CN111076729A - Deep space spacecraft relative position measuring method based on intensity coherent detection array - Google Patents
Deep space spacecraft relative position measuring method based on intensity coherent detection array Download PDFInfo
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Abstract
The invention discloses a deep space spacecraft relative position measuring method based on an intensity coherent detection array, and belongs to the field of spacecraft optical measurement and control. The implementation method of the invention comprises the following steps: building a satellite-borne downlink laser transmitting telescope on a spacecraft to be measured, and transmitting the same-frequency laser to a foundation strength coherent array by each spacecraft simultaneously when the relative position of the spacecraft is required to be measured; constructing an intensity coherent observation unit on the ground; arranging a plurality of foundation strength coherent units into an array to form a foundation strength coherent observation system; acquiring signals by using a short-baseline low-precision foundation observation array, and calculating the photocurrent correlation between every two observation units in the observation array according to the acquired signals; constructing coordinates; acquiring a spatial frequency spectrum of a target retroreflection laser pattern, and drawing an interference pattern according to the acquired spatial frequency spectrum; and calculating to obtain the relative position of the spacecraft according to the interference pattern, and realizing high-precision relative position measurement of the deep space spacecraft. The invention can reduce the requirements of the scale and the precision of the observation array.
Description
Technical Field
The invention relates to a high-precision measurement method for relative positions of spacecrafts, in particular to a high-precision measurement method for relative positions of spacecrafts by using an interference measurement method, which can be applied to measurement of the relative positions of two spacecrafts during butting of the spacecrafts and belongs to the field of optical measurement and control of the spacecrafts.
Background
With the development of aerospace deep space exploration, the measurement and control precision requirements for the spacecraft are higher and higher. Particularly, in the process of rendezvous and docking of the spacecrafts, the relative positions of the spacecrafts need to be measured with high precision. The relative position measurement of the traditional spacecraft adopts the same-beam interference technology to measure interference fringes of radio emitted to the ground by the spacecraft on the ground so as to obtain the relative position of the spacecraft. The radio wavelength is longer, and the interference measurement result is less influenced by atmospheric turbulence and an ionized layer; but use long baseline measurement equipment; in addition, in the transmitting process, the beam width of radio waves is wide, the density of radio energy received on the ground is low, and a large-caliber antenna is required to be used for measuring the spacecraft. The caliber of the existing radio interference array antenna for deep space measurement exceeds 30m meters, the length of a measurement base line exceeds 2000 kilometers, and the difficulty of a measurement system in the construction process is high.
If the visible light can be used for replacing the interference measurement of the target position, the original measurement precision level can be achieved by only using a shorter measurement baseline, and the scale of measurement equipment is greatly reduced; and the diffraction angle of the visible light is far smaller than the width of the transmitting beam of the radio wave, the directivity of the light beam is better in the transmission process, the energy obtained on the ground is more concentrated, and enough signal to noise ratio can be obtained by adopting a smaller receiving aperture. However, the interference measurement result of the optical field amplitude is greatly influenced because the atmospheric turbulence can introduce extra phase errors into the measurement result. Meanwhile, the precision requirement of the measuring equipment is high based on the light field amplitude interference method, the difficulty of building the related measuring equipment is high under the condition of the prior art, and the light field amplitude interference measuring method is difficult to be applied in practice.
The intensity coherent array is a high-precision measuring device for measuring the angular diameter of a fixed star, and the high-frequency detector is used for collecting the light intensity fluctuation of a target at different positions, and the spatial frequency spectrum of the target is obtained by calculating the coherence of the light intensity fluctuation, so that the angular diameter of the target is measured. The intensity coherent array has the advantages of high measurement precision, low equipment precision requirement, insensitivity to atmospheric turbulence and the like.
Disclosure of Invention
The invention discloses a deep space spacecraft relative position measuring method based on an intensity coherent detection array, which aims to solve the technical problems that: the high-precision relative position measurement of the deep space spacecraft is completed by using the short-baseline low-precision foundation observation array, and the scale and precision requirements of the observation array are reduced. The invention can be applied to the fields of intersection and butt joint of deep space spacecraft, track determination and the like.
The purpose of the invention is realized by the following technical scheme.
The invention discloses a method for measuring the relative position of a deep space spacecraft based on an intensity coherent detection array. And constructing an intensity coherent observation unit on the ground. A plurality of foundation strength coherent units are arranged into an array to form a foundation strength coherent observation system, and all observation units are required to be ensured not to be on the same straight line when the array is arranged. Acquiring signals by using a short-baseline low-precision foundation observation array, and calculating the photocurrent correlation between every two observation units in the observation array according to the acquired signals; constructing coordinates; acquiring a complete space spectrum of the retroreflection laser pattern, and drawing interference fringes according to the acquired space spectrum of the retroreflection laser pattern; and realizing high-precision relative position measurement of the deep space spacecraft according to the acquired space spectrum and the drawn interference fringes.
The invention discloses a deep space spacecraft relative position measuring method based on an intensity coherent detection array, which comprises the following steps:
the method comprises the following steps: a satellite-borne downlink laser transmitting telescope is built on a spacecraft to be measured and mainly comprises a direction control part, a laser and a beam expanding telescope. When the position of the spacecraft needs to be measured, the spacecraft simultaneously emits laser with the same frequency to the foundation strength coherent array.
Step two: an intensity coherent observation unit is constructed on the ground, and comprises a light collecting surface, an optical filter, a high-frequency detector, a multiplier and data storage and calculation equipment.
Step three: and arranging the foundation strength coherent units in the second step into an array to form a foundation strength coherent observation system, wherein all observation units are required to be ensured not to be on the same straight line when the array is arranged.
Preferably, the array arrangement mode in the third step is a cross-shaped array or a Y-shaped array. The longest base line of the array is K, and the distance between coherent units in the array is delta K.
Step four: acquiring signals by using a short-baseline low-precision foundation observation array, and calculating the photocurrent correlation between every two observation units in the observation array according to the acquired signals; constructing coordinates; acquiring a complete space spectrum of a retroreflection laser pattern, and drawing interference fringes according to the acquired space spectrum of the laser pattern; and realizing high-precision relative position measurement of the deep space spacecraft according to the acquired space spectrum and the drawn interference fringes.
Step 4.1: and (5) signal acquisition.
When the relative position of the deep space spacecraft needs to be measured, instructions are injected into the spacecraft through the ground remote control equipment, the satellite-borne downlink laser transmitting telescopes of the spacecrafts emit lasers with the same frequency towards the direction of the foundation strength coherent observation system, and the included angle of each spacecraft relative to the observation system needs to be smaller than the view field of each observation unit. Each observation unit in the laser coherent array receives laser emitted by all spacecrafts, the high-frequency detector converts random fluctuation of light intensity obtained by receiving into corresponding high-frequency photocurrent in proportion, the sampling frequency of the high-frequency detector is vHz, and the maximum system error introduced by a system light collecting surface and atmospheric turbulence is required to be smaller than
Step 4.2: and calculating the current correlation.
And calculating the correlation of the high-frequency photocurrents output by A, B two observation units in the array by using a multiplier, wherein the calculation formula of the correlation of the photocurrents at different positions is as follows:
wherein IAAnd (T) is the high-frequency photocurrent of the A position observation unit at each moment, and T is the total observation time length.
And calculating the photocurrent correlation between every two observation units in the observation array according to the formula.
Step 4.3: and (5) constructing coordinates.
Constructing an orthogonal coordinate system in the observation system, taking a certain direction as an x-axis and the normal direction of the x-axis as a y-axis, recording the coordinates of all observation units as (x)k,yk) Where k is the observation unit number.
Preferably, in step 4.3, the weft direction is set to be the x-axis.
Step 4.4: a spatial frequency spectrum of the retroreflected laser pattern is acquired, and interference fringes are drawn according to the acquired spatial frequency spectrum.
Acquiring a space spectrum of a retroreflection laser pattern of the spacecraft according to the relative position between the observation units, wherein the space spectrum is calculated as follows:
γ(l,m)＝SAB；l＝xA-xB,m＝yA-yB。
and acquiring a complete space spectrum of the retroreflected laser by using all the photocurrent correlations acquired in the step 4.2, and drawing interference fringes according to the acquired space spectrum.
Step 4.5: extracting fringe spacing of interference patterns in the l direction and the m direction, recording the fringe spacing in the l direction as a, and the fringe spacing in the m direction as b, and calculating the fringe spacing d and the fringe direction theta according to the fringe spacing, wherein the calculation formula is as follows:
step 4.6: and calculating to obtain the position relation of the two spacecrafts, so that the high-precision relative position measurement of the deep space spacecraft can be realized.
The included angle between the connecting line of the two spacecrafts to be measured and the y-axis of the observation system is theta, and the distance between the two spacecrafts is theta
Wherein L is the orbit height of the spacecraft, lambda is the retro-reflection laser wavelength, the distance of the d-position interference fringes, and omega where the detector array is located is the latitude. Spacecraft spacing measurement error is
Has the advantages that:
1. the invention discloses a method for measuring the relative position of a deep space spacecraft based on an intensity coherent detection array. And constructing an intensity coherent observation unit on the ground. And arranging the foundation strength coherent units in the second step into an array to form a foundation strength coherent observation system, wherein all observation units are required to be ensured not to be on the same straight line when the array is arranged. Acquiring signals by using a short-baseline low-precision foundation observation array, and calculating the photocurrent correlation between every two observation units in the observation array according to the acquired signals; constructing coordinates; acquiring a complete space spectrum of a retroreflection laser pattern, and drawing interference fringes according to the acquired space spectrum; and calculating and acquiring the precise relative position and distance of two or more deep space spacecraft according to the distance and direction of the drawn interference fringes.
2. The existing radio interference measurement needs to construct a large-scale long baseline interference array to measure the position of a spacecraft, the measurement cost is high, and the power of a satellite-borne transmitter is high. The method for measuring the relative position of the deep space spacecraft based on the intensity coherent detection array can obtain higher measurement precision by using a small observation array by using an optical intensity coherent method, has lower precision requirement on observation equipment, can effectively complete high-precision relative position measurement on a space target, and assists the deep space spacecraft to complete tasks such as intersection butt joint, positioning navigation and the like
Drawings
FIG. 1 is a satellite-borne laser emission device of the optical intensity coherence measurement system of the present invention;
FIG. 2 is a ground based intensity correlation unit;
FIG. 3 is an intensity coherence measurement array layout;
FIG. 4 is a flow chart of the method for measuring the relative position of a deep space spacecraft based on an intensity coherent detection array disclosed by the invention;
fig. 5 is an interference fringe pattern.
In the figure: the system comprises a laser 1, a turning reflector 2, a transmitting telescope secondary mirror 3, a transmitting telescope primary mirror 4, an atmospheric turbulence 5 in the process of transmitting to the ground, a low-precision light collecting surface 6, a detector front light collecting lens 7, an optical filter 8, a high-frequency detector 9, a filter 10, a multiplier 11, a data storage device and a calculation device 12, an intensity coherence measurement device 13 and a calculation control center 14.
Detailed Description
For a better understanding of the objects and advantages of the present invention, reference should be made to the following detailed description taken in conjunction with the accompanying drawings and examples.
Example 1: the lunar orbit lunar landing capsule and the relative position measurement of the orbit capsule are taken as an example.
As shown in fig. 4, the method for measuring the relative position of the deep space spacecraft based on the intensity coherent detection array disclosed in this embodiment includes the following specific steps:
the method comprises the following steps: the deep space spacecraft relative position measuring system based on the intensity coherent detection array needs to be respectively provided with a satellite-borne downlink laser transmitting telescope on a lunar chamber and a track chamber, and the schematic diagram of the laser transmitting telescope is shown in figure 1: light emitted by the laser 1 is reflected to the secondary mirror 3 of the transmitting telescope through the deflection reflecting mirror 2, the emitted laser is expanded through the two confocal mirrors of the secondary mirror 3 and the primary mirror 4, the diffraction angle of the laser is reduced by enlarging the diameter of the laser emission, and the energy density of the laser on the ground is ensured. In order to ensure the coverage area of the laser spot on the ground and the laser energy density, the laser emitting caliber is 20cm, the laser divergence angle is set to be 54', the laser wavelength is 1064nm, the laser coherence time is 1ns, and the laser spot has the travel diameter of 100km on the ground. The telescope control system directs the laser beam to the ground intensity coherent observation system by adjusting the deflection reflector 2.
Step two: the deep space spacecraft relative position measuring system based on the intensity coherent detection array needs to construct an intensity coherent observation system on the ground, the intensity coherent observation system consists of a plurality of intensity coherent detection units, and a schematic diagram of the intensity coherent receiving unit is shown in figure 2 and comprises a low-precision light collecting surface 6, an optical filter 8, a high-frequency detector 9, a multiplier 11, data storage and calculation equipment 13 and a calculation control center 14.
The main body of the intensity coherent detection unit is a large-aperture light collecting surface, and the light collecting surface only collects light energy without maintaining the phase of a light field. The focal length of the light collecting surface is 7.6m, the diameter is 1m, the light collecting surface is made by splicing spherical mirrors, and only the energy is required to be collected to a high-frequency detector. The array control system controls the direction of the light collecting surface and ensures that the retro-reflection laser of the spacecraft to be tested can be collected on the light intensity detector by the light collecting surface, and the tracking error of the light collecting surface is less than 3'. The retroreflection laser of the spacecraft passes through the light collecting surface 6 and the collimating lens 7, converts the light beam into parallel light, filters stray light from other light sources through the optical filter 8, and concentrates the light on the high-frequency detector 9. The random fluctuation of the light intensity is collected by a high-frequency detector, the sampling frequency of the high-frequency detector is 1GHz, the light sensing surface of the detector is 10 micrometers, the central wavelength of the pass band of the optical filter is 1064nm, and the width of the pass band is 10 nm. The sum of the optical path error of the surface of the collecting lens and the optical path error of the atmospheric turbulence in the observation process needs to be less than 0.15 m.
Step three: and constructing an intensity coherent observation system in a northern latitude 30-degree region. The observation system is composed of 29 intensity coherent observation unit arrays, and the array configuration diagram is shown in figure 3. The array is arranged in a cross shape, one array is arranged in the center, and 7 arrays are respectively arranged on the other four arms. Wherein two arms are arranged in the north-south direction, and the other two arms are arranged in the east-west direction. The observation units have a pitch of 1.5 m.
Step four: acquiring signals by using a short-baseline low-precision foundation observation array, and calculating the photocurrent correlation between every two observation units in the observation array according to the acquired signals; constructing coordinates; acquiring a complete space spectrum of a retroreflection laser pattern, and drawing interference fringes according to the acquired space spectrum; and realizing high-precision relative position measurement of the deep space spacecraft according to the drawn interference fringes.
Step 4.1: and respectively sending instructions to the lunar landing cabin and the track cabin by using a remote control system, so that the remote control system emits continuous laser to the position of the foundation strength coherent observation system. The ground strength coherent observation system points to the spacecraft and stably tracks the spacecraft, the high-frequency detector continuously converts the random fluctuation of the target light intensity into photocurrent, the amplitude of the photocurrent fluctuation is recorded and stored, and the observation is continuously carried out for 1200 s.
Step 4.2: and calculating the photocurrent correlation between the detectors in the observation array by using the multiplier. And calculating the photocurrent correlation degree between every two observation units according to the recorded photocurrent intensity output by each detector through the following formula. :
wherein IAAnd (T) is the high-frequency photocurrent of the A position observation unit at each moment, and T is the total observation time length.
Step 4.3: and (5) constructing coordinates. Constructing an orthogonal coordinate system in the observation system, taking the east-west direction as an x axis and the south-north direction as a y axis, recording the coordinates of all observation units as (x)k,yk) Where k is the observation unit number.
Step 4.4: acquiring a spatial frequency spectrum: acquiring a space spectrum of the space vehicle retroreflection laser interference according to the relative position between the observation units, wherein the space spectrum is calculated as follows:
γ(l,m)＝SAB；l＝xA-xB,m＝yA-yB。
and acquiring a complete space spectrum of the retroreflected laser pattern by using all the photocurrent correlations acquired in the step 4.2, and drawing interference fringes according to the acquired space spectrum. The interference fringes are shown in figure 5.
Step 4.5: the fringe spacing of the interference pattern in the horizontal direction and the vertical direction is extracted, the fringe spacing is 3.2m in the north-south direction, and the fringe spacing is 2.4m in the east-west direction. And calculating the stripe spacing and the stripe direction of the space spectrum according to the stripe spacing, wherein the stripe spacing is 1.92m, and the included angle between the stripes and the north-south direction is 37 degrees.
Step 4.6: and calculating to obtain the position relation of the two spacecrafts, namely realizing the high-precision relative position measurement of the deep space spacecraft.
Observing that the included angle between the connecting line of the two spacecrafts and the south-north direction of the observation system is 37 degrees, and after the unified measurement and control system obtains the distance between the spacecrafts and the ground observation station, according to a formula
The spacecraft separation distance can be calculated to be 15.7 m. The distance measurement precision of the spacecraft can reach hundreds of meters, and the relative angle measurement precision can reach 40 nrad. The accuracy of the radio interference measurement is equivalent to that of the existing global stationing.
The above detailed description is intended to illustrate the objects, aspects and advantages of the present invention, and it should be understood that the above detailed description is only exemplary of the present invention and is not intended to limit the scope of the present invention, and any modifications, equivalents, improvements and the like made within the spirit and principle of the present invention should be included in the scope of the present invention.
Claims (4)
1. The deep space spacecraft relative position measuring method based on the intensity coherent detection array is characterized by comprising the following steps: comprises the following steps of (a) carrying out,
the method comprises the following steps: a satellite-borne downlink laser transmitting telescope is built on a spacecraft to be measured and mainly comprises a direction control part, a laser and a beam expanding telescope; when the position of the spacecraft is required to be measured, each spacecraft simultaneously emits laser with the same frequency to the foundation strength coherent array;
step two: constructing an intensity coherent observation unit on the ground, wherein the intensity coherent observation unit comprises a light collecting surface, an optical filter, a high-frequency detector, a multiplier and data storage computing equipment;
step three: arranging a plurality of foundation strength coherent units in the second step into an array to form a foundation strength coherent observation system, wherein all observation units are required to be ensured not to be on the same straight line when the array is arranged;
step four: acquiring signals by using a short-baseline low-precision foundation observation array, and calculating the photocurrent correlation between every two observation units in the observation array according to the acquired signals; constructing coordinates; calculating and acquiring the spatial frequency of the retroreflection laser pattern according to the photocurrent correlation, and drawing an interference pattern according to the light field spatial coherence; and calculating the relative position between the spacecrafts according to the drawn interference pattern, and realizing high-precision relative position measurement of the deep space spacecraft.
2. The deep space spacecraft relative position measurement method based on the intensity coherent detection array according to claim 1, characterized in that: the implementation method of the fourth step is that,
step 4.1: signal acquisition;
when the relative position of deep space spacecraft is required to be measured, instructions are injected into the spacecraft through ground remote control equipment, a satellite-borne downlink laser transmitting telescope of each spacecraft transmits laser with the same frequency to the direction of a foundation strength coherent observation system, and the included angle of each spacecraft relative to the observation system is required to be smaller than the view field of each observation unit; each observation unit in the laser coherent array receives laser emitted by all spacecrafts, the high-frequency detector converts random fluctuation of light intensity obtained by receiving into corresponding high-frequency photocurrent in proportion, the sampling frequency of the high-frequency detector is vHz, and the maximum system error introduced by a system light collecting surface and atmospheric turbulence is required to be smaller than
Step 4.2: calculating the current correlation;
and calculating the correlation of the high-frequency photocurrents output by A, B two observation units in the array by using a multiplier, wherein the calculation formula of the correlation of the photocurrents at different positions is as follows:
wherein IA(T) is the high-frequency photocurrent of the A position observation unit at each moment, and T is the total observation duration;
calculating the photocurrent correlation between every two observation units in the observation array according to the formula;
step 4.3: constructing coordinates;
constructing an orthogonal coordinate system in the observation system, taking a certain direction as an x-axis and the normal direction of the x-axis as a y-axis, recording the coordinates of all observation units as (x)k,yk) Wherein k is the observation unit number;
step 4.4: acquiring a spatial frequency spectrum, and drawing an interference pattern according to the acquired spatial frequency spectrum;
acquiring a space spectrum of a retroreflection laser pattern of the spacecraft according to the relative position between the observation units, wherein the space spectrum of the pattern is calculated as follows:
γ(l,m)＝SAB；l＝xA-xB,m＝yA-yB；
acquiring the complete space spectrum of the retroreflection laser pattern by using all the photocurrent correlations acquired in the step 4.2, and drawing an interference pattern according to the acquired space spectrum;
step 4.5: extracting fringe spacing of the interference pattern in the direction I and the direction m, calculating the fringe spacing and the fringe direction of the space spectrum according to the fringe spacing, and calculating the formula as follows:
step 4.6: calculating to obtain the position relation of the two spacecrafts, namely realizing high-precision relative position measurement of the deep space spacecraft;
the included angle of the connecting line of the two spacecrafts and the y axis of the observation system is theta, and the distance between the two spacecrafts is theta
Wherein L is the orbit height of the spacecraft, lambda is the retroreflective laser wavelength, the distance between d interference fringes, and omega where the detector array is located is the latitude; spacecraft spacing measurement error is
3. The deep space spacecraft relative position measurement method based on the intensity coherent detection array according to claim 1 or 2, characterized in that: the array arrangement mode in the third step is a cross-shaped arrangement array or a Y-shaped array; the longest base line of the array is K, and the distance between coherent units in the array is delta K.
4. The deep space spacecraft relative position measurement method based on the intensity coherent detection array according to claim 3, characterized in that: in step 4.3, the weft direction is set to be the x-axis.
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