Full-window self-adaptive guidance method for launching ground fire transfer orbit

Abstract

The invention relates to a full-window self-adaptive guidance method for a launching ground fire transfer orbit, which comprises the following steps of: planning an iterative guidance target orbit based on a quadratic curve according to launching time to obtain an orbital element; estimating the shutdown time of the last active section of rocket flight, and correcting the number of target orbits influenced by the flight time according to the earth rotation angular velocity and the shutdown time; an iterative guidance initial value is established according to the launching time, and iterative guidance calculation is carried out in each guidance period in the last active section of rocket flight to update a program angle by combining the number of target orbits and adopting a single-stage iterative guidance method; and according to the launching time, updating the shutdown amount of the last active section for ending the flight of the active section. According to the method, adaptive adjustment of the active section guidance method is achieved, correction of launching time deviation is achieved, and the purpose of high-precision orbit injection in a full window is achieved.

Description

Technical Field

[0001] The invention belongs to the technical field of launch vehicle control and relates to a full-window adaptive guidance method for launching an Earth-Mars transfer orbit. Background Art

[0002] The target orbit for launching a Mars probe by a carrier rocket is the Earth-Mars transfer orbit. In order to reduce the fuel consumption of the probe's mid-course orbit change and ensure that the probe can enter the Martian gravitational influence sphere, it is necessary to fine-tune the flight trajectory and target orbit of the carrier rocket. In traditional design methods, the launch window is usually discretely divided into several adjacent equally spaced sub-windows, and the corresponding target orbit and trajectory are designed for each sub-window. Therefore, the guidance system needs to design multiple trajectories. Under this design method, the fuel consumption of the probe's orbit change is not optimal. If you want to obtain an Earth-Mars transfer orbit that better matches the launch time, you need to shorten the sub-window duration, which will greatly increase the number of trajectories, which not only increases the workload of ground design, but also is not conducive to command decisions before the rocket launch. Summary of the invention

[0003] The technical problem solved by the present invention is: to overcome the shortcomings of the prior art, to propose a full-window adaptive guidance method for launching an Earth-Mars transfer orbit, and to adjust the program angle and shutdown parameters on the basis of traditional iterative guidance for launch missions of carrier rockets that require high-precision orbit insertion within the full window of the Earth-Mars transfer orbit. Based on the launch time, the iterative guidance program angle is corrected by compensating the iterative guidance target orbit, the iterative guidance estimated remaining flight time, the iterative initial value and the shutdown parameter, so as to realize adaptive adjustment of the active segment guidance method and achieve the purpose of high-precision orbit insertion within the full window.

[0004] The solution to the technical problem of the present invention is: a full-window adaptive guidance method for launching an Earth-Mars transfer orbit, comprising the following steps:

[0005] Plan the target orbit based on the quadratic curve iteration guidance according to the launch time and obtain the orbit elements;

[0006] estimating the shutdown time of the last active segment of the rocket flight, and correcting the target orbital elements affected by the flight time according to the angular velocity of the earth's rotation and the shutdown time;

[0007] The initial value of iterative guidance is established according to the launch time. Combined with the target orbital elements, a single-stage iterative guidance method is used. In the last active phase of the rocket flight, iterative guidance calculation is performed in each guidance cycle to update the program angle.

[0008] According to the launch time, the shutdown amount of the last active segment is updated to end the active segment flight.

[0009] Furthermore, the calculation method of the iterative guidance target trajectory is:

[0010] Define the launch time as T. Before entering the iterative guidance, the iterative guidance target orbit is calculated using a quadratic curve: According to the launch time, the orbital root apogee height H of the target orbit is updated a 、Perigee height H p , orbital inclination i, argument of perigee ω, and longitude of ascending node Ω are as follows:

[0011]

[0012] Among them, k a0 ,k a1 ,k a2 is the coefficient of the quadratic function of the apogee height, k p0 ,k p1 ,k p2 is the coefficient of the quadratic function of perigee height, k i0 ,k i1 ,k i2 is the coefficient of the quadratic function of orbital inclination, k ω0 ,k ω1 ,k ω2 is the coefficient of the quadratic function of the perigee argument, k Ω0 ,k Ω1 ,k Ω2 are the coefficients of the quadratic function of the longitude of the ascending node.

[0013] Furthermore, the shutdown time of the last active phase of the rocket flight is estimated as follows:

[0014] For full-window launch, single-stage iterative guidance is used. Before entering iterative guidance, the shutdown time T of the last active segment of the rocket flight is estimated using a quadratic curve based on the launch time T. C :

[0015] T C =k C0 +k C1 ·T+k C2 ·T 2

[0016] Among them, k C0 ,k C1 ,k C2 is the coefficient of the quadratic function of the shutdown time.

[0017] Furthermore, the target orbit elements affected by the flight time are corrected as follows:

[0018] Ω * =Ω+ω e ·TC

[0019] Among them, ω e is the angular velocity of the Earth's rotation, Ω * is the corrected longitude of the ascending node of the target orbit.

[0020] Furthermore, the initial value of iterative guidance is established according to the launch time, specifically:

[0021] Before entering the iterative guidance, according to the launch time T, the initial value of the iterative guidance based on the quadratic curve is established, including the iterative guidance to estimate the remaining flight time T of the last active segment k , the radius of the earth's center at the point of entry into orbit r k , the RMS velocity at the point of entry into orbit is V k , velocity inclination angle θ at the entry point k :

[0022]

[0023] Among them, k T0 ,k T1 ,k T2 is the coefficient of the quadratic function for estimating the remaining flight time, k r0 ,k r1 ,k r2 is the radius of the Earth's center at the point of entry into orbit r k Coefficients of the quadratic function, k V0 ,k V1 ,k V2 is the RMS velocity at the point of entry into orbit V k Coefficients of the quadratic function, k θ0 ,k θ1 ,k θ2 is the velocity inclination angle θ at the entry point k Coefficients of the quadratic function.

[0024] Furthermore, an iterative guidance calculation is performed in each guidance cycle to update the program angle, specifically:

[0025] Apogee height H based on the number of orbit elements of the iterative guidance target a 、Perigee height H p , orbital inclination i, perigee argument ω and corrected ascending node longitude Ω * , with T k ,r k ,V k ,θ k As the initial value of the iteration, iterative guidance calculation is performed in each guidance cycle to update T k ,r k ,V k ,θ k, and obtain the iterative guidance pitch program angle φ and yaw program angle ψ.

[0026] Furthermore, the shutdown amount of the last active segment is updated, specifically:

[0027] K=k K0 +k K1 T+k K2 T 2

[0028] Among them, K is the shutdown amount of the last active segment, k K0 ,k K1 ,k K2 is the coefficient of the quadratic curve function of the active segment shutdown quantity.

[0029] A computer-readable storage medium stores a computer program, and when the computer program is executed by a processor, the steps of a full-window adaptive guidance method for launching an Earth-Mars transfer orbit are implemented.

[0030] An electronic device comprises a memory, a processor and a computer program stored in the memory and executable on the processor, wherein the processor implements the steps of a full-window adaptive guidance method for launching an Earth-Mars transfer orbit when executing the computer program.

[0031] A computer program product includes a computer program, which, when executed by a processor, implements the steps of a full-window adaptive guidance method for launching an Earth-Mars transfer orbit.

[0032] The beneficial effects of the present invention compared with the prior art are:

[0033] (1) The present invention proposes an iterative guidance program angle adjustment method within the full window of rocket launch, which can enter the Earth-Mars transfer target orbit that satisfies the relative motion relationship between the Earth and Mars at the shutdown time.

[0034] (2) The present invention realizes the correction of the launch time deviation by performing quadratic curve compensation calculation on the iterative guidance target orbit, iterative initial value and shutdown parameters of the last active segment of the rocket.

[0035] (3) The present invention performs iterative guidance calculations based on the corrected target orbit and initial values, and controls the rocket to fly to an accurate Earth-Mars transfer orbit within the full window through adaptive adjustment of the iterative guidance program angle. BRIEF DESCRIPTION OF THE DRAWINGS

[0036] Figure 1 The present invention is a flow chart of a full-window adaptive guidance method for launching an Earth-Mars transfer orbit. DETAILED DESCRIPTION

[0037] The present invention proposes a full-window adaptive guidance method for launching an Earth-Mars transfer orbit. In the last active segment of the flight, the iterative guidance related parameters and the active segment shutdown related parameters are compensated and calculated in a quadratic curve manner according to the rocket launch time. When the rocket launch time changes, the iterative guidance program angle and the shutdown parameters in the full window are adaptively adjusted to send the Mars probe into an Earth-Mars transfer orbit that satisfies the relative motion relationship between the Earth and Mars.

[0038] The present invention will be further described below in conjunction with the embodiments.

[0039] Example 1

[0040] like Figure 1 As shown, a full-window adaptive guidance method for launching an Earth-Mars transfer orbit in this embodiment includes the following steps:

[0041] (1) Plan the target orbit based on the quadratic curve according to the launch time and obtain the orbit elements;

[0042] (2) estimating the shutdown time of the last active segment of the rocket flight, and correcting the target orbital elements affected by the flight time according to the angular velocity of the earth's rotation and the shutdown time;

[0043] (3) Establish the initial value of iterative guidance according to the launch time, combine the target orbit elements, adopt the single-stage iterative guidance method, and perform iterative guidance calculation in each guidance cycle in the last active phase of the rocket flight to update the program angle;

[0044] (4) According to the launch time, the shutdown amount of the last active segment is updated to end the active segment flight.

[0045] In step (1), the calculation method of the iterative guidance target trajectory is:

[0046] Define the launch time as T. Before entering the iterative guidance, the iterative guidance target orbit is calculated using a quadratic curve: According to the launch time, the orbital root apogee height H of the target orbit is updated a 、Perigee height H p , orbital inclination i, argument of perigee ω, and longitude of ascending node Ω are as follows:

[0047]

[0048] Among them, k a0 ,k a1 ,k a2 is the coefficient of the quadratic function of the apogee height, k p0 ,k p1 ,k p2 is the coefficient of the quadratic function of perigee height, ki0 ,k i1 ,k i2 is the coefficient of the quadratic function of orbital inclination, k ω0 ,k ω1 ,k ω2 is the coefficient of the quadratic function of the perigee argument, k Ω0 ,k Ω1 ,k Ω2 are the coefficients of the quadratic function of the longitude of the ascending node.

[0049] In step (2), the shutdown time of the last active segment of the rocket flight is estimated as follows:

[0050] For full-window launch, single-stage iterative guidance is used. Before entering iterative guidance, the shutdown time T of the last active segment of the rocket flight is estimated using a quadratic curve based on the launch time T. C :

[0051] T C =k C0 +k C1 ·T+k C2 ·T 2

[0052] Among them, k C0 ,k C1 ,k C2 is the coefficient of the quadratic function of the shutdown time.

[0053] In step (2), the target orbit elements affected by the flight time are corrected, specifically:

[0054] The longitude of the ascending node Ω is affected by the rotation of the Earth, so the angular velocity of the Earth's rotation ω is used. e and the last active segment shutdown time T C , correct the longitude Ω of the ascending node of the calculated iterative guidance target orbit:

[0055] Ω * =Ω+ω e ·T C

[0056] Among them, Ω * is the corrected longitude of the ascending node of the target orbit.

[0057] In step (3), the initial value of iterative guidance is established according to the launch time, specifically:

[0058] Before entering the iterative guidance, according to the launch time T, the initial value of the iterative guidance based on the quadratic curve is established, including the iterative guidance to estimate the remaining flight time T of the last active segment k , the radius of the earth's center at the point of entry into orbit r k, the RMS velocity at the point of entry into orbit is V k , velocity inclination angle θ at the entry point k :

[0059]

[0060] Among them, k T0 ,k T1 ,k T2 is the coefficient of the quadratic function for estimating the remaining flight time, k r0 ,k r1 ,k r2 is the radius of the Earth's center at the point of entry into orbit r k Coefficients of the quadratic function, k V0 ,k V1 ,k V2 is the RMS velocity at the point of entry into orbit V k Coefficients of the quadratic function, k θ0 ,k θ1 ,k θ2 is the velocity inclination angle θ at the entry point k Coefficients of the quadratic function.

[0061] In step (3), an iterative guidance calculation is performed in each guidance cycle to update the program angle, specifically:

[0062] Apogee height H based on the number of orbit elements of the iterative guidance target a 、Perigee height H p , orbital inclination i, perigee argument ω and corrected ascending node longitude Ω * , with T k ,r k ,V k ,θ k As the initial value of the iteration, iterative guidance calculation is performed in each guidance cycle to update T k ,r k ,V k ,θ k , and obtain the iterative guidance pitch program angle φ and yaw program angle ψ.

[0063] In step (4), the shutdown amount of the last active segment is updated, specifically:

[0064] K=k K0 +k K1 T+k K2 T 2

[0065] Among them, K is the shutdown amount of the last active segment, k K0 ,k K1 ,k K2 is the coefficient of the quadratic curve function of the active segment shutdown quantity.

[0066] The present invention calculates the program angle based on the traditional iterative guidance by iteratively guiding the target orbit, shutdown amount, shutdown time and iteration initial value according to the adjustment of the launch time, and realizes adaptive guidance under the condition of the change of the Mars transfer orbit in the full window, thereby meeting the requirements of orbit insertion accuracy.

[0067] The present application provides a computer-readable storage medium, wherein the computer-readable storage medium stores computer instructions, and when the computer instructions are executed on a computer, the computer executes Figure 1 The method described.

[0068] Those skilled in the art will appreciate that the embodiments of the present application may be provided as methods, systems, or computer program products. Therefore, the present application may adopt the form of a complete hardware embodiment, a complete software embodiment, or an embodiment in combination with software and hardware. Moreover, the present application may adopt the form of a computer program product implemented on one or more computer-usable storage media (including but not limited to disk storage and optical storage, etc.) that contain computer-usable program code.

[0069] The present application is described with reference to the flowcharts and / or block diagrams of the methods, devices (systems), and computer program products according to the embodiments of the present application. It should be understood that each process and / or box in the flowchart and / or block diagram, as well as the combination of the processes and / or boxes in the flowchart and / or block diagram, can be implemented by computer program instructions. These computer program instructions can be provided to a processor of a general-purpose computer, a special-purpose computer, an embedded processor, or other programmable data processing device to generate a machine, so that the instructions executed by the processor of the computer or other programmable data processing device generate instructions for implementing the processes in the flowchart and / or block diagram. Figure 1 A process or multiple processes and / or boxes Figure 1 A device that provides the functions specified in a block or multiple blocks.

[0070] These computer program instructions may also be stored in a computer-readable memory capable of directing a computer or other programmable data processing device to operate in a specific manner, so that the instructions stored in the computer-readable memory produce an article of manufacture comprising an instruction device, which implements the process Figure 1 A process or multiple processes and / or boxes Figure 1 A function specified in one or more boxes.

[0071] These computer program instructions can also be loaded onto a computer or other programmable data processing device so that a series of operating steps are executed on the computer or other programmable device to produce a computer-implemented process, thereby providing instructions for implementing the process. Figure 1 A process or multiple processes and / or boxes Figure 1A step that specifies a function in one or more boxes.

[0072] Obviously, those skilled in the art can make various changes and modifications to the present application without departing from the spirit and scope of the present application. Thus, if these modifications and variations of the present application fall within the scope of the claims of the present application and their equivalents, the present application is also intended to include these modifications and variations.

[0073] The contents not described in detail in the specification of the present invention belong to the common knowledge of those skilled in the art.

Claims

1. A full-window adaptive guidance method for launching an Earth-Mars transfer orbit, characterized in that: The following steps are involved: Plan the target orbit based on the quadratic curve iteration guidance according to the launch time and obtain the orbit elements; estimating the shutdown time of the last active segment of the rocket flight, and correcting the target orbital elements affected by the flight time according to the angular velocity of the earth's rotation and the shutdown time; The initial value of iterative guidance is established according to the launch time. Combined with the target orbital elements, a single-stage iterative guidance method is used. In the last active phase of the rocket flight, iterative guidance calculation is performed in each guidance cycle to update the program angle. According to the launch time, the shutdown amount of the last active segment is updated to end the active segment flight.

2. A full-window adaptive guidance method for launching an Earth-Mars transfer orbit according to claim 1, characterized in that: The calculation method of iterative guidance target trajectory is: Define the launch time as T. Before entering the iterative guidance, the iterative guidance target orbit is calculated using a quadratic curve: According to the launch time, the orbital root apogee height H of the target orbit is updated a 、Perigee height H p , orbital inclination i, argument of perigee ω, and longitude of ascending node Ω are as follows: Among them, k a0 ,k a1 ,k a2 is the coefficient of the quadratic function of the apogee height, k p0 ,k p1 ,k p2 is the coefficient of the quadratic function of perigee height, k i0 ,k i1 ,k i2 is the coefficient of the quadratic function of orbital inclination, k ω0 ,k ω1 ,k ω2 is the coefficient of the quadratic function of the perigee argument, k Ω0 ,k Ω1 ,k Ω2 are the coefficients of the quadratic function of the longitude of the ascending node.

3. A full-window adaptive guidance method for launching an Earth-Mars transfer orbit according to claim 2, characterized in that: The estimated shutdown time for the last active segment of rocket flight is: For full-window launch, single-stage iterative guidance is used. Before entering iterative guidance, the shutdown time T of the last active segment of the rocket flight is estimated using a quadratic curve based on the launch time T. C : T C =k C0 +k C1 ·T+k C2 ·T 2 Among them, k C0 ,k C1 ,k C2 is the coefficient of the quadratic function of the shutdown time.

4. A full-window adaptive guidance method for launching an Earth-Mars transfer orbit according to claim 3, characterized in that: Correct the target orbit elements affected by the flight time, specifically: Oh * =Ω+Ω e ·T C Among them, ω e is the Earth's rotation angular velocity, Ω * is the corrected longitude of the ascending node of the target orbit.

5. A full-window adaptive guidance method for launching an Earth-Mars transfer orbit according to claim 4, characterized in that: The initial value of iterative guidance is established according to the launch time, specifically: Before entering the iterative guidance, according to the launch time T, the initial value of the iterative guidance based on the quadratic curve is established, including the iterative guidance to estimate the remaining flight time T of the last active segment k , the radius of the earth's center at the point of entry into orbit r k , the RMS velocity at the point of entry into orbit is V k , velocity inclination angle θ at the entry point k : Among them, k T0 ,k T1 ,k T2 is the coefficient of the quadratic function for estimating the remaining flight time, k r0 ,k r1 ,k r2 is the radius of the Earth's center at the point of entry into orbit r k Coefficients of the quadratic function, k V0 ,k V1 ,k V2 is the RMS velocity at the point of entry into orbit V k Coefficients of the quadratic function, k θ0 ,k θ1 ,k θ2 is the velocity inclination angle θ at the entry point k Coefficients of the quadratic function.

6. A full-window adaptive guidance method for launching an Earth-Mars transfer orbit according to claim 5, characterized in that: An iterative guidance calculation is performed in each guidance cycle to update the program angle, specifically: Apogee height H based on the number of orbit elements of the iterative guidance target a 、Perigee height H p , orbital inclination i, perigee argument ω and corrected ascending node longitude Ω * , with T k ,r k ,V k ,θ k As the initial value of the iteration, iterative guidance calculation is performed in each guidance cycle to update T k ,r k ,V k ,θ k , and obtain the iterative guidance pitch program angle φ and yaw program angle ψ.

7. A full-window adaptive guidance method for launching an Earth-Mars transfer orbit according to claim 2, characterized in that: Update the shutdown amount of the last active segment, specifically: K=k K0 +k K1 T+k K2 T 2 Among them, K is the shutdown amount of the last active segment, k K0 ,k K1 ,k K2 is the coefficient of the quadratic curve function of the active segment shutdown quantity.

8. A computer-readable storage medium storing a computer program, characterized in that: When the computer program is executed by a processor, the steps of the method according to any one of claims 1 to 7 are implemented.

9. An electronic device comprising a memory, a processor, and a computer program stored in the memory and executable on the processor, characterized in that: When the processor executes the computer program, the steps of the method according to any one of claims 1 to 7 are implemented.

10. A computer program product, comprising a computer program, characterized in that When the computer program is executed by a processor, the steps of the method according to any one of claims 1 to 7 are implemented.

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