Assembleable satellite separation method and system based on optimal separation orbit model

Abstract

The present invention provides a method and system for separating an assembleable satellite based on an optimal separation orbit model, comprising the following steps: Step 1: establishing an optimal separation orbit model; Step 2: performing on-orbit separation of an assembleable satellite according to the optimal separation orbit model. The present invention can design the separation orbit of an assembleable satellite, meet the on-orbit separation requirements of an assembleable satellite, take into account the safety of close-range separation and the global optimum of long-distance transfer, and improve the mission maneuverability of the assembleable satellite.

Description

Technical Field

[0001] The present invention relates to the technical field of satellite separation orbits, and in particular, to a method and system for separating an assembleable satellite based on an optimal separation orbit model. In particular, it preferably relates to an optimal separation orbit design method for an assembleable satellite. Background Art

[0002] In view of the future application needs of space-based information ports, ultra-large remote sensing, and large-capacity communications, there is an urgent need to develop ultra-large satellites. Due to the constraints of the launch envelope, ultra-large satellites cannot be directly launched into orbit by a single launch vehicle, and usually require multiple launches and in-orbit assembly. Assembled satellites are composed of several modules, among which the core module is usually called the main satellite, and other functional modules are usually called sub-satellites.

[0003] In order to achieve satellite module reorganization and effectively avoid space threats, each sub-star module of the assembleable satellite must have the ability to separate in orbit: when the satellite module is reassembled or faces a threat, the docking mechanism between the main satellite and the sub-satellite is unlocked, and the impact generated by the unlocking process will provide the initial separation speed for each sub-star module; under the action of the orbital control engine, each sub-star module reaches the predetermined target orbit respectively.

[0004] Compared with the traditional orbit transfer problem, the separation orbit design problem for assembleable satellites has the following main differences:

[0005] 1. The separation orbit for assembleable satellites must first ensure the safety of the separation process, especially in the close-range separation stage. There are few studies on how to avoid collision between the satellite and the main satellite and prevent the satellite from causing plume interference to the main satellite. "Small Satellite On-orbit Safe Separation Speed ​​Design" (China Space Science and Technology, June 2007) proposed how to design the initial separation speed to keep the relative distance between the satellite and the main satellite not less than a fixed value after a period of separation; but in practical applications, the feasible area of ​​ignition maneuvers can be constrained based on the relative motion law between the satellite and the main satellite, thereby reducing the requirements for the initial separation speed.

[0006] 2. The Lambert problem is the basic problem of dual-pulse orbit transfer. It usually takes the position vector and flight time of two points on a given orbit to find the orbital parameters connecting the two points. The essence is to solve the two-point boundary value problem of differential equations. In order to extend the on-orbit life and reduce consumption, the on-orbit separation of the assembler satellite module can take fuel optimization or energy optimization as the optimization goal without limiting the transfer time. "Research on Limited Thrust Fuel-Efficient Coplanar Circular Orbit Transfer" (Computer Simulation, October 2013) only studied the coplanar circular orbit transfer problem and only considered the fuel-efficient transfer strategy; "Energy Optimal and Fuel Optimal Lambert Rendezvous Problem" (Journal of Beijing University of Aeronautics and Astronautics, September 2018) proposed a general framework for solving the energy-optimal and fuel-optimal Lambert problems based on vector form, but it requires the position vector of the satellite before and after the orbit change to be given in advance, and it is impossible to optimize the selection of transfer points and solve the global optimal transfer orbit. Summary of the invention

[0007] In view of the defects in the prior art, the object of the present invention is to provide a method and system for separating assembleable satellites based on an optimal separation orbit model.

[0008] According to the present invention, a method for separating an assembleable satellite based on an optimal separation orbit model comprises the following steps:

[0009] Step 1: Establish the optimal separation orbit model;

[0010] Step 2: Perform in-orbit separation of the assembleable satellite according to the optimal separation orbit model.

[0011] Preferably, the assembleable satellite includes a main satellite and a sub-satellite, and the separation of the assembleable satellite includes close-range separation; and the step 1 includes the following steps:

[0012] Step 1.1: Solve the relative motion between the satellite and the primary satellite during the close separation process to obtain a free separation orbit;

[0013] Step 1.2: Calculate the orbital elements of the sub-star in the free separation orbit;

[0014] Step 1.3: Calculate the free separation feasible area according to the orbital elements;

[0015] Step 1.4: Select separation strategy;

[0016] Step 1.5: Solve the optimal separation transfer trajectory according to the separation strategy and the free separation feasible region.

[0017] Preferably, the separation of the assembleable satellite further comprises long-distance transfer;

[0018] The close separation stage: after the satellite separates from the unlocking point, the orbit control engine is in a shutdown state, and it only relies on the initial separation speed to deviate from the initial orbit and generate relative motion with the main satellite until the satellite drifts to the starting point;

[0019] The long-distance transfer stage: the sub-satellite starts the orbit control engine at the starting point and performs a pulse maneuver, goes through the separation transfer arc and reaches the end point, starts the orbit control engine and performs a pulse maneuver to enter the target orbit.

[0020] Preferably, the step 1.1 further comprises: when the sub-satellite orbit control engine is not turned on, the relative motion with the main satellite is generated only by the initial separation speed during the close separation period.

[0021] Preferably, the step 1.2 further comprises: the orbital elements of the free separation orbit of the sub-star to be solved include the specific angular momentum modulus h free , orbital inclination i free , eccentricity e free , right ascension of ascending node Ω free 、Arg of perigee ω free and true close angle f free .

[0022] Preferably, the step 1.3 includes: in order to prevent the sub-satellite from colliding with the main satellite, according to the set safe separation distance z safe , constrain the true anomaly f1 corresponding to the starting point on the free separation orbit:

[0023] f1∈[min(f free )1f free )2),max(f free )1f free )2)]

[0024] Among them, f free )1 is the true anomaly angle f free The feasible lower limit of free )2 is the true anomaly angle f free The feasible upper limit f free )2;f free )1.f free )2 two components and θ free The form is as follows:

[0025]

[0026] Where:

[0027]

[0028]

[0029] Preferably, step 1.4 includes: selecting an energy-optimal or fuel-optimal separation strategy according to mission requirements.

[0030] Preferably, step 1.5 includes: making an initial guess for the optimal track change point, then solving the control equation, calculating the objective function, and optimizing the track change point until the cutoff condition is met, thereby searching for the optimal track change point.

[0031] According to the present invention, an assembleable satellite separation system based on an optimal separation orbit model includes the following modules:

[0032] Module M1: Establishing the optimal separation track model;

[0033] Module M2: performing on-orbit separation of the assembleable satellite according to the optimal separation orbit model.

[0034] Preferably, the module M1 includes the following modules:

[0035] Module M1.1: Solve the relative motion between the satellite and the primary satellite during the close separation process to obtain the free separation orbit;

[0036] Module M1.2: Calculate the orbital elements of the satellite in the free separation orbit;

[0037] Module M1.3: Calculate the free separation feasible area based on the orbital elements;

[0038] Module M1.4: Selecting a separation strategy;

[0039] Module M1.5: Solve the optimal separation transfer trajectory based on the separation strategy and the free separation feasible area.

[0040] Compared with the prior art, the present invention has the following beneficial effects:

[0041] 1. The present invention can design the separation orbit of the assembleable satellite, meet the requirements of on-orbit separation of the assembleable satellite, take into account the safety of close-range separation and the global optimization of long-distance transfer, improve the mission maneuverability of the assembleable satellite, and provide a basis for the design of the on-orbit separation orbit of the assembleable satellite and the formulation of the control strategy;

[0042] 2. The present invention meets the working requirements of on-orbit separation of sub-star modules of assembleable satellites, ensures the safety and reliability of the separation process, is conducive to protecting the safety of high-value space assets, and is conducive to improving the evasion capability of assembleable satellites against space threats;

[0043] 3. The present invention designs the objective functions under the energy optimal transfer strategy and the fuel optimal transfer strategy respectively, which can be selected according to actual application requirements. BRIEF DESCRIPTION OF THE DRAWINGS

[0044] Other features, objects and advantages of the present invention will become more apparent from the detailed description of non-limiting embodiments made with reference to the following drawings:

[0045] Figure 1 A schematic diagram of the separation process of the assembleable satellite provided by the present invention;

[0046] Figure 2 A flow chart of the optimal separation orbit design method for an assembleable satellite provided by the present invention;

[0047] Figure 3 The optimal separation transfer trajectory solution flow chart provided by the present invention;

[0048] Figure 4 This is a schematic diagram of an assembleable satellite used in the present invention.

[0049] Figure numerals: 101, unlocking point; 102, initial orbit; 103, free separation orbit; 104, starting point; 105, end point; 106, separation transfer arc; 107, target orbit; 201, main satellite; 202, first sub-satellite; 203, second sub-satellite; 204, third sub-satellite; 205, solar sail panel; 206, deployable truss. DETAILED DESCRIPTION

[0050] The present invention is described in detail below in conjunction with specific embodiments. The following embodiments will help those skilled in the art to further understand the present invention, but are not intended to limit the present invention in any form. It should be noted that, for those of ordinary skill in the art, several changes and improvements can also be made without departing from the concept of the present invention. These all belong to the protection scope of the present invention.

[0051] The embodiment of the present invention discloses a method for separating assembleable satellites based on an optimal separation orbit model, such as Figure 1 As shown, the unlocking point is the point at which the docking mechanism of the main satellite and the sub-satellite module is unlocked. Initial orbit is the orbit of the main satellite and the sub-satellite before unlocking. Free separation orbit is the orbit of the sub-satellite without starting the orbit control engine after unlocking. Starting point is the starting point of the sub-satellite for the first pulse through the orbit control engine. End point is the starting point of the sub-satellite for the second pulse through the orbit control engine. Separation transfer orbit is the transfer orbit of the sub-satellite from the free separation orbit to the target orbit under the action of double pulses. Target orbit is the predetermined orbit that the sub-satellite needs to reach after the pulse maneuver.

[0052] like Figure 1 and Figure 2As shown in the figure, the main satellite and the sub-star module are unlocked at the unlocking point, and the impact generated will provide the sub-star module with an initial separation speed δv0. Through the design of the unlocking device for the assembleable satellite, it can be ensured that when the initial separation speed δv0 is provided to the sub-satellite module, the sub-satellite module can be separated at a given close distance separation time t free There will be no collision between the sub-star and the main star.

[0053] The method includes the following steps: Step 1: Establish an optimal separation orbit model. The assembleable satellite includes a main satellite and a sub-satellite, and the separation of the assembleable satellite includes close-range separation and long-range transfer. Close-range separation stage: After the sub-satellite separates from the unlocking point, the orbit control engine is in a shutdown state, and only relies on the initial separation speed to deviate from the initial orbit and produce relative motion with the main satellite until the sub-satellite drifts to the starting point. Long-range transfer stage: The sub-satellite starts the orbit control engine at the starting point and performs a pulse maneuver, goes through the separation transfer arc to reach the end point, starts the orbit control engine, performs a pulse maneuver, and enters the target orbit.

[0054] Step 1 includes the following steps: Step 1.1: Solve the relative motion between the satellite and the main satellite during the close separation process to obtain the free separation orbit. When the satellite orbit control engine is not turned on, the satellite will be separated from the main satellite during the close separation time t free Within t, the relative motion with the main satellite is generated only by the initial separation velocity. Solve the relative motion during the close separation process. The satellite orbit control engine is not turned on. free In the time period, the relative motion between the satellite and the main star is generated only by the initial separation velocity. In the orbital coordinate system of the main star, the relative motion between the satellite and the main star during the close separation process is solved:

[0055]

[0056]

[0057]

[0058] Where: δr is the relative displacement, which can be expressed in the host star orbit coordinate system as [δx δy δz] T , where δx, δy, and δz are the components of δr in the host orbit coordinate system, and the superscript T indicates transposition; are the first-order derivatives of the corresponding components with respect to time, are the second-order derivatives of the corresponding components with respect to time; the angular velocity of the primary star r p The orbital radius of the main star, μ = 398600.5 km 3 / s 2 ; f is the orbit control thrust, which can be expressed in the host satellite orbit coordinate system as [f x f yf z ] T , where f x 、f y 、f z are the components of f in the primary star orbit coordinate system.

[0059] Step 1.2: Calculate the orbital elements of the sub-star in the free separation orbit. Calculate the orbital elements of the sub-star in the free separation orbit. Solve the orbital elements of the sub-star in the free separation orbit [h free i free e free Ω free ω free θ free ] T The orbital elements of the free separation orbit of the sub-star to be solved include the specific angular momentum modulus h free , orbital inclination i free , eccentricity e free , right ascension of ascending node Ω free 、Arg of perigee ω free and true close angle f free .

[0060] Step 1.3: Calculate the free separation feasible area according to the orbital elements. In order to prevent the satellite from colliding with the primary satellite, according to the set safe separation distance z safe , constrain the true anomaly angle f1 corresponding to the starting point on the free separation orbit. Calculation of the feasible area of ​​free separation. According to step 1.1, the relative distance between the satellite and the main satellite in the z-axis direction of the main satellite orbit coordinate system is a periodically changing trigonometric function. Therefore, it is necessary to pre-set the safe separation distance z in the z-axis direction of the main satellite orbit coordinate system. safe , constrain the true anomaly f1 corresponding to the starting point on the free separation orbit:

[0061] f1∈[min(f free )1f free )2),max(f free )1f free )2)]

[0062] Among them, f free )1 is the true anomaly angle f free The feasible lower limit of free )2 is the true anomaly angle f free The feasible upper limit f free )2. In order to facilitate trigonometric function calculations, f is introduced free )1.f free )2, respectively and θ free , so f free )1.ffree )2 can be expressed in the following component form:

[0063]

[0064] Where:

[0065]

[0066]

[0067] Step 1.4: Select a separation strategy. Common separation strategies include energy-optimal and fuel-optimal. Select the energy-optimal or fuel-optimal separation strategy based on mission requirements.

[0068] Step 1.5: Solve the optimal separation transfer orbit according to the separation strategy and the free separation feasible region. Make an initial guess for the optimal change point, then solve the control equation, calculate the objective function, and optimize the change point until the cutoff condition is met, thereby searching for the optimal change point. Solve the optimal separation transfer orbit. Figure 3 As shown, step 1.5 includes the following steps:

[0069] Initial state step: The initial guess of the velocity inclination angles of the starting point and the end point is θ1 and θ2.

[0070] Step 501: Solve the control equation.

[0071] For the optimal energy transfer strategy, the governing equation is as follows:

[0072] 3C 4 +c4Y 3 -c5Y-2c2=0

[0073] For the optimal fuel transfer strategy, the governing equation is as follows:

[0074] c 12 +c 13 Y+c 14 Y 2 +c 15 Y 3 +c 16 Y 4 +c 17 Y 5 +c 18 Y 6 +c 19 Y 7 +c 20 Y 8 =0

[0075] For the above two types of control equations, the variable Y of the control equation is a function of the eccentricity e of the separation transfer orbit and the true anomaly f2 of the terminal point:

[0076]

[0077] Where: e is the eccentricity of the separation transfer orbit; f2 is the true anomaly angle of the terminal point.

[0078] For the above two types of control equations, the correlation coefficient c of the control equation 2～20 The form is as follows:

[0079]

[0080]

[0081]

[0082]

[0083]

[0084]

[0085]

[0086]

[0087]

[0088]

[0089]

[0090]

[0091]

[0092]

[0093]

[0094]

[0095]

[0096]

[0097]

[0098] Coefficient c 2～20The variables involved include: m=r2 / r1, r1 is the modulus of the position vector of the starting point, r2 is the modulus of the position vector of the end point; Δf is the phase difference between the true anomaly angle of the starting point and the end point; v1 is the velocity of the starting point before the orbit change; v4 is the velocity of the end point after the orbit change; ξ1 is the orbital angle between the free separation orbit of the sub-satellite and the separation transfer orbit; ξ2 is the angle between the separation transfer orbital plane and the target orbital plane; Θ1 is the velocity inclination angle of the starting point; Θ2 is the velocity inclination angle of the end point.

[0099] Step 502: Calculate the objective function.

[0100] For the optimal energy transfer strategy, the objective function Y best The form is as follows:

[0101]

[0102] The coefficient c1 in the objective function is as follows:

[0103]

[0104] For the optimal fuel transfer strategy, the objective function Y best The form is as follows:

[0105]

[0106] According to the energy optimal or fuel optimal transfer strategy, the corresponding objective function is selected; after substituting the roots of the control equation obtained in step 501 into the objective function one by one, the minimum objective function Y is obtained. best .

[0107] Step 503: Optimize the track change point. Based on the particle swarm algorithm, according to Y best According to the change of the velocity inclination angles θ1 and θ2 at the starting point and the end point, the starting point is within the free separation feasible area solved in step 1.3.

[0108] Step 504: Determine the cut-off condition. Determine whether to terminate the optimization of the track change point based on whether the cut-off condition such as the iteration number is met.

[0109] According to the optimal transfer plan obtained in step 1.5, the sub-satellite performs pulse maneuvers at the starting point and the end point to reach the target orbit.

[0110] Step 2: Perform in-orbit separation of the assembleable satellite according to the optimal separation orbit model.

[0111] The present invention adopts a satellite that can be assembled as Figure 4 As shown, the assembleable satellite includes a solar sail panel, a deployable truss, a main satellite and a sub-satellite. The sub-satellites include a first sub-satellite, a second sub-satellite and a third sub-satellite.

[0112] For the first satellite, the optimal separation orbit is designed by the present invention. For example, the main satellite is normally operated in the GEO orbit, the relative speed of the satellite separation is 0.3m / s along the separation direction, the thrust of the satellite orbit control engine is 490N, and the target of the satellite is transferred to the GEO+300km graveyard orbit.

[0113] Execute step 1.1. When the orbit control engine is not turned on, the sub-satellites will separate freely within 100 seconds without collision.

[0114] Execute step 1.2, the number of orbital elements of the sub-star free separation is:

[0115]

[0116] Execute steps 1.3, 1.4 and 1.5. If the energy optimal strategy is adopted, the separation track design is as follows:

[0117]

[0118] If the fuel optimal strategy is adopted, the separation orbit is designed as follows:

[0119]

[0120] Those skilled in the art know that, in addition to realizing the system and its various devices, modules, and units provided by the present invention in a purely computer-readable program code, it is entirely possible to realize the same functions in the form of logic gates, switches, application-specific integrated circuits, programmable logic controllers, and embedded microcontrollers by logically programming the method steps. Therefore, the system and its various devices, modules, and units provided by the present invention can be considered as a hardware component, and the devices, modules, and units included therein for realizing various functions can also be regarded as structures within the hardware component; the devices, modules, and units for realizing various functions can also be regarded as both software modules for realizing the method and structures within the hardware component.

[0121] The above describes the specific embodiments of the present invention. It should be understood that the present invention is not limited to the above specific embodiments, and those skilled in the art can make various changes or modifications within the scope of the claims, which does not affect the essence of the present invention. In the absence of conflict, the embodiments of the present application and the features in the embodiments can be combined with each other arbitrarily.

Claims

1. A method for separating assembleable satellites based on an optimal separation orbit model, characterized in that: The steps include: Step 1: Establish the optimal separation orbit model; Step 2: performing on-orbit separation of the assembleable satellite according to the optimal separation orbit model; The assembleable satellite includes a main satellite and a sub-satellite, and the separation of the assembleable satellite includes close-range separation; the step 1 includes the following steps: Step 1.1: Solve the relative motion between the satellite and the primary satellite during the close separation process to obtain a free separation orbit; Step 1.2: Calculate the orbital elements of the sub-star in the free separation orbit; Step 1.3: Calculate the free separation feasible area according to the orbital elements; Step 1.4: Select separation strategy; Step 1.5: Solve the optimal separation transfer trajectory according to the separation strategy and the free separation feasible region.

2. The method for separating assembleable satellites based on the optimal separation orbit model according to claim 1, characterized in that: The separation of the assembleable satellite also includes long-distance transfer; The close separation stage: after the satellite separates from the unlocking point, the orbit control engine is in a shutdown state, and it only relies on the initial separation speed to deviate from the initial orbit and generate relative motion with the main satellite until the satellite drifts to the starting point; The long-distance transfer stage: the sub-satellite starts the orbit control engine at the starting point and performs a pulse maneuver, goes through the separation transfer arc and reaches the end point, starts the orbit control engine and performs a pulse maneuver to enter the target orbit.

3. The method for separating assembleable satellites based on the optimal separation orbit model according to claim 2, characterized in that: The step 1.1 also includes: when the sub-satellite orbit control engine is not turned on, the relative motion with the main satellite is generated only by the initial separation speed during the close separation period.

4. The method for separating assembleable satellites based on the optimal separation orbit model according to claim 2, characterized in that: The step 1.2 also includes: the orbital elements of the free separation orbit of the sub-star to be solved include the specific angular momentum modulus h free , orbital inclination i free , eccentricity e free , right ascension of ascending node Ω free 、Arg of perigee ω free and true close angle f free .

5. The method for separating assembleable satellites based on the optimal separation orbit model according to claim 2, characterized in that: The step 1.3 includes: in order to prevent the sub-satellite from colliding with the main satellite, according to the set safe separation distance z safe , constrain the true anomaly f1 corresponding to the starting point on the free separation orbit: f1∈[min(f free )1 f free )2),max(f free )1 f free )2)] Among them, f free )1 is the true anomaly angle f free The feasible lower limit of free )2 is the true anomaly angle f free The feasible upper limit f free )2; f free )1.f free )2 two components and θ free The form is as follows: Where:

6. The method for separating assembleable satellites based on the optimal separation orbit model according to claim 1, characterized in that: The step 1.4 includes: selecting an energy-optimal or fuel-optimal separation strategy according to mission requirements.

7. The method for separating assembleable satellites based on the optimal separation orbit model according to claim 1, characterized in that: The step 1.5 includes: making an initial guess for the optimal track change point, then solving the control equation, calculating the objective function, and optimizing the track change point until the cutoff condition is met, thereby searching for the optimal track change point.

8. An assembleable satellite separation system based on an optimal separation orbit model, characterized in that: A method for separating assembleable satellites based on an optimal separation orbit model as described in any one of claims 1 to 7 is applied, comprising the following modules: Module M1: Establishing the optimal separation track model; Module M2: performing on-orbit separation of the assembleable satellite according to the optimal separation orbit model.

9. The assembleable satellite separation system based on the optimal separation orbit model according to claim 8, characterized in that: The module M1 includes the following modules: Module M1.1: Solve the relative motion of the satellite and the primary star during the close separation process to obtain the free separation orbit; Module M1.2: Calculate the orbital elements of the satellite in the free separation orbit; Module M1.3: Calculate the free separation feasible area based on the orbital elements; Module M1.4: Selecting a separation strategy; Module M1.5: Solve the optimal separation transfer trajectory based on the separation strategy and the free separation feasible area.

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