CN110703355B - Calibration method and device of satellite-borne accelerometer - Google Patents
Calibration method and device of satellite-borne accelerometer Download PDFInfo
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Abstract
The invention discloses a calibration method and device of a satellite-borne accelerometer, and belongs to the technical field of satellite gravity detection. The calibration method comprises the following steps: firstly, calculating to obtain the motion acceleration of a carrier by a carrier phase difference method; secondly, determining a non-conservative force calibration reference value according to the difference value of the motion acceleration of the carrier and the conservative force acceleration; and then, establishing a calibration model for the non-conservative force calibration reference value and the non-conservative force observation value, and accurately calibrating the satellite-borne accelerometer. The invention calculates the non-conservative force calibration reference value based on the GPS carrier phase differential method and the Newton second motion law, establishes the calibration model with the non-conservative force observation value to estimate the calibration parameter, and can effectively avoid the error influence of the empirical model method which highly depends on the atmospheric density model and the radiation coefficient, thereby improving the calibration precision of the satellite-borne accelerometer.
Description
Technical Field
The invention relates to the technical field of satellite gravity detection, in particular to a calibration method and device of a satellite-borne accelerometer.
Background
A high-precision three-axis satellite-borne accelerometer is a key load of a new generation of gravity measurement satellite, and can directly measure non-conservative force acting on the satellite. The accelerometer has scale factors and deviation factors due to the influence of space environment in the satellite operation process, so that the research on the calibration method of the satellite-borne accelerometer has important scientific significance for realizing the precise orbit determination and recovering the high-precision high-resolution earth gravitational field model.
However, when the conventional empirical model method is used for calibrating the satellite-borne accelerometer, firstly, an atmospheric density model and an emissivity value with high known precision are required, and then an empirical physical model is established for non-conservative forces such as atmospheric resistance, sunlight pressure and earth radiation pressure in sequence. However, the current mathematical model cannot meet the expected target, so that the atmospheric resistance model at the position along the track has large error and influences the calibration precision. In addition, in the integral method for simultaneously estimating the accelerometer calibration parameters and the earth gravity field parameters during orbit determination and gravity field inversion, the accelerometer calibration parameters and the gravity field parameters generate a coupling effect due to no direct standard, and the prior constraint information is relied on, so that the traditional calibration method of the satellite-borne accelerometer has the problem of limited precision.
Disclosure of Invention
The present invention is directed to a calibration method and apparatus for a satellite-borne accelerometer, so as to solve the problems mentioned in the background art.
In order to achieve the above purpose, the embodiments of the present invention provide the following technical solutions:
a calibration method of a satellite-borne accelerometer comprises the following steps:
acquiring carrier phase original data, geometric orbit original data, accelerometer original data and satellite attitude original data;
calculating the carrier phase original data by using a carrier phase differential method to obtain carrier phase acceleration;
calculating to obtain the carrier motion acceleration of the low-orbit gravity satellite under the inertial System according to the carrier phase acceleration and by combining the geometric relationship between a Global Positioning System (GPS) satellite and the low-orbit gravity satellite;
according to the original data of the geometric orbit and the perturbation force model, calculating to obtain the conservative force acceleration of the low-orbit gravity satellite;
determining a non-conservative force calibration reference value according to the difference value of the motion acceleration of the carrier and the conservative force acceleration;
processing the original data of the satellite attitude by using an interpolation method to obtain an attitude continuous observation value, and calculating to obtain a nonconservative force observation value of the low-orbit gravity satellite under an inertial system according to the attitude continuous observation value and the original data of the accelerometer;
and establishing a calibration model of the satellite-borne accelerometer through the non-conservative force calibration reference value and the non-conservative force observation value.
In an embodiment of the present invention, a calculation formula of carrier phase acceleration is as follows:
in the formula (I), the compound is shown in the specification,
and
respectively representing carrier phase velocity and carrier phase acceleration;
representing the original observed value of the carrier phase without an ionized layer; e.g. of the type T And f T Respectively representing corresponding first-order and second-order difference operators;
representing a carrier phase ionosphere-free combination value; c represents the speed of light; dt P Representing a GPS satellite clock error; rel represents a relativistic effect; ω represents the antenna phase winding error;
representing the geometric distance between the GPS satellite and the low-orbit gravity satellite; dt G Representing the receiver clock error of a low-orbit gravity satellite; n represents the carrier phase integer ambiguity; ε represents the observed random error.
In another preferred scheme adopted by the embodiment of the invention, the calculation formula of the motion acceleration of the carrier is as follows:
in the formula (I), the compound is shown in the specification,
the unit vector of the sight line direction from the GPS satellite to the low-orbit gravity satellite is represented;
and
respectively representing the sight line speed and the sight line acceleration;
and
respectively representing the first difference and the second difference of the clock error of the low-orbit gravity satellite receiver;
and
respectively representing the speed and acceleration of the GPS satellite;
and
respectively representing the motion speed and the motion acceleration of the carrier of the low-orbit gravity satellite.
In another preferred embodiment of the present invention, a calibration formula of the calibration model is:
in the formula (I), the compound is shown in the specification,
representing a conservative force acceleration value;
represents a non-conservative force observation;
representing a scale factor;
representing a deviation factor; (s) x ,s y ,s z ) Respectively representing three components of a scale factor; (b) x ,b y ,b z ) Representing the deviation factor three components, respectively.
The embodiment of the invention also provides a calibration device of the satellite-borne accelerometer, which comprises:
the data acquisition module is used for acquiring carrier phase original data, geometric orbit original data, accelerometer original data and satellite attitude original data;
the carrier phase acceleration calculation module is used for calculating the carrier phase original data by utilizing a carrier phase differential method to obtain carrier phase acceleration; the calculation formula of the carrier phase acceleration calculation module is the same as that of the carrier phase acceleration calculation module;
the carrier motion acceleration calculation module is used for calculating the carrier motion acceleration of the low-orbit gravity satellite under the inertial system according to the carrier phase acceleration and by combining the geometric relationship between the GPS satellite and the low-orbit gravity satellite; the calculation formula of the carrier motion acceleration calculation module is the same as that of the carrier motion acceleration calculation module;
the conservative force acceleration calculation module is used for calculating the conservative force acceleration of the low-orbit gravity satellite according to the original data of the geometric orbit and the perturbation force model;
the calibration reference value determining module is used for determining a non-conservative force calibration reference value according to the difference value of the motion acceleration of the carrier and the conservative force acceleration;
the non-conservative force observation value calculation module is used for processing the original data of the satellite attitude by using an interpolation method to obtain an attitude continuous observation value and calculating a non-conservative force observation value of the low-orbit gravity satellite in the inertial system according to the attitude continuous observation value and the original data of the accelerometer;
the calibration module is used for establishing a calibration model of the satellite-borne accelerometer through the non-conservative force calibration reference value and the non-conservative force observation value; the calibration formula of the calibration module is the same as the calibration formula of the calibration model described above.
Compared with the prior art, the technical scheme provided by the embodiment of the invention has the following technical effects:
according to the embodiment of the invention, the non-conservative force calibration reference value is calculated based on the GPS carrier phase differential method and the Newton second motion law, and the calibration model estimation calibration parameter is established on the non-conservative force observation value, so that the influence of the empirical model method on high dependence on the atmospheric density model and the radiation coefficient error can be effectively avoided, and the calibration precision of the satellite-borne accelerometer can be improved. Compared with the traditional method, the calibration method provided by the embodiment of the invention has the advantages that the principle is simple and visual, and the calibration precision is stable.
Drawings
Fig. 1 is a schematic structural diagram of a calibration apparatus for a satellite-borne accelerometer provided in embodiment 2.
Detailed Description
The technical means of the present invention will be described in further detail with reference to the embodiments.
Example 1
The embodiment provides a calibration method of a satellite-borne accelerometer, which comprises the following steps:
s1, acquiring carrier phase original data, geometric orbit original data, accelerometer original data and satellite attitude original data; in particular, these data sources include challenge Minisatellite payload (CHAMP), earth Gravity Field Recovery and Climate experiments (Gravity Recovery and Climate experiments, grain), earth Gravity Field and Ocean circulation exploration (Gravity Field and Gravity-State area circulation n Explorer, GOCE), earth Gravity Field Recovery and Follow-up of Climate experiments (grain low On, grain-FO-satellite), and other low-rail Gravity satellite raw data sets related to GPS carrier phase observations, geometric orbit observations, simplified dynamic orbit observations, satellite-borne accelerometer observations, satellite attitude observations, and the like. In addition, the observation value data of the satellite-borne accelerometer in the original data of the accelerometer needs to be preprocessed after being acquired, wherein the preprocessing comprises two parts of temperature correction and air injection attitude adjustment: the observation value of the satellite-borne accelerometer is sensitive to temperature, particularly in the normal direction; the temperature control system has obvious influence on the accelerometer calibration deviation factor, and can be corrected through a linear model. In addition, a plurality of burrs exist in the observation value of the satellite-borne accelerometer, which are mainly caused by the action of thrust pulse force generated by the satellite attitude-adjusting ignition jet, so that the calibration deviation factor of the accelerometer is influenced.
Since the orbit data to be distributed is expressed in the earth-fixed system and the calibration model is constructed in the inertial system, the data needs to be subjected to coordinate transformation, and in actual implementation, the data can be realized by using an origin conversion method based on celestial sphere almanac, that is, by sequentially using rotation matrices such as celestial sphere intermediate polar motion, earth rotation, and polar motion.
S2, calculating the carrier phase original data by using a carrier phase differential method to obtain carrier phase acceleration; the calibration of the satellite-borne accelerometer is realized based on a carrier phase differential method, an atmospheric density model and radiation coefficient information are not required to be provided, the priori constraint is not relied on, the algorithm principle is simple and intuitive, and the calibration precision is stable and reliable. And calculating the motion acceleration of the low-orbit satellite carrier by a GPS carrier phase differential method, and deducting all conservative force accelerations established by the perturbation force model to obtain a non-conservative force calibration reference value. Specifically, firstly, carrying out error model correction on an original observation value without an ionized layer of a carrier phase, and simultaneously interpolating GPS satellite clock error information by using an International GPS Service (IGS) precision clock error product to meet time synchronization so as to obtain a carrier phase non-ionized layer combination value; and then carrying out coordinate transformation and quadratic difference on the carrier phase acceleration value to calculate the carrier phase acceleration value. The calculation formula of the carrier phase acceleration is as follows:
in the formula (I), the compound is shown in the specification,
and
respectively representing carrier phase velocity and carrier phase acceleration;
representing the original observed value of the carrier phase without an ionized layer; e.g. of the type T And f T Respectively representing corresponding first-order and second-order difference operators;
representing a carrier phase ionosphere free combination value; c represents the speed of light; dt P Representing a GPS satellite clock error; rel represents the relativistic effect(ii) a ω represents the antenna phase winding error;
representing the geometric distance between the GPS satellite and the low-orbit gravity satellite; dt G Representing the receiver clock error of a low-orbit gravity satellite; n represents the carrier phase integer ambiguity; ε represents the observed random error.
S3, calculating to obtain the carrier motion acceleration of the low-orbit gravity satellite under the inertial system according to the carrier phase acceleration and by combining the geometric relationship between the GPS satellite and the low-orbit gravity satellite; specifically, the motion acceleration of the low-orbit gravity satellite carrier under the inertial system is determined according to a line-of-sight acceleration relation between the GPS satellite and the low-orbit gravity satellite and the GPS satellite position information calculated by IGS precise ephemeris, and the calculation formula of the motion acceleration of the carrier is as follows:
in the formula (I), the compound is shown in the specification,
the unit vector of the sight line direction from the GPS satellite to the low-orbit gravity satellite is represented;
and
respectively representing the sight line speed and the sight line acceleration;
and
respectively representing the clock error of low-orbit gravity satellite receiverDifference and second order difference values;
and
respectively representing the speed and acceleration of the GPS satellite;
and
respectively representing the motion speed and the motion acceleration of the carrier of the low-orbit gravity satellite.
S4, calculating to obtain conservative force acceleration of the low-orbit gravity satellite according to the original data of the geometric orbit and the perturbation force model; specifically, rough difference elimination, interpolation and time synchronization processing are carried out on original data of the geometric orbit by using the simplified dynamic orbit to obtain a continuous high-precision geometric orbit, coordinate transformation is carried out on the continuous high-precision geometric orbit, then, dynamic modeling is respectively carried out on the lunisolar gravity, the solid tide, the sea tide, the polar tide, the relativistic effect and the like according to the geometric orbit information under the inertial system, and all conservative force acceleration of the low-rail gravity satellite is obtained through calculation.
And S5, obtaining a non-conservative force calibration reference value according to the difference value of the motion acceleration of the carrier and the conservative force acceleration by utilizing a Newton second motion law.
S6, processing the original data of the satellite attitude by using an interpolation method to obtain an attitude continuous observation value, and calculating to obtain a non-conservative force observation value of the low-orbit gravity satellite under an inertial system according to the attitude continuous observation value and the original data of the accelerometer; specifically, the method comprises the steps of firstly carrying out interpolation processing on original satellite attitude data to obtain an attitude continuous observation value, then keeping time synchronization between the attitude continuous observation value and an accelerometer correction observation value obtained through preprocessing, and carrying out coordinate transformation on the accelerometer observation value by utilizing four-element information of the attitude data to obtain a non-conservative force observation value of the low-orbit gravity satellite under the inertial system.
And S7, establishing a calibration model of the satellite-borne accelerometer through the non-conservative force calibration reference value and the non-conservative force observation value, and estimating a scale factor and a deviation factor by using a least square method. Specifically, the calibration formula of the calibration model is as follows:
in the formula (I), the compound is shown in the specification,
representing a conservative force acceleration value;
represents a non-conservative force observation;
representing a scale factor;
representing a deviation factor; (s) x ,s y ,s z ) Respectively representing three components of a scale factor; (b) x ,b y ,b z ) Representing the deviation factor three components, respectively.
Example 2
With reference to fig. 1, this embodiment provides a calibration device for a satellite-borne accelerometer, which is used for implementing the above method, and comprises:
the data acquisition module is used for acquiring carrier phase original data, geometric orbit original data, accelerometer original data and satellite attitude original data;
the carrier phase acceleration calculation module is used for calculating the carrier phase original data by utilizing a carrier phase differential method to obtain carrier phase acceleration; the calculation formula of the carrier phase acceleration calculation module is the same as that of the carrier phase acceleration calculation module in the embodiment 1;
the carrier motion acceleration calculation module is used for calculating the carrier motion acceleration of the low-orbit gravity satellite under the inertial system according to the carrier phase acceleration and by combining the geometric relationship between the GPS satellite and the low-orbit gravity satellite; the calculation formula of the carrier motion acceleration calculation module is the same as that of the carrier motion acceleration calculation module in the embodiment 1;
the conservative force acceleration calculation module is used for calculating the conservative force acceleration of the low-orbit gravity satellite according to the original data of the geometric orbit and the perturbation force model;
the calibration reference value determining module is used for determining a non-conservative force calibration reference value according to the difference value of the motion acceleration of the carrier and the conservative force acceleration;
the non-conservative force observation value calculation module is used for processing the original data of the satellite attitude by using an interpolation method to obtain an attitude continuous observation value and calculating a non-conservative force observation value of the low-orbit gravity satellite under an inertial system according to the attitude continuous observation value and the original data of the accelerometer;
the calibration module is used for establishing a calibration model of the satellite-borne accelerometer through the non-conservative force calibration reference value and the non-conservative force observation value; the calibration formula of the calibration module is the same as that of the calibration model of embodiment 1 described above.
The specific implementation method of the calibration apparatus is the same as the calibration method provided in embodiment 1, and is not described herein again.
In summary, the embodiment of the invention calculates the non-conservative force calibration reference value based on the GPS carrier phase differential method and the newton second law of motion, and establishes the calibration model for estimating the calibration parameter from the non-conservative force observation value, so that the influence of the empirical model method on the high dependence on the atmospheric density model and the radiation coefficient error can be effectively avoided, and the calibration accuracy of the satellite-borne accelerometer can be improved.
It should be noted that the above embodiments are only specific and clear descriptions of technical solutions and technical features of the present application. However, to those skilled in the art, aspects or features that are part of the prior art or common general knowledge are not described in detail in the above embodiments.
Of course, the technical solutions of the present application are not limited to the above-mentioned embodiments, and those skilled in the art should take the description as a whole, and the technical solutions in the embodiments may also be appropriately combined, so that other embodiments that may be understood by those skilled in the art may be formed.
Claims (8)
1. A calibration method of a satellite-borne accelerometer is characterized by comprising the following steps:
acquiring carrier phase original data, geometric orbit original data, accelerometer original data and satellite attitude original data;
calculating the carrier phase original data by using a carrier phase differential method to obtain carrier phase acceleration;
calculating to obtain the carrier motion acceleration of the low-orbit gravity satellite under the inertial system according to the carrier phase acceleration and by combining the geometric relationship between the GPS satellite and the low-orbit gravity satellite;
according to the original data of the geometric orbit and the perturbation force model, calculating to obtain the conservative force acceleration of the low-orbit gravity satellite;
determining a non-conservative force calibration reference value according to the difference value of the motion acceleration of the carrier and the conservative force acceleration;
processing the original data of the satellite attitude by using an interpolation method to obtain an attitude continuous observation value, and calculating to obtain a non-conservative force observation value of the low-orbit gravity satellite under an inertial system according to the attitude continuous observation value and the original data of the accelerometer;
and establishing a calibration model of the satellite-borne accelerometer through the non-conservative force calibration reference value and the non-conservative force observation value.
2. The method for calibrating a space-borne accelerometer according to claim 1, wherein in the step, the calculation formula of the carrier phase acceleration is as follows:
in the formula (I), the compound is shown in the specification,
and
respectively representing carrier phase velocity and carrier phase acceleration;
representing the original observed value of the carrier phase without an ionized layer; e.g. of the type T And f T Respectively representing corresponding first-order and second-order difference operators;
representing a carrier phase ionosphere free combination value; c represents the speed of light; dt P Representing a GPS satellite clock error; rel represents relativistic effects; ω represents the antenna phase winding error;
representing the geometric distance between the GPS satellite and the low-orbit gravity satellite; dt G Representing the receiver clock error of a low-orbit gravity satellite; n represents the carrier phase integer ambiguity; ε represents the observed random error.
3. The method for calibrating a satellite-borne accelerometer according to claim 2, wherein in the step, the calculation formula of the motion acceleration of the carrier is as follows:
in the formula (I), the compound is shown in the specification,
the unit vector of the sight direction from the GPS satellite to the low-orbit gravity satellite is represented;
and
respectively representing the sight line speed and the sight line acceleration;
and
respectively representing the first difference and the second difference of the clock error of the low-orbit gravity satellite receiver;
and
respectively representing the speed and acceleration of the GPS satellite;
and
respectively representing the motion speed and the motion acceleration of the carrier of the low-orbit gravity satellite.
4. The method for calibrating the satellite-borne accelerometer according to claim 3, wherein in the step, the calibration formula of the calibration model is as follows:
in the formula (I), the compound is shown in the specification,
representing a conservative force acceleration value;
represents a non-conservative force observation;
representing a scale factor;
representing a deviation factor; (s) x ,s y ,s z ) Respectively representing three components of the scale factor; (b) x ,b y ,b z ) Representing the deviation factor three components, respectively.
5. A calibration device for a satellite-borne accelerometer, comprising:
the data acquisition module is used for acquiring carrier phase original data, geometric orbit original data, accelerometer original data and satellite attitude original data;
the carrier phase acceleration calculation module is used for calculating the carrier phase original data by utilizing a carrier phase differential method to obtain carrier phase acceleration;
the carrier motion acceleration calculation module is used for calculating the carrier motion acceleration of the low-orbit gravity satellite under the inertial system according to the carrier phase acceleration and by combining the geometric relationship between the GPS satellite and the low-orbit gravity satellite;
the conservative force acceleration calculation module is used for calculating the conservative force acceleration of the low-orbit gravity satellite according to the original data of the geometric orbit and the perturbation force model;
the calibration reference value determining module is used for determining a non-conservative force calibration reference value according to the difference value of the motion acceleration of the carrier and the conservative force acceleration;
the non-conservative force observation value calculation module is used for processing the original data of the satellite attitude by using an interpolation method to obtain an attitude continuous observation value and calculating a non-conservative force observation value of the low-orbit gravity satellite under an inertial system according to the attitude continuous observation value and the original data of the accelerometer;
and the calibration module is used for establishing a calibration model of the satellite-borne accelerometer through the non-conservative force calibration reference value and the non-conservative force observation value.
6. The device for calibrating the satellite-borne accelerometer according to claim 5, wherein the calculation formula of the carrier phase acceleration calculation module is as follows:
in the formula (I), the compound is shown in the specification,
and
respectively representing carrier phase velocity and carrier phase acceleration;
representing the original observed value of the carrier phase without an ionized layer; e.g. of the type T And f T Respectively representing corresponding first-order and second-order difference operators;
representing a carrier phase ionosphere free combination value; c represents the speed of light; dt P Representing a GPS satellite clock error; rel represents a relativistic effect; ω represents the antenna phase winding error;
representing the geometric distance between the GPS satellite and the low-orbit gravity satellite; dt is G Representing the receiver clock error of a low-orbit gravity satellite; n represents the carrier phase integer ambiguity; ε represents the observed random error.
7. The calibration device of the satellite-borne accelerometer according to claim 5, wherein the calculation formula of the carrier motion acceleration calculation module is as follows:
in the formula (I), the compound is shown in the specification,
and
respectively representing carrier phase speed and carrier phase acceleration;
representing the geometric distance between the GPS satellite and the low-orbit gravity satellite; c represents the speed of light;
the unit vector of the sight direction from the GPS satellite to the low-orbit gravity satellite is represented;
and
respectively representing the sight line speed and the sight line acceleration;
and
respectively representing the first difference and the second difference of the clock error of the low-orbit gravity satellite receiver;
and
respectively representing the speed and acceleration of the GPS satellite;
and
respectively representing the motion speed and the motion acceleration of the carrier of the low-orbit gravity satellite.
8. The calibration device of the satellite-borne accelerometer according to claim 5, wherein the calibration formula of the calibration module is as follows:
in the formula (I), the compound is shown in the specification,
representing the acceleration of the motion of the carrier of the low-orbit gravity satellite;
representing a conservative force acceleration value;
represents a non-conservative force observation;
representing a scale factor;
representing a deviation factor; (s) x ,s y ,s z ) Respectively representing three components of the scale factor; (b) x ,b y ,b z ) Representing the deviation factor three components, respectively.
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| CN112146663B (en) * | 2020-09-27 | 2021-05-18 | 武汉大学 | GNSS satellite precision orbit determination enhancing method and system based on bias satellite-borne accelerometer |
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Citations (3)
| Publication number | Priority date | Publication date | Assignee | Title |
|---|---|---|---|---|
| CN102323450A (en) * | 2011-05-19 | 2012-01-18 | 中国科学院测量与地球物理研究所 | Satellite-borne accelerometer data calibrating method based on dual-satellite adjacent energy difference principle |
| RU2012127092A (en) * | 2012-06-28 | 2014-01-10 | Федеральное государственное унитарное предприятие "Научно-производственный центр автоматики и приборостроения имени академика Н.А. Пилюгина" (ФГУП "НПЦАП") | METHOD FOR DETERMINING PARAMETERS OF MODEL OF ERRORS OF MEASUREMENTS ACCELEROMETERS OF INERTIAL NAVIGATION SYSTEM BY MEASUREMENTS OF SATELLITE NAVIGATION |
| CN109581523A (en) * | 2018-11-12 | 2019-04-05 | 湖北省地震局 | It is a kind of using satellite to satellite tracking device to the method and system of accelerometer calibration |
Family Cites Families (3)
| Publication number | Priority date | Publication date | Assignee | Title |
|---|---|---|---|---|
| US8311740B2 (en) * | 2010-01-28 | 2012-11-13 | CSR Technology Holdings Inc. | Use of accelerometer only data to improve GNSS performance |
| US8718963B2 (en) * | 2010-04-27 | 2014-05-06 | Memsic, Inc. | System and method for calibrating a three-axis accelerometer |
| US9891245B2 (en) * | 2015-06-29 | 2018-02-13 | CloudNav Inc. | Real-time accelerometer calibration |
-
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- 2019-09-19 CN CN201910887331.4A patent/CN110703355B/en active Active
Patent Citations (3)
| Publication number | Priority date | Publication date | Assignee | Title |
|---|---|---|---|---|
| CN102323450A (en) * | 2011-05-19 | 2012-01-18 | 中国科学院测量与地球物理研究所 | Satellite-borne accelerometer data calibrating method based on dual-satellite adjacent energy difference principle |
| RU2012127092A (en) * | 2012-06-28 | 2014-01-10 | Федеральное государственное унитарное предприятие "Научно-производственный центр автоматики и приборостроения имени академика Н.А. Пилюгина" (ФГУП "НПЦАП") | METHOD FOR DETERMINING PARAMETERS OF MODEL OF ERRORS OF MEASUREMENTS ACCELEROMETERS OF INERTIAL NAVIGATION SYSTEM BY MEASUREMENTS OF SATELLITE NAVIGATION |
| CN109581523A (en) * | 2018-11-12 | 2019-04-05 | 湖北省地震局 | It is a kind of using satellite to satellite tracking device to the method and system of accelerometer calibration |
Non-Patent Citations (1)
| Title |
|---|
| "动力法校准GRACE星载加速度计";邹贤才 等;《武汉大学学报-信息科学版》;20150331;第40卷(第3期);第357-360页 * |
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