阅读说明：本技术 异常分离情况下组合体geo变轨策略生成方法及系统 (Method and system for generating GEO orbital transfer strategy of assembly under abnormal separation condition ) 是由 陈占胜 李楠 成飞 邓武东 潘瑞雪 杨牧 于 2021-07-05 设计创作，主要内容包括：本发明提供了一种异常分离情况下组合体GEO变轨策略生成方法及系统,包括：根据组合体与运载器或组合体之间异常分离时刻状态,确定组合体分离点轨道参数及与组合体轨道转移策略生成相关的平台参数；根据平台参数确定轨道转移策略的设计约束条件；根据所述约束条件以轨道转移所需速度增量最小为原则,进行最优变轨次数估计；根据估计结果,建立描述单次变轨状态变化及多次变轨间联系的数学模型；利用所述数学模型以GEO轨道为目标,通过迭代寻优生成满足设计约束的轨道转移策略。本发明解决了组合体航天器异常分离条件下,传统有限推力复杂设计方法计算量大、计算速度慢等不足,具有一定的工程实用性。(The invention provides a method and a system for generating a GEO orbital transfer strategy of an assembly under the condition of abnormal separation, which comprises the following steps: determining an orbit parameter of a separation point of the assembly and a platform parameter related to the generation of an orbit transfer strategy of the assembly according to the state of the abnormal separation moment between the assembly and the carrier or between the assemblies; determining a design constraint condition of a track transfer strategy according to the platform parameters; according to the constraint conditions, carrying out optimal orbital transfer times estimation on the basis of the principle that the speed increment required by orbital transfer is minimum; establishing a mathematical model for describing the single orbital transfer state change and the connection among multiple orbital transfers according to the estimation result; and generating an orbit transfer strategy meeting design constraints by using the mathematical model and taking the GEO orbit as a target through iterative optimization. The method overcomes the defects of large calculated amount, low calculating speed and the like of the traditional limited thrust complex design method under the condition of abnormal separation of the combined spacecraft, and has certain engineering practicability.)

1. A method for generating a GEO orbital transfer strategy of an assembly under the condition of abnormal separation is characterized by comprising the following steps:

step A: determining an orbit parameter of a separation point of the assembly and a platform parameter related to the generation of an orbit transfer strategy of the assembly according to the state of the abnormal separation moment between the assembly and the carrier or between the assemblies;

and B: determining a design constraint condition of a track transfer strategy according to the platform parameters;

and C: according to the constraint conditions, carrying out optimal orbital transfer times estimation on the basis of the principle that the speed increment required by orbital transfer is minimum;

step D: establishing a mathematical model for describing the single orbital transfer state change and the connection among multiple orbital transfers according to the estimation result;

step E: and generating an orbit transfer strategy meeting design constraints by using the mathematical model and taking the GEO orbit as a target through iterative optimization.

2. The method for generating the GEO orbital transfer strategy of the combined body under the abnormal separation condition according to claim 1, wherein the step a comprises:

step S2.1: the combination body comprises a plurality of cabin sections, each cabin section carries a propulsion system, and the abnormal separation comprises a launching section abnormal and a transfer section abnormal, wherein the launching section abnormal and the transfer section abnormal are separated in advance, the combination body cannot be sent to a preset orbit by a carrier, and the transfer section abnormal are separated in advance; firstly, the orbit state at the moment of abnormal separation is determined, the general form of GEO orbit is adopted for description, and the semimajor axis a₀Angle of inclination i₀Eccentricity e₀Geographic longitude λ₀Geographic latitude η₀；

Step S2.2: defining the relevant platform parameters of the combined spacecraft, main satellite or main satellite and propulsion pod, including the separation moment weight m₀M 'of available Fuel residual quantity'₀Thrust F of engine₀Specific impulse Isp of engine₀。

3. The method for generating the GEO-orbital transfer strategy of the assembly under the abnormal separation condition according to claim 2, wherein in the step C, in order to determine the optimal orbital transfer parameter dimension, the optimal orbital transfer times are estimated according to the following steps based on the principle that the increment of the speed required by the transfer is minimum:

step S4.1: and calculating the minimum velocity increment delta v of the spacecraft from the abnormal separation moment state to the GEO target orbit according to the following formula:

Δv＝v_n-v₀

wherein r is₀Designing the earth center distance of the abnormal separation moment of the object for the transfer orbit, and calculating the earth center distance according to the number of the orbits; a is₀Is a semi-major axis of the track; mu is an earth gravity constant; v. of₀Designing the speed of the abnormal separation moment of the object for the transfer orbit; v. of_nIs the target track speed; a is_nIs a target track semi-major axis;

step S4.2: calculating fuel consumption delta m corresponding to the minimum velocity increment of the orbital transfer according to a rocket formula:

wherein g is 9.80665m/s²Is the acceleration of the earth's gravity;

step S4.3: calculating the second flow dm of the propellant according to the specific impulse of the engine, and calculating the total track change time t by combining the fuel consumption:

step S4.4: single maximum ignition time T combined with thruster_maxCalculating to obtain an optimal orbital transfer frequency estimated value N:

wherein [ x ] is an eave function and represents the minimum integer which is more than or equal to x.

4. The method for generating the GEO orbital transfer strategy of the assembly under the abnormal separation condition according to claim 3, wherein the step D comprises:

step S5.1: according to the variable to be optimized for each ignition: semi-major axis a of orbital transfer target_kBefore track change, offset circle Q_kK is 1, …, N, calculating pre-ignition parameters:

track period T_k：

Drift rate of longitude

Wherein, ω is_e＝7.2921×10^-5rad/s, and pi is the circumference ratio.

Ascending node geographic longitude λ_k：

Step S5.2: according to the semi-major axis a of the target at each orbital transfer_kAnd k is 1, …, N, and the single ignition point and the target track parameter are calculated as follows:

calculating the target track velocity v by the activity formula_k：

Wherein r is_kDesigning the ground center distance of the object at the ignition moment for the transfer orbit, and calculating the ground center distance according to the number of the orbits;

orbital transfer velocity increment Δ v_k：

Wherein, α and β_kCalculated according to the following formula

β_k＝π-i_k-1-α

Track-changing back inclination angle i_k：

Orbital transfer fuel consumption Δ m_k：

Ignition duration t_k：

5. The method for generating the GEO orbital transfer strategy of the combined body under the abnormal separation condition according to claim 4, wherein the step E comprises:

step S6.1: converting the GEO orbit transfer strategy solving problem into a multivariable, multi-target and multi-constraint optimizing problem, wherein an optimization model is described as follows:

an objective function:

wherein the content of the first and second substances,representing a track target penalty function;representing the deviation between the jth orbital parameter orbital transfer final value and a target orbit; Δ Orb ═ (Δ a, Δ i, Δ e, Δ λ, Δ η), and represents the deviation of the tracking result from the target track; gamma represents a weight coefficient;

constraint conditions are as follows:

step S6.2: selecting a multivariable, multi-target and multi-constraint optimization algorithm for iterative optimization aiming at the optimization model;

step S6.3: and outputting an optimization result, and determining a GEO orbit transfer strategy of the assembly spacecraft under the abnormal separation condition.

6. An assembly GEO orbital transfer strategy generation system under the condition of abnormal separation is characterized by comprising:

a module A: determining an orbit parameter of a separation point of the assembly and a platform parameter related to the generation of an orbit transfer strategy of the assembly according to the state of the abnormal separation moment between the assembly and the carrier or between the assemblies;

and a module B: determining a design constraint condition of a track transfer strategy according to the platform parameters;

and a module C: according to the constraint conditions, carrying out optimal orbital transfer times estimation on the basis of the principle that the speed increment required by orbital transfer is minimum;

a module D: establishing a mathematical model for describing the single orbital transfer state change and the connection among multiple orbital transfers according to the estimation result;

and a module E: and generating an orbit transfer strategy meeting design constraints by using the mathematical model and taking the GEO orbit as a target through iterative optimization.

7. The system for generating a GEO-orbital transfer strategy for an abnormal separation situation of claim 6, wherein the module a comprises:

module S2.1: the combination body comprises a plurality of cabin sections, each cabin section carries a propulsion system, and the abnormal separation comprises a launching section abnormal and a transfer section abnormal, wherein the launching section abnormal and the transfer section abnormal are separated in advance, the combination body cannot be sent to a preset orbit by a carrier, and the transfer section abnormal are separated in advance; firstly, the orbit state at the moment of abnormal separation is determined, the general form of GEO orbit is adopted for description, and the semimajor axis a₀Angle of inclination i₀Eccentricity e₀Geographic longitude λ₀Geographic latitude η₀；

Module S2.2: defining the relevant platform parameters of the combined spacecraft, main satellite or main satellite and propulsion pod, including the separation moment weight m₀M 'of available Fuel residual quantity'₀Thrust F of engine₀Specific impulse Isp of engine₀。

8. The system for generating the GEO-orbital transfer strategy for the combination under the abnormal separation condition of claim 7, wherein in the module C, in order to determine the optimal orbital transfer parameter dimension, the optimal orbital transfer times estimation is performed according to the following modules on the principle that the increment of the speed required by the transfer is minimum:

module S4.1: and calculating the minimum velocity increment delta v of the spacecraft from the abnormal separation moment state to the GEO target orbit according to the following formula:

Δv＝v_n-v₀

wherein r is₀Designing the earth center distance of the abnormal separation moment of the object for the transfer orbit, and calculating the earth center distance according to the number of the orbits; a is₀Is a semi-major axis of the track; mu is an earth gravity constant; v. of₀Designing the speed of the abnormal separation moment of the object for the transfer orbit; v. of_nIs the target track speed; a is_nIs a target track semi-major axis;

module S4.2: calculating fuel consumption delta m corresponding to the minimum velocity increment of the orbital transfer according to a rocket formula:

wherein g is 9.80665m/s²Is the acceleration of the earth's gravity;

module S4.3: calculating the second flow dm of the propellant according to the specific impulse of the engine, and calculating the total track change time t by combining the fuel consumption:

module S4.4: single maximum ignition time T combined with thruster_maxCalculating to obtain an optimal orbital transfer frequency estimated value N:

wherein [ x ] is an eave function and represents the minimum integer which is more than or equal to x.

9. The system for generating a GEO-orbital transfer strategy for an abnormal separation situation of claim 8, wherein the module D comprises:

module S5.1: according to the variable to be optimized for each ignition: semi-major axis a of orbital transfer target_kBefore track change, offset circle Q_kK is 1, …, N, calculating pre-ignition parameters:

track period T_k：

Drift rate of longitude

Wherein, ω is_e＝7.2921×10^-5rad/s, and pi is the circumference ratio.

Ascending node geographic longitude λ_k：

Module S5.2: according to the semi-major axis a of the target at each orbital transfer_kAnd k is 1, …, N, and the single ignition point and the target track parameter are calculated as follows:

calculating the target track velocity v by the activity formula_k：

Wherein r is_kDesigning the ground center distance of the object at the ignition moment for the transfer orbit, and calculating the ground center distance according to the number of the orbits;

orbital transfer velocity increment Δ v_k：

Wherein, α and β_kCalculated according to the following formula

β_k＝π-i_k-1-α

Track-changing back inclination angle i_k：

Orbital transfer fuel consumption Δ m_k：

Ignition duration t_k：

10. The system for generating a GEO-orbital transfer strategy for an abnormal separation situation of claim 10, wherein the module E comprises:

module S6.1: converting the GEO orbit transfer strategy solving problem into a multivariable, multi-target and multi-constraint optimizing problem, wherein an optimization model is described as follows:

an objective function:

wherein the content of the first and second substances,representing a track target penalty function;representing the deviation between the jth orbital parameter orbital transfer final value and a target orbit; Δ Orb ═ (Δ a, Δ i, Δ e, Δ λ, Δ η), and represents the deviation of the tracking result from the target track; gamma represents a weight coefficient;

constraint conditions are as follows:

module S6.2: selecting a multivariable, multi-target and multi-constraint optimization algorithm for iterative optimization aiming at the optimization model;

module S6.3: and outputting an optimization result, and determining a GEO orbit transfer strategy of the assembly spacecraft under the abnormal separation condition.

Technical Field

The invention relates to astronavigation aircraft orbital dynamics, in particular to a method and a system for generating a GEO orbital transfer strategy of an assembly under the condition of abnormal separation.

Background

Geostationary orbit spacecraft (GEO) has the characteristic of being relatively static to the earth, and is widely applied to the fields of communication, navigation, relay and the like. Compared with the traditional single spacecraft, the combined spacecraft has the advantages that each cabin section carries the propulsion system, so that on one hand, the combined spacecraft has better robustness when dealing with the abnormal conditions of the active section and the multiple orbit transfer sections, and on the other hand, the orbit transfer strategy needs to be designed for the combined body or each cabin section of the combined body rapidly under the condition that the combined body is separated abnormally.

At present, a plurality of transfer orbit strategy design schemes are applied at home and abroad, but the transfer orbit strategy design schemes are mostly complex integral optimization based on limited thrust or a segmented orbit transfer method based on a hybrid propulsion system. In the Chinese patent of "a method for transferring and converting pulse orbital transfer to limited thrust orbital transfer" (patent document: CN112455725A), such as Liujun Yao, Zhao Jian Wei, Zongshi, etc., firstly, the position of a pulse ignition point is determined, then limited thrust integral optimization is carried out at the position, and the ignition direction is corrected. In the Chinese patent 'small geostationary satellite orbit transfer method and system' (patent document: CN111891396A), Linbao army, Jiang Guo Wei, Fanyuan and the like, a method is provided, wherein a chemical propulsion system is firstly utilized to lift the altitude of a spacecraft at the near site, an electric propulsion system is utilized to adjust the altitude, the inclination angle and the eccentricity ratio of the spacecraft at the near site, and finally the chemical propulsion system is utilized to capture the spacecraft at a fixed point; on one hand, the altitude, the inclination angle and the eccentricity of the near place are adjusted in stages, the fuel consumption is large, and the on-orbit service life of the spacecraft is influenced.

In summary, the design optimization of the orbital transfer strategy method needs to be developed according to the requirement for rapidly generating the GEO orbit transfer strategy of the assembly spacecraft under the abnormal separation condition.

Disclosure of Invention

Aiming at the defects in the prior art, the invention aims to provide a method and a system for generating a GEO orbital transfer strategy of an assembly under the condition of abnormal separation.

The invention provides a method for generating a GEO orbital transfer strategy of an assembly under the condition of abnormal separation, which comprises the following steps:

step A: determining an orbit parameter of a separation point of the assembly and a platform parameter related to the generation of an orbit transfer strategy of the assembly according to the state of the abnormal separation moment between the assembly and the carrier or between the assemblies;

and B: determining a design constraint condition of a track transfer strategy according to the platform parameters;

and C: according to the constraint conditions, carrying out optimal orbital transfer times estimation on the basis of the principle that the speed increment required by orbital transfer is minimum;

step D: establishing a mathematical model for describing the single orbital transfer state change and the connection among multiple orbital transfers according to the estimation result;

step E: and generating an orbit transfer strategy meeting design constraints by using the mathematical model and taking the GEO orbit as a target through iterative optimization.

Preferably, the step a includes:

step S2.1: the combination body comprises a plurality of cabin sections, each cabin section carries a propulsion system, and the abnormal separation comprises a launching section abnormal and a transfer section abnormal, wherein the launching section abnormal and the transfer section abnormal are separated in advance, the combination body cannot be sent to a preset orbit by a carrier, and the transfer section abnormal are separated in advance; firstly, the orbit state at the moment of abnormal separation is determined, the general form of GEO orbit is adopted for description, and the semimajor axis a₀Angle of inclination i₀Eccentricity e₀And the groundLongitude λ₀Geographic latitude η₀；

Step S2.2: defining the relevant platform parameters of the combined spacecraft, main satellite or main satellite and propulsion pod, including the separation moment weight m₀M 'of available Fuel residual quantity'₀Thrust F of engine₀Specific impulse Isp of engine₀。

Preferably, in the step B:

three design constraints are summarized by combining actual engineering: and the ground measurement and control conditions restrict the geographical longitude span and the measurement and control duration of the ignition section, the characteristics of the thruster, the loss of the arc section and the safety protection requirements restrict the single longest ignition time and restrict the available fuel of the rail transfer object.

Preferably, in the step C, in order to determine the dimension of the track transfer optimization parameter, the optimal number of times of track transfer needs to be estimated according to the following steps on the principle that the increment of the speed required by the transfer is minimum:

step S4.1: and calculating the minimum velocity increment delta v of the spacecraft from the abnormal separation moment state to the GEO target orbit according to the following formula:

Δv＝v_n-v₀

wherein r is₀Designing the earth center distance of the abnormal separation moment of the object for the transfer orbit, and calculating the earth center distance according to the number of the orbits; a is₀Is a semi-major axis of the track; mu is an earth gravity constant; v. of₀Designing the speed of the abnormal separation moment of the object for the transfer orbit; v. of_nIs the target track speed; a is_nIs a target track semi-major axis;

step S4.2: calculating fuel consumption delta m corresponding to the minimum velocity increment of the orbital transfer according to a rocket formula:

wherein g is 9.80665 m-s²Is the acceleration of the earth's gravity;

step S4.3: calculating the second flow dm of the propellant according to the specific impulse of the engine, and calculating the total track change time t by combining the fuel consumption:

step S4.4: single maximum ignition time T combined with thruster_maxCalculating to obtain an optimal orbital transfer frequency estimated value N:

wherein [ x ] is an eave function and represents the minimum integer which is more than or equal to x.

Preferably, the step D includes:

step S5.1: according to the variable to be optimized for each ignition: semi-major axis a of orbital transfer target_kBefore track change, offset circle Q_kK is 1, …, N, calculating pre-ignition parameters:

track period T_k：

Drift rate of longitude

Wherein, ω is_e＝7.2921×10^-5rad/s, and pi is the circumference ratio.

Point of intersection of the riseGeographic longitude λ_k：

Step S5.2: according to the semi-major axis a of the target at each orbital transfer_kAnd k is 1, …, N, and the single ignition point and the target track parameter are calculated as follows:

calculating the target track velocity v by the activity formula_k：

Wherein r is_kDesigning the ground center distance of the object at the ignition moment for the transfer orbit, and calculating the ground center distance according to the number of the orbits;

orbital transfer velocity increment Δ v_k：

Wherein, α and β_kCalculated according to the following formula

β_k＝π-i_k-1-α

Track-changing back inclination angle i_k：

Orbital transfer fuel consumption Δ m_k：

Ignition duration t_k：

Preferably, the step E includes:

step S6.1: converting the GEO orbit transfer strategy solving problem into a multivariable, multi-target and multi-constraint optimizing problem, wherein an optimization model is described as follows:

an objective function:

wherein the content of the first and second substances,representing a track target penalty function;representing the deviation between the jth orbital parameter orbital transfer final value and a target orbit; Δ Orb ═ (Δ a, Δ i, Δ e, Δ λ, Δ η), and represents the deviation of the tracking result from the target track; gamma represents a weight coefficient;

constraint conditions are as follows:

step S6.2: selecting a multivariable, multi-target and multi-constraint optimization algorithm for iterative optimization aiming at the optimization model;

step S6.3: and outputting an optimization result, and determining a GEO orbit transfer strategy of the assembly spacecraft under the abnormal separation condition.

The invention provides a system for generating a GEO orbital transfer strategy of an assembly under the condition of abnormal separation, which comprises:

a module A: determining an orbit parameter of a separation point of the assembly and a platform parameter related to the generation of an orbit transfer strategy of the assembly according to the state of the abnormal separation moment between the assembly and the carrier or between the assemblies;

and a module B: determining a design constraint condition of a track transfer strategy according to the platform parameters;

and a module C: according to the constraint conditions, carrying out optimal orbital transfer times estimation on the basis of the principle that the speed increment required by orbital transfer is minimum;

a module D: establishing a mathematical model for describing the single orbital transfer state change and the connection among multiple orbital transfers according to the estimation result;

and a module E: and generating an orbit transfer strategy meeting design constraints by using the mathematical model and taking the GEO orbit as a target through iterative optimization.

Preferably, the module a comprises:

module S2.1: the combination body comprises a plurality of cabin sections, each cabin section carries a propulsion system, and the abnormal separation comprises a launching section abnormal and a transfer section abnormal, wherein the launching section abnormal and the transfer section abnormal are separated in advance, the combination body cannot be sent to a preset orbit by a carrier, and the transfer section abnormal are separated in advance; firstly, the orbit state at the moment of abnormal separation is determined, the general form of GEO orbit is adopted for description, and the semimajor axis a₀Angle of inclination i₀Eccentricity e₀Geographic longitude λ₀Geographic latitude η₀；

Module S2.2: defining the relevant platform parameters of the combined spacecraft, main satellite or main satellite and propulsion pod, including the separation moment weight m₀M 'of available Fuel residual quantity'₀Thrust F of engine₀Specific impulse Isp of engine₀。

In the module B:

three design constraints are summarized by combining actual engineering: and the ground measurement and control conditions restrict the geographical longitude span and the measurement and control duration of the ignition section, the characteristics of the thruster, the loss of the arc section and the safety protection requirements restrict the single longest ignition time and restrict the available fuel of the rail transfer object.

Preferably, in the module C, in order to determine the optimal dimension of the track transfer parameter, the optimal number of times of track transfer needs to be estimated according to the following module based on the principle that the increment of the speed required by the transfer is minimum:

module S4.1: and calculating the minimum velocity increment delta v of the spacecraft from the abnormal separation moment state to the GEO target orbit according to the following formula:

Δv＝v_n-v₀

wherein r is₀Designing the earth center distance of the abnormal separation moment of the object for the transfer orbit, and calculating the earth center distance according to the number of the orbits; a is₀Is a semi-major axis of the track; mu is an earth gravity constant; v. of₀Designing the speed of the abnormal separation moment of the object for the transfer orbit; v. of_nIs the target track speed; a is_nIs a target track semi-major axis;

module S4.2: calculating fuel consumption delta m corresponding to the minimum velocity increment of the orbital transfer according to a rocket formula:

wherein g is 9.80665m/s²Is the acceleration of the earth's gravity;

module S4.3: calculating the second flow dm of the propellant according to the specific impulse of the engine, and calculating the total track change time t by combining the fuel consumption:

module S4.4: single maximum ignition time T combined with thruster_maxCalculating to obtain an optimal orbital transfer frequency estimated value N:

wherein [ x ] is an eave function and represents the minimum integer which is more than or equal to x.

Preferably, the module D comprises:

module S5.1: according to the variable to be optimized for each ignition: semi-major axis a of orbital transfer target_kBefore track change, offset circle Q_kK is 1, …, N, calculating pre-ignition parameters:

track period T_k：

Drift rate of longitude

Wherein, ω is_e＝7.2921×10^-5rad/s, and pi is the circumference ratio.

Ascending node geographic longitude λ_k：

Module S5.2: according to the semi-major axis a of the target at each orbital transfer_kAnd k is 1, …, N, and the single ignition point and the target track parameter are calculated as follows:

calculating the target track velocity v by the activity formula_k：

Wherein r is_kDesigning the ground center distance of the object at the ignition moment for the transfer orbit, and calculating the ground center distance according to the number of the orbits;

orbital transfer velocity increment Δ v_k：

Wherein, α and β_kCalculated according to the following formula

β_k＝π-i_k-1-α

Track-changing back inclination angle i_k：

Orbital transfer fuel consumption Δ m_k：

Ignition duration t_k：

Preferably, said module E comprises:

module S6.1: converting the GEO orbit transfer strategy solving problem into a multivariable, multi-target and multi-constraint optimizing problem, wherein an optimization model is described as follows:

an objective function:

wherein the content of the first and second substances,representing a track target penalty function;denotes the jthDeviation between the track parameter orbital transfer final value and a target track; Δ Orb ═ (Δ a, Δ i, Δ e, Δ λ, Δ η), and represents the deviation of the tracking result from the target track; gamma represents a weight coefficient;

constraint conditions are as follows:

module S6.2: selecting a multivariable, multi-target and multi-constraint optimization algorithm for iterative optimization aiming at the optimization model;

module S6.3: and outputting an optimization result, and determining a GEO orbit transfer strategy of the assembly spacecraft under the abnormal separation condition.

Compared with the prior art, the invention has the following beneficial effects:

the method overcomes the defects of large calculated amount, low calculating speed and the like of the traditional limited thrust complex design method under the condition of abnormal separation of the combined spacecraft, and has certain engineering practicability.

Drawings

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

FIG. 1 is a schematic block diagram of a method for rapidly generating an assembly GEO orbital transfer strategy under an abnormal separation condition;

FIG. 2 is a schematic view of a configuration of an assembled spacecraft (two sections);

FIG. 3 is a schematic view of an abnormally separated (out of tolerance) orbit of an assembled spacecraft from a carrier;

FIG. 4 is a diagram of a GEO orbit transfer strategy simulation of an assembly spacecraft based on abnormal separation conditions.

Detailed Description

The present invention will be described in detail with reference to specific examples. The following examples will assist those skilled in the art in further understanding the invention, but are not intended to limit the invention in any way. It should be noted that it would be obvious to those skilled in the art that various changes and modifications can be made without departing from the spirit of the invention. All falling within the scope of the present invention.

As shown in the attached figure 1, the invention provides a method for quickly generating a GEO (geosynchronous orbit) orbit-changing strategy of an assembly under the condition of abnormal separation. The method specifically comprises the following steps:

step A: and determining the track parameters of the separation point of the assembly and the platform parameters related to the generation of the track transfer strategy of the assembly according to the abnormal separation time state between the assembly and the carrier or between the assemblies.

And B: and determining the design constraint conditions of the track transfer strategy according to factors such as ground measurement and control requirements, single ignition capability and the like.

And C: and (4) carrying out optimal orbital transfer times estimation on the basis of the minimum speed increment required by orbital transfer.

Step D: as shown in fig. 4, a mathematical model describing the relationship between the state change of single orbital transfer and the multiple orbital transfers is established.

Step E: and with the GEO orbit as a target, quickly generating an orbit transfer strategy meeting design constraints through iterative optimization.

The step A comprises the following steps:

step S2.1: as shown in fig. 2, a combed spacecraft typically includes a plurality of sections, and each section carries a propulsion system. As shown in fig. 3, the abnormal separation includes an abnormal launching section in which the vehicle fails to send the assembled spacecraft to a predetermined orbit, i.e., the separation in advance, and an abnormal transfer section in which a cabin section in the assembled spacecraft fails to send the assembled spacecraft to a quasi-geosynchronous orbit, i.e., the separation in advance. Without loss of generality, it is described in the general form of a GEO orbit, i.e. semimajor axis a₀Angle of inclination i₀Eccentricity e₀Geographic longitude λ₀Geographic latitude η₀；

Step S2.2: in addition, there is a need to specify relevant platform parameters for the transfer strategy design object (combined spacecraft, main satellite or main satellite and propulsion pod), typically including the separation time moment weight m₀M 'of available Fuel residual quantity'₀Thrust F of engine₀Specific impulse Isp of engine₀；

And step B, summarizing design constraints in 3 aspects by combining actual engineering requirements: the ground measurement and control conditions are used for constraining the geographical longitude span and the measurement and control duration of the ignition section, the characteristics of a thruster, the loss of an arc section and the safety protection requirements on the longest ignition time of a single ignition and the available fuel of a rail transfer object;

in step C, in order to determine the optimal dimension of the track transfer parameter, the optimal number of times of track transfer needs to be estimated according to the following steps and the principle that the increment of the speed required by transfer is minimum:

step S4.1: and calculating the minimum velocity increment delta v of the spacecraft from the abnormal separation moment state to the GEO target orbit according to the following formula:

Δv＝v_n-v₀

wherein r is₀Designing the ground center distance of the abnormal separation moment of the object for the transfer orbit, and calculating the ground center distance according to the number of the orbits; v. of₀Designing the speed of the abnormal separation moment of the object for the transfer orbit; v. of_nIs the target track speed; a is_nThe GEO orbit is 42164km which is the semimajor axis of the target orbit;

step S4.2: calculating fuel consumption delta m corresponding to the minimum velocity increment of the orbital transfer according to a rocket formula:

wherein g is 9.80665m/s²Is the acceleration of the earth's gravity;

step S4.3: calculating the second flow dm of the propellant according to the specific impulse of the engine, and calculating the total track change time t by combining the fuel consumption:

step S4.4: single maximum ignition time T combined with thruster_maxCalculating to obtain an optimal orbital transfer frequency estimated value N:

wherein [ x ] is an eave function and represents the minimum integer which is more than or equal to x.

The step D comprises the following steps:

step S5.1: according to the variable to be optimized for each ignition, i.e. the semi-major axis a of the target of the orbital transfer_kBefore track change, offset circle Q_kK — 1, …, N calculates the pre-ignition parameters:

track period T_k：

Drift rate of longitude

Wherein, ω is_e＝7.2921×10^-5rad/s

Ascending node geographic longitude λ_k：

Step S5.2: according to the semi-major axis a of the target at each orbital transfer_kAnd k is 1, …, N, and the single ignition point and the target track parameter are calculated as follows:

calculating the target orbit according to the activity formulaVelocity v_k：

Wherein r is_kDesigning the ground center distance of the object at the ignition moment for the transfer orbit, and calculating the ground center distance according to the number of the orbits;

orbital transfer velocity increment Δ v_k：

Wherein, α and β_kCalculated according to the following formula

β_k＝π-i_k-1-α

Track-changing back inclination angle i_k：

Orbital transfer fuel consumption Δ m_k：

Ignition duration t_k：

The step E comprises the following steps:

step S6.1: converting the GEO orbit transfer strategy solving problem into a multivariable, multi-target and multi-constraint optimizing problem, wherein an optimization model is described as follows:

an objective function:

wherein the content of the first and second substances,representing a track target penalty function;representing the deviation between the jth orbital parameter orbital transfer final value and a target orbit; Δ Orb ═ (Δ a, Δ i, Δ e, Δ λ, Δ η), and represents the deviation of the tracking result from the target track; gamma represents a weight coefficient, and a larger positive integer is taken for carrying out large-weight punishment on the parameters which do not meet the precision index;

constraint conditions are as follows:

step S6.2: selecting the existing multivariate, multi-target and multi-constraint optimization algorithm for iterative optimization aiming at the optimization model;

step S6.3: and outputting an optimization result, and determining a GEO orbit transfer strategy of the assembly spacecraft under the abnormal separation condition.

In this embodiment, assuming that an anomaly occurs at the time of separation of the vehicle and the combined spacecraft, the orbit inclination angle is deviated at the time of separation due to the fault of the vehicle, and the separation orbit parameter is determined according to step a:

Orb₀＝(a₀,i₀,e₀,λ₀,η₀)＝(24473.64km,30°,0.731215,-9.21°,30°)

determining platform parameters related to the combination body and the track transfer as follows: weight m at the moment of separation₀3863kg of available residual fuel quantity m'₀1820kg, engine thrust F₀350N, engine specific impulse Isp₀＝315s；

Determining track transfer strategy design constraints according to step B: shortest measurement and control duration t_fireMore than or equal to 30min, ground measurement and control geographic longitude range [30 DEG E, 170 DEG E]Maximum time t of single ignition of engine_engine4050s or less, and most fuel is available for rail transfer

According to step C, the GEO target orbit parameter Orb is combined_nThe number of times of tracking was estimated (42164km,30 °,0,60 °,0 °):

and finally, according to the optimization model established in the step E, combining the step D to update parameters of single orbital transfer state change and multiple orbital transfer, and generating a track transfer strategy which meets design constraints through iterative optimization, wherein the track transfer strategy is shown in the table 1.

TABLE 1 iterative optimization solution for orbital transfer strategy

The invention also provides a system for generating the GEO orbital transfer strategy of the combination under the condition of abnormal separation, which comprises the following steps:

a module A: and determining the track parameters of the separation point of the assembly and the platform parameters related to the generation of the track transfer strategy of the assembly according to the abnormal separation time state between the assembly and the carrier or between the assemblies.

And a module B: and determining the design constraint conditions of the track transfer strategy according to the platform parameters.

And a module C: and according to the constraint conditions, carrying out optimal orbital transfer times estimation on the basis of the principle that the speed increment required by orbital transfer is minimum.

A module D: and establishing a mathematical model for describing the single orbital transfer state change and the connection among multiple orbital transfers according to the estimation result.

And a module E: and generating an orbit transfer strategy meeting design constraints by using the mathematical model and taking the GEO orbit as a target through iterative optimization.

Those skilled in the art will appreciate that, in addition to implementing the system and its various devices, modules, units provided by the present invention as pure computer readable program code, the system and its various devices, modules, units provided by the present invention can be fully implemented by logically programming method steps in the form of logic gates, switches, application specific integrated circuits, programmable logic controllers, embedded microcontrollers and the like. Therefore, the system and various devices, modules and units thereof provided by the invention can be regarded as a hardware component, and the devices, modules and units included in the system for realizing various functions can also be regarded as structures in the hardware component; means, modules, units for performing the various functions may also be regarded as structures within both software modules and hardware components for performing the method.

The foregoing description of specific embodiments of the present invention has been presented. It is to be understood that the present invention is not limited to the specific embodiments described above, and that various changes or modifications may be made by one skilled in the art within the scope of the appended claims without departing from the spirit of the invention. The embodiments and features of the embodiments of the present application may be combined with each other arbitrarily without conflict.