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Pseudo-range/pseudo-range rate tight integration method for inertial/satellite navigation system

A satellite navigation system and satellite positioning system technology, applied in the field of inertial/satellite integrated navigation, can solve the problems of normal positioning, loose position/speed of inertial/satellite integrated navigation system, etc.

Inactive Publication Date: 2013-03-27
BEIJING AUTOMATION CONTROL EQUIP INST
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  • Summary
  • Abstract
  • Description
  • Claims
  • Application Information

AI Technical Summary

Problems solved by technology

[0004] The technical problem to be solved by the present invention is that the inertial / satellite integrated navigation system cannot normally locate in a complex environment by using a loose combination of position / speed

Method used

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  • Pseudo-range/pseudo-range rate tight integration method for inertial/satellite navigation system
  • Pseudo-range/pseudo-range rate tight integration method for inertial/satellite navigation system
  • Pseudo-range/pseudo-range rate tight integration method for inertial/satellite navigation system

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Experimental program
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Effect test

Embodiment 1

[0377] Assuming that within a sampling period of 1 second, the inertial navigation position vector (φ Ins ,λ Ins , H Ins ) is (39.8157, 116.1548, 100), the inertial navigation velocity vector (V X，Ins , V Y，Ins , V Z，Ins ) is (-0.33, 0.26, 0); the current GPS position vector (φ GPS ,λ GPS , H GPS ) is (39.8128, 116.1542, 100), GPS velocity vector (V X，GPS , V Y，GPS , V Z，GPS ) is (0.001, -0.015, 0).

[0378] At present, the receiver has captured 4 GPS satellites in total, and the position and velocity vectors of each GPS satellite are expressed as Pseudorange and pseudorange rate are expressed as in,

[0379] The position and velocity of the first GPS satellite are

[0380] (-8714963, 20689154, 13285714, -188.2085, -1779.6547, 2615.1907)

[0381] The pseudorange and pseudorange rate of the first satellite

[0382] (19822240, -186.5034);

[0383] The position and velocity of the second GPS satellite are

[0384] (-23296408, 12000899, -3312671, 149.9774, -513...

Embodiment 2

[0416] The difference with embodiment 1 in this embodiment is:

[0417] Only use the pseudorange and pseudorange rate observations of the 1st to 3rd GPS satellites in Embodiment 1, according to the state model represented by formula (1)~formula (13), and calculate the observation matrix according to formula (59), then press Equation (29) calculates the estimated value of the available state variables as follows:

[0418] X=[ΔV x ΔV y Δφ Δλ Δk Ф x Ф y Ф z ε x ε y ε z δρ δρ v ] T

[0419] =[-0.098 0.423 0.0023 0 0 0 0 0 0 0 0 7559.5 3.735] T

Embodiment 3

[0421] The difference between this embodiment and the above two embodiments is:

[0422] Only utilize the pseudo-range and pseudo-range rate observations of the first and second GPS satellites in Embodiment 1, according to the state model represented by formula (1) ~ formula (13), and calculate the observation matrix according to formula (59), then press Equation (29) calculates the estimated value of the available state variables as follows:

[0423] X=[ΔV x ΔV y Δφ Δλ Δk Ф x Ф y Ф z ε x ε y ε z δρ δρ v ] T

[0424] =[-0.0983 0.42 0.0053 0 0 0 0 0 0 0 0 7559.52 3.751] T

[0425] Embodiment 2 and Embodiment 3 show that when the number of visible satellites is less than four, at least two, the method of the present invention can still obtain high estimation accuracy for state variables.

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Abstract

The invention belongs to the technical field of inertial / satellite integrated navigation, in particular to a pseudo-range / pseudo-range rate tight integration method for an inertial / satellite navigation system. The method includes: applying differences between the pseudo-range and pseudo-range rate measured by a global positioning system and the range information and rate information, related to a satellite and solved by inertial navigation, as observation; and estimating inertial errors and clock parameter of the global positioning system by Kalman filtering, and calibrating combinational output by output calibration. The method includes the steps of establishing a model; estimating system errors; adaptively adjusting parameters; and calibrating output. The technical problem that the inertial / satellite integrated navigation system with position / speed loose integration fails in normal positioning in complex environments is solved. The use of the pseudo-range / pseudo-range rate tight integration allows normal positioning for the inertial / satellite integrated navigation system in the complex environments.

Description

technical field [0001] The invention belongs to the technical field of inertial / satellite integrated navigation, in particular to an inertial / satellite navigation system pseudo-range / pseudo-range rate tight combination method. Background technique [0002] The inertial / satellite integrated navigation system has penetrated into the military and aerospace fields, and plays an important role in maintaining national security and promoting economic growth. [0003] Since the satellite navigation system is easily affected by the interference of the electromagnetic environment, complex terrain, trees, and buildings, the number of visible satellites is reduced, and it cannot be positioned normally, so that the inertial / satellite integrated navigation system is used in navigation such as electronic warfare battlefields, valleys, and riverbeds, oil and gas The use in complex environments such as surveying and blind guide services between urban buildings is greatly restricted. The trad...

Claims

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Application Information

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IPC IPC(8): G01S19/40G01C21/16
Inventor 李邦清王黎斌刘峰李文耀扈光锋谢仕民
Owner BEIJING AUTOMATION CONTROL EQUIP INST
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