CN111637880A

CN111637880A - Low-cost microminiaturized star sensor and design method thereof

Info

Publication number: CN111637880A
Application number: CN202010397439.8A
Authority: CN
Inventors: 桑晓茹; 刘惟芳; 杨峰; 任维佳
Original assignee: Changsha Tianyi Space Technology Research Institute Co Ltd
Current assignee: Changsha Tianyi Space Technology Research Institute Co Ltd
Priority date: 2020-05-12
Filing date: 2020-05-12
Publication date: 2020-09-08

Abstract

The invention relates to a microminiaturized star sensor and a design method thereof, which is characterized by comprising the following steps: a multi-dimensional adjustment mechanism (103) capable of adjusting angles and/or displacements from five dimensions receives starlight output by an objective lens combination unit (140) in a manner of being disposed at a non-focal plane of the objective lens combination unit (140), and generates a starfield image based on at least one of the starlight. By improving the specific structure of the objective lens combination unit, the weight and the size of the whole star sensor are further reduced. By arranging the star chart camera as the photoelectric sensor array, the size and the weight of the star detection unit are reduced, the power is reduced, and the size and the weight of the star sensor are further reduced. Through the arrangement of the diaphragm unit, a star field image containing stray light radiation is eliminated, the camera attitude can be determined more accurately, and the accurate judgment of the spacecraft attitude is improved.

Description

Low-cost microminiaturized star sensor and design method thereof

Technical Field

The invention relates to the technical field of aerospace, in particular to a microminiaturized star sensor and a design method thereof, and particularly relates to a low-cost microminiaturized star sensor and a design method thereof.

Background

Current conventional systems for determining attitude of an aerospace vehicle include a star sensor and a gyroscope, each of which is used independently to acquire attitude information and to transmit the attitude information to the flight computer of the spacecraft. It is a drawback to use only star sensors to determine the attitude. Under the condition that the slew rate of the spacecraft is high, namely the moving speed is too fast or the spacecraft rolls over, the view field of a star camera in the star sensor can be too fast to focus due to too fast change, the position and the zero degree of a star can become fuzzy and cannot be fully analyzed by the star camera, and further the view field cannot be correctly compared with the star catalogue of the known star. If the star camera does not have a priori knowledge of the spacecraft attitude, it must rely on the application of a "spatial wandering" algorithm, which is computationally more intensive due to the lack of basic information about the previous spacecraft attitude. This can result in poor resolution, such that stars in the star camera field of view do not properly match stars in the star catalogue. The determination of the attitude of the spacecraft by using only the gyroscope is also disadvantageous. Inherent errors associated with gyroscopes include gyroscope "drift", bias and scale factor errors, which affect the accuracy of the gyroscope output data. In the prior art, a Charge Coupled Device (CCD) and a gyroscope are generally adopted to determine the attitude of a spacecraft. The size, weight, control electronics and power requirements of CCDs prohibit their use in small aerospace applications. At the same time, the CCD is also vulnerable to radiation and the conventional gyroscope also has a large size and weight, with the drawback of being sensitive to vibrations and to radiation.

Chinese patent CN 103148853B discloses an attitude determination method for an aerospace vehicle based on a star sensor, which is characterized by comprising the following steps: acquiring attitude information of a plurality of aerospace vehicles by a plurality of star sensors and gyroscopes, wherein each star sensor and each gyroscope are combined into an acquisition unit to acquire corresponding attitude information of the aerospace vehicles; selecting effective attitude information of the aerospace vehicles from the attitude information of the aerospace vehicles, and carrying out local state estimation on the effective attitude information of the aerospace vehicles through a Kalman filter to obtain a plurality of local optimal state estimations; respectively calculating the fusion weight of each local optimal state estimation; respectively carrying out weighted fusion on each fusion weight and the corresponding local optimal state estimation to obtain global optimal estimation; and determining the attitude of the aerospace vehicle according to the global optimal estimation. The patent improves robustness and reconfigurability by changing algorithms, rather than improving gyroscopes and star sensors from the perspective of reducing volume and mass. For a spacecraft less than ten kilograms, the gyroscope and the star sensor cannot be applied to the traditional gyroscope and the star sensor at all because the mass and the volume of the gyroscope and the star sensor are large. Therefore, how to determine the attitude of the spacecraft by using a low-quality, small-volume and high-definition star image and high-precision gyroscope data is an unsolved technical problem which cannot be realized by only changing the algorithm.

Furthermore, on the one hand, due to the differences in understanding to the person skilled in the art; on the other hand, since the inventor has studied a lot of documents and patents when making the present invention, but the space is not limited to the details and contents listed in the above, however, the present invention is by no means free of the features of the prior art, but the present invention has been provided with all the features of the prior art, and the applicant reserves the right to increase the related prior art in the background.

Disclosure of Invention

Aiming at the defects of the prior art, the invention provides a microminiaturized star sensor which is characterized in that a star detection unit of the star sensor at least comprises an objective lens combination unit and a multi-dimensional detection unit, the multi-dimensional detection unit capable of adjusting angles and/or displacements from five dimensions is arranged on a non-focal plane of the objective lens combination unit and has a certain distance with a focal point of the objective lens combination unit, the multi-dimensional detection unit adjusts the receiving angle in a mode of adjusting the angles and/or displacements based on the five dimensions and receives star light output by the objective lens combination unit, and a star field image is generated based on at least one star light.

In the prior art, the detector is arranged directly in the focal plane of the objective lens combination unit, and the residual radiation may be diffused and illuminated by the detection device arranged directly in the focal plane of the star objective lens combination unit. In contrast to the prior art, the present invention arranges a diaphragm unit between the objective lens combination unit and the multi-dimensional detection unit, wherein at least one diaphragm in the diaphragm unit is arranged on the focal plane of the objective lens combination unit, and the multi-dimensional detection unit receives the starlight output by the objective lens combination unit through the diaphragm unit. The aperture unit comprises a field diaphragm and an aperture diaphragm which are arranged on the same optical axis, the aperture diaphragm is arranged on a focal plane of the objective lens combination unit, and the starlight output by the objective lens combination unit is transmitted to the multi-dimensional detection unit through the aperture diaphragm and the field diaphragm in sequence. Through the arrangement of the diaphragm unit, the multi-dimensional detection unit is not directly arranged on the focal plane of the objective lens combination unit, but is at a certain distance from the focal point of the objective lens combination unit, and the field diaphragm is positioned on the focal plane of the star objective lens combination unit. The residual radiation scattered by the wall of the main objective lens in the star objective lens combination unit can be attenuated, a star field image without stray light radiation is not contained, and the accurate determination of the attitude of the camera is facilitated. According to a preferred embodiment, the multi-dimensional detection unit receives the star light output by the diaphragm unit in such a way that the residual radiation is attenuated by the relay optical assembly.

The problem that the imaging precision of a camera is influenced by different mass center position drifts caused by position changes due to the fact that the multi-dimensional detection unit is subjected to mechanical vibration or temperature changes is avoided. The invention sets the multi-dimensional detection unit to be capable of translation or rotation in five adjustment dimensions. The multi-dimensional detection unit comprises a star map camera and an adjusting frame for installing the star map camera, wherein the adjusting frame moves and/or rotates in four dimensions so that the star map camera adjusts star light receiving angles in the four dimensions to reposition an optical axis, a photoelectric sensing array for imaging in the star map camera is translated based on the dimension of the optical axis direction, and therefore a focal plane between the photoelectric sensing array and a lens for receiving star light in the star map camera is repositioned and focus is adjusted. The star map camera in the multi-dimensional detection unit can move and be adjusted in five dimensions, so that the focus of a received image can be adjusted, the camera and the lens formed by the photoelectric sensing array can realize automatic focusing, and the focusing capability is realized. In case the incoming star light is fully collimated, refocusing of the optical image can be achieved by adjusting the focus. The multidimensional adjustment of the multidimensional detection unit can be beneficial to inputting required target starlight into the star chart camera by adjusting the focus point, so that more accurate camera attitude is obtained, and the accurate attitude of the spacecraft is determined by combining the angular rate.

The star map camera in the multi-dimensional detection unit is a photoelectric sensor array, and compared with the traditional imaging device CCD, the size and the weight of the star detection unit can be greatly reduced, the power is reduced, and the size and the weight of the star sensor are further reduced.

Similarly, the center of mass position drift caused by position change caused by mechanical vibration or temperature change of the detector on the image plane is prevented from being different, and the imaging precision of the camera is influenced. The objective lens combination unit comprises a first lens, a second lens, a third lens, a fourth lens, a fifth lens, a sixth lens and a seventh lens which are sequentially arranged along the same optical axis in the star light incidence direction, wherein the combination of the first lens, the second lens and the third lens is positive focal power and satisfies the following relation: phi is less than or equal to 1.75 phi₁₂₃2.65 phi or less, the combination of the fourth lens, the fifth lens and the sixth lens has negative focal power and satisfies the relation: less than or equal to phi 3.55 phi₄₅₆Less than or equal to 4.25 phi; the optical power of the seventh lens satisfies the following relationship: phi is less than or equal to 1.55 phi₇Less than or equal to 2.25 phi, wherein phi₁₂₃Is the combined focal power of the first lens, the second lens and the third lens₄₅₆Is the combined focal power of the fourth lens, the fifth lens and the sixth lens₇Is the focal power of the seventh lens, phi is the overall focal power of the optical system. By improving the specific structure of the objective lens combination unit, the detection spectrum width is improved to 550nm, the entrance pupil diameter phi is enlarged by 12.85mm, the volume of the objective lens combination unit is reduced by 43.8% under the condition of equal optical path length, and the weight and the size of the whole star sensor are further reduced. Moreover, the field of view is constantThe star point target optical signals are transmitted through the seven lenses and imaged on an image plane in an image space telecentric mode, the included angle between the principal ray and the optical axis is not more than 0.5 degrees, the mass center position drift caused by the position change of the multi-dimensional detection unit on the image plane due to mechanical vibration or temperature change is avoided, and the high imaging precision is ensured.

In the prior art, the camera attitude and the gyroscope attitude are respectively determined by using an independent processing system, so that the device is more, the mass is large, and the miniaturization of the star sensor is not facilitated. The invention transmits the star field image data output by the multi-dimensional detection unit and the angular rate data output by the inertial detection unit to an attitude analysis unit in a mode of synchronous integration and isolation in a preset mode so as to determine the inertial attitude of the spacecraft. According to the design method of the microminiaturized star sensor, the attitude of the camera is determined by utilizing the star field image output by the star detection unit, the attitude of the spacecraft is determined by utilizing the attitude of the gyroscope to analyze the attitude of the camera, on the basis of not adding new equipment, the size and the mass of the star sensor are undoubtedly greatly reduced, and the judgment accuracy of the inertial attitude is improved, so that the microminiaturized star sensor is formed and applied to small spacecrafts such as satellites.

The invention also relates to a design method of the microminiaturized star sensor, which is characterized by comprising the following steps: the multi-dimensional detection unit capable of adjusting angles and/or displacements from five dimensions receives starlight output by the objective lens combination unit in a mode of being arranged on a non-focal plane of the objective lens combination unit, generates a starfield image based on at least one starlight, determines a gyroscope attitude based on an angular rate detected by at least one gyroscope in the inertial detection unit, and resolves the camera attitude according to the gyroscope attitude so as to determine the inertial attitude of the spacecraft. The star field image data output by the multi-dimensional detection unit and the angular rate data output by the inertial detection unit are synchronously integrated into a data stream in a predetermined mode and transmitted to a posture analysis unit, and the camera posture data and the gyroscope posture data in the data stream exist in a mode of being isolated from each other. The invention does not add new high-cost relation processing equipment, integrates the camera attitude and the gyroscope attitude into one data stream for calculation processing, reduces partial technical processing equipment and ensures that the integral manufacturing cost of the star sensor is lower.

Drawings

FIG. 1 is a schematic diagram of the overall structure of a star sensor;

FIG. 2 is a schematic view of an optical element of the star sensor;

FIG. 3 is a schematic view of a radial structure of a multi-dimensional adjustment structure in an optical unit;

FIG. 4 is a schematic axial view of a multi-dimensional adjustment structure in an optical unit;

FIG. 5 is a schematic structural view of an objective lens combining unit in the optical unit; and

FIG. 6 is a schematic diagram of the logical structure of the star sensor.

List of reference numerals

40: a housing; 100: a satellite detection unit; 200: an inertia detection unit; 300: a control unit; 400: an attitude analysis unit; 103: a multi-dimensional detection unit; 107: a field stop; 107A: a first light receiving face; 107B: a first light emitting surface; 109: an aperture diaphragm; 109A: a second light receiving face; 109B: a second light emitting surface; 110: a diaphragm unit; 111: a relay optical assembly; 112: a photoelectric sensing array; 114: a lens; 115: an adjusting frame; 116: a vertical plane; 117: a star map camera; 120: a third adjustment dimension; 130: a fifth adjustment dimension; 140: an objective lens combination unit; 141: a first lens; 142: a second lens; 143: a third lens; 144: a fourth lens; 145: a fifth lens; 146: a sixth lens; 147: a seventh lens; 148: an image plane; 149: an entrance aperture diaphragm; 150: a star information data converter; 210: a set of gyroscope sensors; 211: a first axis sensor; 212: a second axis sensor; 213: a third axis sensor; 220: a gyroscope data converter; 410: an image processor; 420: a graphics data memory; 430: a gyroscope data processor; 440: a gyroscope data storage; 450: a pose data corrector; 460: a gesture analysis processor; 461: a first gesture gate; 462: a first self-initializing processor; 463: a second self-initializing processor; 470: an evaluation system; 480: a command control data interface; 481: an instruction processor; 482: a counter; 483: a serial port.

Detailed Description

The following detailed description is made with reference to fig. 1 to 6 of the drawings.

In the present invention, a star field refers to a group of visible stars in an observation field of any size. Such as the field of view of a keplerian telescope. A starfield image refers to an image formed and recorded from detected starlight in the field of view.

The inertial attitude refers to the flight attitude of the spacecraft under the action of inertia. The attitude does not refer to the attitude of the spacecraft, but several data information. For example, output information of an inertial measurement unit (gyroscope) can be used for providing accurate attitude information for the spacecraft.

A non-focal plane is a plane perpendicular to the optical axis and not containing a focal point. When a beam of light parallel to the main optical axis passes through the convex lens and intersects at a point, the point is called the "focal point". The plane passing through the focal point and perpendicular to the optical axis is called the "focal plane".

The focus adjustment means that light is converged at one point by adjusting the distance between the lens and the imaging plane.

The focal power, also known as diopter, is equal to the difference between the image and object beam convergence, and characterizes the ability of the optical system to deflect light. The 1 focal power means that parallel rays form a focus at 1 meter after being refracted by the lens, and the focal length is in m (meter) unit. m is^-1The value is the optical power value. Obviously, the optical power is inversely proportional to the focal length. In both lenses, the negative power has two concave surfaces and the positive power has two convex surfaces.

Example 1

As shown in fig. 1, the present invention provides an ultra-miniaturized star sensor that can be either an initialized star sensor or a low-cost ultra-miniaturized star sensor. The invention also provides a spacecraft with the subminiature star sensor, such as a miniature aerospace vehicle with the subminiature star sensor. The invention also provides a multi-dimensional adjustment star image acquisition device or an optical system for the subminiature star sensor.

A microminiaturized star sensor at least comprises a star probe unit 100, an inertial probe unit 200 and a control unit 300, as shown in FIG. 1. The satellite detection unit 100, the inertia detection unit 200 and the control unit 300 are respectively connected by electric signals, so that the satellite information output by the satellite detection unit 100 and the gyroscope information output by the inertia detection unit 200 are synchronously integrated into a data stream and are sent to the control unit 300.

Preferably, the star probe unit 100, the inertial probe unit 200 and the control unit 300 are electrically and electrically connected with a flight computer in the aerospace vehicle, respectively. An attitude analysis unit in the aerospace vehicle is responsive to the data stream to resolve and determine an attitude of the aerospace vehicle based on the star information and the gyroscope information. Preferably, the control unit 300 adjusts devices in the star detection unit in response to a control instruction of the attitude analysis unit 400, thereby improving the clarity of imaging the stars of the star detection unit 100. The control unit 300 controls and adjusts the gyroscopes in the respective directions in the inertia detection unit in response to the control instruction of the attitude analysis unit 400.

As shown in fig. 1, the star sensing unit 100 is disposed at a star light receiving end of the star sensor, and is used for collecting star light in the universe. The star information is a clear cosmic star image after radiation filtering and correction. The star detection unit 100 includes an objective lens combination unit 140 for starlight input, a star map camera 117 capable of multi-dimensional adjustment to receive star information, and a star information data converter 150. The objective lens combination unit 140 transmits the received starlight to the light receiving end of the star camera 117. The star map camera 117 images the star light to generate a star field image of the stars for comparison with the star catalogue to determine the camera pose of the aerospace vehicle. The star detection unit 100 includes a star information data converter for converting star images into numbers representing star positions and intensities.

The inertial detection unit 200 is disposed at a detection end of the star detection unit 100, and is configured to collect gyroscope information of the aerospace vehicle. The gyroscope information includes gyroscope reference information, angular velocity, and temperature information. The gyroscope is a micro-electromechanical gyroscope. Preferably, the inertial detection unit 200 includes a gyroscope data converter 220 for converting the gyroscope angular rate data into a number representative of the gyroscope angular rate.

Preferably, as shown in fig. 6, the gyro sensor group 210 in the inertia detection unit 200 is a three-axis gyro sensor group for acquiring the inertia in X, Y and Z directions. The three-axis gyroscope sensor group includes at least a first axis sensor 211 for acquiring angular rates along a first axis, a second axis sensor 212 for acquiring angular rates along a second axis, and a third axis sensor 213 for acquiring angular rates along a third axis. Preferably, the gyro sensor group 210 further includes an application specific integrated circuit ASIC for outputting signals. For example, a first asic connected to a first axis sensor outputs a first signal, a second asic connected to a second axis sensor outputs a second signal, and a third asic connected to a third axis sensor outputs a third signal.

Preferably, the first, second and third application specific integrated circuits are connected with the gyroscope data converter 220 and/or the control unit 300. The first signal, the second signal and the third signal are transmitted to the gyroscope data converter 220 and/or the control unit 300.

Preferably, the control unit 300 includes a command circuit. The command circuit selectively synchronizes the output information of the inertial detection unit 200 and the star detection unit 100 according to a predetermined pattern so that the two output information are distinguished from each other in the integration process. Preferably, the command circuit is a programmable logic device, and the programmable logic device is used for realizing the selective synchronous integration of the output information of the inertial detection unit and the star detection unit in a preset mode. The present invention integrates the output information of the inertial detection unit synchronously with the output information of the star detection unit 100 in the data stream. The data stream may be in digital form or in any other data format that can be used, such as analog data and optical data. Preferably, the data transmission of the present invention is quaternion coordinates. The control unit 300 may include a field programmable gate array, a command circuit and an isolation circuit. The isolation circuit isolates the command stream B from the stream of data C. Then, the command stream B is input to the command circuit 202, and a stream of data C is output from the command control circuit. The data stream C comprises the outputs a and G of the star detection unit and the pipe star detection unit, respectively. The command circuit selectively and synchronously integrates the outputs a and G in a predetermined pattern in the data stream C in order to isolate these outputs from each other during their integration. The programmable logic device implements selective synchronous integration of outputs a and G in a predetermined pattern. The predetermined mode is set by information from the command control data interface 470 and the information is contained in the command stream B.

For example, the data stream A output by the star probe unit includes bits C0-C11 interleaved with the gyroscope system output G of bits M0-M3. Bits C0-C11 may or may not represent any camera data, or the beginning of camera data, nor a digital representation of pixel intensity. The M0-M3 bits may not represent any gyroscope data, the beginning of gyroscope data, or gyroscope data itself. The programmable logic device sets a predetermined pattern and based on the predetermined pattern, the data stream packager interleaves the output of the star detection unit a and the output of the pipe star detection unit G into a data stream C.

Command stream B typically includes command, synchronization and clock data. The synchronization data S contained in the command stream B can proceed through the isolation circuit and arrive at the programmable logic device. The synchronization data S in the command stream B instructs the programmable logic device to selectively synchronize integration of the star probe unit output and the inertial probe unit output in the data stream C in a predetermined pattern. Based on the synchronization data S, the programmable logic device directs the gyroscope control apparatus to output a command to the inertial detection unit as part of the signal E, or in the case where a gyroscope data converter is included in the system. The commands control when the inertial detection unit or gyroscope data converter can accept the gyroscope output G so that camera data bits and gyroscope data bits can be inserted in the serial digital data stream C.

Preferably, the command circuit may further comprise at least one acquisition device register for setting an image acquisition rate, power of the star probe unit, and a first control device for controlling the acquisition device in response to the acquisition device register setting. Preferably, the command circuit may further comprise at least one gyro register for setting gyro power and gyro reference information, and a second control device for controlling the inertial detection unit in response to the gyro register. Preferably, the command circuit may further comprise a data stream packer for interleaving the output information of the satellite detection unit and the inertial detection unit into a digital data stream. Preferably, the second control device is capable of controlling the acquisition time at which the inertial detection unit acquires the gyroscope data.

Preferably, as shown in fig. 6, the attitude analysis unit 400 within the aerospace vehicle includes at least one image processor 410 and a graphic data memory 420. The image processor 410 includes an image reader and a star information processor, among others. The image reader is used for reading the stored star field image. The star information processor is used for generating star positions from the read star field images. Preferably, the image processor further comprises a sidereal directory and a position comparison processor. The position comparison processor is used for comparing the generated star position with the star catalogue. A pose analysis processor 460 within the star detection unit generates camera pose information in response to the position information output by the position comparison processor. Preferably, the attitude analysis unit 400 further includes a first self-initializing processor 462 for converting the attitude of the aerospace vehicle to an initial star position. After the attitude analysis unit 400 receives the data in the form of synchronous integrated serial digital data sent by the control unit 300, the image processor 410 outputs the processed image information to the star field image in a parallel manner. The graphics data memory 420 stores the starfield images in parallel digital form.

Preferably, as shown in fig. 6, the attitude analysis unit 400 within the aerospace vehicle includes at least one gyroscope data processor 430 and a gyroscope data memory 440. Preferably, gyroscope data processor 430 comprises a gyroscope rate processor including a gyroscope data reader for reading stored gyroscope angular rate data. The gyroscope rate processor may further comprise: a gyroscope compensator for processing the gyroscope angular rate data and generating a compensated gyroscope rate; and the gyroscope integrator is used for integrating the compensated gyroscope rate and generating the gyroscope attitude. Preferably, the pose analysis unit 400 further comprises a second self-initializing processor 463 for providing camera pose information to the pose analysis processor 460. Preferably, attitude analysis unit 400 within the aerospace vehicle further includes a first attitude gate 461 that prevents attitude analysis processor 460 from receiving gyroscope attitude information when integrated attitude analysis processor 460 is powered up. Preferably, in case the first gesture gate 461 is turned on, the gyro data processor 430 transmits the data to the gesture analysis processor 460. With the first gesture gate 461 turned off, the data of the gyro data processor 430 cannot be transferred to the gesture analysis processor 460. For example, when the posture analysis unit 400 is powered on, N =1, the first posture gate 461 is opened. N is the number of times the sensor of the present invention processes data.

Preferably, as shown in fig. 6, the attitude analysis unit 400 in the aerospace vehicle further includes an attitude analysis processor 460 and an attitude data corrector 450. Attitude analysis processor 460 provides the attitude of the aerospace vehicle in quaternion coordinates. Attitude data corrector 450 is used to estimate aircraft attitude errors to correct for gyroscope errors, including drift, scale factors, and bias errors. Preferably, both the pose analysis processor 460 and the pose data corrector 450 may be kalman filters, in particular, a square root 27 kalman filter and/or a state kalman filter.

Preferably, as shown in fig. 6, the attitude analysis unit 400 within the aerospace vehicle further includes a command and control data interface 480. The command control data interface 480 is connected to the control unit 300, the star probe unit 100, the inertia probe unit 200, and the attitude analysis processor 460, respectively. Command control data interface 480 includes serial port 483, counter 482, and instruction processor 481. Serial port 483 is used to reformat the signal representing the attitude of the aerospace vehicle and the attitude error signal representing the aerospace vehicle. The counter 482 is used to count the number of times the attitude of the aerospace vehicle has been transmitted. Instruction processor 481 is used to distribute commands based on command type.

Preferably, as shown in fig. 6, the attitude analysis unit 400 within the aerospace vehicle further includes an evaluation system 470. The evaluation system 470 is connected to the image processor 410 and the instruction processor 481, respectively. Wherein the image processor 410 transmits the image processing information to the evaluation system 470. Instruction processor 481 transmits the assigned command to evaluation system 470. With the evaluation system 470 activated, the attitude analysis processor 460 generates an aerospace vehicle absolute attitude. The absolute attitude of the aerospace vehicle is determined by the star field images acquired by the star detection unit at high frequency.

The connection relationship in the attitude analysis unit 400 is: the control unit 300 is connected to the image processor 410, the graphic data memory 420, the gyro data processor 430, and the gyro data memory 440 through signal lines, respectively. The image processor 410 and the posture data corrector 450 are connected by a signal line. The gyroscope data processor 430 is in signal connection with the gesture analysis processor 460 via a first gesture gate 461. The first self-initialization processor 462 is connected to the first gesture gate 461 and the gesture analysis processor 460 via signals. The signal input terminal of the second self-initialization processor 463 is connected to the signal input terminal of the pose data corrector 450 and the signal output terminal of the pose analysis processor 460, respectively. The signal output terminal of the second self-initialization processor 463 and the signal output terminal of the gyroscope data processor 430 are connected to the input terminal of the command control data interface 480 through signal lines. An output terminal of the command control data interface 480 is connected to an input terminal of the graphic data memory 420 and a signal input terminal of the control unit 300, respectively. The signal output terminal of the posture data corrector 450 is connected to the signal input terminal of the posture analysis processor 460 and the input terminal of the command control data interface 480 through signal lines, respectively. The signal input of the evaluation system 470 is connected to the signal output of the image processor 410 and the signal output of the instruction processor 481 via signal lines, respectively.

The evaluation system 470 includes an enabler for enabling the evaluation system 470 and generating a self-scoring continuous frequency command for output to the command control data interface 480. Enablers are connected to image processor 410 and instruction processor 481, respectively. The image processor 410 is connected to the data entry of the first comparator. The data outlet of the first comparator is connected with the data inlet of the second comparator. The data outlet of the second comparator is connected to the data inlet of the error handler. The data egress of the error handler is connected to a command control data interface 480.

When the command control data interface 480 receives the self-scoring continuous frequency command, the command control data interface 480 forwards the continuous frequency command to the controller unit 300 to increase the star camera rate for starfield image acquisition to a continuous frequency. The continuous frequency of the particular camera of the present invention is the fastest star field image acquisition rate. The continuous frequency star camera attitude represents the absolute attitude of the aerospace vehicle determined over a limited period of time of continuous star camera star field image acquisition. The continuous frequency star camera pose is determined by the image processor in the same manner as the camera pose. The image processor forwards the continuous frequency star camera poses to the first comparator.

The continuous frequency star camera pose is distinguished from the camera pose in that the continuous frequency star camera pose is determined by continuous view image acquisition of the star camera for a limited time; the camera pose is determined every five minutes. A first comparator compares the time-varying continuous frequency star camera pose with the time-varying aerospace vehicle pose and provides a self-scoring error output. The second comparator compares the self-scoring error output to an expected error threshold and may provide a threshold deviation output. The required error threshold is a preset threshold. To improve accuracy, the error threshold is set to a smaller value. If the self-scoring error output is greater than the desired error threshold, the error processor outputs a self-scoring low frequency command to the command control data interface 480 to increase the star map camera rate at which the star field images are acquired to a greater frequency. For example, less than every five minutes.

Generating the star camera pose involves determining a star position from the observation data and comparing the star position to a star catalogue. In conventional techniques, the process is more computationally intensive (the so-called "space loss" problem) if no a priori knowledge of the attitude of the aerospace vehicle is available. In this case, the generated star location must be compared to the entire star catalogue. This is a problem that conventional systems encounter each time a starfield image is acquired. To avoid this problem, the present invention uses previously determined aerospace vehicle attitude data after initial power up. Since the aerospace vehicle attitude has already been determined (at least once). Therefore, in order to alleviate the "space loss" problem after power-up, i.e., self-initialization is performed. The star information processor receives the aerospace vehicle attitude from the second self-initializing processor 463. In this case, i.e., when N >1, the star information processor converts the aerospace vehicle pose to an initial star position, thereby simplifying the star identification process by narrowing down the areas in the star catalogue that must be searched by the star information processor to find a match.

Preferably, the star probe unit 100, the inertial probe unit 200, the control unit 300, and the attitude analysis unit 400 are respectively provided with a housing 40 adapted to the configuration thereof. The housing 40 may be of any suitable shape. Preferably, the star probe unit 100, the inertial probe unit 200, and the control unit 300 may be connected to the attitude analysis unit 400 through the connection device 10. Preferably, the connection device 10 may be any suitable connector known in the art, including a cable connection providing a serial link.

Preferably, the star information data converter 150 and the gyroscope data converter 220 in the present invention may be a processor, an application specific integrated chip, or a computer having a corresponding data conversion function. The image processor 410, the gyroscope data processor 430, the pose data corrector 450, the pose analysis processor 460, the first self-initialization processor 462, the second self-initialization processor 463, the evaluation system 470, and the instruction processor 481 may each be a processor, an application specific integrated chip, or a computer loaded with a corresponding computer program, capable of performing a corresponding data processing function or an instruction generating function. The graphic data memory 420 and the gyro data memory 440 may be hardware such as a magnetic disk, a hard disk, a memory chip, a computer, etc. having a storage function. The first gesture gate 461 may be a logical gate. Counter 482 is a logic module, processor, or microchip having a counting function.

The main working method of the star sensor comprises the following steps:

when the star probe unit 100, the inertia probe unit 200, and the control unit 300 are powered on, at this time N =1, the first posture gate 461 is turned on, preventing initial gyro posture data of the inertia probe unit from being input to the posture analyzing processor 460. At this time, the attitude of the aerospace vehicle is the same as the attitude of the camera. When N >1, the first gesture gate 461 is turned off and the gesture analysis processor 460 receives and processes the camera gesture and the gyroscope gesture. Namely, the attitude of the camera is analyzed through the attitude of the gyroscope, so that the inertial attitude of the spacecraft is determined. To reduce the spatial loss problem at N =1, the pose analysis processor may utilize a spatial loss pose determination algorithm developed by Mortari. The Mortari algorithm identifies stars in the star field image from a star catalogue of 1500 stars to help identify the line of sight of the star map camera without prior knowledge of the attitude of the aerospace vehicle. The inertial detection unit 200 continuously maintains the attitude information of the spacecraft until a new star field image update of the star detection unit is obtained. The gyroscope attitude of the inertial detection unit 200 drives the aerospace vehicle to stably fly and generate the absolute attitude of the aerospace vehicle. The inertial detection unit samples the inertia or angular velocity at a high frequency of about 320 Hz.

After determining the absolute attitude of the aerospace vehicle, attitude analysis processor 460 communicates to the main computer in real time at a frequency of approximately 5 Hz. The star detection unit system acquires new star field images and determines camera poses for correcting pose errors including but not limited to gyroscope drift, scale factors, bias errors and angular random walk. The attitude of the aerospace vehicle is updated when the star detection unit obtains a new star field image and determines a new camera attitude.

Wherein after the attitude analysis processor 460 computes the spacecraft absolute attitude, the attitude data corrector 450 computes the error of the spacecraft attitude. The second self-initializing processor 463 sends the absolute attitude of the spacecraft to the image processor 410. The first self-initialization processor 462 sends the camera pose to the image processor 410 for an update of the spacecraft pose. Pose analysis processor 460 corrects for gyroscope errors in previously determined gyroscope poses. When pose analysis processor 460 receives the new gyroscope pose, in the event that the spacecraft pose has been determined, pose analysis processor 460 will generate a new spacecraft pose based on the gyroscope pose and spacecraft pose calculations. At this time, the camera attitude is not updated yet, and the spacecraft attitude is determined by the gyroscope attitude.

The star detection unit of the present invention is shown in fig. 2. The star detection unit 100 specifically includes a star objective combination unit 140, a diaphragm unit 110, and a multi-dimensional detection unit 103. The star objective lens combination unit 140 is a star light input unit composed of a plurality of objective lenses, and can input the taken star light. Preferably, the star objective lens combination unit 140 further includes a refractor, a reflector, and a lens. The diaphragm unit 110 includes at least a field diaphragm 107 and an aperture diaphragm 109. Wherein, the field stop 107 is arranged on the optical axis a of the star objective lens combination unit 140. The aperture stop 109 is provided in the exit pupil of the star objective lens combination unit 140, or at the position of the real image. A relay optical assembly 111 is arranged between the field stop 107 and the multi-dimensional detection unit 103. The relay optical assembly 111 transmits the image transmitted by the star objective lens combination unit 140 from the focal plane of the objective lens to the multi-dimensional detection unit 103. Preferably, the image magnification of the relay optical assembly 111 is 1.2 to 2, preferably 1.4 to 1.6. More preferably, the image magnification of the relay optical assembly 111 is about 1.5.

Specifically, the multi-dimensional detection unit 103 is not directly disposed on the focal plane of the objective lens combination unit, but is located at a distance from the focal point of the objective lens combination unit, and the field stop 107 is located on the focal plane of the star objective lens combination unit 140. Namely, the position of the multi-dimensional detecting unit 103 of the present invention is set opposite to that of the conventional star sensor. In the conventional technique, if the detector is directly disposed at the focal plane of the objective lens combination unit, the residual radiation may be diffused and illuminated by the detection device directly disposed at the focal plane of the star objective lens combination unit 140. The multi-dimensional detection unit 103 and the relay optical assembly 111 are combined, so that residual radiation scattered by the wall of the main objective in the star objective combination unit 140 can be attenuated, and the defect that the multi-dimensional detection unit 103 diffuses and illuminates the residual radiation is avoided.

The relay optical assembly 111 in the present invention is capable of substantially reducing the residual radiation reaching the multi-dimensional detection unit 103, which is scattered by the edges of the aperture stop 109, the walls of the main objective and the light shield arranged upstream of the field of view and stops the diffuse transmission. Preferably, a light shield is arranged outside the star detection unit and is in a conical barrel shape. The source radiation outside the detection field of view can be blocked, for example, the solar radiation is blocked and prevented from reaching the multi-dimensional detection unit 103 along the light path, and the multi-dimensional detection unit 103 is prevented from being blind and incapable of working normally. Preferably, the relay optical assembly may be a relay lens.

As shown in fig. 2, the field stop 107 has a first light receiving surface 107A and a first light emitting surface 107B. The aperture stop 109 has a second light receiving surface 109A and a second light emitting surface 109B. Preferably, the first light emitting surface 107A and the second light receiving surface 109A are provided with a high absorption layer. The high absorption layer can be an absorptive reflection layer or an absorptive diffusion layer formed by coating an absorbent. Preferably, the first light receiving surface 107A and the first light emitting surface 107B, and the second light receiving surface 109A and the second light emitting surface 109B are each provided with a high absorption layer. The highly absorbing layer is capable of absorbing a substantial amount of incident radiation with an absorption of more than 90%, preferably 95%. Even more, the absorptivity of the high absorption layer is as high as 98-99%. Preferably, the absorptive reflective layer reflects a portion of light absorbed by the ordinary surface without diffusion. The absorptive diffusion layer is subjected to a diffusion treatment of a portion of the ordinary surface where the non-absorbed radiation is diffused.

Preferably, an absorptive reflecting layer or an absorptive diffusing layer is provided only on the first light emitting surface 107B and the second light emitting surface 109B, with an absorption rate of more than 90%, preferably 95%. This has the advantage that by increasing the degree of absorption, the diffusion of stray radiation through the optical device is limited. It is avoided that stray radiation from outside the field of view may reach the multi-dimensional detection unit 103. According to the invention, through the arrangement of the diaphragm unit 110, stray light radiation is reduced, and the imaging definition of the star map is further improved.

Preferably, the multi-dimensional detection unit 103 can move in multiple dimensions to accommodate the displacement variation of the image received by the starlight. As shown in fig. 2 to 4, the structure of the multi-dimensional detection unit 103 is shown from 3 angles. The multi-dimensional detection unit 103 comprises a star camera 117 and an adjustment block 115. Star camera 117 is fixed on adjustment frame 115 and faces relay optics assembly 111. As shown in fig. 3, the third axis 3 of the adjustment frame is parallel to the optical axis, and the second axis 2 and the first axis 1 are perpendicular to the optical axis, respectively. The adjustment frame 115 can be translated in the first axis direction and the second axis direction, respectively, to perform two-dimensional displacement adjustment. The adjustment frame 115 is arranged obliquely and at an angle to the vertical plane 116. The vertical plane 116 is parallel to a plane formed by the first and second axes. The adjusting frame 115 is provided with an angle adjusting device having multiple degrees of freedom, such as a spherical angle adjuster capable of multi-angle adjustment. Thus, the adjustment frame 115 is movable in rotation about a central axis in the first axial direction to perform the adjustment of the third adjustment dimension 120. Preferably, the angular rotation amplitude is about 120 degrees. The adjustment block 115 can adjust the pitch angle between itself and the vertical plane 116 through the angle adjuster, thereby performing pitch adjustment in the fourth dimension. Preferably, the amplitude of the pitch angle is less than 60 degrees, preferably less than 45 degrees, more preferably less than 30 degrees.

The star camera 117 includes a lens 114 and a photo sensor array 112. Preferably, the camera within the star detection unit 100 is preferably an Active Pixel Sensor (APS). The active pixel sensor is a photosensor array, namely, photosensor array 112. Each sensor has a local amplifier and row and column addressing functions. The size and weight of the camera system can be greatly reduced compared to the conventional imaging device CCD. Because fewer control and drive circuit electronics are required for an APS star map camera than a CCD. Active pixel sensors can integrate analog and digital functions on the same chip or chip. Also, compared to high capacitance CCDs, power can be reduced using active pixel sensors since active pixel sensors typically use standard 5Vdc and 3.3Vdc power supplies. Active pixel sensors are relatively radiation resistant, as they can be fabricated in processes such as silicon on insulator. And the active pixel sensor is insensitive to charge transfer efficiency effects typically associated with radiation-damaged CCDs, resulting in better imaging. Preferably, the lens 114 is disposed upstream of the optical path of the photo sensor array 112, and the focal point of the lens 114 converges on the photo sensor array 112. Preferably, the photo sensor array 112 is connected to at least one displacement device for translational movement in the direction of the third axis 3 for angular adjustment in a fifth adjustment dimension 130. The spherical angle adjuster and the displacement device are connected with a star information data converter 150 in the star detection unit through a communication line. The star information data converter 150 is connected to the control unit 300 through a communication line so that the sphere angle adjuster and the displacement device can adjust the angle and the displacement according to command information of the control unit 300. Preferably, the variables for each angle and displacement of the adjustment block 115 can be determined. For example, the photo sensor array 112 is offset by 2 arcseconds. The displacement adjustment of the adjustment frame 115 has an advantage in that the optical axis of the lens 114 can be repositioned. First, it is beneficial to adjust the focus, so that the camera formed by the photo sensor array 112 and the lens 114 can realize auto-focusing, and have focusing capability, in case the input star light is completely collimated. Refocusing of the optical image may be achieved by adjusting the focus; second, the ability to translate the position of the optical image by focus adjustment significantly enhances the resolution of the optical image (e.g., super-resolution image) received using the following method: such as suitable image post-processing techniques (e.g., phase diversity post-processing techniques).

Preferably, the objective lens combination unit 140 in the star probe unit 100 includes a first lens 141, a second lens 142, a third lens 143, a fourth lens 144, a fifth lens 145, a sixth lens 146, and a seventh lens 147, which are sequentially arranged along the same optical axis in the star light incident direction. The combination of the first lens 141, the second lens 142, and the third lens 143 has positive refractive power, and satisfies the following equation: phi is less than or equal to 1.75 phi₁₂₃Less than or equal to 2.65 phi. The combination of the fourth lens 144, the fifth lens 145, and the sixth lens 146 is negative power, and satisfies the following equation: less than or equal to phi 3.55 phi₄₅₆Less than or equal to 4.25 phi. The power of the seventh lens 147 satisfies the following formula: phi is less than or equal to 1.55 phi₇Less than or equal to 2.25 phi. Wherein phi₁₂₃Is the combined power of the first lens 141, the second lens 142 and the third lens 143, phi₄₅₆Is the combined power of the fourth lens 144, the fifth lens 145 and the sixth lens 146₇Is the power of the seventh lens 147The optical power of the optical system.

Preferably, the first lens 141 is provided with an entrance aperture stop 149. Preferably, the aperture stop 109 is disposed at the position of the image plane 148 of the objective lens combination unit 140, so as to receive the star phase information of the star light more clearly.

When a star point target optical signal passes through the first lens 141 limited by the entrance aperture stop 149, the first lens 141 and the second lens 142 form three separated first lens groups as a positive lens group by the third lens 143, light is converged to mainly correct spherical aberration, coma aberration and axial chromatic aberration, and a certain astigmatism correction capability is realized on two adjacent surfaces of the second lens 142 and the third lens 143; the fourth lens 144, the fifth lens 145 and the sixth lens 146 form a negative lens group, which is mainly used for correcting astigmatism, distortion and the like; the seventh lens 147 has negative focal power, corrects residual distortion, and mainly reduces the principal ray to realize an image-side telecentric optical path. The seven lenses correct curvature of field by a reasonable distribution of optical powers. The out-of-symmetry variation of the system architecture corrects for the vertical axis aberrations introduced by the imaging target being at infinity. The fixed star point target optical signals of different view fields are transmitted through the seven lenses and imaged on the image plane 148 in an image space telecentric mode, the included angle between the principal ray and the optical axis is not more than 0.5 degrees, the mass center position drift caused by the position change caused by the mechanical vibration or temperature change of the multi-dimensional detection unit on the image plane 148 is avoided, and the high imaging precision is ensured.

The working spectrum of the objective lens combination unit 140 in the conventional star sensor is generally 300nm, and data is obtained through experiments, so that the detection spectrum width of the objective lens combination unit 140 reaches 550nm, which is improved by 1.8 times. The detection capability of the optical system of the conventional star sensor is equal to the detection capability of the optical system of the conventional star sensor with the entrance pupil diameter phi of 12.85mm and the entrance pupil diameter phi of 17.24 mm. The volume of the objective lens combination unit 140 is reduced by 43.8% under the condition that the optical path length is equal. The aperture of the optical system is reduced on the premise of ensuring the detection capability, the stray light suppression of the whole star sensor is facilitated, and the weight and the size of the whole star sensor are further reduced. The first lens 141, the second lens 142, the third lens 143, the fourth lens 144, the fifth lens 145, the sixth lens 146 and the seventh lens 147 are all spherical lenses, so that the processing difficulty and the assembly difficulty are reduced, and the manufacturability and the assembly yield of the star sensor objective lens combination unit 140 are facilitated.

Preferably, the first lens is made of quartz, and the second lens is made of lanthanum flint glass or special flint glass. The third lens is made of lanthanum crown glass or heavy lanthanum flint glass, and the fourth lens is made of heavy flint glass. The fifth lens and the seventh lens are made of heavy lanthanum flint glass, and the sixth lens is made of heavy flint glass.

Example 2

The embodiment provides a design method of a microminiaturized star sensor. The content repeated in this embodiment from embodiment 1 is not described again.

A design method of a microminiaturized star sensor comprises the following steps: the multi-dimensional detection unit 103 capable of adjusting angles and/or displacements from five dimensions receives the starlights output from the objective lens combination unit 140 in a manner of being disposed at a non-focal plane of the objective lens combination unit 140, and generates a starfield image based on at least one of the starlights.

Preferably, the multi-dimensional detection unit 103 receives the starlight output by the objective lens combination unit 140 through a diaphragm unit 110, wherein at least one diaphragm in the diaphragm unit 110 is disposed on a focal plane of the objective lens combination unit 140, so as to clearly receive the starlight or the starfield image.

Preferably, the diaphragm unit 110 includes a field diaphragm 107 and an aperture diaphragm 109 arranged on the same optical axis, the aperture diaphragm 109 is disposed on the focal plane of the objective lens combination unit 140, and the star light output by the objective lens combination unit 140 is transmitted to the multi-dimensional detection unit 103 through the aperture diaphragm 109 and the field diaphragm 107 in sequence.

Preferably, the multi-dimensional detection unit 103 receives the star light output by the diaphragm unit 110 in a manner of attenuating the residual radiation by the relay optical component 111.

Preferably, the multi-dimensional detection unit 103 includes a star map camera 117 and an adjusting block 115 for mounting the star map camera 117, the adjusting block 115 moves and/or rotates in four dimensions so as to enable the star map camera 117 to adjust a receiving angle to receive target star light which is not directly input, and the photo sensor array 112 for imaging in the star map camera 117 translates based on the dimension of the optical axis direction, so that the photo sensor array 112 and the lens 114 for receiving star light in the star map camera 117 realize focusing adjustment.

Preferably, the objective lens combination unit 140 includes a first lens 141, a second lens 142, a third lens 143, a fourth lens 144, a fifth lens 145, a sixth lens 146, and a seventh lens 147, which are sequentially arranged along the star light incidence direction with the optical axis. The combination of the first lens 141, the second lens 142, and the third lens 143 has positive refractive power, and satisfies the relationship: phi is less than or equal to 1.75 phi₁₂₃Less than or equal to 2.65 phi. The combination of the fourth lens 144, the fifth lens 145, and the sixth lens 146 is negative power, and satisfies the relationship: less than or equal to phi 3.55 phi₄₅₆Less than or equal to 4.25 phi. The power of the seventh lens 147 satisfies the following relationship: phi is less than or equal to 1.55 phi₇Less than or equal to 2.25 phi. Wherein phi₁₂₃Is the combined power of the first lens 141, the second lens 142 and the third lens 143, phi₄₅₆Is the combined power of the fourth lens 144, the fifth lens 145 and the sixth lens 146₇Is the power of the seventh lens 147, phi is the overall power of the optical system.

Preferably, the camera pose is determined based on the star field image output by the multi-dimensional detection unit 103. The gyro attitude is determined based on the angular rate detected by at least one gyro in the inertial detection unit 200. And resolving the camera attitude according to the gyroscope attitude so as to determine the inertial attitude of the spacecraft.

Preferably, the star field image data output by the multi-dimensional detection unit 103 and the angular rate data output by the inertial detection unit 200 are transmitted to the attitude analysis unit 400 in a manner of being synchronously integrated in a predetermined pattern and isolated from each other to determine the inertial attitude of the spacecraft.

It should be noted that the above-mentioned embodiments are exemplary, and that those skilled in the art, having benefit of the present disclosure, may devise various arrangements that are within the scope of the present disclosure and that fall within the scope of the invention. It should be understood by those skilled in the art that the present specification and figures are illustrative only and are not limiting upon the claims. The scope of the invention is defined by the claims and their equivalents.

Claims

1. A microminiaturized star sensor is characterized in that a star detection unit (100) of the star sensor at least comprises an objective lens combination unit (140) and a multi-dimensional detection unit (103),

a multi-dimensional detection unit (103) capable of adjusting angles and/or displacements from five dimensions is arranged in a non-focal plane of the objective combination unit (140) at a distance from a focal point of the objective combination unit (140),

the multi-dimensional detection unit (103) adjusts the receiving angle in a mode of adjusting the angle and/or displacement based on five dimensions, receives the star light output by the objective lens combination unit (140), and generates a star field image based on at least one star light.

2. The microminiaturized star sensor according to claim 1, characterized in that the star sensor further comprises an aperture unit (110),

the diaphragm unit (110) is arranged between the objective combination unit (140) and the multi-dimensional detection unit (103), wherein,

at least one diaphragm in the diaphragm unit (110) is arranged on a focal plane of the objective lens combination unit (140), and the multi-dimensional detection unit (103) receives starlight output by the objective lens combination unit (140) through the diaphragm unit (110).

3. The microminiaturized star sensor according to claim 2, wherein the aperture unit (110) comprises a field stop (107) and an aperture stop (109) arranged coaxially,

the aperture stop (109) is arranged in the focal plane of the objective combination unit (140),

the starlight output by the objective lens combination unit (140) is transmitted to the multi-dimensional detection unit (103) through the aperture diaphragm (109) and the field diaphragm (107) in sequence.

4. The microminiaturized star sensor according to one of the preceding claims,

the multi-dimensional detection unit (103) receives the star light output by the diaphragm unit (110) in such a way that the residual radiation is attenuated by the relay optical assembly (111).

5. The microminiaturized star sensor according to one of the preceding claims,

the multi-dimensional detection unit (103) comprises a star atlas camera (117) and an adjustment frame (115) for mounting the star atlas camera (117),

the adjustment block (115) moves and/or rotates in four dimensions so that the star map camera (117) adjusts star light reception angles in four dimensions to reposition the optical axis,

the photo-sensing array (112) for imaging within the star camera (117) is translated based on the dimensions of the optical axis direction such that the focal plane between the photo-sensing array (112) and the star light receiving lens (114) within the star camera (117) is repositioned and focus adjusted.

6. The microminiaturized star sensor according to any one of the preceding claims, wherein the objective lens assembly unit (140) comprises a first lens (141), a second lens (142), a third lens (143), a fourth lens (144), a fifth lens (145), a sixth lens (146) and a seventh lens (147) arranged in sequence along the direction of the star light incidence on the optical axis,

the combination of the first lens (141), the second lens (142), and the third lens (143) is positive optical power, and satisfies the relationship: phi 123 is more than or equal to 1.75 phi and less than or equal to 2.65 phi,

the combination of the fourth lens (144), the fifth lens (145) and the sixth lens (146) has negative focal power and satisfies the relationship: phi is more than or equal to 3.55 phi and less than or equal to | phi 456 and less than or equal to 4.25 phi;

the optical power of the seventh lens (147) satisfies the following relationship: phi is more than or equal to 1.55 phi and less than or equal to 7 and less than or equal to 2.25 phi,

wherein phi 123 is the combined focal power of the first lens (141), the second lens (142) and the third lens (143), phi 456 is the combined focal power of the fourth lens (144), the fifth lens (145) and the sixth lens (146), phi 7 is the focal power of the seventh lens (147), and phi is the overall focal power of the optical system.

7. The microminiaturized star sensor according to one of the preceding claims, characterized in that a camera pose is determined based on the image of the star field output by the multi-dimensional detection unit (103),

determining a gyroscope pose based on an angular rate detected by at least one gyroscope in the inertial detection unit (200),

and resolving the camera attitude according to the gyroscope attitude so as to determine the inertial attitude of the spacecraft.

8. Microminiaturized star sensor according to one of the preceding claims, characterized in that the image data of the star field output by the multi-dimensional detection unit (103) and the angular rate data output by the inertial detection unit (200) are fed to an attitude analysis unit (400) in a synchronously integrated and isolated manner in a predetermined pattern for determining the inertial attitude of the spacecraft.

9. A design method of a microminiaturized star sensor is characterized by comprising the following steps:

a multi-dimensional detection unit (103) capable of adjusting angles and/or displacements from five dimensions receives starlight output by the objective lens combination unit (140) in a manner of being arranged at a non-focal plane of the objective lens combination unit (140) and generates a starfield image based on at least one of the starlight,

10. The microminiaturized star sensor related method as recited in claim 9,

star field image data output by the multi-dimensional detection unit (103) and angular rate data output by the inertial detection unit (200) are synchronously integrated into a data stream in a predetermined pattern and transmitted to a pose analysis unit (400), and camera pose data and gyroscope pose data in the data stream exist in an isolated manner from each other.