A method and device for discharging residual propellant of a reentry vehicle

Abstract

The embodiments of the present invention relate to the field of aircraft technology, and in particular to a method and device for discharging the remaining propellant of a reentry vehicle. The method includes determining the relationship between the amount of propellant discharged and the total orbital control velocity increment; determining the additional change in the geographic longitude of the ascending node of the aircraft according to the ground trajectory adjustment requirements of the sub-satellite point of the aircraft; establishing a relationship expression between the additional change in geographic longitude, the number of orbits between different pulse accelerations of the aircraft, and the orbital control velocity increment of different pulse accelerations; iteratively solving the relationship expression with the number of two pulse accelerations and the interval of one orbit between different pulse accelerations as conditions, obtaining the orbital control velocity increment of each pulse acceleration and the specific number of orbits of each pulse, and judging whether each orbital control velocity increment meets the verification conditions. The method for discharging the remaining propellant of a reentry vehicle provided by the embodiment of the present invention can not only meet the fuel discharge requirements of the aircraft, but also complete the propellant discharge method in the orbital plane for the sub-satellite point position adjustment requirements.

Description

Technical Field

[0001] The embodiments of the present invention relate to the field of aircraft technology, and in particular to a method and device for discharging residual propellant of a reentry vehicle. Background Art

[0002] For spacecraft that need to re-enter the atmosphere from orbit, their flight in the atmosphere is an unpowered gliding flight. The lift and drag coefficients determined by the shape of the spacecraft and the mass of the spacecraft jointly determine the flight envelope and range capability of the spacecraft. Therefore, for a spacecraft that uses an integrated design to travel between the earth and the sky, the mass of the remaining propellant needs to be controlled within a certain accuracy range before leaving orbit. Due to the uncertainty of the duration and mission of the spacecraft's on-orbit cruise, the mass of the remaining propellant before completing the on-orbit mission and preparing to leave orbit and return is also uncertain, so the demand for propellant emissions should be determined according to actual conditions.

[0003] To ensure the efficiency of propellant discharge, the propellant consumption is usually achieved by turning on the orbit control engine, and the orbit control engine startup process will inevitably affect the spacecraft orbit. If you want to have a smaller impact on the orbit, you can adjust the inclination and perform a mutually offset orbit control at the ascending and descending nodes to complete the discharge of the remaining propellant. However, for reentry requirements, it is often accompanied by adjustments to the subsatellite point position.

[0004] Therefore, in response to the above problems, there is an urgent need for an orbital propellant discharge method that can meet the fuel discharge requirements of the aircraft and complete the position adjustment requirements of the sub-satellite point. Summary of the invention

[0005] The embodiments of the present invention provide a method, device, electronic device and storage medium for discharging the remaining propellant of a reentry vehicle, which can not only meet the fuel discharge requirements of the vehicle, but also complete the propellant discharge method in the orbital plane to meet the requirements of adjusting the sub-satellite point position.

[0006] In a first aspect, an embodiment of the present invention provides a method for discharging residual propellant of a reentry vehicle, comprising:

[0007] Determine the relationship between propellant discharge and total orbital velocity increment;

[0008] Determine the additional change in geographic longitude of the spacecraft's ascending node based on ground track adjustment requirements for the spacecraft's sub-satellite point;

[0009] Establishing a relationship expression between the additional change in geographic longitude, the number of orbits between different pulse accelerations of the aircraft, and the orbit control velocity increments of different pulse accelerations;

[0010] Taking the number of two pulse accelerations and the interval of one orbital revolution between different pulse accelerations as conditions, the relational expression is iteratively solved to obtain the orbital control speed increment of each pulse acceleration and the specific orbital revolution of each pulse, and it is judged whether each orbital control speed increment meets the verification conditions. If so, the propellant is discharged according to the current orbital control speed increment, the number of pulse accelerations and the orbital revolution between different pulse accelerations; if not, the number of pulse accelerations and the orbital revolution between different pulse accelerations are increased, the solution is iteratively solved again, and the verification is verified again using the verification conditions until the verification conditions are met.

[0011] In a possible design, the verification condition includes that the sum of all the track control speed increments is 0;

[0012] The sum of the absolute values ​​of the track control speed increments is equal to the total track control speed increment.

[0013] In a possible design, the verification condition includes that the sum of all the track control speed increments is 0;

[0014] The difference between the sum of the absolute values ​​of the track control speed increments and the total track control speed increment is less than a preset error amount.

[0015] In a possible design, the determining of the additional change in the geographic longitude of the ascending node of the aircraft according to the ground track adjustment requirement of the sub-satellite point of the aircraft includes:

[0016] According to the nodal period of orbital motion and the precession rate of the ascending node right ascension, an expression for the change of geographic longitude is established;

[0017] According to the first-order partial derivative, the precession rate of the geographical longitude of the ascending node in one orbital period is solved to obtain the change of the semi-major axis;

[0018] The semi-major axis change is used to express additional changes in geographic longitude.

[0019] In a possible design, the relationship expression between the additional change in geographic longitude, the number of orbits between different pulse accelerations of the spacecraft, and the orbital control speed increments of different pulse accelerations includes:

[0020] expressing the semi-major axis change using an orbit control velocity increment;

[0021] Using the orbit control speed increment to express the additional change in geographic longitude, a speed and longitude change expression is obtained;

[0022] The relationship expression is determined according to the velocity longitude variation expression and the number of revolutions between different pulse accelerations of the aircraft.

[0023] In a possible design, the velocity longitude change expression is:

[0024]

[0025] Among them, λ N is the change in geographical longitude for a certain circle, Δλ N is the additional change in geographic longitude for a certain orbit, a0 is the semi-major axis of the nominal orbit, ΔV t is the orbit control velocity increment.

[0026] In a possible design, the relational expression is as follows:

[0027]

[0028] Among them, λ N is the change in geographical longitude for a certain circle, Δλ N is the additional change in geographic longitude for a particular orbit, M t is the number of pulse executions, N is the number of pulse executions, ΔV t is the orbit control velocity increment.

[0029] In a second aspect, an embodiment of the present invention further provides a device for discharging excess propellant of a reentry vehicle, comprising:

[0030] A first calculation unit determines the relationship between the propellant discharge amount and the total orbit control velocity increment;

[0031] a second calculation unit, determining an additional change in the geographic longitude of the ascending node of the aircraft based on a ground track adjustment requirement of the sub-satellite point of the aircraft;

[0032] A third calculation unit is used to establish a relationship expression between the additional change in geographic longitude, the number of orbits between different pulse accelerations of the aircraft, and the orbit control speed increments of different pulse accelerations;

[0033] The judgment unit, taking the number of pulse accelerations between two times and the interval of one revolution between different pulse accelerations as conditions, iteratively solves the relational expression to obtain the orbital control speed increment of each pulse acceleration and the specific number of revolutions of each pulse. The judgment unit, judges whether each orbital control speed increment meets the verification condition. If so, the propellant is discharged according to the current orbital control speed increment, the number of pulse accelerations and the number of revolutions between different pulse accelerations; if not, the number of pulse accelerations and the number of revolutions between different pulse accelerations are increased, the solution is iterated again, and the verification condition is used again for verification until the verification condition is met.

[0034] In a third aspect, an embodiment of the present invention further provides an electronic device, including a memory and a processor, wherein the memory stores a computer program, and when the processor executes the computer program, the method described in any embodiment of this specification is implemented.

[0035] In a fourth aspect, an embodiment of the present invention further provides a computer-readable storage medium having a computer program stored thereon, which, when executed in a computer, enables the computer to execute the method described in any embodiment of this specification.

[0036] The embodiments of the present invention provide a method, device, electronic device and storage medium for discharging the remaining propellant of a reentry vehicle, and the remaining propellant of a reentry vehicle is discharged. The solution provided by the present invention can not only meet the fuel discharge requirements of the aircraft, but also complete the propellant discharge method in the orbital plane to meet the position adjustment requirements of the sub-satellite point. The method first obtains the orbit control speed increment requirement based on the propellant discharge amount requirement, and then based on the east-west adjustment requirements of the geographical longitude of the sub-satellite point, provides an orbit control pulse solution method with the least number of pulses and the fastest completion of the task. BRIEF DESCRIPTION OF THE DRAWINGS

[0037] In order to more clearly illustrate the embodiments of the present invention or the technical solutions in the prior art, the drawings required for use in the embodiments or the description of the prior art will be briefly introduced below. Obviously, the drawings described below are some embodiments of the present invention. For ordinary technicians in this field, other drawings can be obtained based on these drawings without paying creative work.

[0038] Figure 1 This is a flow chart of a method for discharging excess propellant from a reentry vehicle provided by one embodiment of the present invention;

[0039] Figure 2 is a hardware architecture diagram of an electronic device provided by an embodiment of the present invention;

[0040] Figure 3 The present invention is a structural diagram of a reentry vehicle surplus propellant discharge device provided in one embodiment of the present invention. DETAILED DESCRIPTION

[0041] In order to make the purpose, technical solutions and advantages of the embodiments of the present invention clearer, the technical solutions in the embodiments of the present invention will be clearly and completely described below in conjunction with the drawings in the embodiments of the present invention. Obviously, the described embodiments are part of the embodiments of the present invention, rather than all the embodiments. Based on the embodiments in the present invention, all other embodiments obtained by ordinary technicians in this field without making creative work are within the scope of protection of the present invention.

[0042] The specific implementation of the above concept is described below.

[0043] Please refer to Figure 1 The embodiment of the present invention provides a method for discharging remaining propellant of a reentry vehicle, the method comprising:

[0044] Step 100, determining the relationship between the propellant discharge amount and the total orbital control velocity increment;

[0045] Step 102, determining an additional change in the geographic longitude of the ascending node of the aircraft according to the ground track adjustment requirement of the aircraft sub-satellite point;

[0046] Step 104, establishing a relationship expression between the additional change in geographic longitude, the number of orbits between different pulse accelerations of the aircraft, and the orbit control velocity increment of different pulse accelerations;

[0047] Step 106, with the number of pulse accelerations and the interval of one revolution between different pulse accelerations as conditions, the relational expression is iteratively solved to obtain the orbital control speed increment of each pulse acceleration and the specific number of revolutions of each pulse, and it is determined whether each orbital control speed increment meets the verification conditions. If so, the propellant is discharged according to the current orbital control speed increment, the number of pulse accelerations and the number of revolutions between different pulse accelerations; if not, the number of pulse accelerations and the number of revolutions between different pulse accelerations are increased, the solution is iteratively solved again, and the verification is performed again using the verification conditions until the verification conditions are met.

[0048] The solution provided by the present invention can not only meet the fuel discharge requirements of the aircraft, but also complete the propellant discharge method in the orbital plane to meet the requirements of adjusting the position of the sub-satellite point. The method first obtains the orbit control speed increment requirement based on the propellant discharge requirement, and then provides an orbit control pulse solution method with the least number of pulses and the fastest completion of the task based on the east-west adjustment requirements of the geographical longitude of the sub-satellite point.

[0049] Described below Figure 1 How the various steps are performed.

[0050] First, with respect to step 100 , the relationship between the propellant discharge amount and the total orbital control velocity increment is determined.

[0051] Specifically, the aircraft dry weight M0, the aircraft remaining fuel m fuel , Specific impulse I sp , thrust F, the relationship between propellant consumption and orbit control engine speed increment:

[0052]

[0053] It can be obtained that if the fuel mass to be discharged is Δm, the corresponding total speed increment of the orbital control engine is ΔV:

[0054]

[0055] The calculation of the velocity increment using out-of-plane orbital control discharge is relatively simple. When the propellant discharge velocity increment ΔV is determined each time, the velocity increment with an absolute value of ΔV / 2 can be used at the ascending node and the descending node to cancel the inclination effect. This method will not be described in detail.

[0056] Then, with respect to step 102, in one embodiment of the present invention, step 102 includes:

[0057] According to the nodal period of orbital motion and the precession rate of the ascending node right ascension, an expression for the change of geographic longitude is established;

[0058] According to the first-order partial derivative, the precession rate of the geographical longitude of the ascending node in one orbital period is solved to obtain the change of the semi-major axis;

[0059] The semi-major axis change is used to express additional changes in geographic longitude.

[0060] Specifically, the fuel discharge in the plane needs to be combined with the adjustment requirements of the sub-satellite point trajectory. For orbital reentry vehicles, the sub-satellite point adjustment requirements can be described as the adjustment requirements for the precession of the geographic longitude of the ascending node. For a fixed orbit, the orbit has an ideal change in the geographic longitude of the ascending node per revolution, denoted by λ N The ground track adjustment requirement is described as the additional change in the geographic longitude of the ascending node, Δλ N .

[0061] The nominal orbital semi-major axis is a0, the orbital eccentricity is e, and the orbital inclination is i. Then the change in geographical longitude of the sub-satellite point in one circle is λ N for:

[0062]

[0063] In the above formula:

[0064] is the precession rate of the ascending node's right ascension;

[0065] is the nodal period of orbital motion;

[0066] The constant J2 = 0.001082 is the gravitational perturbation coefficient of the Earth, μ = 3.986006 × 10 5 km 3 / s 2 is the gravitational constant, R e =6378.145km is the average radius of the earth.

[0067] For propellant discharge within the orbital plane, ignoring the orbital attenuation caused by the atmosphere, propellant discharge within the orbital plane can be completed by raising or lowering the orbit at the same latitude angle within the orbital plane at intervals of N full cycles.

[0068] According to the first-order partial derivative, the precession rate of the geographic longitude of the ascending node and the semi-major axis of the orbit are solved for one orbital period.

[0069]

[0070] The semi-major axis change Δa can be obtained, and the corresponding change in the geographic longitude of the ascending node after one circle Δλ N for:

[0071]

[0072] Regarding step 104, in one embodiment of the present invention, step 104 includes:

[0073] expressing the semi-major axis change using an orbit control velocity increment;

[0074] Using the orbit control speed increment to express the additional change in geographic longitude, a speed and longitude change expression is obtained;

[0075] The relationship expression is determined according to the velocity longitude variation expression and the number of revolutions between different pulse accelerations of the aircraft.

[0076] Specifically, the change in the semi-major axis is related to the orbit control velocity increment ΔV t The relationship between can be approximated by a small pulse:

[0077]

[0078] In the above formula, n is the orbital angular velocity.

[0079] Then the speed increment ΔV can be obtained t The relationship between the change in the geographical longitude of a circle of ascending nodes is as follows:

[0080]

[0081] The first pulse execution cycle is recorded as the M1th cycle, the second pulse execution cycle is recorded as the M2th cycle, and so on. The Nth pulse execution cycle is recorded as the Mth cycle. N To meet the phase modulation accuracy of the sub-satellite point, the relationship expression needs to be satisfied after N pulses:

[0082]

[0083] For step 106, if N orbit control pulses are designed to discharge propellant, and the orbital shape and the control requirements for the geographical longitude of the sub-satellite point need to be guaranteed after the discharge is completed, the constraints of the N discharge pulses can be written as follows:

[0084]

[0085] Since the propellant discharge is not an exact solution, the constraints are relaxed as follows:

[0086]

[0087] Where ΔV k2 A positive number designed to satisfy the propellant discharge constraint.

[0088] Further solving for the in-plane emission pulse:

[0089] |Δλ N |≥0, firstly, it can be designed as two pulse emissions, with an interval of M cycles between the two pulse emissions. The initial value of M is set to 1, and the iterative solution is:

[0090]

[0091] ΔV2=-ΔV1 (4)

[0092] This solution naturally satisfies the constraints of formulas (1) and (2). Substituting ΔV1 and ΔV2 into formula (3) for verification.

[0093] like This means that the amount of emission control completed by the M circle exceeds the emission demand, so set M = M + 1 and calculate formula (4) again. It can be seen that as M increases Gradually decrease, if ΔV1 and ΔV2 can meet the accuracy range of formula (2) and (3), the calculation is completed. The emission form is:

[0094] The execution number of the first pulse is M1, and the execution number of the second pulse is M2=M1+M;

[0095] The two pulses are:

[0096] If the value is M-1, When the value is M, This means that 2 pulses cannot find a feasible solution. Expand it to 3 pulses:

[0097] The execution number of the first pulse is M1, and the execution number of the second pulse is M2=M1+M;

[0098] In the above formula, when M satisfies M-1, When the value is M

[0099] Let the third pulse execution cycle be M3=M2+N, and initialize N=1;

[0100] The three pulses are:

[0101]

[0102] The three pulses naturally satisfy formulas (1) and (2). If they cannot satisfy the constraints of formula (3), then N = N + 1, and solve formula (5) again. As N increases, the absolute values ​​of ΔV2 and ΔV3 will further decrease. The absolute value of also gradually decreases until the constraint of formula (3) is satisfied.

[0103] like Figure 2 , Figure 3 As shown, an embodiment of the present invention provides a device for discharging excess propellant of a reentry vehicle. The device embodiment can be implemented by software, or by hardware or a combination of software and hardware. From the hardware level, Figure 2 As shown, it is a hardware architecture diagram of an electronic device in which a residual propellant discharge device of a reentry vehicle provided by an embodiment of the present invention is located. Figure 2 In addition to the processor, memory, network interface, and non-volatile memory shown, the electronic device in the embodiment may also include other hardware, such as a forwarding chip responsible for processing messages, etc. Taking software implementation as an example, Figure 3 As shown, as a device in a logical sense, the CPU of the electronic device in which it is located reads the corresponding computer program in the non-volatile memory into the memory and runs it. This embodiment provides a reentry vehicle residual propellant discharge device, including:

[0104] A first calculation unit determines the relationship between the propellant discharge amount and the total orbit control velocity increment;

[0105] a second calculation unit, determining an additional change in the geographic longitude of the ascending node of the aircraft based on a ground track adjustment requirement of the sub-satellite point of the aircraft;

[0106] A third calculation unit is used to establish a relationship expression between the additional change in geographic longitude, the number of orbits between different pulse accelerations of the aircraft, and the orbit control speed increments of different pulse accelerations;

[0107] The judgment unit, taking the number of pulse accelerations between two times and the interval of one revolution between different pulse accelerations as conditions, iteratively solves the relational expression to obtain the orbital control speed increment of each pulse acceleration and the specific number of revolutions of each pulse. The judgment unit, judges whether each orbital control speed increment meets the verification condition. If so, the propellant is discharged according to the current orbital control speed increment, the number of pulse accelerations and the number of revolutions between different pulse accelerations; if not, the number of pulse accelerations and the number of revolutions between different pulse accelerations are increased, the solution is iterated again, and the verification condition is used again for verification until the verification condition is met.

[0108] In one embodiment of the present invention, the verification condition includes that the sum of all the track control speed increments is 0;

[0109] The sum of the absolute values ​​of the track control speed increments is equal to the total track control speed increment.

[0110] In one embodiment of the present invention, the verification condition includes that the sum of all the track control speed increments is 0;

[0111] The difference between the sum of the absolute values ​​of the track control speed increments and the total track control speed increment is less than a preset error amount.

[0112] In one embodiment of the present invention, the second computing unit is further configured to perform the following operations:

[0113] According to the nodal period of orbital motion and the precession rate of the ascending node right ascension, an expression for the change of geographic longitude is established;

[0114] According to the first-order partial derivative, the precession rate of the geographical longitude of the ascending node in one orbital period is solved to obtain the change of the semi-major axis;

[0115] The semi-major axis change is used to express additional changes in geographic longitude.

[0116] In one embodiment of the present invention, the third computing unit is further configured to perform the following operations:

[0117] The expression for establishing the relationship between the additional change in geographic longitude, the number of orbits between different pulse accelerations of the aircraft, and the orbit control velocity increments of different pulse accelerations includes:

[0118] expressing the semi-major axis change using an orbit control velocity increment;

[0119] Using the orbit control speed increment to express the additional change in geographic longitude, a speed and longitude change expression is obtained;

[0120] The relationship expression is determined according to the velocity longitude variation expression and the number of revolutions between different pulse accelerations of the aircraft.

[0121] In one embodiment of the present invention, the velocity longitude change expression is:

[0122]

[0123] Among them, λ N is the change in geographical longitude for a certain circle, Δλ N is the additional change in geographic longitude for a certain orbit, a0 is the semi-major axis of the nominal orbit, ΔV t is the orbit control velocity increment.

[0124] In one embodiment of the present invention, the relational expression is as follows:

[0125]

[0126] Among them, λ N is the change in geographical longitude for a certain circle, Δλ N is the additional change in geographic longitude for a particular orbit, M t is the number of pulse executions, N is the number of pulse executions, ΔV t is the orbit control velocity increment.

[0127] It is to be understood that the structure illustrated in the embodiment of the present invention does not constitute a specific limitation on a reentry vehicle excess propellant discharge device. In other embodiments of the present invention, a reentry vehicle excess propellant discharge device may include more or fewer components than those illustrated, or combine certain components, or separate certain components, or arrange the components differently. The illustrated components may be implemented in hardware, software, or a combination of software and hardware.

[0128] The information interaction, execution process and other contents between the modules in the above-mentioned device are based on the same concept as the embodiment of the method of the present invention. For the specific contents, please refer to the description in the embodiment of the method of the present invention, and no further description is given here.

[0129] An embodiment of the present invention further provides an electronic device, including a memory and a processor, wherein the memory stores a computer program, and when the processor executes the computer program, a method for discharging residual propellant of a reentry vehicle in any embodiment of the present invention is implemented.

[0130] An embodiment of the present invention further provides a computer-readable storage medium having a computer program stored thereon. When the computer program is executed by a processor, the processor executes a method for discharging residual propellant of a reentry vehicle in any embodiment of the present invention.

[0131] Specifically, a system or device equipped with a storage medium can be provided, on which software program code that implements the functions of any of the above-mentioned embodiments is stored, and a computer (or CPU or MPU) of the system or device can be enabled to read and execute the program code stored in the storage medium.

[0132] In this case, the program code itself read from the storage medium can realize the function of any one of the above-mentioned embodiments, and thus the program code and the storage medium storing the program code constitute a part of the present invention.

[0133] The storage medium embodiments for providing the program code include a floppy disk, a hard disk, a magneto-optical disk, an optical disk (such as CD-ROM, CD-R, CD-RW, DVD-ROM, DVD-RAM, DVD-RW, DVD+RW), a magnetic tape, a non-volatile memory card, and a ROM. Alternatively, the program code can be downloaded from a server computer by a communication network.

[0134] In addition, it should be clear that the functions of any of the above embodiments can be implemented not only by executing the program code read by the computer, but also by enabling an operating system operating on the computer to complete part or all of the actual operations based on instructions from the program code.

[0135] In addition, it can be understood that the program code read from the storage medium is written to a memory provided in an expansion board inserted into the computer or to a memory provided in an expansion module connected to the computer, and then based on the instructions of the program code, a CPU installed on the expansion board or expansion module is enabled to perform part or all of the actual operations, thereby realizing the functions of any of the above-mentioned embodiments.

[0136] It should be noted that, in this article, relational terms such as first and second 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 statement "comprise a ..." do not exclude the presence of other identical factors in the process, method, article or device including the elements.

[0137] A person of ordinary skill in the art can understand that all or part of the steps of implementing the above method embodiments can be completed by hardware related to program instructions, and the aforementioned program can be stored in a computer-readable storage medium. When the program is executed, it executes the steps of the above method embodiments; and the aforementioned storage medium includes: ROM, RAM, magnetic disk or optical disk, etc., various media that can store program codes.

[0138] Finally, it should be noted that the above embodiments are only used to illustrate the technical solutions of the present invention, rather than to limit it. Although the present invention has been described in detail with reference to the aforementioned embodiments, those skilled in the art should understand that they can still modify the technical solutions described in the aforementioned embodiments, or make equivalent replacements for some of the technical features therein. However, these modifications or replacements do not deviate the essence of the corresponding technical solutions from the spirit and scope of the technical solutions of the embodiments of the present invention.

Claims

1. A method for discharging excess propellant from a reentry vehicle, characterized in that: include: Determine the relationship between propellant discharge and total orbital velocity increment; Determine the additional change in geographic longitude of the spacecraft's ascending node based on ground track adjustment requirements for the spacecraft's sub-satellite point; Establishing a relationship expression between the additional change in geographic longitude, the number of orbits between different pulse accelerations of the aircraft, and the orbit control velocity increments of different pulse accelerations; Taking the number of two pulse accelerations and the interval of one orbital revolution between different pulse accelerations as conditions, the relational expression is iteratively solved to obtain the orbital control speed increment of each pulse acceleration and the specific orbital revolution of each pulse, and it is judged whether each orbital control speed increment meets the verification conditions. If so, the propellant is discharged according to the current orbital control speed increment, the number of pulse accelerations and the orbital revolution between different pulse accelerations; if not, the number of pulse accelerations and the orbital revolution between different pulse accelerations are increased, the solution is iteratively solved again, and the verification is verified again using the verification conditions until the verification conditions are met.

2. The method according to claim 1, characterized in that The verification condition includes that the sum of all the track control speed increments is 0; The sum of the absolute values ​​of the track control speed increments is equal to the total track control speed increment.

3. The method according to claim 1, characterized in that The verification condition includes that the sum of all the track control speed increments is 0; The difference between the sum of the absolute values ​​of the track control speed increments and the total track control speed increment is less than a preset error amount.

4. The method according to claim 1, characterized in that: The additional change in the geographic longitude of the ascending node of the aircraft is determined according to the ground track adjustment requirement of the sub-satellite point of the aircraft, including: According to the nodal period of orbital motion and the precession rate of the ascending node right ascension, an expression for the change of geographic longitude is established; According to the first-order partial derivative, the precession rate of the geographical longitude of the ascending node in one orbital period is solved to obtain the change of the semi-major axis; The semi-major axis change is used to express additional changes in geographic longitude.

5. The method according to claim 4, characterized in that The expression for establishing the relationship between the additional change in geographic longitude, the number of orbits between different pulse accelerations of the aircraft, and the orbit control velocity increments of different pulse accelerations includes: expressing the semi-major axis change using an orbit control velocity increment; Using the orbit control speed increment to express the additional change in geographic longitude, a speed and longitude change expression is obtained; The relationship expression is determined according to the velocity longitude variation expression and the number of revolutions between different pulse accelerations of the aircraft.

6. The method according to claim 4, characterized in that The velocity longitude change expression is: Among them, λ N is the change in geographical longitude for a certain circle, Δλ N is the additional change in geographic longitude for a certain orbit, a0 is the semi-major axis of the nominal orbit, N is the number of pulse executions, ΔV t is the orbit control velocity increment.

7. The method according to claim 1, characterized in that The relational expression is as follows: Among them, λ N is the change in geographical longitude for a certain circle, Δλ N is the additional change in geographic longitude for a particular orbit, M t is the number of pulse executions, N is the number of pulse executions, ΔV t is the orbit control velocity increment.

8. A device for discharging excess propellant of a reentry vehicle, characterized in that: include: A first calculation unit determines the relationship between the propellant discharge amount and the total orbit control velocity increment; a second calculation unit, determining an additional change in the geographic longitude of the ascending node of the aircraft based on a ground track adjustment requirement of the sub-satellite point of the aircraft; A third calculation unit is used to establish a relationship expression between the additional change in geographic longitude, the number of orbits between different pulse accelerations of the aircraft, and the orbit control speed increments of different pulse accelerations; The judgment unit, taking the number of pulse accelerations between two times and the interval of one revolution between different pulse accelerations as conditions, iteratively solves the relational expression to obtain the orbital control speed increment of each pulse acceleration and the specific number of revolutions of each pulse. The judgment unit, judges whether each orbital control speed increment meets the verification condition. If so, the propellant is discharged according to the current orbital control speed increment, the number of pulse accelerations and the number of revolutions between different pulse accelerations; if not, the number of pulse accelerations and the number of revolutions between different pulse accelerations are increased, the solution is iterated again, and the verification condition is used again for verification until the verification condition is met.

9. An electronic device, comprising a memory and a processor, wherein the memory stores a computer program, and when the processor executes the computer program, the method according to any one of claims 1 to 7 is implemented.

10. A computer-readable storage medium having a computer program stored thereon, which, when executed in a computer, causes the computer to execute the method according to any one of claims 1 to 7.

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