CN112649007A - Integrated design method of attitude sensor - Google Patents
Integrated design method of attitude sensor Download PDFInfo
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- CN112649007A CN112649007A CN202110046406.3A CN202110046406A CN112649007A CN 112649007 A CN112649007 A CN 112649007A CN 202110046406 A CN202110046406 A CN 202110046406A CN 112649007 A CN112649007 A CN 112649007A
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- 238000000034 method Methods 0.000 title claims abstract description 19
- 238000013461 design Methods 0.000 title claims abstract description 15
- 238000012545 processing Methods 0.000 claims abstract description 31
- 239000013307 optical fiber Substances 0.000 claims description 23
- 239000000835 fiber Substances 0.000 claims description 12
- 238000007499 fusion processing Methods 0.000 claims description 5
- 238000001914 filtration Methods 0.000 claims description 3
- 238000003384 imaging method Methods 0.000 claims description 3
- 230000014759 maintenance of location Effects 0.000 claims description 2
- 238000009434 installation Methods 0.000 abstract description 10
- 230000004927 fusion Effects 0.000 abstract description 8
- 230000002035 prolonged effect Effects 0.000 abstract 1
- 238000005259 measurement Methods 0.000 description 8
- 238000009825 accumulation Methods 0.000 description 3
- 238000004364 calculation method Methods 0.000 description 3
- 230000000712 assembly Effects 0.000 description 2
- 238000000429 assembly Methods 0.000 description 2
- 230000009286 beneficial effect Effects 0.000 description 2
- 230000000694 effects Effects 0.000 description 2
- 230000007774 longterm Effects 0.000 description 2
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- 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
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- G—PHYSICS
- G01—MEASURING; TESTING
- G01C—MEASURING DISTANCES, LEVELS OR BEARINGS; SURVEYING; NAVIGATION; GYROSCOPIC INSTRUMENTS; PHOTOGRAMMETRY OR VIDEOGRAMMETRY
- G01C19/00—Gyroscopes; Turn-sensitive devices using vibrating masses; Turn-sensitive devices without moving masses; Measuring angular rate using gyroscopic effects
- G01C19/58—Turn-sensitive devices without moving masses
- G01C19/64—Gyrometers using the Sagnac effect, i.e. rotation-induced shifts between counter-rotating electromagnetic beams
- G01C19/72—Gyrometers using the Sagnac effect, i.e. rotation-induced shifts between counter-rotating electromagnetic beams with counter-rotating light beams in a passive ring, e.g. fibre laser gyrometers
- G01C19/721—Details
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- 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
- G01C21/02—Navigation; Navigational instruments not provided for in groups G01C1/00 - G01C19/00 by astronomical means
- G01C21/025—Navigation; Navigational instruments not provided for in groups G01C1/00 - G01C19/00 by astronomical means with the use of startrackers
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Abstract
The invention provides an integrated design method of an attitude sensor, which relates to the technical field of aerospace navigation systems and comprises the steps of integrally designing the mechanical structures of a star sensor and a fiber-optic gyroscope; integrally designing the star sensor and the electronic structure of the fiber-optic gyroscope; and fusing the data collected by the star sensor and the fiber-optic gyroscope. The problems that mechanical deviation, data fusion and time delay are caused by installation reference and structural heat distortion existing in split installation of the attitude sensor in the prior art, and a large amount of original information is lost at a subsystem level and the tail end of the data fusion are solved at least partially, and the accuracy of attitude determination and angular speed determination is improved. Meanwhile, the weight and power consumption of the system are effectively reduced, the reliability of the sensor is improved, the service life of the sensor is prolonged, the installation of the spacecraft is facilitated, and on-satellite processing resources are saved.
Description
Technical Field
The invention relates to the technical field of aerospace navigation systems, in particular to an integrated design method of an attitude sensor.
Background
With the increase of the demand of laser communication satellites, the demand of high-precision and high-stability control of spacecrafts is increasing. The high-precision and high-dynamic attitude sensor is a key guarantee for ensuring that the spacecraft meets the index requirements. In consideration of the limitations of cost and satellite resources, the difficulties and obstacles faced by the solution from a single machine level are difficult to overcome, so that the current difficulties are solved in the design of optimizing the attitude sensor from a system level and a use level, and the superiority and the robustness of system indexes are ensured.
The attitude sensor adopted in the prior art is introduced:
the fiber-optic gyroscope is a passive inertial navigation system, does not depend on external information, and completely depends on the movement of a sensitive carrier of an inertial measurement unit under an inertial system, so that the problems that the navigation precision is influenced by the zero offset of an instrument and the navigation error is accumulated along with time exist. In performing attitude measurement, the fiber optic gyroscope is capable of outputting angular velocity at a high frequency and with high accuracy, with high accuracy of attitude angular velocity measurement in a short time, but there is a long-term accumulation of errors, and initial attitude information is required if it is desired to output attitude information.
The star sensor is a near-autonomous starlight navigation system, a starry sky is taken as a working object, a fixed star is taken as an observation target, and high-precision attitude information of a carrier relative to an inertial space is determined through the sensitive fixed star of the sensor. When the attitude measurement is carried out, the star sensor has high static attitude measurement accuracy, but has small view field and poor dynamic performance, and the performance of the star sensor is seriously deteriorated under the condition of large angular speed.
In the prior art, the two attitude sensors are installed on the satellite in a split mode, and the respective problems of the two attitude sensors are partially solved; but still has the problem that split type installation leads to the mechanical deviation that two installation benchmark and structure thermal distortion brought, the calculation error that data fusion and time delay brought. Meanwhile, in the prior art, when information after attitude measurement is output, the output data of the two attitude sensors are combined, so that the use and the precision are improved to a certain extent; however, since data is combined at the subsystem level and the end, a large amount of original information is lost due to errors and system influence, and the effect of improving the overall performance of the system is limited.
Disclosure of Invention
The invention provides an integrated design method of an attitude sensor, which aims to solve the problems that mechanical deviation is caused by installation reference and structural heat distortion in split installation of the attitude sensor in the prior art, calculation errors are caused by data fusion and time delay, and a large amount of original information is lost by the data fusion at a subsystem level and at the tail end.
Specifically, the invention provides an integrated design method of an attitude sensor, which comprises the following steps:
the star sensor and the mechanical structure of the optical fiber gyroscope are integrally designed; integrally designing the star sensor and the electronic structure of the fiber-optic gyroscope; and fusing the data collected by the star sensor and the fiber-optic gyroscope.
The attitude sensor comprises a star sensor and a fiber-optic gyroscope; the star sensor comprises a detector, a light shield and a power supply and processing circuit part, and image data acquisition is carried out; the optical fiber gyroscope comprises an optical fiber gyroscope head part and a power supply and processing circuit part, and image data acquisition and temperature data acquisition are carried out.
The integrated design of the star sensor and the mechanical structure of the optical fiber gyroscope comprises the following steps:
the head of the optical fiber gyroscope is arranged at the periphery of the lower part of the star sensor, the star sensor detector and the light shield are arranged at the head of the attitude sensor, and the power supply and processing circuit parts of the optical fiber gyroscope and the star sensor are arranged at the lower middle part of the attitude sensor. The design effectively combines the structural characteristics of the star sensor and the fiber-optic gyroscope in the prior art, and is beneficial to saving space so as to facilitate the installation of the attitude sensor.
The integrated design of the star sensor and the electronic structure of the fiber-optic gyroscope comprises the following steps:
the electronic structure of the attitude sensor comprises a star sensor imaging circuit module, a fiber-optic gyroscope circuit module and a power supply module; the star sensor and the optical fiber gyroscope share a first processing chip for image data acquisition; the image data receiving and the optical fiber gyroscope temperature data acquisition share a second processing chip; and the CPU reads and writes the data to complete the fusion processing of the data.
The fusion processing of the data collected by the star sensor and the fiber-optic gyroscope comprises the following steps:
in software processing, the optical fiber gyroscope is used for assisting the star sensor to extract the star point mass center and match the star points, the characteristic of the optical fiber gyroscope under error accumulation in a short time is used, attitude constraint among a plurality of star images is established, and the matching speed with navigation stars in a star bank is improved; the constant drift of the fiber optic gyroscope is calibrated on line by using the star sensor, and the accurate attitude information output by the star sensor is utilized; and the precision of the attitude sensor is improved by utilizing EKF filtering. Through the data fusion processing, high-precision attitude angle and angular velocity information can be output at high frequency, so that the attitude determination output characteristics of the star sensor and the fiber-optic gyroscope are fully combined to exert respective advantages, and the use and precision are greatly improved.
The integrated design method of the attitude sensor provided by the invention at least has the following beneficial effects: through the integrated design of the structure, the control system and the information fusion, the mechanical deviation caused by the installation reference of the split type structure and the thermal distortion of the structure and the calculation error caused by the data fusion and the time delay are solved, and the precision of the attitude determination and the angular speed determination is improved. Meanwhile, the integrated design can effectively reduce the weight and power consumption of the system, improve the reliability and service life of the sensor, facilitate the installation of the spacecraft and save on-board processing resources.
Drawings
FIG. 1 illustrates the advantage of the method of the present invention for attitude sensor design over the prior art.
FIG. 2 shows the structural layout of the attitude sensor in one embodiment of the method of the present invention.
FIG. 3 shows a flow chart of attitude sensor internal information in one embodiment of the method of the present invention.
Detailed Description
It should be noted that the components in the figures may be exaggerated and not necessarily to scale for illustrative purposes. In the figures, identical or functionally identical components are provided with the same reference symbols.
In the present invention, "disposed on …", "disposed over …" and "disposed over …" do not exclude the presence of an intermediate therebetween, unless otherwise specified. Further, "disposed on or above …" merely indicates the relative positional relationship between two components, and may also be converted to "disposed below or below …" and vice versa in certain cases, such as after reversing the product direction.
In the present invention, the embodiments are only intended to illustrate the aspects of the present invention, and should not be construed as limiting.
In the present invention, the terms "a" and "an" do not exclude the presence of a plurality of elements, unless otherwise specified.
It is further noted herein that in embodiments of the present invention, only a portion of the components or assemblies may be shown for clarity and simplicity, but those of ordinary skill in the art will appreciate that, given the teachings of the present invention, required components or assemblies may be added as needed in a particular scenario. Furthermore, features from different embodiments of the invention may be combined with each other, unless otherwise indicated. For example, a feature of the second embodiment may be substituted for a corresponding or functionally equivalent or similar feature of the first embodiment, and the resulting embodiments are likewise within the scope of the disclosure or recitation of the present application.
It is also noted herein that, within the scope of the present invention, the terms "same", "equal", and the like do not mean that the two values are absolutely equal, but allow some reasonable error, that is, the terms also encompass "substantially the same", "substantially equal". By analogy, in the present invention, the terms "perpendicular", "parallel" and the like in the directions of the tables also cover the meanings of "substantially perpendicular", "substantially parallel".
The numbering of the steps of the methods of the present invention does not limit the order of execution of the steps of the methods. Unless specifically stated, the method steps may be performed in a different order.
The invention is further elucidated with reference to the drawings in conjunction with the detailed description.
And the star sensor and the mechanical structure of the optical fiber gyroscope are integrally designed.
In the prior art, a star sensor power supply and processing circuit part is mainly concentrated on the lower part of an attitude sensor structure, and a detector and a light shield are concentrated on the upper part of the attitude sensor; the power supply and processing circuit part of the fiber-optic gyroscope is mainly concentrated in the middle of the attitude sensor structure, and the head part of the optical fiber is concentrated in the periphery of the attitude sensor structure. According to the characteristics, the star sensor and the structure characteristics of the optical fiber gyroscope are combined, when the mechanical structure of the star sensor and the optical fiber gyroscope is integrally designed, the head of the optical fiber gyroscope can be arranged on the periphery of the lower part of the star sensor, the head of the star sensor still keeps the state of the former star sensor, and the power supply and processing circuit parts of the optical fiber gyroscope and the star sensor are arranged in the lower middle part of the attitude sensor. The structural scheme is favorable for saving space and is convenient for realizing and installing the attitude sensor. The specific structural layout is shown in fig. 2.
The star sensor and the electronic structure of the fiber-optic gyroscope are integrally designed. The attitude sensor electronic structure mainly comprises a fiber optic gyroscope circuit module, a star sensor imaging circuit and a power supply module. The data acquisition of the fiber-optic gyroscope and the star sensor share a first processing chip, so that the data acquisition synchronism of the fiber-optic gyroscope and the star sensor is ensured; the data image receiving and the temperature data acquisition of the fiber-optic gyroscope share a second processing chip, so that the signal collection and interaction synchronism are ensured; and finally, finishing the fusion algorithm of the attitude determination and the angular velocity determination of the attitude sensor by using the CPU.
In this embodiment, the internal information flow of the attitude sensor is shown in fig. 3, which specifically includes the following contents: the processing chip can be an FPGA chip, and the first processing chip FPGA1 collects image signals of the fiber-optic gyroscope and the star sensor from a CMOS active pixel image sensor (APS) and sends control signals to the image signals; the first processing chip FPGA1 accesses image data in the image memory SRAM; the second processing chip FPGA2 collects the temperature data of the fiber-optic gyroscope from the temperature collection AD converter and sends a control signal to the temperature data; the first processing chip FPGA1 is connected with the second processing chip FPGA2 through two interface chips; the CPU sends control signals to the second processing chip FPGA2 and reads and writes data mutually; the CPU reads the data in the two FLASH program storages; the CPU and the program operation memory SDRAM read and write data mutually.
And fusing the data collected by the star sensor and the fiber-optic gyroscope.
The star sensor has high static attitude measurement accuracy, but has small field of view and poor dynamic performance, and the performance of the star sensor is seriously deteriorated under the condition of large angular speed; the fiber optic gyroscope is capable of outputting angular velocity at high frequency and high accuracy, has high attitude angular velocity measurement accuracy in a short time, but has long-term error accumulation, and requires initial attitude information if it is desired to output the attitude information. Therefore, the two attitude determination output characteristics are fully combined to exert respective advantages, and the use and the precision can be greatly improved.
Conventionally, data of a star sensor and a fiber optic gyroscope are fused at a system level and a tail end, a large amount of original information can be lost due to errors and system influence, the effect on improving the overall performance of a system is limited, and therefore the data are fused in an attitude sensor urgently needed. In software processing, on one hand, the characteristic that the accumulated error of the fiber optic gyroscope is small in a short time is utilized, the star sensor assembly is assisted to extract the mass center of a star point and match the star point, attitude constraint among a plurality of star images is established, and the matching speed of the star images with navigation stars in a star database is improved; on the other hand, the accurate attitude information output by the star sensor component can carry out on-line calibration on the constant drift of the fiber optic gyroscope component. By carrying out fusion processing on the information, high-precision attitude angle and angular velocity information is output at high frequency, and the precision of an attitude sensor can be improved by utilizing EKF filtering.
Claims (6)
1. An integrated design method of an attitude sensor, the attitude sensor comprises a star sensor and a fiber-optic gyroscope, which is characterized in that,
the star sensor and the mechanical structure of the optical fiber gyroscope are integrally designed;
integrally designing the star sensor and the electronic structure of the fiber-optic gyroscope;
and fusing the data collected by the star sensor and the fiber-optic gyroscope.
2. The inventive method of claim 1, wherein: the star sensor comprises a detector, a light shield and a power supply and processing circuit part, and image data acquisition is carried out; the optical fiber gyroscope comprises an optical fiber gyroscope head part and a power supply and processing circuit part, and image data acquisition and temperature data acquisition are carried out.
3. The invention method of claim 2, wherein the integrating the star sensor with the mechanical structure of the fiber optic gyroscope comprises:
the head of the optical fiber gyroscope is arranged at the periphery of the lower part of the star sensor, the star sensor detector and the light shield are arranged at the head of the attitude sensor, and the power supply and processing circuit parts of the optical fiber gyroscope and the star sensor are arranged at the lower middle part of the attitude sensor.
4. The invention method of claim 2, wherein the integrating the star sensor with the electronic structure of the fiber optic gyroscope comprises:
the electronic structure of the attitude sensor comprises a star sensor imaging circuit module, a fiber-optic gyroscope circuit module and a power supply module; the star sensor and the optical fiber gyroscope share a first processing chip for image data acquisition; the image data receiving and the optical fiber gyroscope temperature data acquisition share a second processing chip; and the CPU reads and writes the data to complete the fusion processing of the data.
5. The invention method according to one of claims 1 and 4, wherein the fusing of the data collected by the star sensor and the fiber-optic gyroscope comprises: extracting the mass center of a star point and matching the star point by using a fiber optic gyroscope to assist a star sensor; the star sensor is used for carrying out on-line calibration on the constant drift of the fiber optic gyroscope; and the precision of the attitude sensor is improved by utilizing EKF filtering.
6. The invention method as claimed in claim 1, wherein the integrating the star sensor with the electronic structure of the fiber optic gyroscope further comprises: the processing chip is an FPGA chip, and the first processing chip collects image signals of the fiber-optic gyroscope and the star sensor from the APS and sends control signals to the APS; the first processing chip accesses the image data in an image memory SRAM; the second processing chip collects the temperature data of the fiber-optic gyroscope from the temperature collection AD converter and sends a control signal to the temperature data; the first processing chip is connected with the second processing chip through two interface chips; the CPU sends control signals to the second processing chip and reads and writes data mutually; the CPU reads the data in the two FLASH program storages; the CPU and the program operation memory SDRAM read and write data mutually.
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Citations (5)
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|---|---|---|---|---|
| CN101881619A (en) * | 2010-06-25 | 2010-11-10 | 哈尔滨工程大学 | Ship's inertial navigation and astronomical positioning method based on attitude measurement |
| CN103674023A (en) * | 2013-12-26 | 2014-03-26 | 中国人民解放军国防科学技术大学 | Method for dynamically measuring attitude of star sensor based on top accurate angle relevance |
| CN107529374B (en) * | 2012-11-09 | 2015-07-15 | 上海航天控制工程研究所 | A kind of inertial attitude sensor |
| CN108519086A (en) * | 2018-03-20 | 2018-09-11 | 北京航天控制仪器研究所 | A kind of inertial platform spherical shape stage body multifunction structure |
| CN108827310A (en) * | 2018-07-12 | 2018-11-16 | 哈尔滨工程大学 | A kind of star sensor secondary gyroscope online calibration method peculiar to vessel |
-
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- 2021-01-13 CN CN202110046406.3A patent/CN112649007A/en active Pending
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|---|---|---|---|---|
| CN101881619A (en) * | 2010-06-25 | 2010-11-10 | 哈尔滨工程大学 | Ship's inertial navigation and astronomical positioning method based on attitude measurement |
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