CN110849318A - Method and device for acquiring sun altitude angle of spacecraft subsatellite point and imaging method - Google Patents
Method and device for acquiring sun altitude angle of spacecraft subsatellite point and imaging method Download PDFInfo
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Abstract
The invention discloses a method and a device for acquiring a sun altitude angle of a satellite lower point of a spacecraft. The invention can be directly calculated by the spacecraft loading system, does not need to be solved by a ground system and then injected into the spacecraft system through an uplink channel, and can overcome the defects of large workload, low efficiency and long-term support of ground workers. The invention also provides an optical imaging load imaging method of the spacecraft.
Description
Technical Field
The invention relates to the technical field of spaceflight, in particular to a method and a device for acquiring a sun altitude angle of a spacecraft subsatellite point. The invention also relates to an optical imaging load imaging method of the spacecraft.
Background
In the working process of the optical imaging load of the spacecraft along with the on-orbit operation of the spacecraft, the change range of the incident light energy of the optical imaging load is very large due to the change of conditions such as the solar altitude angle, the ground scene reflectivity and the like, so that the situations that the target of interest in the obtained remote sensing image is saturated and cannot be distinguished can occur. Therefore, in order to achieve the best imaging effect, the optical imaging load needs to have on-track dimming capability, wherein the solar altitude is one of the important parameters for dimming reference.
In the prior art, a method for determining the solar altitude is to calculate the solar altitude within a certain time by a ground system and then inject the solar altitude into a control system of an optical imaging load through an uplink channel. However, the method has the disadvantages of heavy workload, low efficiency, relatively short time for each calculation, and long-term support of ground workers during the operation of the spacecraft.
Disclosure of Invention
In view of this, the present invention provides a method and a device for acquiring a solar altitude angle of a spacecraft substellar point, which can overcome the defects in the prior art. The invention also provides an optical imaging load imaging method of the spacecraft.
In order to achieve the purpose, the invention provides the following technical scheme:
a spacecraft intersatellite point solar altitude angle obtaining method comprises the following steps:
calculating a solar vector under a satellite inertial coordinate system in orbit operation of the spacecraft according to the corresponding moment of the solar altitude angle of the satellite lower point of the spacecraft to be obtained and a pre-established relation between the solar vector under the satellite inertial coordinate system in orbit operation of the spacecraft and time;
acquiring position coordinates of the spacecraft under a satellite inertial coordinate system at the moment corresponding to the acquired sun altitude of the spacecraft subsatellite point;
and calculating to obtain the solar altitude angle of the satellite lower point of the spacecraft according to the obtained solar vector of the satellite inertial coordinate system of the spacecraft in orbit and the position coordinate of the spacecraft.
Preferably, the method for establishing the relation between the solar vector and the time under the planetary inertial coordinate system in the orbiting process of the spacecraft comprises the following steps:
constructing an n-order Fourier approximation function model for describing the relation between the solar vector and the time under the satellite inertial coordinate system of the spacecraft in orbit, wherein n is a positive integer greater than or equal to 1;
collecting a plurality of solar vector data under a planet inertia coordinate system in the orbit operation process of the spacecraft;
and fitting the constructed n-order Fourier approximation function model by using the obtained plurality of solar vector data, determining the model parameters of the n-order Fourier approximation function model, and obtaining the relation between the solar vector and the time of the spacecraft in the orbit running in the planet inertia coordinate system.
Preferably, the n-order fourier approximation function model for describing the relation between the solar vector and the time in the satellite inertial coordinate system of the spacecraft in orbit is described as follows:
x_sun＝xa0+xa1·cos(t·xw)+xb1·sin(t·xw)+…+xan·cos(nt·xw)+xbn·sin(nt·xw)；
y_sun＝ya0+ya1·cos(t·yw)+yb1·sin(t·yw)+…+yan·cos(nt·yw)+ybn·sin(nt·yw)；
z_sun＝za0+za1·cos(t·zw)+zb1·sin(t·zw)+…+zan·cos(nt·zw)+zbn·sin(nt·zw)；
wherein x _ sun represents the x coordinate of the sun vector of the spacecraft in the orbit running under the planetary inertial coordinate system, and y _ sun represents the spacecraft in the orbit runningThe y coordinate of the sun vector under the planetary inertial coordinate system in orbit running, z _ sun represents the z coordinate of the sun vector under the planetary inertial coordinate system in orbit running of the spacecraft, t represents the time difference from the starting time and the ending time of the sun vector data participating in fitting, xa0、xai、xbi、xw、ya0、yai、ybi、yw、za0、zai、zbi、zw(i ═ 1, …, n) represents model parameters, respectively.
Preferably, n is 4 or more.
Preferably, n is 8, and the 8 th order fourier approximation function model constructed for describing the relation between the sun vector and the time in the satellite inertial coordinate system of the spacecraft in orbit is described as follows:
preferably, the obtained plurality of solar vector data are fitted by adopting a least square method, and model parameters of the n-order Fourier approximation function model are determined.
Preferably, the obtaining of the solar altitude angle of the intersatellite point of the spacecraft by calculating according to the obtained solar vector of the spaceborne inertial coordinate system in the orbiting process of the spacecraft and the position coordinate of the spacecraft comprises: calculated according to the following formula:
wherein (x, y, z) represents the sun vector coordinate of the spacecraft in the orbit running under the satellite inertial coordinate system, and (x _ sat, y _ sat, z _ sat) represents the position coordinate of the spacecraft in the satellite inertial coordinate system.
A spacecraft intersatellite point solar altitude angle acquisition device is used for executing the spacecraft intersatellite point solar altitude angle acquisition method.
A spacecraft optical imaging payload imaging method comprising:
calculating the sun altitude angle of the satellite lower point of the spacecraft by using the method for acquiring the sun altitude angle of the satellite lower point of the spacecraft;
and controlling an optical imaging load shooting image of the spacecraft according to the obtained sun altitude angle of the intersatellite point of the spacecraft so as to enable the obtained image to meet the preset requirement.
Preferably, the step of controlling the optical imaging load shooting image of the spacecraft according to the obtained sun altitude angle of the intersatellite point of the spacecraft comprises the following steps: and setting the integration series, the integration time or the gain of the time delay integration charge-coupled device of the optical imaging load of the spacecraft according to the obtained sun altitude of the intersatellite point of the spacecraft so as to control the optical imaging load of the spacecraft to shoot images.
According to the technical scheme, firstly, the solar vector of the spacecraft in the orbit running planet inertial coordinate system is calculated according to the corresponding time of the acquired spacecraft satellite lower point solar altitude and the pre-established relation between the solar vector and time of the spacecraft in the orbit running planet inertial coordinate system, the position coordinate of the spacecraft in the orbit inertial coordinate system at the corresponding time of the spacecraft lower point solar altitude to be acquired is acquired, and then the spacecraft lower point solar altitude is calculated according to the acquired solar vector of the spacecraft in the orbit running planet inertial coordinate system and the position coordinate of the spacecraft. The method and the device for acquiring the sun altitude angle of the satellite lower point of the spacecraft can be directly calculated by a spacecraft loading system, do not need to be firstly calculated by a ground system and then injected into the spacecraft system through an uplink channel, and can overcome the defects of large workload, low efficiency and long-term support of ground workers in the prior art.
The optical imaging load imaging method for the spacecraft, provided by the invention, can achieve the beneficial effects.
Drawings
In order to more clearly illustrate the embodiments of the present invention or the technical solutions in the prior art, the drawings used in the description of the embodiments or the prior art will be briefly described below, it is obvious that the drawings in the following description are only some embodiments of the present invention, and for those skilled in the art, other drawings can be obtained according to the drawings without creative efforts.
Fig. 1 is a flowchart of a method for acquiring a solar altitude angle of a spacecraft substellar point according to an embodiment of the present invention;
FIG. 2 is a flowchart of a method for establishing a relationship between a solar vector and time in an inertial coordinate system of a planet in orbiting of a spacecraft in an embodiment of the present invention;
fig. 3 is a flowchart of a spacecraft optical imaging load imaging method according to an embodiment of the present invention.
Detailed Description
In order to make those skilled in the art better understand the technical solution of the present invention, the technical solution in the embodiment of the present invention will be clearly and completely described below with reference to the drawings in the embodiment of the present invention, and it is obvious that the described embodiment is only a part of the embodiment of the present invention, and not all embodiments. All other embodiments, which can be derived by a person skilled in the art from the embodiments given herein without making any creative effort, shall fall within the protection scope of the present invention.
Referring to fig. 1, fig. 1 is a flowchart of a method for acquiring a solar altitude angle of a spacecraft substellar point according to an embodiment of the present invention. As can be seen, the method comprises the following steps:
s10: and calculating the solar vector of the spacecraft in the orbit running planet inertia coordinate system according to the corresponding moment of the acquired spacecraft in-orbit point solar altitude angle and the pre-established relation between the solar vector and time of the spacecraft in the orbit running planet inertia coordinate system.
According to the pre-established relation between the sun vector and the time under the planetary inertial coordinate system of the spacecraft in orbit operation, the sun vector under the planetary inertial coordinate system of the spacecraft at each moment in orbit operation can be obtained. For example, when the sun altitude of the interstellar point of the spacecraft at the time t1 is to be acquired, the sun vector of the planet inertial coordinate system at the time t1 is obtained according to the time t1 and the pre-established relationship.
S11: and acquiring the position coordinates of the spacecraft under a satellite inertial coordinate system at the corresponding moment of the sun altitude of the spacecraft subsatellite point to be acquired.
If the sun altitude of the intersatellite point of the spacecraft at the time t1 is to be acquired, the position coordinates of the spacecraft at the time t1 in the inertial coordinate system of the spacecraft are acquired in the step.
S12: and calculating to obtain the solar altitude angle of the satellite lower point of the spacecraft according to the obtained solar vector of the satellite inertial coordinate system of the spacecraft in orbit and the position coordinate of the spacecraft.
The method for acquiring the solar altitude angle of the spacecraft intersatellite point can be directly calculated by a spacecraft loading system, and does not need to be firstly calculated by a ground system and then injected into the spacecraft system through an uplink channel, so that the defects of large workload, low efficiency and long-term support of ground workers in the prior art can be overcome.
The method for acquiring the solar altitude angle of the intersatellite point of the spacecraft is described in detail in the following with reference to specific embodiments. The method for acquiring the solar altitude angle of the spacecraft subsatellite point comprises the following steps:
s10: and calculating the solar vector of the spacecraft in the orbit running planet inertia coordinate system according to the corresponding moment of the acquired spacecraft in-orbit point solar altitude angle and the pre-established relation between the solar vector and time of the spacecraft in the orbit running planet inertia coordinate system.
In one embodiment, referring to fig. 2, a method for establishing a relationship between a solar vector and time in an inertial coordinate system of a planet in orbit of a spacecraft mainly includes the following steps:
s100: and constructing an n-order Fourier approximation function model for describing the relation between the solar vector and the time under the satellite inertial coordinate system of the spacecraft in orbit, wherein n is a positive integer greater than or equal to 1.
In the method, an n-order Fourier approximation function is used for describing the relation between the solar vector and time in the planet inertial coordinate system in the orbiting process of the spacecraft, firstly, an n-order Fourier approximation function model is constructed, and n is a positive integer greater than or equal to 1. Specifically, the n-order fourier approximation function model for describing the relation between the solar vector and time in the satellite inertial coordinate system of the spacecraft in orbit is described as follows:
x_sun＝xa0+xa1·cos(t·xw)+xb1·sin(t·xw)+…+xan·cos(nt·xw)+xbn·sin(nt·xw)；
y_sun＝ya0+ya1·cos(t·yw)+yb1·sin(t·yw)+…+yan·cos(nt·yw)+ybn·sin(nt·yw)；
z_sun＝za0+za1·cos(t·zw)+zb1·sin(t·zw)+…+zan·cos(nt·zw)+zbn·sin(nt·zw)；
wherein x _ sun represents the x coordinate of the sun vector of the spacecraft in the orbit running under the planet inertia coordinate system, y _ sun represents the y coordinate of the sun vector of the spacecraft in the orbit running under the planet inertia coordinate system, z _ sun represents the z coordinate of the sun vector of the spacecraft in the orbit running under the planet inertia coordinate system, t represents the time difference from the starting and stopping time of the sun vector data participating in fitting, xa0、xai、xbi、xw、ya0、yai、ybi、yw、za0、zai、zbi、zw(i ═ 1, …, n) represents model parameters, respectively.
In practical application, according to multiple test results, when n is greater than or equal to 4, the accuracy of the solar vector under the satellite inertial coordinate system of the spacecraft in orbit running calculated according to the n-order Fourier approximation function model obtained through fitting can meet requirements, wherein the larger the value of n is, the higher the accuracy of the calculation result is, but the corresponding calculation amount is increased, and the calculation time is increased, so that the requirements on the calculation accuracy and the calculation amount are comprehensively considered in practical application to determine the value of n. Preferably, in a specific embodiment, the calculation accuracy and the calculated quantity n are considered to be 8, that is, an 8-order fourier approximation function is used to describe the relationship between the sun vector and the time in the satellite inertial coordinate system during the orbiting of the spacecraft, and a specifically established function model is as follows:
s101: and collecting a plurality of solar vector data under a planet inertia coordinate system in the orbit operation process of the spacecraft.
The method comprises the steps of collecting a plurality of sun vector data under a planet inertia coordinate system in the orbit operation process of the spacecraft, determining the accuracy of finally calculating the solar altitude angle according to the collected data, and determining the quantity of the collected data in combination with the requirements on calculation accuracy and calculation amount in practical application. For example, data may be collected for a spacecraft operating in orbit for 1 year at 10 minute intervals, i.e., 52560 sets of data are collected for fitting.
Alternatively, a Satellite Toolkit (STK) may be used to acquire the sun vector coordinates (x, y, z) in the planetary coordinate system during orbiting of the spacecraft.
S102: and fitting the constructed n-order Fourier approximation function model by using the obtained plurality of solar vector data, determining the model parameters of the n-order Fourier approximation function model, and obtaining the relation between the solar vector and the time of the spacecraft in the orbit running in the planet inertia coordinate system.
And fitting the acquired plurality of solar vector data, optionally fitting the acquired plurality of solar vector data by adopting a least square method, and determining model parameters of the established n-order Fourier approximation function model so as to obtain the relation between the solar vector and the time of the spacecraft in the orbit running under the planet inertia coordinate system.
S11: and acquiring the position coordinates of the spacecraft under a satellite inertial coordinate system at the corresponding moment of the sun altitude of the spacecraft subsatellite point to be acquired.
In practical application, the position coordinates of the spacecraft in the satellite inertial coordinate system can be directly given by the guidance, navigation and control system of the spacecraft, or the position coordinates of the spacecraft in the satellite inertial coordinate system can be calculated according to orbit parameters of the spacecraft in orbit, which are given by the guidance, navigation and control system of the spacecraft, and different spacecrafts can obtain the position coordinates of the spacecraft in the satellite inertial coordinate system in a corresponding mode.
S12: and calculating to obtain the solar altitude angle of the satellite lower point of the spacecraft according to the obtained solar vector of the satellite inertial coordinate system of the spacecraft in orbit and the position coordinate of the spacecraft.
Specifically, the solar altitude of the intersatellite point of the spacecraft can be calculated based on a spherical triangle edge cosine formula and a celestial coordinate system parameter, and the calculation formula is described as follows:
wherein Sun _ angle represents the solar altitude angle of the satellite lower point of the spacecraft, (x, y, z) represents the solar vector coordinate of the spacecraft in the satellite inertial coordinate system in orbit operation, and (x _ sat, y _ sat, z _ sat) represents the position coordinate of the spacecraft in the satellite inertial coordinate system.
The method for acquiring the sun altitude angle of the satellite lower point of the spacecraft can realize on-orbit real-time calculation, and the precision of a calculation result can meet the application requirement.
Correspondingly, the embodiment of the invention also provides a spacecraft intersatellite point solar altitude angle acquisition device which is used for executing the spacecraft intersatellite point solar altitude angle acquisition method.
The device for acquiring the solar altitude angle of the satellite lower point of the spacecraft, which is used for calculating and acquiring the solar altitude angle of the satellite lower point of the spacecraft, can be directly calculated by a spacecraft loading system, and does not need to be solved by a ground system and then injected into the spacecraft system through an uplink channel, so that the defects of large workload, low efficiency and long-term support of ground workers in the prior art can be overcome.
Correspondingly, an embodiment of the present invention further provides a spacecraft optical imaging payload imaging method, please refer to fig. 3, including the following steps:
s20: and calculating the sun altitude angle of the satellite lower point of the spacecraft by using the method for acquiring the sun altitude angle of the satellite lower point of the spacecraft.
S21: and controlling an optical imaging load shooting image of the spacecraft according to the obtained sun altitude angle of the intersatellite point of the spacecraft so as to enable the obtained image to meet the preset requirement.
The obtained image meets the preset requirement, namely the brightness of the shot image is within a preset range, the image contrast is within the preset range, the image is not too bright or too dark, the image brightness is not saturated, and a better imaging effect is achieved.
In specific implementation, the integration stage number, the integration time or the gain of the time delay integration charge-coupled device of the optical imaging load of the spacecraft can be set according to the obtained sun altitude of the intersatellite point of the spacecraft, so as to control the optical imaging load of the spacecraft to shoot images.
In the working process of the optical imaging load of the spacecraft running along with the spacecraft in orbit, the change range of the entrance pupil amplitude brightness of the optical imaging load is large due to the change of conditions such as solar altitude, ground scene reflectivity and the like, so that the optical imaging load generally needs to meet the requirement of saturated radiance in order to achieve the optimal imaging effect and avoid the phenomenon that the interested target in the obtained remote sensing image is saturated and cannot be distinguished, namely, the output image cannot be saturated when the entrance pupil radiance of the optical imaging load is smaller than the saturated radiance. In the prior art, a small fixed gain is usually adopted to avoid image saturation, but the obtained image has the problems of overall darkness, insufficient detail information level and the like. Therefore, the method can realize on-orbit real-time calculation of the sun altitude of the intersatellite point of the spacecraft, and real-time adjustment of the optical imaging load of the spacecraft according to the sun altitude, so that the optical imaging load of the spacecraft can obtain a better imaging effect.
The method, the device and the imaging method for acquiring the sun altitude angle of the satellite lower point of the spacecraft provided by the invention are described in detail above. The principles and embodiments of the present invention are explained herein using specific examples, which are presented only to assist in understanding the method and its core concepts. It should be noted that, for those skilled in the art, it is possible to make various improvements and modifications to the present invention without departing from the principle of the present invention, and those improvements and modifications also fall within the scope of the claims of the present invention.
Claims (10)
1. A spacecraft substellar point solar altitude angle obtaining method is characterized by comprising the following steps:
calculating a solar vector under a satellite inertial coordinate system in orbit operation of the spacecraft according to the corresponding moment of the solar altitude angle of the satellite lower point of the spacecraft to be obtained and a pre-established relation between the solar vector under the satellite inertial coordinate system in orbit operation of the spacecraft and time;
acquiring position coordinates of the spacecraft under a satellite inertial coordinate system at the moment corresponding to the acquired sun altitude of the spacecraft subsatellite point;
and calculating to obtain the solar altitude angle of the satellite lower point of the spacecraft according to the obtained solar vector of the satellite inertial coordinate system of the spacecraft in orbit and the position coordinate of the spacecraft.
2. A spacecraft intersatellite point solar altitude angle acquisition method according to claim 1, wherein the method for establishing the relation between the solar vector and the time in the satellite inertial coordinate system in the orbit operation of the spacecraft comprises the following steps:
constructing an n-order Fourier approximation function model for describing the relation between the solar vector and the time under the satellite inertial coordinate system of the spacecraft in orbit, wherein n is a positive integer greater than or equal to 1;
collecting a plurality of solar vector data under a planet inertia coordinate system in the orbit operation process of the spacecraft;
and fitting the constructed n-order Fourier approximation function model by using the obtained plurality of solar vector data, determining the model parameters of the n-order Fourier approximation function model, and obtaining the relation between the solar vector and the time of the spacecraft in the orbit running in the planet inertia coordinate system.
3. A spacecraft subsatellite point solar altitude angle acquisition method according to claim 2, wherein an n-th order fourier approximation function model for describing the relation of a solar vector and time under a spacecraft in-orbit satellite inertial coordinate system is constructed and described as follows:
x_sun＝xa0+xa1·cos(t·xw)+xb1·sin(t·xw)+…+xan·cos(nt·xw)+xbn·sin(nt·xw)；
y_sun＝ya0+ya1·cos(t·yw)+yb1·sin(t·yw)+…+yan·cos(nt·yw)+ybn·sin(nt·yw)；
z_sun＝za0+za1·cos(t·zw)+zb1·sin(t·zw)+…+zan·cos(nt·zw)+zbn·sin(nt·zw)；
wherein x _ sun represents the x coordinate of the sun vector of the spacecraft in the orbit running under the planet inertia coordinate system, y _ sun represents the y coordinate of the sun vector of the spacecraft in the orbit running under the planet inertia coordinate system, z _ sun represents the z coordinate of the sun vector of the spacecraft in the orbit running under the planet inertia coordinate system, t represents the space participating in fittingTime difference of start and stop time of positive vector data, xa0、xai、xbi、xw、ya0、yai、ybi、yw、za0、zai、zbi、zw(i ═ 1, …, n) represents model parameters, respectively.
4. A spacecraft substellar solar altitude angle acquisition method as claimed in claim 3, wherein n is greater than or equal to 4.
5. The method for acquiring the sun altitude angle at the satellite earth point of the spacecraft according to claim 3, wherein n is 8, and an 8-order Fourier approximation function model for describing the relation between the sun vector and the time in the satellite inertial coordinate system in the orbital operation of the spacecraft is described as follows:
6. the spacecraft subsatellite point solar altitude angle acquisition method according to claim 2, wherein a least square method is adopted to fit a plurality of acquired solar vector data to determine model parameters of the n-order Fourier approximation function model.
7. A spacecraft intersatellite point solar altitude angle acquisition method according to any one of claims 1-6, wherein the step of calculating and obtaining the solar altitude angle of the intersatellite point of the spacecraft according to the obtained solar vector of the spacecraft in an in-orbit running planet inertial coordinate system and the position coordinate of the spacecraft comprises the following steps: calculated according to the following formula:
wherein (x, y, z) represents the sun vector coordinate of the spacecraft in the orbit running under the satellite inertial coordinate system, and (x _ sat, y _ sat, z _ sat) represents the position coordinate of the spacecraft in the satellite inertial coordinate system.
8. A spacecraft intersatellite point solar altitude angle acquisition apparatus for performing the spacecraft intersatellite point solar altitude angle acquisition method according to any one of claims 1 to 7.
9. A spacecraft optical imaging payload imaging method, comprising:
calculating an off-satellite point solar altitude of the spacecraft using the off-satellite point solar altitude acquisition method of any one of claims 1 to 7;
and controlling an optical imaging load shooting image of the spacecraft according to the obtained sun altitude angle of the intersatellite point of the spacecraft so as to enable the obtained image to meet the preset requirement.
10. A spacecraft optical imaging load imaging method according to claim 9, wherein controlling the captured optical imaging load image of the spacecraft according to the obtained intersatellite point solar altitude of the spacecraft comprises: and setting the integration series, the integration time or the gain of the time delay integration charge-coupled device of the optical imaging load of the spacecraft according to the obtained sun altitude of the intersatellite point of the spacecraft so as to control the optical imaging load of the spacecraft to shoot images.
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