A virtual orbit construction method for satellite emergency orbit change

Abstract

The invention discloses a virtual orbit construction method for satellite emergency orbit change, comprising: step S1: determining the inclination and ascending node right ascension of the virtual orbit according to a specified ignition direction; step S2: determining the perigee argument of the virtual orbit according to the virtual geocentric direction and the ignition direction; step S3: selecting the semi-major axis of the virtual orbit according to a predetermined ignition duration, on-board software constraints and quasi-inertial requirements, and calculating the mean anomaly of the virtual orbit before the orbit change. The invention ensures the quasi-inertial ignition attitude of the orbit system of the virtual orbit, selects the ignition arc segment as the apogee of the virtual orbit, constructs a semi-major axis as large as possible within the range of the semi-major axis and eccentricity allowed by the on-board software, and finally constructs the virtual orbit parameters at the specified time. The invention has the advantage that the ignition in any inertial direction can be realized by a satellite orbit holding mode program-controlled orbit control thruster and the constraints of the body axis and the geocentric direction are met.

Description

Technical Field

[0001] The present invention relates to the technical field of spacecraft orbit control, and in particular to a virtual orbit construction method for satellite emergency orbit change. Background Art

[0002] For high-orbit satellites or large elliptical satellites, due to the low perigee and inclination constraints of the launch vehicle's rocket-satellite separation, it is usually necessary to ignite the orbit control thruster to change the orbit to reach the final orbit. If before reaching the target orbit, due to large deviations from the initial launch orbit or abnormalities in the propulsion system or battery problems in the energy system, an emergency orbit change is required. In order to cope with various orbit change requirements, the velocity increment direction corresponding to the orbit change is variable, and the thruster implementing the orbit change may be a large thrust main engine or a small thrust orbit control engine. The small thrust engine usually jets the satellite program in the orbit control mode, but the satellite attitude reference in the orbit holding mode is usually taken as the orbit system, the satellite attitude is non-inertially stable, and the ignition direction changes. Therefore, a set of virtual orbits is constructed according to the ignition requirements and the constraints of each axis, and the emergency orbit change is performed using the orbit system attitude reference, which can ensure that the ignition direction of the orbit change is stable in the inertial space, and the orbit change attitude is quasi-inertially maintained, which has strong engineering applicability. Summary of the invention

[0003] The purpose of the present invention is to provide a virtual orbit construction method for satellite emergency orbit change. This method aims to solve the problem of how to use the orbit system attitude reference to perform emergency orbit change when constructing a set of virtual orbits in the prior art, ensure that the orbit change ignition direction is stable in the inertial space, and the orbit change attitude is quasi-inertially maintained.

[0004] To achieve the above object, the present invention provides a virtual orbit construction method for satellite emergency orbit change, comprising:

[0005] Step S1: Determine the inclination angle and right ascension of the ascending node of the virtual orbit according to the specified ignition direction;

[0006] Step S2: determining the perigee argument of the virtual orbit based on the virtual geocentric direction and the ignition direction;

[0007] Step S3: According to the predetermined ignition duration, onboard software constraints and quasi-inertial requirements, the semi-major axis of the virtual orbit is selected to calculate the mean anomaly angle of the virtual orbit before the orbit change. Preferably, in step S1, the inclination of the virtual orbit is determined according to the specified ignition direction, and the calculation includes:

[0008] The ignition direction is specified as V x , then the expression of the inclination angle of the virtual track is:

[0009] i=acos(V x (3)).

[0010] Preferably, in the step S1, the right ascension Ω of the ascending node of the virtual orbit is determined according to the specified ignition direction, which is specifically:

[0011] Ω=atan2(V x (1),-V x (2)).

[0012] Preferably, in step S2, the perigee argument of the virtual orbit is determined by the virtual geocentric direction and the ignition direction, and the specific steps are:

[0013] Step S2.1: Calculating the node azimuth vector of the virtual orbit according to the ascending node equator of the virtual orbit;

[0014] Step S2.2: Calculating the dimension argument of the virtual orbit according to the node azimuth vector of the virtual orbit and the virtual geocentric direction;

[0015] Step S2.3: Calculate the argument of perigee of the virtual orbit according to the argument of the dimension of the virtual orbit.

[0016] Preferably, in step S2.1, the node azimuth vector V of the virtual orbit is calculated according to the ascending node equator of the virtual orbit. n , whose expression is:

[0017] V n =[cosΩ-sinΩcosi sinΩsini] T .

[0018] Preferably, in step S2.2, the dimension argument u of the virtual orbit is calculated according to the node orientation vector of the virtual orbit and the virtual geocentric direction, and its expression is:

[0019]

[0020] Among them, V z Indicates the direction of the virtual center of the earth.

[0021] Preferably, in step S2.3, the argument of perigee ω of the virtual orbit is calculated according to the argument of latitude u of the virtual orbit, and its expression is:

[0022] ω=u-π.

[0023] Preferably, in step S3, according to the predetermined ignition duration, onboard software constraints and quasi-inertial requirements, the semi-major axis of the virtual orbit is selected, and the mean anomaly angle M of the virtual orbit before the orbit change is calculated, and its expression is:

[0024] T=6500+2ΔT

[0025]

[0026] Wherein: ΔT is the ignition duration, μ is the gravitational constant, a is the semi-major axis of the virtual orbit, and T is the time from perigee to apogee.

[0027] Compared with the prior art, the present invention has the following beneficial effects:

[0028] The present invention ensures that the orbital system ignition attitude of the virtual orbit is quasi-inertial, selects the ignition arc segment as the apogee of the virtual orbit, and constructs a large semi-major axis a as much as possible within the range of the semi-major axis and eccentricity allowed by the on-board software, and the corresponding apogee instantaneous angular velocity is small, and the attitude reference change during ignition is small. The virtual orbit inclination and ascending node right ascension are constructed from the specified inertial ignition direction, the virtual orbit perigee argument is constructed from the virtual geocentric direction, and the virtual orbit mean anomaly angle is determined according to the time before ignition, and finally the virtual orbit parameters at the specified time are constructed, and the ignition in any inertial direction is realized through the satellite orbit holding mode program-controlled orbit control thruster and the constraints of the body axis and the geocentric direction are met. BRIEF DESCRIPTION OF THE DRAWINGS

[0029] In order to more clearly illustrate the technical solution of the present invention, the following briefly introduces the drawings required for the description. Obviously, the drawings described below are an embodiment of the present invention. For those skilled in the art, other drawings can be obtained based on these drawings without creative work:

[0030] Figure 1 A schematic flow chart of a method for constructing a virtual orbit for emergency satellite orbit change according to an embodiment of the present invention. DETAILED DESCRIPTION

[0031] The following is combined with Figure 1 The virtual orbit construction method for emergency satellite orbit change proposed in the present invention is further described in detail with specific implementation methods. According to the following description, the advantages and features of the present invention will be clearer. It should be noted that the drawings are in a very simplified form and use non-precise proportions, which are only used to conveniently and clearly assist in explaining the purpose of the implementation methods of the present invention. In order to make the purposes, features and advantages of the present invention more obvious and easy to understand, please refer to the drawings. It should be noted that the structures, proportions, sizes, etc. illustrated in the drawings of this specification are only used to match the contents disclosed in the specification for people familiar with this technology to understand and read, and are not used to limit the limiting conditions for the implementation of the present invention, so they have no technical substantive significance. Any structural modification, change in proportional relationship or adjustment of size, without affecting the effects that can be produced by the present invention and the purposes that can be achieved, should still fall within the scope of the technical content disclosed by the present invention.

[0032] In order to construct a set of virtual orbits, an orbit system attitude reference is used to perform emergency orbit change, ensuring that the orbit change ignition direction is stable in the inertial space and the orbit change attitude is quasi-inertially maintained. This embodiment provides a virtual orbit construction method for satellite emergency orbit change, including:

[0033] Step S1: Determine the inclination angle and right ascension of the ascending node of the virtual orbit according to the specified ignition direction;

[0034] According to the specified ignition direction, the inclination of the virtual track is determined, and the calculation includes:

[0035] The ignition direction is specified as V x , then the expression of the inclination angle of the virtual track is:

[0036] i=acos(V x (3)) (1)

[0037] According to the specified ignition direction, the right ascension Ω of the ascending node of the virtual orbit is determined, which is specifically:

[0038] Ω=atan2(V x (1),-V x (twenty two)

[0039] Step S2: determining the perigee argument of the virtual orbit based on the virtual geocentric direction and the ignition direction;

[0040] The specific steps of determining the perigee argument of the virtual orbit from the virtual geocentric direction and the ignition direction are as follows:

[0041] Step S2.1: Calculating the node azimuth vector of the virtual orbit according to the ascending node equator of the virtual orbit;

[0042] Step S2.2: Calculating the dimension argument of the virtual orbit according to the node azimuth vector of the virtual orbit and the virtual geocentric direction;

[0043] Step S2.3: Calculate the argument of perigee of the virtual orbit according to the argument of the dimension of the virtual orbit.

[0044] According to the ascending node equator of the virtual orbit, the node azimuth vector V of the virtual orbit is calculated. n , whose expression is:

[0045] V n =[cosΩ -sinΩcosi sinΩsini] T (3)

[0046] According to the node azimuth vector of the virtual orbit and the virtual geocentric direction, the dimension argument u of the virtual orbit is calculated, and its expression is:

[0047]

[0048] Among them, V z Indicates the direction of the virtual center of the earth.

[0049] According to the dimensional argument u of the virtual orbit, the perigee argument ω of the virtual orbit is calculated, and its expression is:

[0050] ω=u-π (5)

[0051] Step S3: According to the predetermined ignition duration, onboard software constraints and quasi-inertial requirements, the semi-major axis of the virtual orbit is selected, and the mean anomaly angle of the virtual orbit before the orbit change is calculated.

[0052] According to the predetermined ignition duration, onboard software constraints and quasi-inertial requirements, the semi-major axis of the virtual orbit is selected, and the mean anomaly angle M of the virtual orbit before the orbit change is calculated, and its expression is:

[0053] T=6500+2ΔT(6)

[0054]

[0055] Wherein: ΔT is the ignition duration, μ is the gravitational constant, a is the semi-major axis of the virtual orbit, and T is the time from perigee to apogee.

[0056] Assuming that the ignition duration is 5400s, the virtual orbit mean anomaly angle 1 hour before ignition is constructed as:

[0057] T = 6500 + 2 * 5400,

[0058] This embodiment ensures that the orbital system ignition attitude of the virtual orbit is quasi-inertial, selects the ignition arc segment as the apogee of the virtual orbit, and constructs as large a semi-major axis a as possible within the range of the semi-major axis and eccentricity allowed by the on-board software. The corresponding apogee instantaneous angular velocity is small, and the attitude reference change during ignition is small. The virtual orbit inclination and ascending node right ascension are constructed by the specified inertial ignition direction, the virtual orbit perigee argument is constructed by the virtual geocentric direction, and then the virtual orbit mean anomaly is determined according to the time before ignition, and finally the virtual orbit parameters at the specified time are constructed. The advantage is that the ignition in any inertial direction can be realized by the satellite orbit holding mode programmable orbit control thruster and the constraints of the main body axis and the geocentric direction. This embodiment constructs the inclination and ascending node right ascension of the virtual orbit according to the specified inertial ignition direction, constructs the perigee argument of the virtual orbit by the constrained inertial axis, considers the semi-major axis and eccentricity allowed by the on-board software, constructs the largest possible virtual orbit semi-major axis, and then calculates the virtual orbit mean anomaly before the implementation of the orbit change. The virtual track construction method has the advantages of simple calculation and strong versatility, and can be applied to various rescue emergency track change situations.

[0059] It should be noted that, in this article, relational terms such as first and second, etc. are only used to distinguish one entity or operation from another entity or operation, and do not necessarily require or imply any such actual relationship or order between these entities or operations. Moreover, the terms "include", "comprise" or any other variants thereof are intended to cover non-exclusive inclusion, so that a process, method, article or device including a series of elements includes not only those elements, but also other elements not explicitly listed, or also includes elements inherent to such process, method, article or device. In the absence of further restrictions, the elements defined by the sentence "comprise a ..." do not exclude the existence of other identical elements in the process, method, article or device including the elements.

[0060] It should be noted that the devices and methods disclosed in the embodiments of this article can also be implemented in other ways. The device implementation described above is only schematic. For example, the flowcharts and block diagrams in the accompanying drawings show the possible architecture, functions and operations of the devices, methods and computer program products according to multiple embodiments of this article. In this regard, each box in the flowchart or block diagram can represent a part of a module, program or code, and the module, program segment or part of the code contains one or more executable instructions for implementing the specified logical function, and the module, program segment or part of the code contains one or more executable instructions for implementing the specified logical function. It should also be noted that in some alternative implementations, the functions marked in the box can also occur in a different order from the order marked in the accompanying drawings. For example, two consecutive boxes can actually be executed substantially in parallel, and they can sometimes be executed in the opposite order, depending on the functions involved. It should also be noted that each box in the block diagram and / or flowchart, and the combination of boxes in the block diagram and / or flowchart can be implemented by a dedicated hardware-based system for performing a specified function or action, or can be implemented by a combination of dedicated hardware and computer instructions. Although the content of the present invention has been described in detail through the above preferred embodiments, it should be appreciated that the above description should not be considered as a limitation of the present invention. After reading the above content, it will be apparent to those skilled in the art that various modifications and substitutions of the present invention will occur. Therefore, the protection scope of the present invention should be limited by the appended claims.

Claims

1. A method for constructing a virtual orbit for satellite emergency orbit change, characterized in that: include: Step S1: Determine the inclination angle and right ascension of the ascending node of the virtual orbit according to the specified ignition direction. Specifically, according to the specified ignition direction, the inclination of the virtual track is determined, and the calculation includes: the specified ignition direction is V x , then the expression of the inclination angle of the virtual track is: i=acos(V x (3)) According to the specified ignition direction, the right ascension Ω of the ascending node of the virtual orbit is determined, which is specifically: Ω=atan2(V x (1),-V x (2)); Step S2: Determine the perigee argument of the virtual orbit based on the virtual geocentric direction and the ignition direction, and the specific steps are as follows: Step S2.1: Calculate the node azimuth vector V of the virtual orbit according to the ascending node equator of the virtual orbit n , whose expression is: In n =[cosΩ -sinΩcosi sinΩsini] T ; Step S2.2: Calculate the dimension argument u of the virtual orbit according to the node orientation vector of the virtual orbit and the virtual geocentric direction, and its expression is: Among them, V z Indicates the virtual geocentric direction; Step S2.3: Calculate the argument of perigee of the virtual orbit according to the dimensional argument of the virtual orbit; Step S3: According to the predetermined ignition duration, onboard software constraints and quasi-inertial requirements, the semi-major axis of the virtual orbit is selected, and the mean anomaly angle of the virtual orbit before the orbit change is calculated.

2. The method for constructing a virtual orbit for satellite emergency orbit change according to claim 1, characterized in that: In step S2.3, the argument of perigee ω of the virtual orbit is calculated according to the argument of latitude u of the virtual orbit, and its expression is: ω=u-π.

3. The method for constructing a virtual orbit for satellite emergency orbit change according to claim 2, characterized in that: In step S3, according to the predetermined ignition duration, onboard software constraints and quasi-inertial requirements, the semi-major axis of the virtual orbit is selected, and the mean anomaly angle M of the virtual orbit before the orbit change is calculated, and its expression is: T=6500+2ΔT Wherein: ΔT is the ignition duration, μ is the gravitational constant, a is the semi-major axis of the virtual orbit, and T is the time from perigee to apogee.

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