Satellite in-orbit propellant consumption optimization method

Abstract

The invention relates to a satellite in-orbit propellant consumption optimization method, which is suitable for a satellite in a long-term operation mode, and comprises the following steps: S1, determining a satellite orbit condition, and calculating an earth shadow duration T actually experienced by the satellite; s2, determining a satellite load working condition, and analyzing and determining the maximum earth orientation duration T'meeting an energy balance requirement; s3, adjusting a shadow inlet point and a shadow outlet point based on the earth shadow duration T and the maximum earth orientation duration T '; s4, adjusting the image entering point and the image exiting point according to a preset step length n based on the performance requirement and the thermal control balance requirement of the storage battery, and repeatedly executing the step S3 to obtain the optimized image entering point and image exiting point; and step S5, completing satellite in-orbit attitude adjustment based on the optimized image entering point and image exiting point. According to the invention, the problem of propellant shortage of the satellite in various scenes such as long-year solar activity and late service life of the satellite is solved, and the attenuation of the satellite orbit can be delayed, so that the fuel consumption is reduced.

Description

Technical Field

[0001] The present invention relates to the field of spacecraft control technology, and in particular to a method for optimizing satellite on-orbit propellant consumption. Background Art

[0002] Once a satellite is in orbit, propellant cannot be replenished. If propellant resources are insufficient, it will directly lead to the failure to complete the mission, with unacceptable consequences.

[0003] When designing a satellite plan, the fuel required for the satellite during its on-orbit lifespan is calculated. This is done to ensure, on the one hand, that the satellite carries enough fuel to meet the on-orbit lifespan requirements, and on the other hand to avoid unnecessary surplus and reduce the weight of the satellite.

[0004] Satellite propellant consumption mainly includes: 1. Adjustment of the initial orbital semi-major axis and orbital inclination; 2. Attitude control when the control subsystem is operating in other modes other than the normal operating mode; 3. Orbit maintenance required during the life cycle. Orbit maintenance includes compensating for orbital decay caused by the atmosphere, maintaining constellation formation flight, compensating for other perturbations, and performing collision avoidance. According to statistics, for low-orbit satellites, fuel consumption caused by natural orbital decay accounts for the vast majority.

[0005] Although strict estimates of propellant requirements are made during the satellite design phase, there are many situations where a satellite may still be running low on propellant:

[0006] (1) If the orbital deviation is too large, the initial orbit adjustment of the satellite will require the consumption of extra-budgetary propellant.

[0007] (2) Solar activity cannot be accurately predicted. Atmospheric drag is closely related to the atmospheric density at the orbital altitude. The atmospheric density ρ at the orbital altitude is greatly affected by solar activity. The atmospheric density can vary by more than an order of magnitude between high and low solar activity years. The solar activity cycle is eleven years, but actual solar activity cannot be accurately predicted.

[0008] (3) Other unexpected reasons. Taking the constellation satellites maintaining the formation configuration as an example, the difference in the face value ratio between multiple satellites and the failure of the orbital control of member satellites will all lead to unexpected orbital control and propellant consumption.

[0009] Therefore, proposing a method to optimize satellite on-orbit propellant consumption is an urgent problem to be solved. Summary of the invention

[0010] In order to solve the technical problems existing in the above-mentioned prior art, the purpose of the present invention is to provide a method for optimizing satellite on-orbit propellant consumption, which can reduce the fuel consumption during the orbit maintenance phase and extend the life of on-orbit satellites with insufficient fuel.

[0011] To achieve the above-mentioned object of the invention, the present invention provides a method for optimizing satellite on-orbit propellant consumption, which is suitable for satellites in long-term operation mode, and comprises the following steps:

[0012] Step S1, determine the satellite orbit conditions and calculate the actual earth shadow duration T experienced by the satellite;

[0013] Step S2, determining the satellite load condition, analyzing and determining the maximum earth orientation time T' that meets the energy balance requirement;

[0014] Step S3, adjusting the shadow entry point and the shadow exit point based on the ground shadow duration T and the maximum ground orientation duration T';

[0015] Step S4: Based on the battery performance requirements and thermal control balance requirements, the shadow entry point and the shadow exit point are adjusted according to a preset step length n, and the step S3 is repeatedly executed to obtain optimized shadow entry point and shadow exit point;

[0016] Step S5: Based on the optimized shadow entry point and shadow exit point, the satellite's on-orbit attitude adjustment is completed;

[0017] Among them, the long-term operation mode refers to the operation mode of "cruising towards the sun in the sunny area and orienteering towards the ground in the shadow area".

[0018] According to a technical solution of the present invention, the step S2 specifically includes:

[0019] Step S21, confirming the satellite load condition, calculating the satellite long-term load power and the satellite short-term load power under different operating modes;

[0020] Step S22: Conduct energy balance analysis to determine the worst energy balance load condition.

[0021] Step S23: Based on the worst energy balance condition, determine the maximum ground orientation time T' that meets the energy balance requirement;

[0022] Among them, energy balance means that the satellite's energy input can meet the energy output demand.

[0023] According to a technical solution of the present invention, step S3 specifically includes:

[0024] Step S31, let α be the modification amount of the shadow entry point and the shadow exit point, α=(T'-T) / 2.

[0025] Step S32, confirming a new shadow entry point A', which is α ahead of the shadow entry point A;

[0026] Step S33, confirm the new exit point B', which is delayed by α relative to the exit point B.

[0027] According to a technical solution of the present invention, the step S4 specifically includes:

[0028] Step S41, judging whether the new shadow entry point A' and the new shadow exit point B' meet the battery performance requirements, if yes, executing step S42, otherwise, executing step S43;

[0029] Step S42, judging whether the new shadow entry point A' and the new shadow exit point B' meet the thermal control balance requirement;

[0030] Step S43, gradually reduce the ground shadow duration modification amount according to the preset step length n, expressed as: α'=α-n, repeat the step S3, obtain the corrected shadow entry point and the corrected shadow exit point, iterate until the battery performance requirements and thermal control balance requirements are met, and obtain the optimized shadow entry point and shadow exit point.

[0031] According to a technical solution of the present invention, the judgment condition of step S41 is:

[0032] The satellite attitude conversion is controlled according to the new shadow entry point A' and the new shadow exit point B'. When the satellite is in different working modes, the battery power consumption does not exceed the maximum discharge depth.

[0033] According to a technical solution of the present invention, the judgment condition of step S42 is:

[0034] The satellite attitude conversion is controlled according to the new entry point A' and the new exit point B'. During the satellite operation, when the internal and external thermal environments can reach a stable state, all instruments and subsystems on the satellite can operate normally within the specified operating temperature range.

[0035] Compared with the prior art, the present invention has the following beneficial effects:

[0036] The present invention proposes a method for optimizing the on-orbit propellant consumption of satellites. The method reduces the fuel consumption in the orbit maintenance phase by adjusting the satellite attitude, solves the problem of propellant shortage in various scenarios such as years of high solar activity and the late years of satellite life, reduces the time when the sailboard faces the direction of flight, and reduces the equivalent frontal area, which can delay the satellite orbit decay and thus reduce fuel consumption.

[0037] The present invention can reduce propellant consumption by using this method for on-orbit satellites with sufficient power reserve, thereby helping the satellite to extend its on-orbit lifespan; for satellites with insufficient propellant, the method can reduce propellant consumption by changing the satellite use strategy or shortening the satellite use time.

[0038] The invention is easy to implement in engineering and has great practical value. BRIEF DESCRIPTION OF THE DRAWINGS

[0039] In order to more clearly illustrate the embodiments of the present invention or the technical solutions in the prior art, the drawings required for use in the embodiments are briefly introduced below. Obviously, the drawings described below are only some embodiments of the present invention, and for ordinary technicians in this field, other drawings can be obtained based on these drawings without creative work.

[0040] Figure 1 A flowchart of a method for planning narrow-strip area mosaic observation of an agile imaging satellite provided in an implementation case of the present invention is schematically shown;

[0041] Figure 2 A schematic diagram of attitude pointing of a satellite in a long-term operation mode according to an implementation case of the present invention is schematically shown;

[0042] Figure 3 The flowchart of a method for planning narrow-strip area mosaic observation of an agile imaging satellite provided in another implementation example of the present invention is schematically shown. DETAILED DESCRIPTION

[0043] In order to make the purpose, technical solution and advantages of the embodiments of the present invention clearer, the technical solution in the embodiments of the present invention will be clearly and completely described below in conjunction with the drawings in the embodiments of the present invention. Obviously, the described embodiments are part of the embodiments of the present invention, not all of the embodiments. Based on the embodiments of the present invention, all other embodiments obtained by ordinary technicians in this field without creative work are within the scope of protection of the present invention.

[0044] The fuel required by the satellite during its in-orbit period consists of three parts: initial orbit adjustment, orbit maintenance and attitude control. Among them, the fuel consumption caused by orbit maintenance accounts for the vast majority, mainly used to compensate for the orbital attenuation caused by the atmosphere.

[0045] For low-orbit satellites, the orbital atmospheric drag has a significant impact. The calculation formula for the semi-major axis attenuation rate of the satellite orbit caused by atmospheric drag is:

[0046]

[0047] In the formula, C d is the aerodynamic drag coefficient; ρ is the atmospheric density; S is the satellite's frontal area; m is the satellite's mass; n is the satellite's average angular velocity; and f is the satellite's true anomaly.

[0048] Considering that the satellite is in a near-circular orbit and the eccentricity is close to 0, the semi-major axis attenuation rate can be simplified to:

[0049]

[0050] The calculation formula of the satellite's semi-major axis attenuation rate shows that the semi-major axis attenuation rate is proportional to the satellite's windward area. Adjusting the satellite's flight strategy and reducing the windward area to reduce orbital attenuation is a way to save fuel.

[0051] In the long-term operation state, the satellite adopts the mode of "cruising towards the sun in the sunlit area and orienting towards the earth in the shadowed area". In the sunlit area, the whole satellite cruises towards the sun, the -Z axis of the satellite points to the direction of the sun, the cross product of the Z axis and the normal direction of the orbital plane is the X axis, and the cross product of the Z axis and the X axis is the Y axis. In the shadowed area, the whole satellite is oriented towards the earth with three axes, and the +Z axis of the satellite points to the direction of the earth. Figure 2 As shown, point A is the point where the shadow enters the ground, point B is the point where the shadow exits the ground, and the area between points A and B is the shadow area.

[0052] When the satellite moves from the shadow area to the point before exiting the earth's shadow, it will perform an attitude maneuver, changing from three-axis earth orientation to sun cruise, and complete the attitude maneuver before exiting the earth's shadow point B; the satellite in the sunlit area will maintain sun cruise, and when there is a payload mission, it will perform an attitude maneuver according to the command and adjust to the payload working attitude; after the satellite moves from the sunlit area to the point A entering the earth's shadow, it will perform an attitude maneuver, adjusting from sun cruise to three-axis earth orientation; the satellite in the shadow area will maintain a three-axis earth orientation attitude. The time of entering the earth's shadow point A and exiting the earth's shadow point B is calculated autonomously by the control subsystem.

[0053] Assume that the cross-sectional area of ​​the satellite body is a square meters, and the area of ​​the sailboard is b square meters. The satellite adopts the attitude pointing strategy of the shadow area facing the ground and the sunlit area facing the sun. In the attitude facing the ground, the satellite is facing the wind on the +X side, and the windward area is about a square meter of the cross-sectional area of ​​the satellite body. At this time, the windward area is the smallest. In the attitude facing the sun, at both ends of the orbital arc segment illuminated by the sun, the satellite is facing the wind on the +Z side, and the windward area is about a+b square meters. At this time, the windward area is the largest. The average of the satellite's windward area within one orbital period is taken to obtain the equivalent windward area.

[0054] In summary, the on-orbit propellant consumption can be optimized by lengthening the earth's shadow, reducing the time that the sailboard faces the flight direction, reducing the equivalent frontal area, and delaying orbital attenuation.

[0055] like Figures 1 to 3 As shown, a satellite on-orbit propellant consumption optimization method of the present invention is suitable for satellites in long-term operation mode, and the optimization method includes the following steps:

[0056] Step S1, determine the satellite orbit conditions and calculate the actual earth shadow duration T experienced by the satellite;

[0057] Step S2: Determine the satellite load condition, analyze and determine the maximum earth orientation time T' that meets the energy balance requirement, specifically including:

[0058] Step S21, confirming the satellite load condition, calculating the satellite long-term load power and the satellite short-term load power under different operating modes;

[0059] Step S22: Conduct energy balance analysis to determine the worst energy balance load condition.

[0060] Step S23: Based on the worst energy balance condition, determine the maximum ground orientation time T' that meets the energy balance requirement;

[0061] Among them, energy balance means that the satellite's energy input can meet the energy output demand.

[0062] Step S3, adjusting the shadow entry point and the shadow exit point based on the ground shadow duration T and the maximum ground orientation duration T', specifically includes:

[0063] Step S31, let α be the modification amount of the shadow entry point and the shadow exit point, α=(T'-T) / 2.

[0064] Step S32, confirming a new shadow entry point A', which is α ahead of the shadow entry point A;

[0065] Step S33, confirm the new exit point B', which is delayed by α relative to the exit point B.

[0066] Step S4, based on the battery performance requirements and thermal control balance requirements, adjusting the shadow entry point and the shadow exit point according to the preset step length n, and repeatedly executing step S3 to obtain the optimized shadow entry point and shadow exit point, specifically including:

[0067] Step S41, judging whether the new shadow entry point A' and the new shadow exit point B' meet the battery performance requirements, if yes, executing step S42, otherwise, executing step S43;

[0068] To ensure the stable performance of the battery pack on track, the battery must meet the maximum discharge depth requirement and leave a certain margin. The judgment condition of step S41 is:

[0069] The satellite attitude conversion is controlled according to the new shadow entry point A' and the new shadow exit point B'. When the satellite is in different working modes, the battery power consumption does not exceed the maximum discharge depth.

[0070] Step S42, judging whether the new shadow entry point A' and the new shadow exit point B' meet the thermal control balance requirement;

[0071] The judgment condition of step S42 is:

[0072] The satellite attitude conversion is controlled according to the new entry point A' and the new exit point B'. During the satellite operation, when the internal and external thermal environments can reach a stable state, all instruments and subsystems on the satellite can operate normally within the specified operating temperature range.

[0073] Step S43, gradually reduce the ground shadow duration modification amount according to the preset step length n, expressed as: α'=α-n, repeat the step S3, obtain the corrected shadow entry point and the corrected shadow exit point, iterate until the battery performance requirements and thermal control balance requirements are met, and obtain the optimized shadow entry point and shadow exit point.

[0074] Step S5: Based on the optimized shadow entry point and shadow exit point, the satellite's on-orbit attitude adjustment is completed;

[0075] Based on the optimized entry and exit points, on-orbit modifications are completed and the attitude of the on-orbit satellite is adjusted to lengthen the earth's shadow, reduce the time that the sailboard faces the direction of flight, reduce the equivalent frontal area, and delay orbital attenuation, thereby optimizing the satellite's on-orbit propellant consumption.

[0076] In the present invention, the terms "include", "comprises" or any other variants thereof are intended to cover non-exclusive inclusion, so that a process, method, article or terminal device including a series of elements includes not only those elements, but also includes other elements not explicitly listed, or also includes elements inherent to such process, method, article or terminal device. In the absence of further restrictions, an element defined by the sentence "comprises a ..." does not exclude the existence of other identical elements in the process, method, article or terminal device including the element.

[0077] Finally, it should be noted that the above is a preferred embodiment of the present invention. It should be pointed out that although the preferred embodiment of the present invention has been described, for those skilled in the art, once the basic creative concept of the present invention is known, several improvements and modifications can be made without departing from the principles of the present invention, and these improvements and modifications should also be regarded as the protection scope of the present invention. Therefore, the attached claims are intended to be interpreted as including the preferred embodiment and all changes and modifications that fall within the scope of the embodiments of the present invention.

Claims

1. A method for optimizing satellite on-orbit propellant consumption, characterized in that: Adapting a satellite to long-term operation mode includes the following steps: Step S1, determine the satellite orbit conditions and calculate the actual earth shadow duration T experienced by the satellite; Step S2, determining the satellite load condition, analyzing and determining the maximum earth orientation time T' that meets the energy balance requirement; Step S3, adjusting the shadow entry point and the shadow exit point based on the ground shadow duration T and the maximum ground orientation duration T'; Step S4: Based on the battery performance requirements and thermal control balance requirements, the shadow entry point and the shadow exit point are adjusted according to a preset step length n, and the step S3 is repeatedly executed to obtain optimized shadow entry point and shadow exit point; Step S5: Based on the optimized shadow entry point and shadow exit point, the satellite's on-orbit attitude adjustment is completed; Among them, the long-term operation mode refers to the operation mode of "cruising towards the sun in the sunny area and orienteering towards the ground in the shadow area".

2. The satellite on-orbit propellant consumption optimization method according to claim 1, characterized in that: The step S2 specifically includes: Step S21, confirming the satellite load condition, calculating the satellite long-term load power and the satellite short-term load power under different operating modes; Step S22: Conduct energy balance analysis to determine the worst energy balance load condition. Step S23: Based on the worst energy balance condition, determine the maximum ground orientation time T' that meets the energy balance requirement; Among them, energy balance means that the satellite's energy input can meet the energy output demand.

3. The satellite on-orbit propellant consumption optimization method according to claim 1, characterized in that: The step S3 specifically includes: Step S31, let α be the modification amount of the shadow entry point and the shadow exit point, α=(T'-T) / 2. Step S32, confirming a new shadow entry point A', which is α ahead of the shadow entry point A; Step S33, confirm the new exit point B', which is delayed by α relative to the exit point B.

4. The satellite on-orbit propellant consumption optimization method according to claim 3, characterized in that: The step S4 specifically includes: Step S41, judging whether the new shadow entry point A' and the new shadow exit point B' meet the battery performance requirements, if yes, executing step S42, otherwise, executing step S43; Step S42, judging whether the new shadow entry point A' and the new shadow exit point B' meet the thermal control balance requirement; Step S43, gradually reduce the ground shadow duration modification amount according to the preset step length n, expressed as: α'=α-n, repeat the step S3, obtain the corrected shadow entry point and the corrected shadow exit point, iterate until the battery performance requirements and thermal control balance requirements are met, and obtain the optimized shadow entry point and shadow exit point.

5. The method for optimizing satellite on-orbit propellant consumption according to claim 4, characterized in that: The judgment condition of step S41 is: The satellite attitude conversion is controlled according to the new shadow entry point A' and the new shadow exit point B'. When the satellite is in different working modes, the battery power consumption does not exceed the maximum discharge depth.

6. The satellite on-orbit propellant consumption optimization method according to claim 4, characterized in that: The judgment condition of step S42 is: The satellite attitude conversion is controlled according to the new entry point A' and the new exit point B'. During the satellite operation, when the internal and external thermal environments can reach a stable state, all instruments and subsystems on the satellite can operate normally within the specified operating temperature range.

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