Design method of launch window for low-inclination GEO satellite with the most economical propellant throughout its life cycle

Abstract

The present invention discloses a method for designing a launch window for a low-inclination GEO satellite with the least propellant in the whole life cycle. The method comprises the following steps: obtaining input conditions; obtaining the right ascension Ω0 of the ascending node of the transfer orbit corresponding to the launch time t; obtaining the total mass of the orbit transfer propellant and the north-south orbit maintaining propellant; judging whether the total mass of the orbit transfer propellant and the north-south orbit maintaining propellant obtained in step 3 reaches the minimum, and if so, recording the corresponding total amount of propellant and the corresponding launch time t; judging the time t, if t=T f , select the launch time corresponding to the minimum total mass of the orbit transfer propellant and the north-south orbit maintenance propellant as the window front; and obtain the window rear edge according to the preset width of the launch window. The present invention can effectively improve the satellite carrying capacity and extend the service life of the satellite.

Description

Technical Field

[0001] The invention belongs to the technical field of aerospace vehicle design, and in particular relates to a method for designing a launch window for a small-angle GEO satellite with the most life-cycle propellant savings. Background Art

[0002] GEO satellites are usually sent into large elliptical orbits by carriers, and then enter the target working orbit through the satellite's own orbit change. In the process from the separation of the rocket and the satellite entering the working orbit, the satellite is subject to the limitations of various subsystems and requires ground measurement and control to track the satellite in certain specific orbital segments. At the same time, the satellite, the sun and the earth must meet certain geometric relationships, which limits the time when the satellite can be launched. The purpose of the launch window calculation is to select and find these time sets available for launch. Only when the satellite is launched during this period of time can the various conditions in the satellite transfer orbit operation meet the satellite requirements. It can be seen that the launch window design is a prerequisite for the satellite to smoothly enter the working orbit after separation from the rocket.

[0003] In 2007, my country issued Q / W 324A-2007 "Spacecraft Launch Window Design Guidelines". This standard provides general guidelines for spacecraft launch window design and is only of guiding significance for GEO satellite launch window design.

[0004] In recent years, with the deepening of the application of geostationary satellites in the fields of communications, meteorology, navigation, etc. and the continuous emergence of new tasks and new demands, the requirements for satellites are getting higher and higher. This is mainly reflected in the complex and diversified functions of satellites and the long life of satellites. This requires satellites to reduce the demand for propellants as much as possible. However, the existing methods are mainly to optimize and determine the orbit transfer propellant first, and then consider other constraints to design the launch window. In this case, the consumption of propellant during the life of the satellite is a local optimal solution. Usually it is much greater than the global optimal solution. This reduces the carrying capacity of the satellite and limits the service life of the satellite. Summary of the invention

[0005] The technical problem solved by the present invention is: to overcome the shortcomings of the prior art and provide a method for designing a launch window for a low-inclination GEO satellite with the most economical propellant throughout its life cycle, taking the launchable time that meets the constraints as input and consuming the least propellant during the life cycle as the goal, and adopting an exhaustive method to carry out launch window design, which can effectively improve the satellite's carrying capacity and extend the satellite's service life.

[0006] The object of the present invention is achieved by the following technical solution: a method for designing a launch window for a low-inclination geostationary orbit satellite with the least propellant in the entire life cycle, the method comprising the following steps: step 1: obtaining input conditions, wherein the input conditions include: a satellite launch start time T0 and an end time T1 that meet the constraint conditions; f; Step 2: In the time interval [T0,T f ], starting from the launch time t=T0, the right ascension Ω0 of the ascending node of the transfer orbit corresponding to the launch time t is obtained; Step 3: According to the right ascension Ω0 of the ascending node of the transfer orbit corresponding to the launch time t obtained in step 2, a series of different initial right ascensions Ω of the ascending node of the working orbit are selected i , simulate and obtain the total mass of multiple orbit transfer propellants and north-south orbit maintenance propellants; Step 4: select the minimum value of the total mass of multiple orbit transfer propellants and north-south orbit maintenance propellants and the corresponding launch time t; Step 5: judge the time t in step 2, if t = T f , select the launch time corresponding to the minimum value of multiple orbit transfer propellant masses and the total mass of north-south orbit maintained propellant as the window front; Step six: obtain the window back edge according to the preset width of the launch window.

[0007] In the above-mentioned method for designing a launch window for a low-inclination GEO satellite that saves the most propellant throughout its life cycle, in step three, the total mass of the orbit transfer propellant and the north-south orbit maintaining propellant is the mass of the orbit transfer propellant plus the mass of the north-south orbit maintaining propellant.

[0008] In the above-mentioned method for designing a launch window for a low-inclination GEO satellite with the most economical propellant throughout its life cycle, the mass of the orbit transfer propellant is obtained by the following method: the mass of the orbit change propellant is obtained according to the mass of the satellite before the orbit change, the specific impulse of the orbit change engine, the gravitational acceleration and the orbit change velocity increment; the mass of the orbit circularization propellant is obtained according to the specific impulse of the orbit circularization thruster, the efficiency of the orbit circularization thruster and the orbit circularization velocity increment; the mass of the orbit transfer propellant is obtained according to the mass of the orbit change propellant and the mass of the orbit circularization propellant.

[0009] In the above-mentioned method for designing the launch window of a low-inclination GEO satellite with the least propellant throughout its life, the formula for the mass of the propellant for orbit change is as follows: Where dm1 is the mass of the propellant for orbit change; m0 is the mass of the satellite before orbit change; I sp is the specific impulse of the orbital maneuvering engine; g is the acceleration due to gravity; Δv i is the track change speed increment.

[0010] In the above-mentioned method for designing the launch window of a low-inclination GEO satellite with the least propellant throughout its life, the formula for the mass of the orbit circularization propellant is as follows:

[0011]

[0012] Among them, I sp0 is the specific impulse of the orbit circularization thruster; η0 is the efficiency of the orbit circularization thruster; Δv e is the orbit circularization velocity increment; m0 is the satellite mass before orbit change; dm1 is the orbit change propellant mass; dm2 is the orbit circularization propellant mass.

[0013] In the above-mentioned method for designing the launch window of a low-inclination GEO satellite with the least propellant throughout its life, the formula for the mass of the orbit transfer propellant is as follows:

[0014] m1=dm1+dm2;

[0015] Among them, m1 is the mass of orbit transfer propellant, dm1 is the mass of orbit change propellant, and dm2 is the mass of orbit circularization propellant.

[0016] In the above-mentioned low-inclination GEO satellite launch window design method with the most economical full-life propellant, the mass of the north-south orbit maintaining propellant is obtained by the following method: the mass of the north-south orbit maintaining propellant is obtained according to the specific impulse of the north-south orbit maintaining thruster, the efficiency of the north-south orbit maintaining thruster, the north-south orbit maintaining velocity increment, the gravitational acceleration, the satellite mass before the orbit change and the mass of the orbit transfer propellant.

[0017] In the above-mentioned method for designing the launch window of a low-inclination GEO satellite with the least propellant throughout its life, the formula for maintaining the propellant mass in the north-south orbit is as follows:

[0018]

[0019] Among them, I sp1 is the north-south orbit holding thruster specific impulse; η is the north-south orbit holding thruster efficiency; Δv ns is the velocity increment for maintaining the north-south orbit; m1 is the mass of the orbit transfer propellant; m0 is the mass of the satellite before the orbit change; g is the gravitational acceleration; m2 is the mass of the north-south orbit maintaining propellant.

[0020] Compared with the prior art, the present invention has the following beneficial effects:

[0021] (1) The present invention aims to save the most propellant during the entire life cycle and adopts an exhaustive method to design the window. The design result of the propellant consumption during the life cycle is the global optimal solution. Compared with the local optimal solution of the prior art, it can reduce the amount of propellant carried and improve the satellite carrying capacity.

[0022] (2) The present invention aims to save the most propellant during the entire life cycle and adopts an exhaustive method to design the window. The design result of the propellant consumption during the life cycle is the global optimal solution. Compared with the local optimal solution of the prior art, it can reduce the propellant consumption and extend the service life of the satellite.

[0023] (3) The present invention uses the launchable time that satisfies the constraint conditions as input to perform window design, taking into account the satellite launch window constraint conditions, and the design result meets the actual requirements of the project. BRIEF DESCRIPTION OF THE DRAWINGS

[0024] Various other advantages and benefits will become apparent to those of ordinary skill in the art by reading the detailed description of the preferred embodiments below. The accompanying drawings are only for the purpose of illustrating the preferred embodiments and are not to be considered as limiting the present invention. Moreover, the same reference symbols are used throughout the accompanying drawings to represent the same components. In the accompanying drawings:

[0025] Figure 1 This is a flow chart of a method for designing a launch window for a low-inclination geostationary orbit satellite with the most economical full-life propellant provided by an embodiment of the present invention. DETAILED DESCRIPTION

[0026] Exemplary embodiments of the present disclosure will be described in more detail below with reference to the accompanying drawings. Although exemplary embodiments of the present disclosure are shown in the accompanying drawings, it should be understood that the present disclosure can be implemented in various forms and should not be limited by the embodiments set forth herein. On the contrary, these embodiments are provided in order to enable a more thorough understanding of the present disclosure and to be able to fully convey the scope of the present disclosure to those skilled in the art. It should be noted that, in the absence of conflict, the embodiments of the present invention and the features described in the embodiments can be combined with each other. The present invention will be described in detail below with reference to the accompanying drawings and in combination with the embodiments.

[0027] Figure 1 This is a flow chart of a method for designing a launch window for a low-inclination geostationary orbit satellite that uses the least amount of propellant throughout its life cycle, provided by an embodiment of the present invention. Figure 1 As shown, the method comprises the following steps:

[0028] Step 1: Obtain input conditions, which include: satellite launch start time T0 and end time T1 that meet the constraint conditions. f ;

[0029] Step 2: Based on T0 and T obtained in step 1 f , in the time interval [T0,T f ], starting from the launchable time t = T0, determine the right ascension Ω0 of the ascending node of the transfer orbit corresponding to the launch time t;

[0030] Step 3: Based on the Ω0 obtained in step 2, select a series of different working orbit initial ascending node right ascension Ω i , multiple orbit transfer propellants and the total amount of north-south orbit maintenance propellants are obtained through simulation. The specific process is as follows:

[0031] The orbit transfer propellant includes two parts: orbit change propellant and orbit circularization propellant. The specific calculation process is as follows:

[0032] Calculation of trajectory change propellant: Where: m0 is the mass of the satellite before the orbit change; I sp is the specific impulse of the orbital maneuvering engine; g is the acceleration due to gravity; Δvi is the track change speed increment.

[0033] Orbit circularization propellant calculation: Where: I sp0 is the specific impulse of the orbit circularization thruster; η0 is the efficiency of the orbit circularization thruster; Δv e is the orbit circularization velocity increment.

[0034] Orbital transfer propellant calculation: m1=dm1+dm2.

[0035] North-South orbit maintenance propellant calculation: Where: I sp1 is the north-south orbit holding thruster specific impulse; η is the north-south orbit holding thruster efficiency; Δv ns is the north-south track velocity increment.

[0036] The total amount of orbit transfer propellant and north-south orbit maintaining propellant: dm=m1+m2.

[0037] Step 4: Based on the total mass of the orbit transfer propellant and the north-south orbit maintaining propellant obtained in step 3, by comparing the series of propellant total masses obtained in step 3, select the minimum value in the series of propellant total masses, and record the corresponding total propellant amount and the corresponding launchable time t;

[0038] Step 5: Determine the time t in step 2. If t = T f , select the launch time corresponding to the minimum total amount of orbit transfer propellant and north-south orbit maintenance propellant from the results recorded in step 4 as the window front. If t <T f , get the next launch time and jump to step 2;

[0039] Step 6: According to the emission window width requirement, determine the window trailing edge and obtain the emission window.

[0040] Specifically, the first step is to obtain input conditions, which include: satellite design life Tlife, initial mass m0, initial inclination of working orbit is, orbit change engine specific impulse Isp, north-south thruster specific impulse Isp0, satellite launch start time T0 and end time T0 that meet the constraints. f ;

[0041] The second step is to find the emission time interval [T0,T f ], starting from the launchable time t = T0, determine the right ascension Ω0 of the ascending node of the transfer orbit corresponding to the launch time t;

[0042] The third step is to select the right ascension Ω of the initial ascending node of the working orbit i , the total amount of orbit transfer propellant and north-south orbit maintenance propellant is simulated and calculated. The specific process is as follows:

[0043] Calculation of track change speed increment: Where: v1, v2 are the speeds before and after the track change.

[0044] β is the angle between the velocity vectors before and after the track change, calculated as follows:

[0045] cosβ=cos i0 cos i s +sin i0 sin i s cos(Ω i -Ω0);

[0046] Where: i0, Ω0 are the transfer orbit inclination and the right ascension of the ascending node respectively; Ω i It is the initial inclination of the working orbit and the right ascension of the ascending node.

[0047] Calculation of trajectory change propellant:

[0048] Where: m0 is the mass of the satellite before the orbit change; I sp is the specific impulse of the trajectory change engine; g is the acceleration due to gravity.

[0049] Calculation of orbit circularization velocity increment: Δv e =1537.3e; where: e is the orbital eccentricity after the orbit change.

[0050] Orbit circularization propellant calculation: Where: I sp0 is the orbit circularization thruster specific impulse; η0 is the orbit circularization thruster efficiency.

[0051] Orbital transfer propellant calculation: m1=dm1+dm2.

[0052] Calculation of the north-south track speed increment:

[0053] Δv ns =53.66Δi c ;

[0054] Δi c =(T life -T d )D i ;

[0055] Where: T life is the satellite lifespan, in years; T d D is the time (in years) that the inclination vector drifts freely in the inclination holding loop from the satellite fixed point. i is the annual drift rate of the inclination vector (unit: ° / year).

[0056] North-South orbit maintenance propellant calculation:

[0057]

[0058] Where: I sp1 is the north-south orbit holding thruster specific impulse; η is the north-south orbit holding thruster efficiency.

[0059] Total amount of orbit transfer propellant and north-south orbit maintenance propellant: dm = m1 + m2

[0060] Step 4: If the orbit transfer propellant and the north-south orbit keep the total amount of propellant as minimum, record the corresponding total amount of propellant and the corresponding launch time t, otherwise jump to step 3;

[0061] Step 5 If t = T f , select the launch time corresponding to the minimum total amount of orbit transfer propellant and north-south orbit maintenance propellant from the fourth step record results as the window front. If t <T f , take the next launch time and jump to the second step;

[0062] The sixth step is to determine the window trailing edge according to the emission window width requirement to obtain the emission window.

[0063] The present invention aims to minimize propellant consumption throughout the life cycle and adopts an exhaustive method to perform window design. The design result of propellant consumption during the life cycle is a global optimal solution. Compared with the local optimal solution results of the prior art, the propellant carrying amount can be reduced and the satellite carrying capacity can be improved. The present invention aims to minimize propellant consumption throughout the life cycle and adopts an exhaustive method to perform window design. The design result of propellant consumption during the life cycle is a global optimal solution. Compared with the local optimal solution results of the prior art, the propellant consumption can be reduced and the service life of the satellite can be extended. The present invention uses the launchable time that meets the constraints as input for window design, takes into account the satellite launch window constraints, and the design result meets the actual requirements of the engineering.

[0064] Although the present invention has been disclosed as above in the form of a preferred embodiment, it is not intended to limit the present invention. Any person skilled in the art may make possible changes and modifications to the technical solution of the present invention by using the methods and technical contents disclosed above without departing from the spirit and scope of the present invention. Therefore, any simple modifications, equivalent changes and modifications made to the above embodiments based on the technical essence of the present invention without departing from the content of the technical solution of the present invention shall fall within the protection scope of the technical solution of the present invention.

Claims

1. A method for designing a launch window for a low-angle GEO satellite with the least propellant throughout its life cycle, characterized in that: The method comprises the following steps: Step 1: Obtain input conditions, wherein the input conditions include: the satellite launch start time that meets the constraint conditions and end time ; Step 2: In the time interval From the launch time Start, get the launch time Corresponding transfer orbit ascending node right ascension ; Step 3: Based on the launch time obtained in step 2 Corresponding transfer orbit ascending node right ascension , select a series of different working orbit initial ascending node right ascension , the total mass of multiple orbit transfer propellants and north-south orbit maintenance propellants is obtained through simulation; Step 4: Select the minimum value of multiple orbit transfer propellant masses and the total mass of north-south orbit maintenance propellant and the corresponding launch time ; Step 5: Time in step 2 Make a judgment, if , select the launch time corresponding to the minimum value of multiple orbit transfer propellant masses and the total mass of north-south orbit maintenance propellant as the window front; Step 6: Obtain the window rear edge according to the preset width of the transmission window; In step 3, the total mass of the orbit transfer propellant and the north-south orbit maintaining propellant is the mass of the orbit transfer propellant plus the mass of the north-south orbit maintaining propellant; The mass of orbital transfer propellant is obtained by the following method: The mass of the orbit change propellant is obtained according to the mass of the satellite before the orbit change, the specific impulse of the orbit change engine, the gravitational acceleration and the orbit change speed increment; The orbit circularization propellant mass is obtained according to the orbit circularization thruster specific impulse, orbit circularization thruster efficiency and orbit circularization velocity increment; The mass of orbit transfer propellant is obtained according to the mass of orbit change propellant and the mass of orbit circularization propellant; The formula for the mass of the trajectory change propellant is as follows: ； in, is the mass of the trajectory change propellant; is the satellite mass before orbit change; It is the specific impulse of the trajectory change engine; is the acceleration due to gravity; is the track change speed increment; The formula for orbital circularization propellant mass is as follows: ； in, is the orbit circularization thruster specific impulse; is the orbit circularization thruster efficiency; is the orbit circularization velocity increment; is the satellite mass before orbit change; is the mass of the trajectory change propellant; is the orbit circularization propellant mass.

2. The method for designing a launch window for a low-angle GEO satellite with the least propellant throughout the life cycle according to claim 1, characterized in that: The formula for orbital transfer propellant mass is as follows: ； in, is the orbit transfer propellant mass, is the mass of the trajectory change propellant, is the orbit circularization propellant mass.

3. The method for designing a launch window for a low-angle GEO satellite with the most economical full-life propellant according to claim 1, characterized in that: The north-south orbit-maintaining propellant is obtained by the following method: The mass of the north-south orbit maintaining propellant is obtained according to the specific impulse of the north-south orbit maintaining thruster, the efficiency of the north-south orbit maintaining thruster, the north-south orbit maintaining velocity increment, the gravitational acceleration, the satellite mass before the orbit change and the orbit transfer propellant.

4. The method for designing a launch window for a low-angle GEO satellite with the least propellant throughout the life cycle according to claim 3, characterized in that: The formula for maintaining the propellant mass in the north-south orbit is as follows: ； in, is the north-south orbit-keeping thruster specific impulse; The north-south track maintains thruster efficiency; is the north-south track maintaining velocity increment; transfer of propellant mass to orbit; is the satellite mass before orbit change; is the acceleration due to gravity; Maintain propellant mass for north-south orbits.

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