns in the Δ t time are followed successively by roll shaft angle increment Delta θ successively _{x}, pitch axis angle increment Δ θ _{y}, azimuth axis angle increment Δ θ _{z}, the angle that then turns over is expressed as with matrixstyle:
The attitude matrix of t carrier constantly is
So t+ Δ t constantly attitude matrix by
On the basis of attitude algorithm,, obtain the acceleration of motion a in the local geographic coordinate system according to current attitude and specific force _{N}, a _{E}, a _{D}, wherein, a _{N}Be the acceleration of motion of north orientation, a _{E}The acceleration of motion of east orientation, a _{D}Ground to acceleration of motion.The speed of micro air vehicle and position are obtained by the following formula recursion, V _{N}(t) speed of expression t moment micro air vehicle north orientation, V _{E}(t) speed of expression t moment micro air vehicle east orientation, V _{D}(t) expression t constantly micro air vehicle ground to speed, V _{N}(t+ Δ t), V _{E}(t+ Δ t), V _{D}(t+ Δ t) represent respectively t+ Δ t constantly north orientation, east orientation and ground to speed, R represents earth radius, with symbol latitude L _{I}, longitude λ _{I}With height h _{I}Represent the position of the micro air vehicle that obtained by the inertial navigation algorithm respectively, then the speed of inertial navigation and position can be calculated as follows.
(4) obtain the metrical information of position and attitude
Read position (the latitude L of GPS _{G}, longitude λ _{G}With height h _{G}) and course angle ψ _{G}, utilize the ratio force signal that constantly collects recently simultaneously, the measuring principle of pressing the accelerometer obliquity sensor is calculated the roll angle γ of micro air vehicle _{A}With pitching angle theta _{A}
(5) the selfadaptation setup procedure of Kalman filter observing matrix
The roll angle γ of the expectation that obtains with step (1) _{e}Be independent variable, by a piecewise function, the variance R of horizontal attitude in the observation noise matrix of t+1 realtime computer card Thalmann filter constantly _{T+1}, keeping multiple relation in real time in the observation noise matrix of the Kalman filter of calculating between the variance of horizontal attitude and its initial value, this multiple relation has parabolic, and this multiple is the dependent variable of piecewise function, and piecewise function has following form:
The concrete funtcional relationship of piecewise function is determined R by the flight test of reality _{0}Be the initial value of horizontal attitude variance, determine k according to the measurement noise size of accelerometer _{R}Be the public sector of the multiple relation of parabolic under the different conditions, obtain by test.
(6) combined filter and correction step
Utilize step (5) to adjust the Kalman filter of observing matrix parameter in real time, the metrical information that step (4) is obtained is handled, the error of attitude, speed and position that estimating step (2) inertial navigation calculates, in attitude, speed and position that step (2) obtains, deduct the error that this step estimates, and attitude, speed and the positional information of replacement step (2).
This Kalman filter is characterised in that:
The observation noise matrix of wave filter determines in real time that by step (5) state variable X comprises the north orientation platform error angle φ of strapdown inertial navitation system (SINS) _{N}, east orientation platform error angle φ _{E}With ground to platform error angle φ _{D}, north orientation velocity error δ v _{N}, east orientation velocity error δ v _{E}With ground to velocity error δ v _{D}, latitude error δ L, longitude error δ λ and height error δ h, totally 9, i.e. X=[φ _{N}φ _{E}φ _{D}δ v _{N}δ v _{E}δ v _{D}δ L δ λ δ h] ^{T}, wave filter is observed quantity with platform error angle and site error, observational variable is 6.
Handle as follows to the metrical information that step (4) is obtained:
If t+1 is the filtering moment, this moment is by GPS course angle ψ _{G}, the roll angle γ that obliquity sensor records _{A}And pitching angle theta _{A}, can obtain the attitude matrix C that attitude measurement information is determined _{N "} ^{b}
By attitude matrix C _{N "} ^{b}Three platform error angles of the mathematical platform of determining are designated as north orientation platform error angle φ _{NA}, east orientation platform error angle φ _{EA}, ground is to platform error angle φ _{DG}
Utilize this attitude matrix C of inertial navigation constantly _{n} ^{b}(t+1) transposition and C _{N "} ^{b}Multiply each other, obtain the mathematical platform error angle [φ of the attitude battle array correspondence of strapdown inertial navitation system (SINS) _{N}φ _{E}φ _{D}] ^{T}The attitude and the corresponding mathematical platform error angle [φ of attitude battle array that measure with step (4) _{NA}φ _{EA}φ _{DG}] ^{T}Between poor, promptly
The platform error angle observed quantity of Kalman filtering is
Matrix is measured in definition
Then the variance battle array of each filtering State Estimation constantly and the variance battle array of status predication can recursion obtain.Promptly at the filtering P of t constantly _{t}On the basis,, obtain the filtering variance battle array P of the status predication of t+1 constantly by following Recursive Filtering equation _{T+1t}, the filter gain matrix K _{T+1}, the state estimation value And the variance battle array P of State Estimation _{T+1}Can calculate by following formula.
Q _{t}＝Q _{0}
K _{t+1}＝P _{t+1t}H _{t+1} ^{T}(H _{t+1}P _{t+1t}H _{t+1} ^{T}+R _{t+1}) ^{1}
P _{t+1}＝(IK _{t+1}H _{t+1})P _{t+1t}
The state estimation value of aforementioned calculation Be the error of the resultant attitude that obtains of step (4) inertial navigation, speed and position, in the resulting result of step (4), deduct estimated value In the amount of correspondence, can improve the precision of the attitude of little inertia combined navigation system, reach the purpose of determining the micro air vehicle attitude.
(7) feedback step of attitude, positional information
The attitude information that step (6) obtains is input to the control module of micro air vehicle, and control module changes by the attitude of the rudder face control micro air vehicle of micro air vehicle; Position and course information that step (6) obtains are input to the system guide module, are used for completing steps (1).
Like this, the method of the observation noise matrix by the little inertia combined navigation system Kalman filter of dynamic adjustment, improve the precision of little inertia combined navigation system attitude, weaken the dynamic obliquity sensor inflight of micro air vehicle to the correction of system's attitude, weaken the dynamically attitude correction in Navigation, Guidance and Control loop equivalence interference inflight, improve the flight quality of micro air vehicle.
Among Fig. 1, solid arrow each module and that connect has been represented the basic logic connecting relation of Navigation, Guidance and Control closed loops, from the dotted arrow representative of the little inertia combined navigation system of sensing of drawing between micro air vehicle guidance algorithm module and the micro air vehicle flight control algorithm be expectation of the present invention roll angle transmission and utilize relation.
Fig. 2 is the further refinement to Fig. 1, micro air vehicle guidance algorithm module in position comparison module and the course control module pie graph 1, expectation pitching module, pitch control subsystem module and roll control module are formed the micro air vehicle flight control algoritic module among Fig. 1 jointly.Identical among dotted arrow and Fig. 1, shown the source and the transmission of information more clearly.
Fig. 3 is the further refinement of the little inertia combined navigation system of MEMS among Fig. 2, has shown the roll angle γ of the expectation of dotted line representative among Fig. 2 _{e}To little inertia combined navigation system Kalman filter observation noise matrix R _{T+1}The control adjustment.
Fig. 4 is the roll angle data and curves when adopting minute vehicle left side orbit under the prior art situation, transverse axis is represented the time, the longitudinal axis is represented roll angle, be in left in 315 seconds to 330 seconds always and spiral, the fluctuation range of roll angle2 is spent32 degree, the transversal wave movement scope is big inflight, has reached 30 degree.
Fig. 5 is after using the present invention, minute vehicle repeats left and right data and curves of spiraling, the flight course of curve reflection: after 80 seconds, 5 times spiral in a left side, and 4 times spiral in the right side, left and right spiraling mutually alternately, after using the present invention, in the orbit of a left side, the fluctuation range of roll angle is spent30 degree about18, and the fluctuation range 18 of roll angle is spent to 30 degree during right orbit.After using the present invention, fluctuation range obviously reduces, about 12 degree, and error has dwindled 60%.
Six, the effect of invention
The present invention starts with from the angle of Navigation, Guidance and Control closedloop system, and is average weighted based on the Kalman filtering optimum Essence is differentiated in conjunction with the state of flight to micro air vehicle, by dynamic adjustment minitype combined navigation system Kalman filter The method of the observation noise R of ripple device improves the precision of little inertia combined navigation system attitude, weakens micro air vehicle Obliquity sensor weakens the attitude correction equivalence in GNC loop and disturbs the observation correction of system's attitude in the dynamic flying, Improve flight quality.
The orbit test effect:
For the autonomous aloft little inertia combined navigation system of micro air vehicle, the micro air vehicle turning flight that spirals It is the worst state of flight that will often experience. Adopt in the orbit test before and after the present invention two groups of roll angles pair Than curve such as Fig. 4 and Fig. 5, from curve, can find out, before employing the present invention, in the orbit, micro air vehicle Attitude round excursion 30 degree of desired value, adopt the present invention after, the autonomous aloft appearance of micro air vehicle Less than 12 degree, fluctuation obviously reduces attitude round the fluctuation range of desired value, between the curve of flight and the desired value partially Difference obviously reduces, and error has dwindled 60%.
The present invention is directed to the measure error that obliquity sensor is introduced in the dynamic flying, utilize and dynamically adjust the Kalman filter sight Survey the method for noise, improve the performance of little inertia combined navigation system, reduce measure error, the raising micro air vehicle Degree of having a smooth flight.
A large amount of Flight Test results show: the present invention is applied to minute vehicle, consists of Navigation, Guidance and Control closed loop system System, minute vehicle can be realized autonomous attitude stabilization under the fitful wind environment of 56 meter per second and the navigation flight of way point, The autonomous flying quality of micro air vehicle has reached the level of expectation. The present invention has very strong engineering using value.
Claims (3)
Priority Applications (1)
Application Number  Priority Date  Filing Date  Title 

CN200710021401A CN101033973B (en)  20070410  20070410  Attitude determination method of miniaircraft inertial integrated navigation system 
Applications Claiming Priority (1)
Application Number  Priority Date  Filing Date  Title 

CN200710021401A CN101033973B (en)  20070410  20070410  Attitude determination method of miniaircraft inertial integrated navigation system 
Publications (2)
Publication Number  Publication Date 

CN101033973A true CN101033973A (en)  20070912 
CN101033973B CN101033973B (en)  20100519 
Family
ID=38730634
Family Applications (1)
Application Number  Title  Priority Date  Filing Date 

CN200710021401A CN101033973B (en)  20070410  20070410  Attitude determination method of miniaircraft inertial integrated navigation system 
Country Status (1)
Country  Link 

CN (1)  CN101033973B (en) 
Cited By (43)
Publication number  Priority date  Publication date  Assignee  Title 

CN100575877C (en) *  20071212  20091230  南京航空航天大学  Spacecraft shading device combined navigation methods based on many information fusion 
CN101319902B (en) *  20080718  20100908  哈尔滨工程大学  Lowcost combination type positioning and orienting device and combined positioning method 
CN101825468A (en) *  20100423  20100908  东南大学  Strapdown inertial navigation method of dual quaternion based on frequency domain analysis method 
CN101413800B (en) *  20080118  20100929  南京航空航天大学  Navigating and steady aiming method of navigation / steady aiming integrated system 
CN101950174A (en) *  20100930  20110119  清华大学  Method for adjusting parameters of unmanned aerial vehicle (UAV) controller 
CN101968353A (en) *  20100929  20110209  清华大学  Laser probing and image identification based terrain tracking method for unmanned helicopter 
CN101419080B (en) *  20080613  20110420  哈尔滨工程大学  Mini quickconnecting inertia measurement system zero speed correcting method 
CN102235862A (en) *  20100423  20111109  北京航空航天大学  Strapdown inertial navigation device based on micro mechanical gyroscopes 
CN102313546A (en) *  20110414  20120111  南京航空航天大学  Motion platform gesture sensing method based on polarized electromagnetic wave information chain 
CN102323990A (en) *  20110920  20120118  西安费斯达自动化工程有限公司  Pneumatic model for rigid body space motion 
CN102353377A (en) *  20110712  20120215  北京航空航天大学  High altitude long endurance unmanned aerial vehicle integrated navigation system and navigating and positioning method thereof 
CN102419168A (en) *  20110902  20120418  北京航空航天大学  Device for measuring and displaying horizontal angle of inclination of airplane and measurement and display method thereof 
CN102508986A (en) *  20110831  20120620  微迈森惯性技术开发（北京）有限公司  Method and system for tracing cascade rigid motion and walking processes 
CN102519461A (en) *  20111125  20120627  西北工业大学  Euler angle Walsh index approximate output method based on angular velocity 
CN102519462A (en) *  20111125  20120627  西北工业大学  Angular velocity based Euler angle exponent output method 
CN102607562A (en) *  20120412  20120725  南京航空航天大学  Micro inertial parameter adaptive attitude determination method based on carrier flight mode judgment 
CN102607593A (en) *  20120228  20120725  西安费斯达自动化工程有限公司  Aircraft attitude triangle correction model based on accelerometer 
CN102692225A (en) *  20110324  20120926  北京理工大学  Attitude heading reference system for lowcost small unmanned aerial vehicle 
CN102706360A (en) *  20120611  20121003  北京航空航天大学  Method utilizing optical flow sensors and rate gyroscope to estimate state of air vehicle 
CN102955477A (en) *  20121026  20130306  南京信息工程大学  Attitude control system and control method of fourrotor aircraft 
CN102997920A (en) *  20121211  20130327  东南大学  Optimization method of construction frequency domain strapdown inertial navigation attitude based on angular rate input 
CN103246366A (en) *  20120207  20130814  穆克波有限公司  Dynamic offset calibration 
CN103363993A (en) *  20130706  20131023  西北工业大学  Airplane angular rate signal reconstruction method based on unscented kalman filter 
CN103471594A (en) *  20130924  20131225  成都市星达通科技有限公司  Fine alignment algorithm based on AHRS (Attitude and Heading Reference System) 
CN104132662A (en) *  20140725  20141105  辽宁工程技术大学  Closedloop Kalman filter inertial positioning method based on zero velocity update 
CN104199440A (en) *  20140820  20141210  中国运载火箭技术研究院  Fourunit threebus redundancy heterogeneous GNC (guidance navigation control) system 
CN104769496A (en) *  20120817  20150708  展望机器人有限公司  Flying camera with string assembly for localization and interaction 
CN104792336A (en) *  20150331  20150722  北京航空航天大学  Measurement method and device of flying state 
CN104802697A (en) *  20150330  20150729  西北工业大学  Micro inertial measurement unit and measurement unit based adaptive headlamp control method 
CN104880189A (en) *  20150512  20150902  西安克拉克通信科技有限公司  Lowcost tracking antijamming method of antenna for satellite communication in motion 
CN105094138A (en) *  20150715  20151125  东北农业大学  Lowaltitude autonomous navigation system for rotarywing unmanned plane 
CN105115518A (en) *  20150728  20151202  中国运载火箭技术研究院  Inertial navigation system and GPS double antenna course deflection calibration method 
WO2015180171A1 (en) *  20140530  20151203  SZ DJI Technology Co., Ltd.  Aircraft attitude control methods 
CN105445763A (en) *  20140917  20160330  上海新跃仪表厂  Target reconstruction method based on trackingpointing information 
CN107063181A (en) *  20161223  20170818  北京航空航天大学  The measuring method and device of the level inclination of Multifunctional adjustment table under complex environment 
CN107704106A (en) *  20171017  20180216  宁波视睿迪光电有限公司  Attitude positioning method, device and electronic equipment 
CN108037658A (en) *  20171115  20180515  东莞市松迪智能机器人科技有限公司  A kind of method for correcting error of the robot kinematic error based on navigation system 
WO2018214227A1 (en) *  20170522  20181129  深圳市靖洲科技有限公司  Unmanned vehicle realtime posture measurement method 
WO2018214226A1 (en) *  20170522  20181129  深圳市靖洲科技有限公司  Unmanned vehicle realtime posture measurement method 
CN109000682A (en) *  20180727  20181214  中科宇达（北京）科技有限公司  The method of aerial coarse alignment based on rectilinear path 
US10197587B2 (en)  20120317  20190205  MCube Inc.  Device and method for using time rate of change of sensor data to determine device rotation 
US10324108B2 (en)  20120207  20190618  Mcube, Inc.  Dynamic offset correction for calibration of MEMS sensor 
CN110021931A (en) *  20190428  20190716  河海大学  It is a kind of meter and model uncertainty electric system assist predicted state estimation method 
Family Cites Families (3)
Publication number  Priority date  Publication date  Assignee  Title 

CN100356139C (en) *  20050121  20071219  清华大学  Miniature assembled gesture measuring system for minisatellite 
CN100381785C (en) *  20060327  20080416  北京航空航天大学  A Light small type intertia measuring unit 
CN100593689C (en) *  20060526  20100310  南京航空航天大学  Gasture estimation and interfusion method based on strapdown inertial nevigation system 

2007
 20070410 CN CN200710021401A patent/CN101033973B/en not_active IP Right Cessation
Cited By (68)
Publication number  Priority date  Publication date  Assignee  Title 

CN100575877C (en) *  20071212  20091230  南京航空航天大学  Spacecraft shading device combined navigation methods based on many information fusion 
CN101413800B (en) *  20080118  20100929  南京航空航天大学  Navigating and steady aiming method of navigation / steady aiming integrated system 
CN101419080B (en) *  20080613  20110420  哈尔滨工程大学  Mini quickconnecting inertia measurement system zero speed correcting method 
CN101319902B (en) *  20080718  20100908  哈尔滨工程大学  Lowcost combination type positioning and orienting device and combined positioning method 
CN102235862A (en) *  20100423  20111109  北京航空航天大学  Strapdown inertial navigation device based on micro mechanical gyroscopes 
CN101825468A (en) *  20100423  20100908  东南大学  Strapdown inertial navigation method of dual quaternion based on frequency domain analysis method 
CN101968353A (en) *  20100929  20110209  清华大学  Laser probing and image identification based terrain tracking method for unmanned helicopter 
CN101950174A (en) *  20100930  20110119  清华大学  Method for adjusting parameters of unmanned aerial vehicle (UAV) controller 
CN102692225A (en) *  20110324  20120926  北京理工大学  Attitude heading reference system for lowcost small unmanned aerial vehicle 
CN102692225B (en) *  20110324  20150311  北京理工大学  Attitude heading reference system for lowcost small unmanned aerial vehicle 
CN102313546B (en) *  20110414  20121226  南京航空航天大学  Motion platform gesture sensing method based on polarized electromagnetic wave information chain 
CN102313546A (en) *  20110414  20120111  南京航空航天大学  Motion platform gesture sensing method based on polarized electromagnetic wave information chain 
CN102353377A (en) *  20110712  20120215  北京航空航天大学  High altitude long endurance unmanned aerial vehicle integrated navigation system and navigating and positioning method thereof 
CN102508986A (en) *  20110831  20120620  微迈森惯性技术开发（北京）有限公司  Method and system for tracing cascade rigid motion and walking processes 
CN102508986B (en) *  20110831  20150930  微迈森惯性技术开发（北京）有限公司  A kind of cascade rigid motion tracking, gait processes method for tracing and system 
CN102419168A (en) *  20110902  20120418  北京航空航天大学  Device for measuring and displaying horizontal angle of inclination of airplane and measurement and display method thereof 
CN102419168B (en) *  20110902  20130918  北京航空航天大学  Device for measuring and displaying horizontal angle of inclination of airplane and measurement and display method thereof 
CN102323990A (en) *  20110920  20120118  西安费斯达自动化工程有限公司  Pneumatic model for rigid body space motion 
CN102323990B (en) *  20110920  20141119  西安费斯达自动化工程有限公司  Method for modeling pneumatic model for rigid body space motion 
CN102519461B (en) *  20111125  20140205  西北工业大学  Euler angle Walsh index approximate output method based on angular velocity 
CN102519462B (en) *  20111125  20131225  西北工业大学  Angular velocity based Euler angle exponent output method 
CN102519462A (en) *  20111125  20120627  西北工业大学  Angular velocity based Euler angle exponent output method 
CN102519461A (en) *  20111125  20120627  西北工业大学  Euler angle Walsh index approximate output method based on angular velocity 
CN103246366A (en) *  20120207  20130814  穆克波有限公司  Dynamic offset calibration 
US10324108B2 (en)  20120207  20190618  Mcube, Inc.  Dynamic offset correction for calibration of MEMS sensor 
CN103246366B (en) *  20120207  20171114  矽立科技有限公司  Dynamic offset calibration 
CN102607593B (en) *  20120228  20140423  西安费斯达自动化工程有限公司  Aircraft attitude triangle correction model based on accelerometer 
CN102607593A (en) *  20120228  20120725  西安费斯达自动化工程有限公司  Aircraft attitude triangle correction model based on accelerometer 
US10197587B2 (en)  20120317  20190205  MCube Inc.  Device and method for using time rate of change of sensor data to determine device rotation 
CN102607562B (en) *  20120412  20141029  南京航空航天大学  Micro inertial parameter adaptive attitude determination method based on carrier flight mode judgment 
CN102607562A (en) *  20120412  20120725  南京航空航天大学  Micro inertial parameter adaptive attitude determination method based on carrier flight mode judgment 
CN102706360A (en) *  20120611  20121003  北京航空航天大学  Method utilizing optical flow sensors and rate gyroscope to estimate state of air vehicle 
CN102706360B (en) *  20120611  20141126  北京航空航天大学  Method utilizing optical flow sensors and rate gyroscope to estimate state of air vehicle 
CN104769496B (en) *  20120817  20190315  展望机器人股份公司  Flight video camera with the rope component for positioning and interacting 
CN104769496A (en) *  20120817  20150708  展望机器人有限公司  Flying camera with string assembly for localization and interaction 
CN102955477B (en) *  20121026  20150114  南京信息工程大学  Attitude control system and control method of fourrotor aircraft 
CN102955477A (en) *  20121026  20130306  南京信息工程大学  Attitude control system and control method of fourrotor aircraft 
CN102997920A (en) *  20121211  20130327  东南大学  Optimization method of construction frequency domain strapdown inertial navigation attitude based on angular rate input 
CN102997920B (en) *  20121211  20160120  东南大学  Based on the structure frequency domain strapdown inertial navigation attitude optimization method of angular speed input 
CN103363993A (en) *  20130706  20131023  西北工业大学  Airplane angular rate signal reconstruction method based on unscented kalman filter 
CN103363993B (en) *  20130706  20160420  西北工业大学  A kind of aircraft angle rate signal reconstructing method based on Unscented kalman filtering 
CN103471594B (en) *  20130924  20160120  成都市星达通科技有限公司  Based on the fine alignment algorithm of AHRS 
CN103471594A (en) *  20130924  20131225  成都市星达通科技有限公司  Fine alignment algorithm based on AHRS (Attitude and Heading Reference System) 
EP3111286A4 (en) *  20140530  20170412  SZ DJI Technology Co., Ltd.  Aircraft attitude control methods 
US9958874B2 (en)  20140530  20180501  SZ DJI Technology Co., Ltd  Aircraft attitude control methods 
WO2015180171A1 (en) *  20140530  20151203  SZ DJI Technology Co., Ltd.  Aircraft attitude control methods 
CN104132662A (en) *  20140725  20141105  辽宁工程技术大学  Closedloop Kalman filter inertial positioning method based on zero velocity update 
CN104132662B (en) *  20140725  20200117  辽宁工程技术大学  Closed loop Kalman filtering inertial positioning method based on zerospeed correction 
CN104199440B (en) *  20140820  20170503  中国运载火箭技术研究院  Fourunit threebus redundancy heterogeneous GNC (guidance navigation control) system 
CN104199440A (en) *  20140820  20141210  中国运载火箭技术研究院  Fourunit threebus redundancy heterogeneous GNC (guidance navigation control) system 
CN105445763A (en) *  20140917  20160330  上海新跃仪表厂  Target reconstruction method based on trackingpointing information 
CN104802697A (en) *  20150330  20150729  西北工业大学  Micro inertial measurement unit and measurement unit based adaptive headlamp control method 
CN104802697B (en) *  20150330  20161116  西北工业大学  Micro inertial measurement unit and adaptive front lamp control method based on this measuring unit 
CN104792336A (en) *  20150331  20150722  北京航空航天大学  Measurement method and device of flying state 
CN104880189A (en) *  20150512  20150902  西安克拉克通信科技有限公司  Lowcost tracking antijamming method of antenna for satellite communication in motion 
CN104880189B (en) *  20150512  20190129  西安克拉克通信科技有限公司  A kind of antenna for satellite communication in motion low cost tracking antiinterference method 
CN105094138A (en) *  20150715  20151125  东北农业大学  Lowaltitude autonomous navigation system for rotarywing unmanned plane 
CN105115518B (en) *  20150728  20171222  中国运载火箭技术研究院  One kind is used for inertial navigation system and GPS double antennas course drift angle scaling method 
CN105115518A (en) *  20150728  20151202  中国运载火箭技术研究院  Inertial navigation system and GPS double antenna course deflection calibration method 
CN107063181B (en) *  20161223  20190319  北京航空航天大学  The measurement method and device of the level inclination of Multifunctional adjustment table under complex environment 
CN107063181A (en) *  20161223  20170818  北京航空航天大学  The measuring method and device of the level inclination of Multifunctional adjustment table under complex environment 
WO2018214226A1 (en) *  20170522  20181129  深圳市靖洲科技有限公司  Unmanned vehicle realtime posture measurement method 
WO2018214227A1 (en) *  20170522  20181129  深圳市靖洲科技有限公司  Unmanned vehicle realtime posture measurement method 
CN107704106A (en) *  20171017  20180216  宁波视睿迪光电有限公司  Attitude positioning method, device and electronic equipment 
CN108037658A (en) *  20171115  20180515  东莞市松迪智能机器人科技有限公司  A kind of method for correcting error of the robot kinematic error based on navigation system 
CN109000682B (en) *  20180727  20200317  中科宇达（北京）科技有限公司  Air coarse alignment method based on linear track 
CN109000682A (en) *  20180727  20181214  中科宇达（北京）科技有限公司  The method of aerial coarse alignment based on rectilinear path 
CN110021931A (en) *  20190428  20190716  河海大学  It is a kind of meter and model uncertainty electric system assist predicted state estimation method 
Also Published As
Publication number  Publication date 

CN101033973B (en)  20100519 
Similar Documents
Publication  Publication Date  Title 

CN103611324B (en)  A kind of unmanned helicopter flight control system and control method thereof  
Chee et al.  Control, navigation and collision avoidance for an unmanned aerial vehicle  
US8219267B2 (en)  Wind estimation for an unmanned aerial vehicle  
CN106708066B (en)  Viewbased access control model/inertial navigation unmanned plane independent landing method  
Wang et al.  Online highprecision probabilistic localization of robotic fish using visual and inertial cues  
Hong  Fuzzy logic based closedloop strapdown attitude system for unmanned aerial vehicle (UAV)  
Yun et al.  Testing and evaluation of an integrated GPS/INS system for small AUV navigation  
CN1270162C (en)  Attitude estimation in tiltable body using modified quaternion data representation  
CN103076017B (en)  Method for designing Mars entry phase autonomous navigation scheme based on observability degree analysis  
US10337883B2 (en)  Acceleration corrected attitude estimation systems and methods  
KR101320035B1 (en)  Location and pathmap generation data acquisition and analysis systems  
CN106249745B (en)  The control method of four axis unmanned planes  
Ahn et al.  Fast alignment using rotation vector and adaptive Kalman filter  
CN103776446B (en)  A kind of pedestrian's independent navigation computation based on double MEMSIMU  
CN102645933B (en)  Method for implementing flexible combined overload control for aircraft in large airspace  
CN101833338B (en)  Autonomous underwater vehicle vertical plane underactuated motion control method  
Salazar‐Cruz et al.  Embedded control system for a four‐rotor UAV  
Bevly  Global positioning system (GPS): A lowcost velocity sensor for correcting inertial sensor errors on ground vehicles  
US9715234B2 (en)  Multiple rotors aircraft and control method  
CN105929842A (en)  Underactuated UUV plane trajectory tracking control method based on dynamic speed adjustment  
US7355549B2 (en)  Apparatus and method for carrier phasebased relative positioning  
CN100585602C (en)  Inertial measuring system error model demonstration test method  
Bryson et al.  Vehicle model aided inertial navigation for a UAV using lowcost sensors  
JP4951061B2 (en)  System and method for automatically controlling airfoil flight of a drive wing  
Stančić et al.  The integration of strapdown INS and GPS based on adaptive error damping 
Legal Events
Date  Code  Title  Description 

PB01  Publication  
C06  Publication  
SE01  Entry into force of request for substantive examination  
C10  Entry into substantive examination  
GR01  Patent grant  
C14  Grant of patent or utility model  
CF01  Termination of patent right due to nonpayment of annual fee 
Granted publication date: 20100519 Termination date: 20160410 

CF01  Termination of patent right due to nonpayment of annual fee 