Method for optimizing orbit transfer strategy for large elliptical frozen orbit satellite orbit insertion

Abstract

The application provides an orbit transfer strategy optimization method and system for a large-ellipse frozen orbit satellite orbiting, comprising the following steps: acquiring a satellite-rocket separation orbit and a station parameter, calculating an elevation angle of the satellite and the station, and then determining a range of a true anomaly angle of pulse ignition meeting a measurement and control requirement; in combination with the range of the true anomaly angle of pulse ignition, generating a fuel most-probable pulse orbit change strategy of the satellite from an initial orbit to a target orbit, determining an orbit change speed increment, an ignition position and an ignition frequency; determining ignition-on and ignition-off conditions of each orbit change and a drift circle number between adjacent two orbit changes; according to the pulse orbit change strategy, respectively determining an initial value of an ignition time length, an ignition time and a thrust direction of each orbit change, and determining the final ignition time length, the ignition time and the thrust direction of each orbit change. The application adopts a non-apogee double-pulse orbit change strategy to design a fuel optimal orbit transfer strategy of a large-ellipse frozen orbit satellite orbiting, and solves the design method problem of the orbit transfer strategy.

Description

TECHNICAL FIELD

[0001] The present application relates to the field of aerospace technology, in particular, to an orbit transfer strategy optimization method and system for a large-ellipse frozen orbit satellite. BACKGROUND

[0002] Due to the constraints of China's launch capacity and measurement and control deployment conditions, China's existing rockets cannot directly transport satellites to large-ellipse frozen orbits. After the launch vehicle transports the satellite to the star-rocket separation orbit, the satellite itself needs to rely on the propellant carried by the satellite to enter the large-ellipse frozen orbit through multiple orbit changes. At present, China has no precedent for launching and orbit transfer of large-ellipse frozen orbit satellites.

[0003] Patent document CN112629543A discloses a large-ellipse orbit and small-inclination circular orbit orbit planning method, which belongs to the technical field of orbit planning. According to the selected target area information, the constraint conditions of orbit regression characteristics, the constraint conditions of reconnaissance load, the launch deployment parameters and the accurate earth motion model, etc., the planning and design of large-ellipse orbit and small-inclination circular orbit for different types of target area detection are completed. The large-ellipse orbit and small-inclination circular orbit design traverses the angle values in the effective inclination angle range at a specific interval, calculates the longitude value of the satellite orbit subsatellite point trajectory on the multiple transit target point latitude circles corresponding to each inclination angle value i, and calculates the difference between the target point geocentric longitude and the geocentric longitude value when the subsatellite point trajectory transits the target point latitude circle.

[0004] However, patent document CN112629543A is in the form of direct launch into orbit, does not need to design an orbit transfer strategy, and does not consider the launch capacity constraints.

[0005] The document "Research on Autonomous Orbit Correction Method in Large Elliptical Orbit Transfer Process", Master's Thesis of Harbin Institute of Technology, He Wei, 2010, uses an optimal control approach, which needs to generate a nominal transfer orbit in advance, then implement feedback correction control relative to the nominal transfer orbit, and does not consider various engineering constraints such as measurement and control.

[0006] The document "Research on Multi-Revolution Lambert Problem in Spacecraft Rendezvous", Journal of National University of Defense Technology, 2010, No. 5, mainly studies the optimal solution method of multi-revolution Lambert problem. The article gives a general solution to this type of Lambert problem, and the optimization objective is only fuel optimization without considering the orbit transfer time, so it needs to convert the single-revolution Lambert problem into the multi-revolution Lambert problem for research, which is relatively complex.

[0007] Therefore, there is a need for an orbit transfer strategy optimization method and system for large-ellipse frozen orbit satellite orbit insertion, which can serve as a technical basis for future same type satellite mission design. SUMMARY

[0008] In view of the defects in the prior art, the purpose of the present application is to provide an orbit transfer strategy optimization method and system for large-ellipse frozen orbit satellite orbit insertion.

[0009] According to the present application, an orbit transfer strategy optimization method for large-ellipse frozen orbit satellite orbit insertion is provided, which comprises:

[0010] Step S1: obtaining the satellite-rocket separation orbit and station parameters;

[0011] Step S2: calculating the satellite-station elevation angle according to the satellite-rocket separation orbit and station parameters, and then determining the range of the true anomaly angle of pulse ignition that meets the requirements of measurement and control;

[0012] Step S3: combining the range of the true anomaly angle of pulse ignition, generating the fuel most-probable pulse orbit transfer strategy of the satellite from the initial orbit to the target orbit, and determining the orbit transfer speed increment, ignition position and ignition times;

[0013] Step S4: determining the engine start-stop conditions of each orbit transfer and the drift circle number between adjacent two orbit transfers through the orbit transfer speed increment, ignition position and ignition times;

[0014] Step S5: determining the initial values of the ignition time length, ignition time and thrust direction of each orbit transfer according to the pulse orbit transfer strategy;

[0015] Step S6: determining the final ignition time length, ignition time and thrust direction of each orbit transfer according to the initial values.

[0016] Preferably, step S3 comprises:

[0017] Step S3.1: determining the search range of the to-be-optimized variable;

[0018] Step S3.2: constructing the Lambert problem from the initial orbit to the target orbit and solving the Lambert problem in the search range, obtaining the pulse speed vectors Δv1 and Δv2 required for 2-pulse non-apogee orbit transfer, i.e. the orbit transfer speed increment;

[0019] Step S3.3: traversing all values in the search range, selecting the corresponding value with the minimum Δv=||Δv1||+||Δv2||, and generating the fuel most-probable pulse orbit transfer strategy of the satellite from the initial orbit to the target orbit;

[0020] Step S3.4: the first ignition position is the position vector on the initial orbit with a length of a first impulse velocity vector Δv1; the second firing position is a position vector with a length of on the target orbit;

[0021] Step S3.5: Calculate the propellant mass Δm1 required for the maneuver from the initial orbit to the intermediate orbit, and the propellant mass Δm2 required for the maneuver from the intermediate orbit to the target orbit, according to the following formulas:

[0022] Δm1 = m0(1 - exp(-Δv1 / I sp / g))

[0023] Δm2 = (m0-Δm1)(1 - exp(-Δv2 / I sp / g))

[0024] where m0 represents the takeoff mass, I sp represents the specific impulse of the engine, and g represents the gravitational acceleration constant;

[0025] Step S3.6: Calculate the engine working time dt1 required for the maneuver from the initial orbit to the intermediate orbit, and the engine working time dt2 required for the maneuver from the intermediate orbit to the target orbit, according to the following formulas:

[0026] dt1 = Δm1 / (F / I sp / g)

[0027] dt2 = Δm2 / (F / I sp / g)

[0028] where F represents the engine thrust;

[0029] Step S3.7: Obtain the number of orbit changes N1 required for the maneuver from the initial orbit to the intermediate orbit according to the engine working time dt1 required for the maneuver from the initial orbit to the intermediate orbit, and obtain the number of orbit changes N2 required for the maneuver from the intermediate orbit to the target orbit according to the engine working time dt2 required for the maneuver from the intermediate orbit to the target orbit;

[0030] where the average engine working time is controlled within an index T Eng , and the index where T max is the maximum single firing time of the engine.

[0031] Preferably, the engine shutdown condition of each orbit change is determined by the semi-major axis change Δa i caused by each orbit change, the impulse velocity vector, and the orbit parameters before firing;

[0032] The drift number between two adjacent orbit transfers is determined by the geographical longitude at the time of star-rocket separation and the longitude drift rate.

[0033] Preferably, the initial value of the ignition duration Δt i , and the calculation formula is as follows:

[0034] Δt i = dt * Δa i / da

[0035] Wherein, dt represents the required engine working time corresponding to the Lambert orbit transfer impulse velocity of the ignition, such as the position of the ignition coincides with the impulse velocity vector Δv1, then dt = dt1; such as the position of the ignition coincides with the impulse velocity vector Δv2, then dt = dt2; da represents the semi-major axis change amount caused by the Lambert orbit transfer impulse velocity vector, when dt = dt1, da represents the difference between the semi-major axis of the intermediate orbit and the initial orbit; when dt = dt2, da is the difference between the semi-major axis of the target orbit and the intermediate orbit; Δa i represents the semi-major axis change amount caused by the i-th orbit transfer, and i represents the ignition serial number;

[0036] The initial value of the ignition time t i , and the calculation formula is as follows:

[0037] t i = t * - Δt i / 2

[0038] Wherein, t * represents the time corresponding to the impulse ignition position;

[0039] The initial value of the thrust direction α i is determined according to the impulse velocity vector; such as the position of the ignition coincides with the impulse velocity vector Δv1, then α i = Δv1 / || Δv1 ||; such as the position of the ignition coincides with the impulse velocity vector Δv2, then α i = Δv2 / || Δv2 ||.

[0040] Preferably, step S6 includes:

[0041] Step S6.1: taking the ignition duration Δt i , the ignition time t i and the thrust direction α i as the to-be-optimized parameters, and defining an optimization objective function G(X), and the formula is as follows:

[0042]

[0043] Wherein, Δm irepresents the propellant required for the i-th orbit transfer, Δλ Ω represents the difference between the orbit ascending node geographical longitude after the orbit transfer and the target orbit ascending node geographical longitude, β is a weighting coefficient, and N1+N2 represents the number of orbit transfers required to complete the orbit transfer;

[0044] Step S6.2: According to the initial values of the ignition time length, ignition time and thrust direction of each orbit transfer, a general numerical optimization algorithm is used to minimize the optimization objective function G(X), so as to determine the ignition time length, ignition time and thrust direction of each orbit transfer.

[0045] According to the orbit transfer strategy optimization system for large-ellipse frozen orbit satellite orbiting provided by the application, the following technical effects are achieved:

[0046] Module M1: Obtain the satellite-rocket separation orbit and station parameters;

[0047] Module M2: Calculate the satellite-azimuth angle of the station according to the satellite-rocket separation orbit and station parameters, and further determine the range of the true anomaly angle of pulse ignition meeting the requirements of the measurement and control;

[0048] Module M3: Combine the range of the true anomaly angle of pulse ignition, generate the fuel-saving pulse orbit transfer strategy of the satellite from the initial orbit to the target orbit, and determine the orbit transfer speed increment, ignition position and ignition times;

[0049] Module M4: Determine the ignition and shutdown conditions of each orbit transfer and the drift circle number between adjacent two orbit transfers through the orbit transfer speed increment, ignition position and ignition times;

[0050] Module M5: According to the pulse orbit transfer strategy, determine the initial values of the ignition time length, ignition time and thrust direction of each orbit transfer, respectively;

[0051] Module M6: Determine the final ignition time length, ignition time and thrust direction of each orbit transfer according to the initial values.

[0052] Preferably, module M3 comprises:

[0053] Module M3.1: Determine the search range of the optimization variable;

[0054] Module M3.2: Within the search range, construct the Lambert problem from the initial orbit to the target orbit and solve the Lambert problem to obtain the pulse speed vectors Δv1 and Δv2 required for 2-pulse non-apogee orbit transfer, i.e., the orbit transfer speed increment;

[0055] Module M3.3: Traverse all values in the search range, select the corresponding value with the minimum Δv=||Δv1||+||Δv2||, and generate the fuel-saving pulse orbit transfer strategy of the satellite from the initial orbit to the target orbit.

[0056] Module M3.4: the first ignition position is a position on the initial orbit with a position vector length of , and the second ignition position is a position on the target orbit with a position vector length of ;

[0057] Module M3.5: calculate the propellant mass Δm1 required for the maneuver from the initial orbit to the intermediate orbit, and the propellant mass Δm2 required for the maneuver from the intermediate orbit to the target orbit, according to the following formulas:

[0058] Δm1 = m0(1 - exp(-Δv1 / I sp / g))

[0059] Δm2 = (m0-Δm1)(1 - exp(-Δv2 / I sp / g))

[0060] wherein m0 represents the takeoff mass, I sp represents the specific impulse of the engine, and g represents the gravitational acceleration constant;

[0061] Module M3.6: calculate the engine working time dt1 required for the maneuver from the initial orbit to the intermediate orbit, and the engine working time dt2 required for the maneuver from the intermediate orbit to the target orbit, according to the following formulas:

[0062] dt1 = Δm1 / (F / I sp / g)

[0063] dt2 = Δm2 / (F / I sp / g)

[0064] wherein F represents the engine thrust;

[0065] Module M3.7: obtain the number of orbit changes N1 required for the maneuver from the initial orbit to the intermediate orbit according to the engine working time dt1 required for the maneuver from the initial orbit to the intermediate orbit, and obtain the number of orbit changes N2 required for the maneuver from the intermediate orbit to the target orbit according to the engine working time dt2 required for the maneuver from the intermediate orbit to the target orbit;

[0066] wherein the average engine working time is controlled within an index T Eng , and the index wherein T max is the maximum single ignition time of the engine.

[0067] Preferably, the engine shutdown condition of each orbit change is determined by the semi-major axis change Δa iThe impulse velocity vector and the orbit parameter before ignition are determined jointly;

[0068] The drift circle number between two adjacent orbit transfers is determined by the geographical longitude at the time of star-rocket separation and the longitude drift rate.

[0069] Preferably, the initial value of the ignition duration Δt i is calculated according to the following formula:

[0070] Δt i = dt * Δa i / da

[0071] wherein dt represents the required engine working time corresponding to the Lambert orbit transfer impulse velocity at the ignition position, such as dt = dt1 when the ignition position is consistent with the impulse velocity vector Δv1, or dt = dt2 when the ignition position is consistent with the impulse velocity vector Δv2; da represents the semi-major axis variation caused by the Lambert orbit transfer impulse velocity vector, wherein da represents the difference between the semi-major axis of the intermediate orbit and the initial orbit when dt = dt1, and da represents the difference between the semi-major axis of the target orbit and the intermediate orbit when dt = dt2; Δa i represents the semi-major axis variation caused by the i-th orbit transfer, and i represents the ignition sequence number;

[0072] The initial value of the ignition time t i is calculated according to the following formula:

[0073] t i = t * - Δt i / 2

[0074] wherein t * represents the time corresponding to the impulse ignition position;

[0075] The initial value of the thrust direction α i is determined according to the impulse velocity vector, such as α i = Δv1 / || Δv1 || when the ignition position is consistent with the impulse velocity vector Δv1, or α i = Δv2 / || Δv2 || when the ignition position is consistent with the impulse velocity vector Δv2.

[0076] Preferably, the module M6 comprises:

[0077] Module M6.1: defining an optimization objective function G(X) with the ignition duration Δt i , the ignition time t i and the thrust direction α i as the parameters to be optimized, according to the following formula:

[0078]

[0079] wherein, Δm i represents the propellant required for the i-th orbit transfer, Δλ Ω represents the difference between the orbit ascending node geographical longitude after the orbit transfer and the target orbit ascending node geographical longitude, β is a weighting coefficient, and N1+N2 represents the number of orbit transfers required to complete the orbit transfer;

[0080] Module M6.2: According to the initial values of the ignition time length, ignition time and thrust direction of each orbit transfer, a general numerical optimization algorithm is used to minimize the optimization objective function G(X), so as to determine the ignition time length, ignition time and thrust direction of each orbit transfer.

[0081] Compared with the prior art, the present application has the following beneficial effects:

[0082] 1. The present application adopts a non-apogee double-pulse orbit transfer strategy to design a fuel-optimal orbit transfer strategy for a large-ellipse frozen orbit satellite entering an orbit, and solves the design method problem of the orbit transfer strategy.

[0083] 2. The present application does not need to generate a nominal transfer orbit, but can directly generate a subsequent orbit transfer strategy according to the current orbit, and at the same time ensures the control optimality and the satisfaction degree of the engineering constraints. The engineering feasibility is high, and can be used to guide the task implementation.

[0084] 3. The present application solves the orbit transfer strategy that meets the least fuel consumption in the orbit transfer process by considering the satellite orbit transfer engine working capacity, the measurement and control arc segment constraint in the orbit transfer process, the drift circle number after each orbit transfer of the satellite, the ascending node geographical longitude of the satellite entering the large-ellipse frozen orbit, and the like. The present application provides a design basis for the overall design work such as satellite fuel budget and flight program design. BRIEF DESCRIPTION OF DRAWINGS

[0085] Other features, objects and advantages of the present application will become more apparent from the following detailed description of non-limiting embodiments, made with reference to the accompanying drawings:

[0086] Figure 1 The figure is a schematic diagram of the work flow of the present application. DETAILED DESCRIPTION

[0087] The present application will be described in detail below with reference to specific embodiments. The following embodiments will help those skilled in the art to further understand the present application, but do not limit the present application in any form. It should be noted that those skilled in the art can make several changes and improvements without departing from the concept of the present application. These all belong to the protection scope of the present application.

[0088] According to the application, an orbit transfer strategy optimization method for large-elliptical frozen orbit satellite orbiting is provided, as shown in the specification, comprising: Figure 1

[0089] Step S1: obtaining satellite-rocket separation orbit and station parameters.

[0090] Step S2: calculating the elevation angle of the satellite and the station according to the satellite-rocket separation orbit and the station parameters, and determining the range of the true anomaly angle of pulse ignition meeting the requirements of the measurement and control. Specifically, according to the satellite-rocket separation orbit and the station parameters, the elevation angle of the satellite and the station is calculated by taking different orbit true anomaly angles f, and the range of the true anomaly angle of pulse ignition meeting the requirements of the measurement and control is determined according to the elevation angle constraint of the station [f L ,f U ].

[0091] Step S3: combining the range of the true anomaly angle of pulse ignition, adopting non-apogee double-pulse orbit transfer, generating the fuel-saving pulse orbit transfer strategy of the satellite from the initial orbit to the target orbit, and determining the orbit transfer speed increment, ignition position and ignition times. Step S3 comprises:

[0092] Step S3.1: determining the search range of the to-be-optimized variable. The search range comprises the range of the ascending node right ascension difference ΔΩ between the initial orbit and the target orbit [ΔΩ L ,ΔΩ U ], the range of the time Δt for transferring from the initial orbit to the target orbit by 2-pulse non-apogee orbit transfer [Δt L ,Δt U ], and the range of the pulse ignition position f obtained in the step S2 [f L ,f U ].

[0093] Step S3.2: constructing the Lambert problem from the initial orbit to the target orbit and solving the Lambert problem in the search range, to obtain the pulse speed vectors Δv1 and Δv2 required for 2-pulse non-apogee orbit transfer, i.e., the orbit transfer speed increment. That is, different ΔΩ, Δt and f are taken in the above search range, the Lambert problem from the initial orbit to the target orbit is constructed, and the pulse speed vectors Δv1 and Δv2 required for 2-pulse non-apogee orbit transfer can be obtained by solving the problem.

[0094] Step S3.3: traversing all values in the search range, selecting the corresponding value with the minimum Δv = ||Δv1|| + ||Δv2||, and generating the fuel-saving pulse orbit transfer strategy of the satellite from the initial orbit to the target orbit. Specifically, all ΔΩ, Δt and f in the search range given in the step S3.2 are traversed, the corresponding Lambert problem is solved, and the ΔΩ * ,Δt​* , f1 * and The fuel-saving impulse orbit transfer strategy from the initial orbit to the target orbit can be obtained, wherein ||Δv1|| and ||Δv2|| are the magnitudes of Δv1 and Δv2, respectively, i.e., the orbit transfer velocity increments,

[0095] Step S3.4: the first firing position is a position on the initial orbit with a position vector length of , and the second firing position is a position on the target orbit with a position vector length of . The first firing position is determined according to the initial orbit and f1 * , and the second firing position is determined according to the target orbit and .

[0096] Step S3.5: calculate the propellant mass Δm1 required for the maneuver from the initial orbit to the intermediate orbit, and the propellant mass Δm2 required for the maneuver from the intermediate orbit to the target orbit, according to the following formulas:

[0097] Δm1 = m0(1-exp(-Δv1 / I sp / g))

[0098] Δm2 = (m0-Δm1)(1-exp(-Δv2 / I sp / g))

[0099] wherein m0 represents the takeoff mass, I sp represents the specific impulse of the engine, and g represents the gravitational acceleration constant.

[0100] Step S3.6: calculate the engine working time dt1 required for the maneuver from the initial orbit to the intermediate orbit, and the engine working time dt2 required for the maneuver from the intermediate orbit to the target orbit, according to the following formulas:

[0101] dt1 = Δm1 / (F / I sp / g)

[0102] dt2 = Δm2 / (F / I sp / g)

[0103] wherein F represents the engine thrust.

[0104] Step S3.7: according to the engine working time dt1 required for the maneuver from the initial orbit to the intermediate orbit, obtain the number of orbit transfers N1 required for the maneuver from the initial orbit to the intermediate orbit, and according to the engine working time dt2 required for the maneuver from the intermediate orbit to the target orbit, obtain the number of orbit transfers N2 required for the maneuver from the intermediate orbit to the target orbit.

[0105] Among these measures, while minimizing the number of track changes, the average engine operating time is controlled within the target range T. Eng Within, indicators Where T max This is the longest single ignition time for the engine.

[0106] Step S4: Determine the engine shutdown conditions for each track change by using the track change speed increment, ignition position, and number of ignitions (a i i i ,ω i And the number of drift cycles between two adjacent trajectory changes. In the shutdown conditions, a i i is the semi-major axis of the orbit. i Let ω be the orbital inclination angle. i Let be the perigee angle of the orbit. The engine shutdown conditions for each orbit change are determined by the semi-major axis change Δa caused by each orbit change. i The pulse velocity vector and the orbital parameters before ignition are used to determine the orbital parameters. The number of drift cycles between two adjacent orbital maneuvers is determined by the geographical longitude and longitude drift rate at the moment of satellite-launch separation.

[0107] The first pulse orbital change is divided into N1 orbital changes, and the second pulse orbital change is divided into N2 orbital changes. A total of N1 + N2 orbital changes are needed to complete the orbital transfer. (The remaining text appears to be incomplete and requires further context.) The drift number Q1, λ0 and... of the first orbital change can be determined. These represent the geographical longitude and longitude drift rate at the moment of star-rocket separation, λ0 and All of these can be directly determined from the orbital parameters at the moment of satellite-rocket separation. Q1 is the minimum number of drift cycles required to keep λ1 within the telemetry and control range, λ1 is the geographical longitude of the first orbital maneuver, and the length from orbital prediction to position vector is... The position is determined. The drift number and shutdown conditions for each subsequent orbit change are determined.

[0108] According to the constraints, by λ1, λ i and Determine the drift number and shutdown conditions for each subsequent orbit change; It is the longitude drift rate after the first orbital change, λ. i It is the geographical longitude of the i-th orbital change. It is the longitude drift rate after the i-th orbital change; the constraint condition refers to the longitude drift rate caused by the orbital change, provided that the single engine operating time does not exceed the maximum operating time limit. i The larger the value, the fewer the number of orbits between two adjacent orbit changes, and the orbit change point should be within the ground control range; where Δa i It is the change in the semi-major axis caused by each track change, shutdown condition (a) i ii ,ω i ) can be determined by Δa i , impulse velocity vector Δv1 or Δv2 and orbit parameters before ignition.

[0109] Step S5: according to the impulse orbit transfer strategy, respectively determine the initial value of ignition time, ignition time and thrust direction of each orbit transfer. The initial value of ignition time Δt i , the calculation formula is as follows:

[0110] Δt i = dt * Δa i / da

[0111] Wherein, dt represents the required engine working time of Lambert orbit transfer impulse velocity corresponding to the ignition, such as the position of the ignition is consistent with the impulse velocity vector Δv1, then dt = dt1; such as the position of the ignition is consistent with the impulse velocity vector Δv2, then dt = dt2. Da represents the semi-major axis change caused by Lambert orbit transfer impulse velocity vector, when dt = dt1, da is the difference between the semi-major axis of the intermediate orbit and the initial orbit; when dt = dt2, da is the difference between the semi-major axis of the target orbit and the intermediate orbit. Δa i represents the semi-major axis change caused by the i-th orbit transfer, and i is the ignition number.

[0112] The initial value of ignition time t i , the calculation formula is as follows:

[0113] t i = t * - Δt i / 2

[0114] Wherein, t * represents the corresponding time of impulse ignition position, t * is determined according to orbit prediction.

[0115] The initial value of thrust direction α i is determined according to the impulse velocity vector. Such as the position of the ignition is consistent with the impulse velocity vector Δv1, then α i = Δv1 / || Δv1||; such as the position of the ignition is consistent with the impulse velocity vector Δv2, then α i = Δv2 / || Δv2||.

[0116] Step S6: according to the initial value, determine the final ignition time, ignition time and thrust direction of each orbit transfer. Step S6 includes:

[0117] Step S6.1: take the ignition time Δt i , ignition time t i and thrust direction αi For the to-be-optimized parameters, an optimization objective function G(X) is defined, and the formula is as follows:

[0118]

[0119] Where, Δm i represents the propellant required for the i-th orbit transfer, Δλ Ω represents the difference between the orbit ascending node geographical longitude after the orbit transfer and the target orbit ascending node geographical longitude, β is a weighting coefficient, and N1+N2 represents the number of orbit transfers required to complete the orbit transfer.

[0120] Step S6.2: According to the initial values of the ignition time length, ignition time and thrust direction of each orbit transfer, the general numerical optimization algorithm is used to minimize the optimization objective function G(X), so as to determine the ignition time length, ignition time and thrust direction of each orbit transfer.

[0121] The application also provides an orbit transfer strategy optimization system for a large-ellipse frozen orbit satellite entering an orbit, which can be realized by performing the flow steps of the orbit transfer strategy optimization method for the large-ellipse frozen orbit satellite entering the orbit, that is, the orbit transfer strategy optimization method for the large-ellipse frozen orbit satellite entering the orbit can be understood as the preferred embodiment of the orbit transfer strategy optimization system for the large-ellipse frozen orbit satellite entering the orbit by those skilled in the art.

[0122] According to the orbit transfer strategy optimization system for a large-ellipse frozen orbit satellite entering an orbit provided by the application, the following are included:

[0123] Module M1: Obtain the satellite-rocket separation orbit and station parameters.

[0124] Module M2: Calculate the elevation angle of the satellite and the station according to the satellite-rocket separation orbit and the station parameters, and then determine the range of the true anomaly angle of pulse ignition that meets the requirements of the measurement and control.

[0125] Module M3: In combination with the range of the true anomaly angle of pulse ignition, generate the fuel-saving pulse orbit transfer strategy of the satellite from the initial orbit to the target orbit, and determine the orbit transfer speed increment, ignition position and ignition times. Module M3 includes:

[0126] Module M3.1: Determine the search range of the to-be-optimized variables.

[0127] Module M3.2: Within the search range, construct the Lambert problem from the initial orbit to the target orbit and solve the Lambert problem to obtain the pulse speed vectors Δv1 and Δv2 required for 2-pulse non-apogee orbit transfer, that is, the orbit transfer speed increment

[0128] Module M3.3: Select the corresponding value of Δv = ||Δv1|| + ||Δv2|| minimum in all values in the search range, and generate the impulse maneuver strategy of the satellite from the initial orbit to the target orbit with the least fuel consumption.

[0129] Module M3.4: The first firing position is the position with the position vector length of on the initial orbit, which is used to apply the first impulse velocity vector Δv1; the second firing position is the position with the position vector length of on the target orbit, which is used to apply the second impulse velocity vector Δv2.

[0130] Module M3.5: Calculate the propellant mass Δm1 required for the maneuver from the initial orbit to the intermediate orbit, and the propellant mass Δm2 required for the maneuver from the intermediate orbit to the target orbit, and the calculation formula is as follows:

[0131] Δm1 = m0(1 - exp(-Δv1 / I sp / g))

[0132] Δm2 = (m0-Δm1)(1 - exp(-Δv2 / I sp / g))

[0133] Wherein, m0 represents the take-off mass, I sp represents the specific impulse of the engine, and g represents the gravitational acceleration constant.

[0134] Module M3.6: Calculate the working time dt1 required for the engine from the initial orbit to the intermediate orbit, and the working time dt2 required for the engine from the intermediate orbit to the target orbit, and the calculation formula is as follows:

[0135] dt1 = Δm1 / (F / I sp / g)

[0136] dt2 = Δm2 / (F / I sp / g)

[0137] Wherein, F represents the engine thrust.

[0138] Module M3.7: According to the working time dt1 required for the engine from the initial orbit to the intermediate orbit, the number of orbit maneuvers N1 required from the initial orbit to the intermediate orbit is obtained, and according to the working time dt2 required for the engine from the intermediate orbit to the target orbit, the number of orbit maneuvers N2 required from the intermediate orbit to the target orbit is obtained. Wherein, the average working time of the engine is controlled within the index T Eng , and the index Where T max is the maximum single firing time of the engine.

[0139] Module M4: Determine the engine on-off condition of each orbit transfer and the number of drift orbits between two adjacent orbit transfers by the orbit transfer velocity increment, the firing position and the firing number. The engine on-off condition of each orbit transfer is determined by the semi-major axis variation Δa i , the impulse velocity vector and the orbit parameter before firing. The number of drift orbits between two adjacent orbit transfers is determined by the geographical longitude at the moment of star-rocket separation and the longitude drift rate.

[0140] Module M5: Determine the initial value of the firing duration, the firing time and the thrust direction of each orbit transfer according to the impulse orbit transfer strategy. The initial value of the firing duration Δt i , the calculation formula is as follows: Δt i = dt * Δa i / da, wherein dt represents the required engine working time of the Lambert orbit transfer impulse velocity corresponding to the current firing, such as dt = dt1 when the current firing position is consistent with the impulse velocity vector Δv1; such as dt = dt2 when the current firing position is consistent with the impulse velocity vector Δv2. da represents the semi-major axis variation caused by the Lambert orbit transfer impulse velocity vector, which is the difference between the semi-major axis of the intermediate orbit and the initial orbit when dt = dt1; which is the difference between the semi-major axis of the target orbit and the intermediate orbit when dt = dt2. Δa i represents the semi-major axis variation caused by the i-th orbit transfer, and i is the firing number. The initial value of the firing time t i , the calculation formula is as follows: t i = t * - Δt i / 2, wherein t * represents the corresponding time of the impulse firing position. The initial value of the thrust direction α i is determined according to the impulse velocity vector, and the calculation formula is as follows: The initial value of the thrust direction α i is determined according to the impulse velocity vector. Such as α i = Δv1 / ||Δv1|| when the current firing position is consistent with the impulse velocity vector Δv1; such as α i = Δv2 / ||Δv2|| when the current firing position is consistent with the impulse velocity vector Δv2.

[0141] Module M6: Determine the final firing duration, firing time and thrust direction of each orbit transfer according to the initial value. Module M6 includes:

[0142] Module M6.1: Take the firing duration Δt i , the firing time t i and the thrust direction α i as the optimization parameters, and define the optimization objective function G(X) as follows:

[0143]

[0144] wherein, Δm i represents the propellant required for the i-th orbit transfer, Δλ Ω represents the difference between the orbital ascending node geographical longitude after the orbit transfer and the target orbital ascending node geographical longitude, β is a weighting coefficient, and N1+N2 represents the number of orbit transfers required to complete the orbit transfer;

[0145] Module M6.2: According to the initial values of the ignition time length, ignition time and thrust direction of each orbit transfer, a general numerical optimization algorithm is used to minimize the optimization objective function G(X), so as to determine the ignition time length, ignition time and thrust direction of each orbit transfer.

[0146] Those skilled in the art know that, in addition to implementing the system and each device, module and unit thereof provided by the present application in the form of pure computer readable program code, the system and each device, module and unit thereof provided by the present application can also be implemented in the form of logic gates, switches, application specific integrated circuits, programmable logic controllers and embedded microcontrollers, etc. by logically programming the method steps to achieve the same functions. Therefore, the system and each device, module and unit thereof provided by the present application can be considered as a hardware component, and the devices, modules and units included therein for achieving various functions can also be considered as structures within the hardware component; the devices, modules and units for achieving various functions can also be considered as both software modules for implementing methods and structures within hardware components.

[0147] The specific embodiments of the present application are described above. It needs to be understood that the present application 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 essential content of the present application. The embodiments of the present application and the features in the embodiments can be arbitrarily combined with each other without conflict.

Claims

1. An orbit transfer strategy optimization method for a large elliptical frozen orbit satellite, characterized in that, Comprising: Step S1: obtaining the satellite-rocket separation orbit and station parameters; Step S2: calculating the elevation angle of the satellite and the station according to the satellite-rocket separation orbit and station parameters, and further determining the pulse ignition true anomaly angle range meeting the requirements of the measurement and control; Step S3: combining the pulse ignition true anomaly angle range, generating the fuel most-probable pulse orbit transfer strategy of the satellite from the initial orbit to the target orbit, and determining the orbit transfer speed increment, ignition position and ignition times; Step S4: determining the engine start-stop conditions of each orbit transfer and the drift orbit number between the adjacent two orbit transfers through the orbit transfer speed increment, ignition position and ignition times; Step S5: determining the ignition time length, ignition time and initial value of the thrust direction of each orbit transfer according to the pulse orbit transfer strategy; Step S6: determining the final ignition time length, ignition time and thrust direction of each orbit transfer according to the initial value.

2. The method for orbit transfer strategy optimization for large elliptical frozen orbit satellite orbit insertion according to claim 1, characterized in that, Step S3 comprises: Step S3.1: determining the search range of the to-be-optimized variable; Step S3.2: constructing the Lambert problem from the initial orbit to the target orbit and solving the Lambert problem in the search range, obtaining the pulse speed vectors Δv1 and Δv2 required by the 2-pulse non-apogee orbit transfer, i.e. the orbit transfer speed increment; Step S3.3: traversing all values in the search range, selecting the corresponding value with the minimum Δv=||Δv1||+||Δv2||, and generating the fuel most-probable pulse orbit transfer strategy of the satellite from the initial orbit to the target orbit; Step S3.4: the first firing position is a position on the initial orbit with a position vector length of for applying the first impulse velocity vector Δvi; the second firing position is a position on the target orbit with a position vector length of for applying the second impulse velocity vector Δv2; Step S3.5: calculating the propellant mass Δm1 required for the maneuver from the initial orbit to the intermediate orbit, and the propellant mass Δm2 required for the maneuver from the intermediate orbit to the target orbit, and the calculation formula is as follows: Δm1 = m0(1 - exp(-Δv1 / I sp / g)) Am2 = (m0 - Am1) (1 - exp(-Av2 / I sp / g)) where m0represents the takeoff mass, I sp represents the engine specific impulse, and g represents the gravitational acceleration constant; Step S3.6: calculating the engine working time dt1 required for the maneuver from the initial orbit to the intermediate orbit, and the engine working time dt2 required for the maneuver from the intermediate orbit to the target orbit, and the calculation formula is as follows: dt1 = Δm1 / (F / I sp / g) dt2 = Δm2 / (F / I sp / g) Wherein, F represents the engine thrust; Step S3.7: obtaining the orbit transfer times N1 required for the maneuver from the initial orbit to the intermediate orbit according to the engine working time dt1 required for the maneuver from the initial orbit to the intermediate orbit, and obtaining the orbit transfer times N2 required for the maneuver from the intermediate orbit to the target orbit according to the engine working time dt2 required for the maneuver from the intermediate orbit to the target orbit; wherein the average engine operating time is controlled to be within a target T Eng wherein the target T wherein T max is the maximum single engine firing time.

3. The method for orbit transfer strategy optimization for large elliptical frozen orbit satellite orbit insertion according to claim 2, characterized in that, The engine start and stop conditions of each orbit change are determined by the semi-major axis variation Δa caused by each orbit change i The impulse velocity vector and the orbit parameters before ignition are determined jointly; The drift orbit number between the adjacent two orbit transfers is determined by the geographic longitude at the satellite-rocket separation time and the longitude drift rate.

4. The method for orbit transfer strategy optimization for large elliptical frozen orbit satellite orbit insertion according to claim 1, characterized in that, the initial value of the ignition duration Δt i The calculation formula is as follows: Δt i = dt * Δa i / da Wherein, dt represents the required engine working time of the Lambert orbit transfer impulse velocity corresponding to the current ignition, such as the position of the current ignition is consistent with the impulse velocity vector Δv1, then dt = dt1; such as the position of the current ignition is consistent with the impulse velocity vector Δv2, then dt = dt2; da represents the semi-major axis change caused by the Lambert orbit transfer impulse velocity vector, when dt = dt1, da represents the difference between the semi-major axis of the intermediate orbit and the initial orbit; when dt = dt2, da is the difference between the semi-major axis of the target orbit and the intermediate orbit; Δa i represents the semi-major axis change caused by the i-th orbit transfer, i represents the ignition number; Initial value of ignition timing t i The calculation formula is as follows: t i = t * - Δt i / 2 wherein t * denotes the instant corresponding to the position of the spark plug Initial value α of the thrust direction i The value is determined based on the pulse velocity vector; if the location of the current ignition coincides with the pulse velocity vector Δv1, then α... i =Δv1 / ||Δv1||; If the location of the current ignition coincides with the pulse velocity vector Δv2, then α i =Δv2 / ||Δv2‖.

5. The method for orbit transfer strategy optimization for large elliptical frozen orbit satellite orbit insertion according to claim 1, wherein, Step S6 comprises: Step S6.1: defining the optimization objective function G(X) with the ignition time length Δt i , the ignition time t i , and the thrust direction α i as the parameters to be optimized, as follows: where Δm i represents the propellant required for the i-th orbit transfer, Δλ Ω represents the difference between the orbital ascending node geographical longitude after the orbit transfer and the target orbital ascending node geographical longitude, β is a weighting factor, and N1+N2 represents the number of orbit transfers required to complete the orbit transfer; Step S6.2: according to the initial value of the ignition time length, ignition time and thrust direction of each orbit transfer, using a general numerical optimization algorithm to make the optimization objective function G(X) reach the minimum value, so as to determine the ignition time length, ignition time and thrust direction of each orbit transfer.

6. An orbit transfer strategy optimization system for satellites entering highly elliptical frozen orbits, characterized in that, Comprising: Module M1: obtaining the satellite-rocket separation orbit and station parameters; Module M2: calculating the elevation angle of the satellite and the station according to the satellite-rocket separation orbit and station parameters, and further determining the pulse ignition true anomaly angle range meeting the requirements of the measurement and control; Module M3: combining the pulse ignition true anomaly angle range, generating the fuel most-probable pulse orbit transfer strategy of the satellite from the initial orbit to the target orbit, and determining the orbit transfer speed increment, ignition position and ignition times; Module M4: determining the engine shutdown condition of each time of orbit transfer and the number of drift orbits between two adjacent times of orbit transfer by the orbit transfer velocity increment, the ignition position and the ignition times; Module M5: determining the initial values of the ignition duration, the ignition time and the thrust direction of each time of orbit transfer according to the pulse orbit transfer strategy; Module M6: determining the final ignition duration, the ignition time and the thrust direction of each time of orbit transfer according to the initial values.

7. The orbit transfer strategy optimization system for large elliptical frozen orbit satellite on-orbiting according to claim 6, characterized in that, Module M3 includes: Module M3.1: determining the search range of the to-be-optimized variable; Module M3.2: constructing and solving the Lambert problem from the initial orbit to the target orbit in the search range to obtain the pulse velocity vectors Δv1 and Δv2 required by 2-pulse non-apogee orbit transfer, i.e. the orbit transfer velocity increment; Module M3.3: traversing all values in the search range, selecting the corresponding value with the minimum Δv = ||Δv1|| + ||Δv2||, and generating the pulse orbit transfer strategy with the least fuel consumption for the satellite to transfer from the initial orbit to the target orbit; Module M3.4: The first ignition position is a position on the initial orbit with a position vector length of for applying the first impulse velocity vector Δvl; the second ignition position is a position on the target orbit with a position vector length of for applying the second impulse velocity vector Δv2; Module M3.5: calculating the propellant mass Δm1 required for the satellite to transfer from the initial orbit to the intermediate orbit, and the propellant mass Δm2 required for the satellite to transfer from the intermediate orbit to the target orbit, the calculation formula being as follows: Δm1 = m0(1 - exp(-Δv1 / I sp / g)) Am2 = (m0 - Am1) (1 - exp(-Av2 / I sp / g)) where m0represents the takeoff mass, I sp represents the engine specific impulse, and g represents the gravitational acceleration constant; Module M3.6: calculating the working time dt1 required for the engine to transfer from the initial orbit to the intermediate orbit, and the working time dt2 required for the engine to transfer from the intermediate orbit to the target orbit, the calculation formula being as follows: dt1 = Δm1 / (F / I sp / g) dt2 = Δm2 / (F / I sp / g) Wherein, F represents the engine thrust; Module M3.7: obtaining the number of times N1 of orbit transfer required for the satellite to transfer from the initial orbit to the intermediate orbit according to the working time dt1 required for the engine to transfer from the initial orbit to the intermediate orbit, and obtaining the number of times N2 of orbit transfer required for the satellite to transfer from the intermediate orbit to the target orbit according to the working time dt2 required for the engine to transfer from the intermediate orbit to the target orbit; wherein the average engine operating time is controlled to be within a target T Eng wherein the target wherein T max is the maximum single engine firing time.

8. The orbit transfer strategy optimization system for large elliptical frozen orbit satellite on-orbiting according to claim 7, characterized in that, The engine start and stop conditions of each orbit change are determined by the semi-major axis variation Δa caused by each orbit change i The impulse velocity vector and the orbit parameters before ignition are determined jointly; The number of drift orbits between two adjacent times of orbit transfer is determined by the geographical longitude at the time of satellite-rocket separation and the longitude drift rate.

9. The orbit transfer strategy optimization system for large elliptical frozen orbit satellite on-orbiting according to claim 6, characterized in that, the initial value of the ignition duration Δt i The calculation formula is as follows: Δt i = dt * Δa i / da Wherein, dt represents the required engine working time of the Lambert orbit transfer impulse velocity corresponding to the current ignition, such as the position of the current ignition is consistent with the impulse velocity vector Δv1, then dt = dt1; such as the position of the current ignition is consistent with the impulse velocity vector Δv2, then dt = dt2; da represents the semi-major axis change caused by the Lambert orbit transfer impulse velocity vector, when dt = dt1, da represents the difference between the semi-major axis of the intermediate orbit and the initial orbit; when dt = dt2, da is the difference between the semi-major axis of the target orbit and the intermediate orbit; Δa i represents the semi-major axis change caused by the i-th orbit transfer, i represents the ignition number; Initial value of ignition timing t i The calculation formula is as follows: t i = t * - Δt i / 2 wherein t * denotes the instant corresponding to the position of the spark plug Initial value α of the thrust direction i The value is determined based on the pulse velocity vector; if the location of the current ignition coincides with the pulse velocity vector Δv1, then α... i =Δv1 / ||Δv1||; If the location of the current ignition coincides with the pulse velocity vector Δv2, then α i =Δv2 / ||Δv2||.

10. The orbit transfer strategy optimization system for large elliptical frozen orbit satellite on-orbiting according to claim 6, characterized in that, Module M6 includes: Module M6.1 : defining the ignition time length Δt i , the ignition time t i and the thrust direction a i as the parameters to be optimized, defining an optimization objective function G(X) as follows: where Δm i represents the propellant required for the i-th orbit transfer, Δλ Ω represents the difference between the orbital ascending node geographical longitude after the orbit transfer and the target orbital ascending node geographical longitude, β is a weighting factor, and N1+N2 represents the number of orbit transfers required to complete the orbit transfer. Module M6.2: according to the initial values of the ignition duration, the ignition time and the thrust direction of each time of orbit transfer, a general numerical optimization algorithm is used to minimize the optimization objective function G(X) so as to determine the ignition duration, the ignition time and the thrust direction of each time of orbit transfer.

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