CN116822176A - Satellite full system coupling orbit design method, system, terminal and medium - Google Patents

Abstract

The invention provides a satellite full system coupling orbit design method and system, which calculates a quasi-sun synchronous drift circular orbit; based on the quasi-sun synchronous drift circular orbit, in view of the observation requirements of the star observation mission and the sun observation mission, the load of the star observation mission is And design the time-sharing observation orbit inclination of the payload of the solar observation mission, and supplement the mission time of the star observation mission in the two-dimensional shadow area of the sun during the extended service period; based on the design results of the observation orbit inclination, judge again Whether the adjusted orbital mission observation meets the observation requirements: If not, adjust the drift rate at the descending node through the offset of the orbital inclination i. If it meets, determine whether the launch vehicle has an unimplementable state: if so, Then it ends. If not, fine-tune the offset of the orbital inclination i to complete the orbit design of the satellite's full system coupling and achieve the most cost-effective satellite development and design basis.

Description

Satellite full system coupling orbit design method, system, terminal and medium

Technical field

The present invention relates to the technical field of satellite system design. Specifically, it relates to a satellite full system coupling orbit design method and system for large-span solar altitude angles. It also provides a corresponding computer terminal and computer-readable storage medium.

Background technique

Traditional satellites are divided into experimental and operational types. In response to a variety of different needs, satellites are usually customized according to special tasks. It is difficult to take into account multiple mission requirements on one satellite. Satellite orbit design takes mission requirements as input, while orbit design As the design input of other subsystems, there are fewer iterations with other subsystems; the input of traditional satellite orbit design is mission requirements, which generally consider more load requirements and rarely take into account load structure installation, attitude orbit control difficulty, and optical sensitivity. Coupled design factors such as device installation, thermal control and heat dissipation surface design and selection, energy acquisition and other coupling design factors make it difficult to achieve the optimal system design. Generally, satellite orbit design is based on the consideration of being able to perform all tasks during the entire period of time in orbit, although this can improve the timing of task execution. The choice is more flexible, but it will also bring huge waste of resources, causing the cost of satellites to rise or the system design to be complex and reliability reduced.

To sum up, traditional satellites usually have the following technical problems in satellite orbit design:

Satellite orbit design considers a single type of mission;

Satellite orbit design considers mission requirements more than the coupling iteration of the orbit and other subsystems;

The satellite orbit design lacks the concept of mission execution window (time-sharing execution based on orbit status).

Contents of the invention

In view of the above-mentioned deficiencies in the prior art, the present invention provides a satellite full-system coupling orbit design method and system for a large span of solar altitude angles, and also provides a corresponding computer terminal and computer-readable storage medium.

According to one aspect of the present invention, a satellite full system coupling orbit design method is provided, including:

Calculation of quasi-sun-synchronous drifting circular orbits;

Based on the quasi-sun synchronous drifting circular orbit, according to the observation requirements of the star observation mission and the sun observation mission, the payload of the star observation mission and the time-sharing observation orbit inclination of the payload of the sun observation mission are designed, and in During the extended service period, the two-dimensional solar shadow area in the sunlit area supplements the mission time of the star observation mission;

According to the design results of the observation orbit inclination angle, judge again whether the adjusted orbit mission observation meets the observation requirements: if not, adjust the drift rate at the descending node through the offset of the orbit inclination angle i. If it meets, judge whether the carrier Whether the rocket has an unimplementable state: if so, end. If not, fine-tune the offset of the orbital inclination i to complete the orbit design of the satellite's full system coupling.

Preferably, the observation requirements of the star observation mission include: the observation requirements of the first load and the second payload with the highest solar altitude angle greater than or equal to 45° in the star observation mission; the observation requirements of the sun observation mission include: In the solar observation mission, the observation requirements for the third payload with the highest solar altitude angle within the closed range of [17°, 45°].

Preferably, design a drift orbit with the local time of the descending node at [10:00, 2:00], and use the period between [10:00, 8:00] and [4:00, 2:00] as the star observation mission Window, [8:00, 4:00] serves as the mission window for observing the sun that has just risen.

Preferably, the calculation of the quasi-sun synchronous drift circular orbit includes:

Obtain the characteristic parameters required to calculate the quasi-sun synchronous drift circular orbit, the characteristic parameters include: orbit semi-major axis a, orbit eccentricity e and orbit inclination angle i;

According to the average angle of the sun's daily motion relative to the earth, the corresponding relationship between the orbital semi-major axis a, the orbital eccentricity e and the orbital inclination angle i is constructed, the orbital height is given, and the orbital eccentricity e is set =0, the orbital inclination angle i is obtained. According to the observation requirements of different missions, the negative offset orbital inclination angle i is calculated to obtain a quasi-sun-synchronous drifting circular orbit.

Preferably, based on the quasi-sun synchronous drifting circular orbit, and based on the observation requirements of the star observation mission and the sun observation mission, the time-sharing observation orbits of the load of the star observation mission and the load of the sun observation mission are designed, include:

Conduct simulation calculations based on the quasi-sun synchronous drift circular orbit to obtain the observable time intervals of star observation missions and sun observation missions with different launch into orbit times, and decide whether to adjust the orbital inclination i according to the mission time satisfaction;

The offset of the orbital inclination i is adjusted according to the mission observation requirements, and simulation calculations are performed based on the adjusted quasi-sun synchronous drift circular orbit to obtain the orbital period, sunshade time distribution, solar vector and various aspects of the satellite body for different launch into orbit times. The relationship data of the given satellite attitude control, thermal control, and energy subsystems are analyzed to determine whether there is a risk of not meeting the subsystem index requirements; if there is a risk, the orbit inclination offset is reduced by the set degree, if not, continue Next step.

Preferably, the decision of whether the orbit inclination angle i needs to be adjusted based on the mission time satisfaction includes:

The task time satisfaction includes:

During the lifetime observation mission, whether the star sensor of the attitude control subsystem has the time when the sun enters the field of view;

Whether the temperature of a single machine controlled by the thermal control subsystem during the observation mission during the life cycle exceeds the set proportion;

Whether the energy subsystem meets the single-day energy balance during the observation mission during the life cycle;

When the sun enters the field of view during the observation mission within the lifetime, exceeds the standard setting ratio, and/or cannot meet the single-day energy balance, the orbital inclination angle i needs to be adjusted.

Preferably, the two-dimensional solar shadow area in the sunlit area during the extended service period supplements the mission time of the star observation mission, including:

When the satellite enters the sunlit area, adjust the two-dimensional attitude towards the sun so that the normal direction of the solar sail panel is aligned with the direction of sunlight;

When the satellite enters the shadow area, it adjusts its vertical attitude toward the earth to carry out a short-term star observation mission, and achieves a full life cycle of time drift at the descending node, which can make full use of satellite resources to complete the set mission.

Preferably, based on the design results of the observation orbit inclination angle, it is again judged whether the adjusted orbit mission observation meets the observation requirements, including:

Obtain fully iterative observation orbit inclination design results for each subsystem;

Based on the design results of the observed orbit inclination angle and the simulation calculation results of the attitude control subsystem, thermal control subsystem and energy subsystem, comprehensive mission time satisfaction and subsystem index satisfaction, a decision is made whether to adopt the offset of the current orbit inclination i, At the same time, the observation orbit inclination design results are fed back to the launch vehicle side for calculation, and a decision is made whether to fine-tune the offset of the orbit inclination i;

The satellite orbit is finally determined based on the orbital inclination iteration results.

Preferably, the fine adjustment means that the orbit inclination angle i is adjusted within the range of ±0.1°.

According to another aspect of the present invention, a satellite full system coupled orbit design system is provided, including:

Quasi-sun synchronous drift circular orbit calculation module, which is used to calculate quasi-sun synchronous drift circular orbit;

Orbital inclination design module, which is based on the quasi-sun synchronous drift circular orbit, and is based on the observation requirements of the star observation mission and the sun observation mission, and the time-sharing observation orbit for the load of the star observation mission and the load of the sun observation mission. The design work is carried out based on the inclination angle, and the two-dimensional sun shadow area in the sunlit area supplements the mission time of the star observation mission during the extended service period;

Orbital inclination adjustment module, this module judges again whether the adjusted orbital mission observation meets the observation requirements based on the design results of the observed orbital inclination: If not, the drift rate at the descending node is adjusted by the offset of the orbital inclination i , if satisfied, then determine whether the launch vehicle has an unimplementable state: if so, it ends; if not, fine-tune the offset of the orbital inclination i to complete the orbit design of the satellite's full system coupling.

According to a third aspect of the present invention, a computer terminal is provided, including a memory, a processor, and a computer program stored in the memory and executable on the processor. When the processor executes the program, it can be used to execute the above-mentioned The method described in any of the above, or, running the system described above.

According to a fourth aspect of the present invention, there is provided a computer-readable storage medium on which a computer program is stored. When executed by a processor, the program can be used to perform any of the above methods, or to run the above methods. system described in .

Due to the adoption of the above technical solutions, compared with the prior art, the present invention has at least one of the following beneficial effects:

The satellite full system coupling orbit design method, system, terminal and medium provided by the present invention can change the local time drift of the orbit descending node by setting the orbital inclination angle i offset of the sun-synchronous orbit according to the requirement that the observation mission can be carried out in time periods. Rate, that is, based on the relative positional relationship between the on-orbit mission and the sun and the requirements of the on-orbit test time interval, the drift rate of different orbit descending node local times is designed through the size of the inclination angle offset, the inclination angle offset and the orbit descending node local time The relationship of the drift law is a linear relationship. During the inclination offset design process, it is iterated with the attitude control, thermal control, and energy subsystems. That is, the quasi-sun-synchronous orbit with inclination offset is first designed according to the mission requirements (at a certain rate when the descending node is in place). drift), the orbit subsystem provides relevant design input data for attitude control, thermal control, and energy needs. Subsystem simulations for attitude control, thermal control, and energy are performed to provide feedback to the orbit subsystem, giving drift rate correction suggestions, and then modifying the inclination angle. The offset is used to correct the drift rate at the descending intersection point. After multiple iterations, it meets the requirements of project implementation, observation task time and sub-system indicators.

It should be noted that the task on which this invention is based has a certain range of requirements for the maximum solar altitude angle: Task 1 (observation of stars) requires the maximum solar altitude angle to be as large as possible, with the purpose of being able to traverse as many areas as possible during one mission. In the solar illumination angle range, the proposed maximum solar altitude angle requirement is not less than 45°; Task 2 (observation of the sun) is to test the sun with a solar altitude angle of [-17°, -13°] for as long as possible, hoping that the maximum solar altitude The angle is in the range of [17°, 45°], and the purpose is to enable long-term observation during a mission.

The satellite full system coupling orbit design method, system, terminal and medium provided by the present invention takes mission requirements as the first goal, takes the satellite full system design optimization as the second goal, and fully considers satellite attitude control, thermal control, and satellite service software. The coupling design of , energy and other subsystems enables the orbit design to achieve the top-level system optimization and achieve the most cost-effective basis for satellite development and design.

Description of the drawings

Other features, objects and advantages of the present invention will become more apparent by reading the detailed description of the non-limiting embodiments with reference to the following drawings:

Figure 1 is a work flow chart of a satellite full system coupled orbit design method in a preferred embodiment of the present invention.

Figure 2 is a schematic diagram of the components of the satellite full system coupling orbit design system in a preferred embodiment of the present invention.

Figure 3 is a schematic diagram of the maximum solar altitude angle time of a sun-synchronous orbit in which the maximum solar altitude angle is about 66° and the descending node local time is 10:30 in a specific application example of the present invention.

Figure 4 is a schematic diagram of the maximum solar altitude angle time in a sun-synchronous orbit where the maximum solar altitude angle is about 17° and the descending node local time is 6:30 in a specific application example of the present invention.

Figure 5 is a graph showing the change curve of the orbital solar angle β of a manned spacecraft orbit near the inclination angle i=42° in a specific application example of the present invention.

Figure 6 is a schematic diagram of a specific application example of the present invention in which the descending node is near the meridian, the solar wing is fixedly installed, the normal orientation of the cells is the negative normal direction of the orbital surface, and the solar wing is deployed along the ±X direction of the orbital system.

Figure 7 is a schematic diagram of the relationship between the sun-synchronous orbit height and the orbital inclination angle i in a specific application example of the present invention.

Figure 8 is a schematic diagram of the relationship between the descending node of the sun-synchronous orbit and the sun's orientation in a specific application example of the present invention.

Figure 9 is a schematic diagram showing the direction of the solar wing sail panel in the sunlit area pointing to the sun during the drift process of the descending node local time from 2:00 to 0:00 to 22:00 within 8 months in a specific application example of the present invention.

Figure 10 is a work flow chart of a satellite full system coupled orbit design method for a large span of solar altitude angles in a specific application example of the present invention. In Figure 10, according to the qualitative requirements put forward by the user, when the drift descending node is taken, the incident direction of the sun should be as wide as possible, which means: corresponding to the incident angle range of the sun [0°, 90°] (the incident angle of the sun is defined as The angle between the direction of the satellite pointing to the sun and the direction of the optical axis of the observation load; the observation time of the sun should be as long as possible, which means: the duration of each observation [5min, 10min]; the absolute value of the orbital solar angle |β| should be as small as possible, which means: corresponding |β| range [0°, 45°]; the absolute value of orbital solar angle |β| should be as large as possible, which means: corresponding to |β| range [45°, 90°].

Detailed ways

The following is a detailed description of the embodiments of the present invention: This embodiment is implemented on the premise of the technical solution of the present invention, and detailed implementation modes and specific operating procedures are given. It should be noted that, for those of ordinary skill in the art, several modifications and improvements can be made without departing from the concept of the present invention, and these all belong to the protection scope of the present invention.

One embodiment of the present invention provides a method for designing a satellite full-system coupled orbit. The method takes orbit design as the starting point and systematically considers the overall design of the satellite to achieve a satellite full-system coupled orbit design scheme suitable for long-span solar altitude angle requirements.

As shown in Figure 1, this embodiment provides a satellite full system coupled orbit design method, which may include:

S1, calculate the quasi-sun synchronous drift circular orbit;

S2, based on the quasi-sun synchronous drifting circular orbit, based on the observation requirements of the star observation mission and the sun observation mission, the time-sharing observation orbit inclination angle of the load of the star observation mission and the load of the sun observation mission is designed, and during the extended service period The two-dimensional sun shadow area in the inner sunlit area supplements the mission time of the star observation mission;

S3. Based on the design results of the observation orbit inclination angle, judge again whether the adjusted orbit mission observation meets the observation requirements: if not, adjust the drift rate at the descending node through the offset of the orbit inclination angle i. If it meets, then Determine whether the launch vehicle has an unimplementable state: if so, end it; if not, fine-tune the offset of the orbital inclination i to complete the orbit design of the satellite's full system coupling.

In a preferred embodiment of S2, the observation requirements of the star observation mission include: the observation requirements of the first load and the second payload with the highest solar altitude angle greater than or equal to 45° in the star observation mission; the observation requirements of the sun observation mission, Including: In the solar observation mission, the observation requirements for the third payload with the highest solar altitude angle within the closed range of [17°, 45°].

In a preferred embodiment of S3, the drift orbit where the descending intersection point is [10:00,2:00] is designed, and the drift orbit between [10:00,8:00] and [4:00,2:00] As the mission window for observing stars, [8:00, 4:00] is used as the mission window for observing the sun that has just risen.

In a preferred embodiment of S1, calculating the quasi-sun synchronous drift circular orbit may include:

S11, obtain the characteristic parameters required to calculate the quasi-sun synchronous drift circular orbit. The characteristic parameters include: orbital semi-major axis a, orbital eccentricity e, and orbital inclination angle i;

S12. Based on the average angle of the sun's daily motion relative to the earth, construct the corresponding relationship between the orbital semi-major axis a, orbital eccentricity e and orbital inclination angle i. Given the orbital height and setting the orbital eccentricity e=0, we get The orbital inclination angle i, according to the observation requirements of different missions, and the negative offset orbital inclination angle i, are calculated to obtain a quasi-sun synchronous drift circular orbit.

In a preferred embodiment of S2, based on the quasi-sun synchronous drifting circular orbit, according to the observation requirements of the star observation mission and the sun observation mission, the time-sharing observation orbits of the load of the star observation mission and the load of the sun observation mission are designed, Can include:

S21, conduct simulation calculations based on the quasi-sun synchronous drift circular orbit, and obtain the observable time intervals of star observation missions and sun observation missions with different launch into orbit times. The overall designer decides whether to adjust the orbital inclination angle i based on the mission time satisfaction. ;

S22, adjust the offset of the orbital inclination i according to the mission observation requirements of the overall designer, conduct simulation calculations based on the adjusted quasi-sun synchronous drift circular orbit, and obtain the orbital period, sunshade time distribution, solar Based on the relationship data between vectors and various aspects of the satellite body, the given satellite attitude control, thermal control, and energy subsystems are analyzed to determine whether there is a risk of not meeting the subsystem index requirements; if there is a risk, the orbital inclination offset is reduced by the set degree. , if not, continue to the next step.

In a preferred embodiment of S21, deciding whether to adjust the orbital inclination i according to the mission time satisfaction may include:

Task time satisfaction includes:

During the lifetime observation mission, whether the star sensor of the attitude control subsystem has the time when the sun enters the field of view;

Whether the temperature of a single machine controlled by the thermal control subsystem during the observation mission during the life cycle exceeds the set proportion;

Whether the energy subsystem meets the single-day energy balance during the observation mission during the life cycle;

When the sun enters the field of view during the lifetime observation mission, exceeds the standard set ratio, and/or cannot meet the single-day energy balance, the orbital inclination angle i needs to be adjusted.

In another preferred embodiment of S2, during the extended service period, the two-dimensional solar shadow area in the sun illuminated area supplements the mission time of the star observation mission, which may include:

s21, when the satellite enters the sunlit area, adjust the two-dimensional attitude towards the sun so that the normal direction of the solar sail panel is aligned with the direction of sunlight;

s22, when the satellite enters the shadow area, it adjusts its vertical attitude toward the earth to carry out a short-term star observation mission, achieving a full life cycle of time drift at the descending node, and making full use of satellite resources to complete the set mission.

In another preferred embodiment of S3, based on the design results of the observation orbit inclination angle, it is again judged whether the adjusted orbit mission observation meets the observation requirements, including:

S31: Feed back the fully iterative observation orbit inclination design results of each subsystem to the overall designer;

S32, based on the observation orbit inclination design results and the simulation calculation results of the attitude control subsystem, thermal control subsystem and energy subsystem, and comprehensively the mission time satisfaction and subsystem index satisfaction, the overall designer decides whether to adopt the offset of the current orbit inclination i Set the measurement, and at the same time, the observation orbit inclination design result is fed back to the launch vehicle for calculation, and the decision is made whether to fine-tune the offset of the orbit inclination i;

S33, the overall designer finally determines the satellite orbit based on the orbital inclination iteration results of the satellite developer and launch developer.

In another preferred embodiment of S32, fine adjustment means: the adjustment of the orbital inclination angle i is within the range of ±0.1°.

One embodiment of the present invention provides a satellite full system coupling orbit design system. It should be noted that the steps in the method provided by the present invention can be implemented by using corresponding modules, devices, units, etc. in the system. Technology in the art Personnel can refer to the technical solutions of the method to realize the composition of the system, that is, the embodiments in the method can be understood as preferred examples of constructing the system, which will not be described again here.

As shown in Figure 2, the satellite full system coupled orbit design system provided by this embodiment may include:

Quasi-sun synchronous drift circular orbit calculation module, which is used to calculate quasi-sun synchronous drift circular orbit;

Orbital inclination design module. This module is based on the quasi-sun synchronous drifting circular orbit. According to the observation requirements of the star observation mission and the sun observation mission, it designs the time-sharing observation orbit inclination angle of the load of the star observation mission and the load of the sun observation mission, and During the extended service period, the two-dimensional solar shadow area in the sunlit area supplements the mission time of the star observation mission;

Orbital inclination adjustment module, this module judges again whether the adjusted orbital mission observation meets the observation requirements based on the design results of the observed orbital inclination: If not, the drift rate at the descending node is adjusted by the offset of the orbital inclination i , if satisfied, then determine whether the launch vehicle has an unimplementable state: if so, it ends; if not, fine-tune the offset of the orbital inclination i to complete the orbit design of the satellite's full system coupling.

The technical solution and its design principle provided by the above embodiments of the present invention will be further described in detail below with reference to a specific application example.

1. The input of the whole-system coupling orbit design method for satellites for large-span solar altitude angles provided by the above embodiments of the present invention mainly has two major mission requirements:

(1) Observe stars, 2-month window optional: Observation needs to be done when the solar altitude angle changes greatly, that is, the maximum solar altitude angle is as large as possible to ensure that the incident direction of the sun is as wide as possible when observing stars. Select a sun-synchronous orbit with the maximum solar altitude angle corresponding to the initial orbit being around 66° and the descending node time being 10:30 as one of the requirements for the long-span solar altitude angle, as shown in Figure 3.

(2) Observe the sun, with at least a 2-month window optional: it is necessary to observe when the change in the sun's altitude angle is small, that is, the maximum solar altitude angle should be as small as possible to ensure that the observation is within the range of [-17°, -13°] Spend as much time in the sun as possible. A sun-synchronous orbit corresponding to the maximum solar altitude angle of about 17° and the descending node local time of 6:30 is another requirement for the long-span solar altitude angle requirement, as shown in Figure 4.

It should be noted that the concept of the descending node local time mentioned above is the time when the satellite passes the descending node, that is, the local local time of the satellite's subsatellite point corresponding to the moment when the satellite crosses the equator from north to south.

2. Preliminary design of track

Based on the above two contradictory mission requirements, there are two orbit design options, as shown in Figure 10:

Option 1: Non-sun synchronous orbit

The sun moves periodically on both sides of the satellite's orbital plane. The angle between the sun's vector and the orbital plane, the β angle, changes periodically. That is, the sun moves periodically on both sides of the satellite's orbital plane. The β angle is the angle between the sun's vector and the orbital plane. Angle, the normal direction of the sun on the orbital plane is positive, and vice versa; the period of β angle is related to the orbital inclination angle i (the angle between the orbital plane and the equatorial plane). Figure 3 shows the orbital inclination angle i = around 42° The β angle change of the manned spacecraft orbit.

As shown in Figure 5, there are approximately 7 cycles in a year. By consulting the data, we can see that the β angle changes cycle is 54 days, and the change range is [-64°, 63°].

If the above-mentioned non-sun-synchronous orbit is used, a solar wing drive mechanism SADA with at least one-dimensional maneuverability needs to be installed on the structure. The following two installation methods can be used:

The first method: the solar wing is deployed along the ±Y direction, the normal direction of the sail is along the +Z direction, a one-dimensional SADA is installed in the ±Y direction, and rotates around the Y axis to solve the energy problem. The time interval during which the sun is in the vertical direction of the satellite's orbital plane Within (β is larger: within the interval [-90°, -12°)U[12°, 90°)]), the satellite rolling attitude is adjusted to complete the two-dimensional sun alignment or not perform the mission, the satellite platform and the one-dimensional SADA Cooperate to complete the tracking of the sun vector;

The second method: the solar wing is deployed along the ± 12°] interval and the sign changes), the attitude performs a yaw maneuver of 180°, so that the normal direction of the sailboard rotates 180° to receive the sunlight that has drifted to the other side of the orbit.

The two methods described above have the following disadvantages:

The first method will increase the funding cost, and the coupling control of attitude control and SADA will also bring control difficulty and even loss of accuracy. At the same time, if the load observation relies on the turntable mechanism to realize the coupling control of the three, it will be particularly complicated, reducing the reliability of the system.

In the second method, although the control system has reduced the expensive SADA rotating mechanism (about 2 million yuan) and the hardware configuration has been greatly simplified, the energy on the star is obviously tight during the period when the β angle is small. The reason is that the solar vector and the sail The angle between the normal direction of the plate (90°-β) is too large, resulting in the failure to carry out the load task almost 1/3 of the time. Moreover, because the attitude control system yaws 180°, it will bring new challenges to payload installation, optical sensor installation, communication antenna installation, and heat dissipation surface selection. It also requires the star service software to be able to adapt to the payload tasks in the two reference attitudes. Guidance increases the complexity of software system design and the difficulty of other sub-system design.

Option 2: Drift sun-synchronous orbit

According to the two contradictory mission requirements, based on the time span of the mission requirements, consider using a sun-synchronous orbit that drifts in the morning when the initial descending node time is near meridian, and the solar wing is fixedly installed at the same time. The normal orientation of the solar cells is the negative normal direction of the orbital plane of the satellite, and the solar wings are deployed along the ±X direction of the orbital system, as shown in Figure 6.

The so-called sun-synchronous orbit is an orbit in which the angle between the orbital plane and the solar vector is approximately constant. The mathematical formula is expressed as follows:

The calculation formula for the precession rate that satisfies the right ascension of the ascending node is as follows:

Considering circular orbit satellites, the above equation can be simplified to:

In the formula:

is the right ascension change rate of the ascending node, unit is °/day;

R _e is the average radius of the Earth's equator, which is 6378.137km;

a is the semi-major axis of the orbit, the unit is km;

i is the orbital inclination angle, unit is °;

e is the orbital eccentricity, dimensionless;

From the above formula, we can get several characteristics of sun-synchronous orbit:

When satellites pass through the same latitude and in the same direction of movement (from south to north or from north to south), the local time of the subsatellite point is the same;

The orbital semi-major axis a and the orbital inclination angle i of the sun-synchronous circular orbit satisfy a fixed relationship;

The orbital inclination of the sun-synchronous orbit must be greater than 95.675°. The altitude of the sun-synchronous orbit with the orbital inclination i=95.675° is 0, which is a retrograde orbit;

The orbital altitude of the sun-synchronous orbit has an upper limit, which does not exceed 5959km. At this time, the orbital inclination angle i is 180°.

The relationship between the orbital altitude and the orbital inclination angle i of a sun-synchronous orbit with an eccentricity of 0 is shown in Figure 7.

Descending node local time is a very important design indicator for sun-synchronous orbits. It directly represents the solar illumination of the satellite in orbit, as shown in Figure 8. For example, to ensure that the satellite has good energy conditions, you can choose a sun-synchronous orbit at dawn and dusk with the descending node at 6 o'clock. At this time, the direction of sunlight is approximately perpendicular to the satellite orbital plane. Keep the sail expansion direction within the orbital plane to ensure Energy conditions are always good. For earth remote sensing observation satellites, try to choose a meridian sun-synchronous orbit with the descending node local time close to 12 o'clock. At this time, the direction of sunlight is basically parallel to the satellite orbital plane, ensuring that the satellite has good along-light observation conditions when imaging the ground.

The drifting sun-synchronous orbit design is to change the fixed relationship between the orbital semi-major axis α and the orbital inclination angle i, so that the relationship between the orbital surface and the sun is no longer synchronous, but has a slow drift process. For this invention, consider designing a drift orbit where the descending node is [10:00, 2:00], and set the period between [10:00, 8:00] and [4:00, 2:00] as task one ( (Observing stars), [8:00, 4:00] is used as the task window for Task 2 (observing the sun that has just risen). According to the circular sun-synchronous orbit design formula Selecting an orbital altitude of 280km requires an orbital inclination angle i of 3° negative relative to the nominal sun-synchronous orbit to ensure that the local time drift of the descending node is 54min for 4 weeks.

Table 1 The relationship between orbital inclination angle i (nominal 96.67°) offset and orbital drift

The specific calculation is derived as follows:

The influence of semi-major axis deviation and inclination angle deviation on the right ascension of the ascending node is:

In the above formula, Δa and Δi represent the deviation values of the actual semi-major axis a and the actual orbit inclination angle i from the nominal semi-major axis a ^* and the nominal orbit inclination angle i ^* of the sun-synchronous orbit respectively; the unit of Δa is km, and the unit of Δi is °;

In the formula, is the change in the right ascension change rate of the ascending node;

Δa＝aa ^* (5.57)

Δi＝ii ^* (5.58)

The so-called nominal semi-major axis and nominal inclination angle satisfy the fixed relationship between the semi-major axis and inclination angle of the nominal sun-synchronous orbit. It should be noted here that formula (5.56) is applicable when Δa and Δi are small quantities (less than 1/10 of their own values), and it should be noted that the unit of Δi is radians.

The influence of the semi-major axis deviation and inclination angle deviation on the right ascension of the ascending node is converted into the drift amount of the descending node local time, and considered separately, the change of the drift rate of the descending node local time is obtained:

In the formula:

Δt _a-1day is the drift amount of the descending node local time caused by the offset of a relative to the nominal value in one day, unit min/day.

It is the drift amount of the right ascension of the ascending node due to the offset of a relative to the nominal value in one day, unit °/day.

Δt _i-1day is the drift of the descending node local time in one day due to the offset of i relative to the nominal value, in min/day.

It is the drift amount of the local time of the descending node due to the offset of i relative to the nominal value in one day, unit °/day.

From formulas (5.59) and (5.60) we can know:

When Δa>0, Δt _a-1day <0, the time at the descending node drifts westward;

When Δa<0, Δt _a-1day >0, the time at the descending node drifts eastward;

When Δi>0, Δt _i-1day >0, the time at the descending node drifts eastward;

When Δi<0, Δt _i-1day <0, the time at the descending node drifts westward.

Orbit design iteration: provide the orbit design results according to the mission requirements, provide the observation window satisfaction of mission one and mission two at different launch times based on the orbit design results, and determine each subsystem based on the requirements of mission one and mission two for the entire satellite system. If the mission requirements, observation requirements and sub-system requirements for the mission are met, it will stop. If the observation requirements and sub-system requirements for the mission are not met, feedback requirements will be given to the orbit design, and the offset of the orbit inclination i will be corrected for further iterations. Specifically:

According to the preliminary orbit design, first determine whether the mission requirements are met, and provide preliminary statistical results of different typical descending node local times, which are used to obtain the observable descending node local times for tasks one and two during the descending node local time drift process. value, giving the mapping relationship. The preliminary statistical results are as follows:

Table 2 Angle and average statistics of the highest solar altitude angle

The satellite launch window generally ranges from 10min to 30min, which means that the descending node local time of the satellite may be delayed by 10min to 30min due to the uncertainty of the launch time. Therefore, the descending node local time range that the satellite should adapt to after landing should be [ 10:30,2:00], the following will take this interval as an example to illustrate the problem.

Table 3 shows the observable duration of mission 2 corresponding to different orbital solar angles β, and provides the mapping relationship between different β angles corresponding to different solar altitude angles and the corresponding durations.

Table 3 Different β angles correspond to different solar altitude angles and corresponding durations (half orbit, unit s)

Note: The required solar altitude angle range for observing the sun is [-17°, -13°].

The present invention selects different periods to arrange different tasks according to the situation of orbital drift. Among them: in the early stage of orbit entry, under the premise that the orbital solar angle β is less than 25°, according to the energy on-orbit status, the load startup frequency of the star observation mission is reasonably arranged (for example, β is [ In the range of 20°, 30°], the power-on time per day is reduced to 3 hours, and in the range of [55°, 60°], the power-on time per day is 6 hours. Since the multiple relationship of the average illumination angle of the sailboard is sin(60°)/ sin(25°)=2.05); after setting time (for example, 2 months), after the local time of the descending node drifts to less than 8:30, the orbital solar angle β becomes larger to the range of [45°, 63°], and the sun The energy output capacity of the cells will be improved, and solar observation missions requiring observation times longer than [5min, 10min] will be carried out.

Although the orbit designed based on the mission can ensure the completion of the mission, the high reliability requirements of aerospace often cause the service time of satellites to be much longer than the design time. Therefore, this specific application example considers that after 8 months (8 months is the time for satellite orbit maintenance The upper limit of the capability, which can also be said to be the life requirement of the satellite) During the drift process from 2:00 to 0:00 to 22:00 at the descending node, the direction of the solar wing sail panel in the sunlit area points to the sun, as shown in Figure 9 Show.

At the same time, considering capacity expansion, star observation missions can be arranged in the shadow area. The satellite in the sunlit area adjusts its attitude so that the normal direction of the sailboard is aligned with the sun to charge the battery to ensure energy security.

One embodiment of the present invention provides a computer terminal, which includes a memory, a processor, and a computer program stored in the memory and executable on the processor. When the processor executes the program, it can be used to execute any of the above embodiments of the present invention. The method, or the system that runs any one of the above embodiments of the present invention.

Optionally, memory is used to store programs; memory may include volatile memory (English: volatile memory), such as random access memory (English: random-access memory, abbreviation: RAM), such as static random access memory ( English: static random-access memory, abbreviation: SRAM), Double Data Rate Synchronous Dynamic Random Access Memory (English: Double Data Rate Synchronous Dynamic Random Access Memory, abbreviation: DDR SDRAM), etc.; memory can also include non-volatile Memory (English: non-volatile memory), such as flash memory (English: flash memory). The memory is used to store computer programs (such as application programs, functional modules, etc. that implement the above methods), computer instructions, etc. The above-mentioned computer programs, computer instructions, etc. can be stored in one or more memories in partitions. And the above-mentioned computer programs, computer instructions, data, etc. can be called by the processor.

The above-mentioned computer programs, computer instructions, etc. can be stored in one or more memories in partitions. And the above-mentioned computer programs, computer instructions, data, etc. can be called by the processor.

The processor is configured to execute the computer program stored in the memory to implement each step in the method or each module of the system involved in the above embodiments. For details, please refer to the relevant descriptions in the previous method and system embodiments.

The processor and memory can be independent structures or integrated structures integrated together. When the processor and memory are independent structures, the memory and processor can be connected through bus coupling.

An embodiment of the present invention also provides a computer-readable storage medium on which a computer program is stored. The program, when executed by a processor, can be used to perform any of the methods of the above embodiments of the present invention, or to run the present invention. The system of any of the above embodiments.

The satellite full system coupling orbit design method, system, terminal and medium provided by the above embodiments of the present invention are used for the orbit design of the satellite full system coupling with a large span of solar altitude angles. The orbit drift rate is designed according to the mission requirements, that is, according to the on-orbit mission. According to the window requirements and the time requirement of the on-orbit test, different orbit offset rates are designed through the size of the inclination offset. The relationship between the inclination offset and the local time drift law of the orbit descending node is a linear relationship. In the inclination offset design process Iterate attitude control, orbit control, thermal control, and energy to meet project implementation requirements; take mission requirements as the first goal, and optimize the design of the entire satellite system as the second goal, fully considering satellite attitude control, thermal control, and satellite operations The coupling design of software, energy and other sub-systems enables the orbit design to achieve top-level system optimization and achieve the most cost-effective basis for satellite development and design.

Those skilled in the art know that in addition to implementing the system and its various devices provided by the present invention in the form of pure computer-readable program codes, the system and its various devices provided by the present invention can be completely programmed with logic gates, The same function can be achieved in the form of switches, application-specific integrated circuits, programmable logic controllers, and embedded microcontrollers. Therefore, the system and its various devices provided by the present invention can be regarded as a hardware component, and the devices included in it for implementing various functions can also be regarded as structures within the hardware components; The means for implementing various functions are considered to be either software modules implementing methods or structures within hardware components.

Matters not mentioned in the above embodiments of the present invention are all well-known technologies in the art.

Specific embodiments of the present invention have been described above. It should be understood that the present invention is not limited to the specific embodiments described above. Those skilled in the art can make various variations or modifications within the scope of the claims, which does not affect the essence of the present invention.

Claims (10)

1. A satellite full system coupled orbit design method, which is characterized by including: Calculation of quasi-sun-synchronous drifting circular orbits; Based on the quasi-sun synchronous drifting circular orbit, according to the observation requirements of the star observation mission and the sun observation mission, the payload of the star observation mission and the time-sharing observation orbit inclination of the payload of the sun observation mission are designed, and in During the extended service period, the two-dimensional solar shadow area in the sunlit area supplements the mission time of the star observation mission; According to the design results of the observation orbit inclination angle, judge again whether the adjusted orbit mission observation meets the observation requirements: if not, adjust the drift rate at the descending node through the offset of the orbit inclination angle i. If it meets, judge whether the carrier Whether the rocket has an unimplementable state: if so, end. If not, fine-tune the offset of the orbital inclination i to complete the orbit design of the satellite's full system coupling. 2. The satellite full system coupled orbit design method according to claim 1, characterized in that it also includes any one or more of the following: -The observation requirements of the star observation mission include: the observation requirements for the first payload and the second payload with the highest solar altitude angle greater than or equal to 45° in the star observation mission; the observation requirements of the sun observation mission include: observing the sun Observation requirements for the third payload with the highest solar altitude angle within the closed range of [17°, 45°] during the mission; - Design the drift orbit of the descending node at [10:00, 2:00], and use the period between [10:00, 8:00] and [4:00, 2:00] as the mission window for the star observation mission. [8:00, 4:00] serves as the mission window to observe the sun just rising. 3. The satellite full system coupled orbit design method according to claim 1, characterized in that the calculation of the quasi-sun synchronous drift circular orbit includes: Obtain the characteristic parameters required to calculate the quasi-sun synchronous drift circular orbit, the characteristic parameters include: orbit semi-major axis a, orbit eccentricity e and orbit inclination angle i; According to the average angle of the sun's daily motion relative to the earth, the corresponding relationship between the orbital semi-major axis a, the orbital eccentricity e and the orbital inclination angle i is constructed, the orbital height is given, and the orbital eccentricity e is set =0, the orbital inclination angle i is obtained. According to the observation requirements of different missions, the negative offset orbital inclination angle i is calculated to obtain a quasi-sun-synchronous drifting circular orbit. 4. The satellite full system coupling orbit design method according to claim 1, characterized in that, based on the quasi-sun synchronous drift circular orbit, in view of the observation requirements of the star observation mission and the sun observation mission, the observation requirements of the star observation mission are The payload and the time-sharing observation orbit of the payload of the solar observation mission are designed, including: Conduct simulation calculations based on the quasi-sun synchronous drift circular orbit to obtain the observable time intervals of star observation missions and sun observation missions with different launch into orbit times, and decide whether to adjust the orbital inclination i according to the mission time satisfaction; The offset of the orbital inclination i is adjusted according to the mission observation requirements, and simulation calculations are performed based on the adjusted quasi-sun synchronous drift circular orbit to obtain the orbital period, sunshade time distribution, solar vector and various aspects of the satellite body for different launch into orbit times. The relationship data of the given satellite attitude control, thermal control, and energy subsystems are analyzed to determine whether there is a risk of not meeting the subsystem index requirements; if there is a risk, the orbit inclination offset is reduced by the set degree, if not, continue Next step. 5. The satellite full system coupled orbit design method according to claim 4, characterized in that the decision of whether to adjust the orbit inclination i according to the mission time satisfaction includes: The task time satisfaction includes: During the lifetime observation mission, whether the star sensor of the attitude control subsystem has the time when the sun enters the field of view; Whether the temperature of a single machine controlled by the thermal control subsystem during the observation mission during the life cycle exceeds the set proportion; Whether the energy subsystem meets the single-day energy balance during the observation mission during the life cycle; When the sun enters the field of view during the observation mission within the lifetime, exceeds the standard setting ratio, and/or cannot meet the single-day energy balance, the orbital inclination angle i needs to be adjusted. 6. The satellite full system coupling orbit design method according to claim 1, characterized in that the mission time of the two-dimensional solar shadow area supplementary observation star mission in the sunlit area during the extended service period includes: When the satellite enters the sunlit area, adjust the two-dimensional attitude towards the sun so that the normal direction of the solar sail panel is aligned with the direction of sunlight; When the satellite enters the shadow area, it adjusts its vertical attitude toward the earth to carry out a short-term star observation mission, and achieves a full life cycle of time drift at the descending node, which can make full use of satellite resources to complete the set mission. 7. The satellite full system coupling orbit design method according to claim 1, characterized in that, based on the design results of the observation orbit inclination angle, it is again judged whether the adjusted orbit mission observation meets the observation requirements, including: Obtain fully iterative observation orbit inclination design results for each subsystem; Based on the design results of the observed orbit inclination angle and the simulation calculation results of the attitude control subsystem, thermal control subsystem and energy subsystem, comprehensive mission time satisfaction and subsystem index satisfaction, a decision is made whether to adopt the offset of the current orbit inclination i, At the same time, the observation orbit inclination design results are fed back to the launch vehicle for calculation and decision-making whether to fine-tune the offset of the orbit inclination i; the fine-tuning means: the adjustment of the orbit inclination i is within the range of ±0.1°; The satellite orbit is finally determined based on the orbital inclination iteration results. 8. A satellite full system coupled orbit design system, which is characterized by including: Quasi-sun synchronous drift circular orbit calculation module, which is used to calculate quasi-sun synchronous drift circular orbit; Orbital inclination design module, which is based on the quasi-sun synchronous drift circular orbit, and is based on the observation requirements of the star observation mission and the sun observation mission, and the time-sharing observation orbit for the load of the star observation mission and the load of the sun observation mission. The design work is carried out based on the inclination angle, and the two-dimensional sun shadow area in the sunlit area supplements the mission time of the star observation mission during the extended service period; Orbital inclination adjustment module, this module judges again whether the adjusted orbital mission observation meets the observation requirements based on the design results of the observed orbital inclination: If not, the drift rate at the descending node is adjusted by the offset of the orbital inclination i , if satisfied, then determine whether the launch vehicle has an unimplementable state: if so, it ends; if not, fine-tune the offset of the orbital inclination i to complete the orbit design of the satellite's full system coupling. 9. A computer terminal, including a memory, a processor, and a computer program stored in the memory and executable on the processor, characterized in that when the processor executes the program, it can be used to execute claims 1-7 The method of any one of the above, or running the system of claim 8. 10. A computer-readable storage medium with a computer program stored thereon, characterized in that, when executed by a processor, the program can be used to perform the method of any one of claims 1-7, or to run the method of claim 1 The system described in 8.