CN110132286B - X-ray pulsar navigation method considering spacecraft dynamic effect and system deviation - Google Patents

X-ray pulsar navigation method considering spacecraft dynamic effect and system deviation Download PDF

Info

Publication number
CN110132286B
CN110132286B CN201910160452.9A CN201910160452A CN110132286B CN 110132286 B CN110132286 B CN 110132286B CN 201910160452 A CN201910160452 A CN 201910160452A CN 110132286 B CN110132286 B CN 110132286B
Authority
CN
China
Prior art keywords
observation
phase
spacecraft
doppler frequency
pulsar
Prior art date
Legal status (The legal status is an assumption and is not a legal conclusion. Google has not performed a legal analysis and makes no representation as to the accuracy of the status listed.)
Active
Application number
CN201910160452.9A
Other languages
Chinese (zh)
Other versions
CN110132286A (en
Inventor
薛梦凡
彭冬亮
郭云飞
申屠晗
骆吉安
陈志坤
Current Assignee (The listed assignees may be inaccurate. Google has not performed a legal analysis and makes no representation or warranty as to the accuracy of the list.)
Up Guangzhou Electronics Co ltd
Original Assignee
Hangzhou Dianzi University
Priority date (The priority date is an assumption and is not a legal conclusion. Google has not performed a legal analysis and makes no representation as to the accuracy of the date listed.)
Filing date
Publication date
Application filed by Hangzhou Dianzi University filed Critical Hangzhou Dianzi University
Priority to CN201910160452.9A priority Critical patent/CN110132286B/en
Publication of CN110132286A publication Critical patent/CN110132286A/en
Application granted granted Critical
Publication of CN110132286B publication Critical patent/CN110132286B/en
Active legal-status Critical Current
Anticipated expiration legal-status Critical

Links

Images

Classifications

    • G—PHYSICS
    • G01—MEASURING; TESTING
    • G01C—MEASURING DISTANCES, LEVELS OR BEARINGS; SURVEYING; NAVIGATION; GYROSCOPIC INSTRUMENTS; PHOTOGRAMMETRY OR VIDEOGRAMMETRY
    • G01C21/00—Navigation; Navigational instruments not provided for in groups G01C1/00 - G01C19/00
    • G01C21/24—Navigation; Navigational instruments not provided for in groups G01C1/00 - G01C19/00 specially adapted for cosmonautical navigation

Landscapes

  • Engineering & Computer Science (AREA)
  • Radar, Positioning & Navigation (AREA)
  • Remote Sensing (AREA)
  • Physics & Mathematics (AREA)
  • Astronomy & Astrophysics (AREA)
  • Automation & Control Theory (AREA)
  • General Physics & Mathematics (AREA)
  • Position Fixing By Use Of Radio Waves (AREA)
  • Navigation (AREA)

Abstract

The invention discloses an X-ray pulsar navigation method considering a spacecraft dynamic effect and system deviation, which uses a pulsar pulse phase, Doppler frequency and phase difference as navigation observed quantities, utilizes a spacecraft orbit dynamics model propagation orbit, uses a combined probability density function of photons reaching a spacecraft as a likelihood function to estimate a pair of phase offset and Doppler frequency caused by spacecraft initial position error and speed error, simultaneously uses the phase, Doppler frequency and phase difference of adjacent observation periods as navigation observed quantities, and updates the position and speed of the spacecraft through UKF. The influence of the dynamic effect of the spacecraft orbit can be effectively reduced by jointly estimating the phase and the Doppler frequency; the introduction of the phase difference observation quantity can effectively reduce the influence of the pulsar angular position error, the pulsar distance error and the satellite-borne atomic clock difference on the navigation precision.

Description

X-ray pulsar navigation method considering spacecraft dynamic effect and system deviation
Technical Field
The invention belongs to the field of autonomous navigation of spacecrafts, and relates to an X-ray pulsar navigation method considering dynamic effect and system deviation of a spacecraft.
Background
Deep space exploration has attracted great attention from countries in the world as an important feature and sign of the state scientific and technical development level and comprehensive national strength. At present, 3-dimensional navigation of a deep space probe is realized mainly by matching a very long baseline interferometry technology with ranging and Doppler velocity measurement, but the technology needs the probe to establish a communication link with a ground observation station, needs global station arrangement, has limited transmission data, and increases communication errors along with the increase of the distance between a spacecraft and the earth. Therefore, there is a need to improve autonomous navigation capabilities of spacecraft.
X-ray pulsar navigation (XPNAV) is a novel autonomous navigation mode, the navigation precision of the X-ray pulsar navigation is not influenced by the relative position of a spacecraft and a celestial body, and the X-ray pulsar navigation is the most potential deep space navigation mode in the future. However, the accuracy of the XPNAV is limited by the influence of the spacecraft orbit dynamic effect, and because the current measurement level is limited, there are a pulsar angular position error, a pulsar distance error and a satellite-borne atomic clock error, which all reduce the navigation accuracy of the XPNAV.
Disclosure of Invention
In order to solve the problems faced by XPNAV and improve the navigation precision, the invention provides an X-ray pulsar navigation method considering the dynamic effect of a spacecraft and the system deviation. The method uses the pulse phase, Doppler frequency and phase difference of pulsar as navigation observed quantity, uses a spacecraft orbit dynamics model propagation orbit, uses the joint probability density function of photon arrival to spacecraft as a likelihood function to estimate a pair of phase offset and Doppler frequency caused by spacecraft initial position error and speed error, and uses the phase, Doppler frequency and phase difference of adjacent observation period as navigation observed quantity to update the position and speed of spacecraft by UKF. The influence of the dynamic effect of the spacecraft orbit can be effectively reduced by jointly estimating the phase and the Doppler frequency; the introduction of the phase difference observation quantity can effectively reduce the influence of the pulsar angular position error, the pulsar distance error and the satellite-borne atomic clock difference on the navigation precision.
The method comprises the following steps:
step (1) establishing a state equation based on a spacecraft orbit dynamics model
Establishing a state equation based on a spacecraft orbit dynamics model by taking the position r and the speed v of a spacecraft in a solar system centroid coordinate system as state quantities to be estimated for navigation filtering
Figure GDA0002765393360000021
Wherein: x ═ rT vT]TF (-) represents an orbit dynamics function for a state vector of the spacecraft; w (t) is a disturbance noise vector, which is modeled as a zero-mean gaussian white noise process. And (3) carrying out integral calculation on the formula (1) to obtain the motion state of the spacecraft at any moment.
Step (2) calculating a phase value corresponding to photon arrival time by utilizing the estimated track information
Order to
Figure GDA0002765393360000022
Representing the photon arrival time sequence at the spacecraft over the observation period. Knowing the initial estimated position and speed, using the equation of state to propagate the orbit, and firstly obtaining all the positions of the spacecraft
Figure GDA0002765393360000023
Estimated position of time relative to solar system centroid SSB
Figure GDA0002765393360000024
Then calculate all
Figure GDA0002765393360000025
Phase value corresponding to time:
Figure GDA0002765393360000026
wherein: phi is aSSBIs a phase model of the pulsar at SSB; n is the direction vector of the pulsar; c is the speed of light;
Figure GDA0002765393360000027
Figure GDA0002765393360000028
and
Figure GDA0002765393360000029
annual parallax, Charulo retardation and Einstein retardation, respectively.
Step (3) performing joint estimation of pulse phase and Doppler frequency
In the time sequence of photon arrival
Figure GDA00027653933600000210
The joint probability density function is a likelihood function, a pulse phase delta p introduced by a position error of an observation initial moment and a Doppler frequency delta f introduced by a speed error of the observation initial moment are estimated by using a maximum likelihood method, and the cost function is as follows:
Figure GDA00027653933600000211
wherein alpha isSSBAnd betaSSBThe mean rates of arrival of pulsar radiation photons and X-ray background photons at the spacecraft, tsIs the initial observation time of the observation period, h [ [ alpha ] ]]Indicating the standard profile of the pulsar. The modeling and estimation of the Doppler frequency deltaf can eliminate the influence of the dynamic effect of the spacecraft orbit on the navigation precision to a great extent.
Step (4), establishing an observation equation of the navigation system
Let the observation period be [ t ]s,te]With teAnd (3) establishing an observation equation of the pulse phase by taking the moment as the state quantity updating moment:
Figure GDA0002765393360000031
wherein T isobs＝te-tsIs the filter period, up(te) In order to observe the noise in the phase,
Figure GDA0002765393360000032
i.e. the phase observations.
And (5) obtaining an observation equation of the Doppler frequency by derivation of the formula (4):
Figure GDA0002765393360000033
wherein u isf(te) For observing noise at Doppler frequency, frequency observation
Figure GDA0002765393360000034
Calculated from the following formula:
Figure GDA0002765393360000035
the phase difference of adjacent observation periods can counteract most of system deviation caused by pulsar angular position error, pulsar distance error and satellite-borne atomic clock difference, so that the influence of the system deviation on navigation precision can be weakened by introducing phase difference observation quantity. Measuring the phase observation quantity of the current observation period
Figure GDA0002765393360000036
Phase observation from previous observation period
Figure GDA0002765393360000037
Subtracting to obtain a phase difference observation equation:
Figure GDA0002765393360000038
step (5) performing fusion filtering by using an unscented Kalman filter
Writing a system state equation into a matrix form
Xk＝f(Xk-1)+Wk (8)
Wherein XkIs a system state vector, WkIs a disturbance noise vector. The pulse phase and doppler frequency observation equations are also written in matrix form:
Pk＝h1(Xk)+Vk (9)
Fk＝h2(Xk)+Uk (10)
wherein P iskAnd FkRespectively phase and Doppler frequency observation vectors, h1(Xk)、h2(Xk) Respectively, phase and Doppler frequency, VkAnd UkNoise is observed for phase and doppler frequencies, respectively. Satisfies the following conditions:
Figure GDA0002765393360000041
Figure GDA0002765393360000042
e (-) is the mean function.
Because the state vector X of the previous filtering period is used in the phase difference observation equationk-1The assumed condition of Kalman filtering is not satisfied, so that the state vector X of the current filtering period iskWritten as state vector estimate of previous filter cycle
Figure GDA0002765393360000043
With its error vector Δ Xk-1And further obtaining a phase difference observation equation:
Figure GDA0002765393360000044
wherein:
Figure GDA0002765393360000045
Γk＝-Hk-1△Xk-1+Vk-Vk-1 (13)
Dkis a phase difference observation vector. Satisfy the requirement of
Figure GDA0002765393360000046
Mean square error matrix of phase difference noiseComprises the following steps:
Figure GDA0002765393360000047
wherein Kk-1Is a gain matrix.
And (4) updating the position and the speed of the spacecraft by using the equation (8) as a state equation and the equations (9) to (11) as observation equations and by using standard unscented Kalman filtering.
The invention has the beneficial effects that:
the key point of the invention is to use the pulse phase, the Doppler frequency and the phase difference of adjacent observation periods as the observed quantity of the X-ray pulsar navigation. Wherein: the phase and the Doppler frequency are jointly estimated and simultaneously used as navigation observation quantities, so that the influence of the spacecraft orbit dynamic effect can be effectively reduced; the introduction of the phase difference observation quantity can effectively reduce the influence of the pulsar angular position error, the pulsar distance error and the satellite-borne atomic clock difference on the navigation precision.
Drawings
FIG. 1 is a flow chart of an X-ray pulsar navigation method of the present invention considering spacecraft dynamic effects and system biases
FIG. 2 is a schematic diagram of an X-ray pulsar navigation method considering spacecraft dynamic effects and system deviations according to the present invention
Detailed Description
The invention is further described with reference to specific examples.
In this embodiment, taking a high earth orbit satellite as an example, the following steps are specifically included in the navigation performed by using the proposed spacecraft autonomous navigation method based on the X-ray pulsar pulse phase/doppler frequency/phase difference joint observation, as shown in fig. 1 and 2:
step (1) establishing a state equation based on a spacecraft orbit dynamics model
Establishing a state equation based on a spacecraft orbit dynamics model by taking the position x and the speed epsilon of a spacecraft in a J2000.0 earth mass center inertial coordinate system as state quantities to be estimated by navigation filtering
Figure GDA0002765393360000051
Wherein:
Figure GDA0002765393360000052
modeling a zero mean Gaussian white noise process for a disturbance noise vector; the acceleration a of the spacecraft in an ECI coordinate system is
a＝aearth+asun+amoon+aJ2 (2)
Wherein: a isearthAcceleration of gravity of two bodies of the earth; a issunAnd amoonThree body gravitational accelerations of the sun and moon, respectively;
Figure GDA0002765393360000053
j for global non-spherical gravitational perturbation acceleration2An item.
Step (2) calculating a phase value corresponding to photon arrival time by utilizing the estimated track information
Order to
Figure GDA0002765393360000054
Representing the photon arrival time sequence at the spacecraft during the current filtering period. The filtering result of the above one filtering period is used as the initial position and speed, and the orbit is propagated by using the state equation 1 to obtain all the orbit information of the spacecraft
Figure GDA0002765393360000055
Estimated position of time
Figure GDA0002765393360000056
Then calculate all
Figure GDA0002765393360000057
Phase value corresponding to time:
Figure GDA0002765393360000058
wherein: phi is aSSBIs a phase model of the pulsar at SSB; n is the direction vector of the pulsar; c is the speed of light;
Figure GDA0002765393360000059
and
Figure GDA00027653933600000510
annual parallax, Charulo retardation and Einstein retardation, respectively.
Step (3) performing joint estimation of pulse phase and Doppler frequency
Estimating a pulse phase delta p introduced by an initial position error and a Doppler frequency delta f introduced by an initial speed error by using a maximum likelihood method, wherein a cost function is as follows:
Figure GDA00027653933600000511
wherein alpha isSSBAnd betaSSBThe mean rates of arrival of pulsar radiation photons and X-ray background photons at the spacecraft, tsIs the starting time of the current filter period, h [ [ alpha ] ]]Indicating the standard profile of the pulsar.
Step (4), establishing an observation equation of the navigation system
Let teAnd establishing an observation equation of the pulse phase as the end time of the current filtering period:
Figure GDA0002765393360000061
wherein T isobs＝te-tsIs the filter period, up(te) Noise is observed for phase.
Establishing a Doppler frequency observation equation:
Figure GDA0002765393360000062
wherein u isf(te) Observing noise for Doppler frequencies
Measuring the phase observation quantity of the current observation period
Figure GDA0002765393360000063
Phase observation from previous observation period
Figure GDA0002765393360000064
Subtracting to obtain a phase difference observation equation:
Figure GDA0002765393360000065
step (5) performing fusion filtering by using an unscented Kalman filter
Expressing the system state equation as a discretization matrix form
Xk＝f(Xk-1)+Wk (8)
Wherein XkIs a system state vector, WkF (-) represents an orbit dynamics function for the disturbance noise vector; the pulse phase, doppler frequency and phase difference observation equations for the selected navigation pulsar are also represented in the form of a matrix as follows:
Pk＝h1(Xk)+Vk (9)
Fk＝h2(Xk)+Uk (10)
Figure GDA0002765393360000066
wherein P isk、FkAnd DkRespectively phase, Doppler frequency and phase difference observation vectors, VkAnd UkObserving noise vectors, Γ, for phase and Doppler frequencies, respectivelyk＝-Hk-1△Xk-1+Vk-Vk-1For observation of phase differenceNoise vector, h1(Xk)、h2(Xk) Respectively the observed functions of phase and doppler frequency,
Figure GDA0002765393360000071
is the state vector estimated value of the previous filtering period
Figure GDA0002765393360000072
E (-) is the mean function, and the mean square error matrix of the phase difference noise is:
Figure GDA0002765393360000073
wherein Kk-1The filter gain matrix for the previous filter period.
And (3) taking the formula (8) as a state equation and the formulas (9) to (11) as observation equations, and obtaining the posterior state estimation value of the spacecraft in the current filtering period by using a standard unscented Kalman filtering algorithm
Figure GDA0002765393360000074
Error covariance matrix with a posteriori state estimates
Figure GDA0002765393360000075
Returning to the step (2), let k be k +1, ts＝te，te＝te+TobsRepeating the steps (2) to (5) to obtain the posterior state estimated value of the next filtering period
Figure GDA0002765393360000076
And its error covariance matrix
Figure GDA0002765393360000077
The above embodiments are not intended to limit the present invention, and the present invention is not limited to the above embodiments, and all embodiments are within the scope of the present invention as long as the requirements of the present invention are met.
Those skilled in the art will appreciate that the invention may be practiced without these specific details.

Claims (1)

1. An X-ray pulsar navigation method taking spacecraft dynamic effects and system deviations into consideration is characterized by comprising the following steps:
step (1) establishing a state equation based on a spacecraft orbit dynamics model
Establishing a state equation based on a spacecraft orbit dynamics model by taking the position r and the speed v of a spacecraft in a solar system centroid coordinate system as state quantities to be estimated for navigation filtering
Figure FDA0002765393350000011
Wherein: x ═ rT vT]TF (-) represents an orbit dynamics function for a state vector of the spacecraft; w (t) is a disturbance noise vector, which is modeled as a zero-mean Gaussian white noise process; integral calculation is carried out on the formula (1), and the motion state of the spacecraft at any moment can be obtained;
step (2) calculating a phase value corresponding to photon arrival time by utilizing the estimated track information
Order to
Figure FDA0002765393350000012
Representing a photon arrival time sequence at the spacecraft over an observation period; knowing the initial estimated position and speed, using the equation of state to propagate the orbit, and firstly obtaining all the positions of the spacecraft
Figure FDA0002765393350000013
Estimated position of time relative to solar system centroid SSB
Figure FDA0002765393350000014
Then calculate all
Figure FDA0002765393350000015
Phase value corresponding to time:
Figure FDA0002765393350000016
wherein: phi is aSSBIs a phase model of the pulsar at SSB; n is the direction vector of the pulsar; c is the speed of light;
Figure FDA0002765393350000017
and
Figure FDA0002765393350000018
annual parallax, Charulo retardation and Einstein retardation, respectively;
step (3) performing joint estimation of pulse phase and Doppler frequency
In the time sequence of photon arrival
Figure FDA0002765393350000019
The joint probability density function is a likelihood function, a pulse phase delta p introduced by a position error of an observation initial moment and a Doppler frequency delta f introduced by a speed error of the observation initial moment are estimated by using a maximum likelihood method, and the cost function is as follows:
Figure FDA00027653933500000110
wherein alpha isSSBAnd betaSSBThe mean rates of arrival of pulsar radiation photons and X-ray background photons at the spacecraft, tsIs the initial observation time of the observation period, h [ [ alpha ] ]]A standard profile representing a pulsar;
step (4), establishing an observation equation of the navigation system
Let the observation period be [ t ]s,te]With teAnd (3) establishing an observation equation of the pulse phase by taking the moment as the state quantity updating moment:
Figure FDA0002765393350000021
wherein T isobs＝te-tsIs the filter period, up(te) In order to observe the noise in the phase,
Figure FDA0002765393350000022
namely the phase observed quantity;
and (5) obtaining an observation equation of the Doppler frequency by derivation of the formula (4):
Figure FDA0002765393350000023
wherein u isf(te) For observing noise at Doppler frequency, frequency observation
Figure FDA0002765393350000024
Calculated from the following formula:
Figure FDA0002765393350000025
measuring the phase observation quantity of the current observation period
Figure FDA0002765393350000026
Phase observation from previous observation period
Figure FDA0002765393350000027
Subtracting to obtain a phase difference observation equation:
Figure FDA0002765393350000028
step (5) performing fusion filtering by using an unscented Kalman filter
Writing a system state equation into a matrix form
Xk＝f(Xk-1)+Wk (8)
Wherein XkIs a system state vector, WkIs a disturbance noise vector; the pulse phase and doppler frequency observation equations are also written in matrix form:
Pk＝h1(Xk)+Vk (9)
Fk＝h2(Xk)+Uk (10)
wherein P iskAnd FkRespectively phase and Doppler frequency observation vectors, h1(Xk)、h2(Xk) Respectively, phase and Doppler frequency, VkAnd UkPhase and doppler frequency observation noise, respectively; satisfies the following conditions:
Figure FDA0002765393350000031
Figure FDA0002765393350000032
e (-) is a mean function;
the state vector X of the current filtering periodkWritten as state vector estimate of previous filter cycle
Figure FDA0002765393350000033
With its error vector Δ Xk-1And further obtaining a phase difference observation equation:
Figure FDA0002765393350000034
wherein:
Figure FDA0002765393350000035
Γk＝-Hk-1△Xk-1+Vk-Vk-1 (13)
Dka phase difference observation vector is obtained; satisfy the requirement of
Figure FDA0002765393350000036
The mean square error matrix of the phase difference noise is:
Figure FDA0002765393350000037
wherein Kk-1Is a gain matrix;
and (4) updating the position and the speed of the spacecraft by using the equation (8) as a state equation and the equations (9) to (11) as observation equations and by using standard unscented Kalman filtering.
CN201910160452.9A 2019-03-04 2019-03-04 X-ray pulsar navigation method considering spacecraft dynamic effect and system deviation Active CN110132286B (en)

Priority Applications (1)

Application Number Priority Date Filing Date Title
CN201910160452.9A CN110132286B (en) 2019-03-04 2019-03-04 X-ray pulsar navigation method considering spacecraft dynamic effect and system deviation

Applications Claiming Priority (1)

Application Number Priority Date Filing Date Title
CN201910160452.9A CN110132286B (en) 2019-03-04 2019-03-04 X-ray pulsar navigation method considering spacecraft dynamic effect and system deviation

Publications (2)

Publication Number Publication Date
CN110132286A CN110132286A (en) 2019-08-16
CN110132286B true CN110132286B (en) 2021-01-12

Family

ID=67568502

Family Applications (1)

Application Number Title Priority Date Filing Date
CN201910160452.9A Active CN110132286B (en) 2019-03-04 2019-03-04 X-ray pulsar navigation method considering spacecraft dynamic effect and system deviation

Country Status (1)

Country Link
CN (1) CN110132286B (en)

Families Citing this family (7)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
CN110686702B (en) * 2019-09-27 2022-10-11 中国科学院上海光学精密机械研究所 Pulsar photon flow simulation system and method based on light field high-order correlation
CN110940332B (en) * 2019-11-19 2021-06-01 杭州电子科技大学 Pulsar signal phase delay estimation method considering spacecraft orbit dynamic effect
CN112394223B (en) * 2020-11-10 2022-04-01 南京大学 Joint estimation method for signal component frequency and initial phase
CN113074741B (en) * 2021-03-18 2023-04-25 中国人民解放军火箭军工程大学 Augmentation state algorithm for pulsar azimuth error estimation
CN113375697B (en) * 2021-06-23 2022-09-23 西安电子科技大学 X-ray pulsar signal time delay estimation method based on maximum posterior estimation
CN114608587B (en) * 2022-03-16 2023-02-03 中国人民解放军国防科技大学 Template-independent single-pulse-satellite spacecraft orbit determination method
CN117308970B (en) * 2023-11-28 2024-01-30 北京航空航天大学 Double-differential X-ray pulsar navigation method based on phase and Doppler frequency shift

Citations (4)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
CN104296755A (en) * 2014-10-23 2015-01-21 中国空间技术研究院 Determination method of pulse TOA of X-ray pulsar-based navigation
CN107024212A (en) * 2017-06-22 2017-08-08 北京航空航天大学 A kind of astronomical doppler combined navigation method of deep space probe X-ray pulsar/time difference
CN107328409A (en) * 2017-07-28 2017-11-07 北京控制工程研究所 A kind of X-ray pulsar navigation method based on dynamic pulse accumulation window
CN107421533A (en) * 2017-06-22 2017-12-01 北京航空航天大学 A kind of deep space probe X-ray pulsar TOA/DTOA Combinated navigation methods

Family Cites Families (1)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
US10935381B2 (en) * 2016-09-16 2021-03-02 The Charles Stark Draper Laboratory, Inc. Star tracker-aided airborne or spacecraft terrestrial landmark navigation system

Patent Citations (4)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
CN104296755A (en) * 2014-10-23 2015-01-21 中国空间技术研究院 Determination method of pulse TOA of X-ray pulsar-based navigation
CN107024212A (en) * 2017-06-22 2017-08-08 北京航空航天大学 A kind of astronomical doppler combined navigation method of deep space probe X-ray pulsar/time difference
CN107421533A (en) * 2017-06-22 2017-12-01 北京航空航天大学 A kind of deep space probe X-ray pulsar TOA/DTOA Combinated navigation methods
CN107328409A (en) * 2017-07-28 2017-11-07 北京控制工程研究所 A kind of X-ray pulsar navigation method based on dynamic pulse accumulation window

Non-Patent Citations (3)

* Cited by examiner, † Cited by third party
Title
An improved augmented X-ray pulsar navigation algorithm based on the norm of pulsar direction error;Xu Qiang,etc;《ADVANCES IN SPACE RESEARCH》;20181231;第62卷(第11期);第3187-3198页 *
X-ray pulsar-based navigation using pulse phase and Doppler frequency measurements;Xue Mengfan,etc;《SCIENCE CHINA-INFORMATION SCIENCES》;20151231;第58卷(第12期);第1-14页 *
X射线脉冲星导航的时延和频偏参量估计;贺珍妮，等;《装甲兵工程学院学报》;20130430;第27卷(第2期);第92-96页 *

Also Published As

Publication number Publication date
CN110132286A (en) 2019-08-16

Similar Documents

Publication Publication Date Title
CN110132286B (en) X-ray pulsar navigation method considering spacecraft dynamic effect and system deviation
Fang et al. Predictive iterated Kalman filter for INS/GPS integration and its application to SAR motion compensation
CN103487820B (en) A kind of vehicle-mounted strapdown/satellite tight integration seamless navigation method
CN110514203B (en) Underwater integrated navigation method based on ISR-UKF
CN110940333B (en) Deep space probe angle measurement and time delay combined navigation method based on online estimation
CN102679985A (en) Spacecraft constellation decentralized autonomous navigation method using inter-satellite tracking
CN110006427B (en) BDS/INS tightly-combined navigation method in low-dynamic high-vibration environment
CN106597507A (en) High-precision rapid filtering and smoothing algorithm of GNSS/SINS tight combination
CN106997061B (en) A method of gravitational field inversion accuracy is improved based on relative velocity between disturbance star
CN103542853B (en) The absolute Navigation method of a kind of estimated acceleration meter drift
CN113375659A (en) Pulsar navigation method based on starlight angular distance measurement information
CN115356754A (en) Combined navigation positioning method based on GNSS and low-orbit satellite
CN111487660B (en) High-precision real-time micro-nano satellite cluster navigation method
CN113008229A (en) Distributed autonomous integrated navigation method based on low-cost vehicle-mounted sensor
CN107421533B (en) A kind of deep space probe X-ray pulsar TOA/DTOA Combinated navigation method
Gong et al. Precise real-time navigation of the small TJU-1 satellite using GPS, GLONASS and BDS
Abd-Allah Assessment of GOCE models in Egypt
CN110554443A (en) Method for determining earth gravity field based on carrier phase observed value and point acceleration method
CN107830856B (en) Formation-flight-oriented solar TDOA (time difference of arrival) measurement method and integrated navigation method
CN111290008A (en) Dynamic self-adaptive extended Kalman filtering fault-tolerant algorithm
Bagci et al. Integrated NRM/EKF for LEO satellite GPS based orbit determination
Chen et al. Autonomous navigation for lunar final approach based on gravity gradient measurements
Liu et al. Integrated navigation model based on TDCP constrained algorithm
CN111551172A (en) Lunar probe angle measurement/time differential distance/speed combined navigation method
Fang et al. Integrating SINS sensors with odometer measurements for land vehicle navigation system

Legal Events

Date Code Title Description
PB01 Publication
PB01 Publication
SE01 Entry into force of request for substantive examination
SE01 Entry into force of request for substantive examination
GR01 Patent grant
GR01 Patent grant
TR01 Transfer of patent right

Effective date of registration: 20210719

Address after: 230000 B-1014, 10 floor, business office building, Wo Ye garden, Shushan District, Hefei, Anhui.

Patentee after: HEFEI WISDOM LOTUT INTELLECTUAL PROPERTY Co.,Ltd.

Address before: 310018 No. 2 street, Xiasha Higher Education Zone, Hangzhou, Zhejiang

Patentee before: HANGZHOU DIANZI University

TR01 Transfer of patent right
CP03 Change of name, title or address

Address after: 230000 floor 1, building 2, phase I, e-commerce Park, Jinggang Road, Shushan Economic Development Zone, Hefei City, Anhui Province

Patentee after: Dragon totem Technology (Hefei) Co.,Ltd.

Address before: 230000 B-1014, 10 floor, business office building, Wo Ye garden, Shushan District, Hefei, Anhui.

Patentee before: HEFEI WISDOM LOTUT INTELLECTUAL PROPERTY Co.,Ltd.

CP03 Change of name, title or address
TR01 Transfer of patent right

Effective date of registration: 20230707

Address after: Floor 2, Building 1, No. 7, Chaosheng Road, Xinya Street, Huadu District, Guangzhou, Guangdong 510000

Patentee after: UP (GUANGZHOU) ELECTRONICS Co.,Ltd.

Address before: 230000 floor 1, building 2, phase I, e-commerce Park, Jinggang Road, Shushan Economic Development Zone, Hefei City, Anhui Province

Patentee before: Dragon totem Technology (Hefei) Co.,Ltd.

TR01 Transfer of patent right