Planning and correction fused online re-planning method under low-orbit task dynamic fault

Abstract

The invention discloses a planning and correction fused online re-planning method under a low-orbit task dynamic fault. The method comprises the following steps: solving a fuel-saving trajectory planning problem of an original target orbit; judging whether a target orbit of the original low-orbit transportation task can be entered or not; if not, solving the reachable orbit maximization problem in the current orbit plane; judging whether the highest reachable orbit is greater than the perigee height of the original target orbit; if yes, solving an orbital plane similarity problem; judging whether the track surface meets task constraints or not; and if yes, solving an in-plane orbit shape maximization task planning problem of the original target orbit, and completing target orbit parameter planning. According to the method, for a rocket executing a low-orbit transportation task, in a hierarchical task planning solution framework, a trajectory correction strategy with key orbit parameters as iteration quantities is introduced, direct method planning solution precision effective compensation and conservative degree weakening represented by convex programming are achieved, and the original target orbit approximation degree and the task reconstruction success rate are improved.

Description

Technical Field

[0001] The present invention belongs to the technical field of aircraft dynamics and control, and particularly relates to an online replanning method for low-orbit missions under power failures by integrating planning and correction. Background Art

[0002] Low-orbit transportation missions usually use launch vehicles to transport payloads from the ground to near-Earth elliptical orbits or circular orbits. During the flight process, the vehicle needs to continuously overcome the Earth's gravity through the action of the power system to reach the required position and speed. If the power system fails and the thrust decreases, and the vehicle fails to respond in time, then under the action of the Earth's gravity, there will be a risk that the payload cannot enter the orbit or even the mission fails. Therefore, for low-orbit transportation missions, it is of great significance to carry out online mission replanning under power failures.

[0003] Most early studies on mission replanning under power failures used indirect methods for trajectory redesign. The design results had better fuel optimality but poor real-time performance and were difficult to perform online operations. To achieve online mission replanning, on the one hand, in response to this problem, a step-by-step degradation planning and solving framework with orbital altitude, orbital plane, and orbital shape as replanning objectives has gradually emerged, disassembling the problem from a strategy perspective to reduce the complexity of solving. On the other hand, with the development of the primal-dual interior-point algorithm, convex optimization technology with it as the core solving algorithm has the advantage of converging in a finite number of iterative steps when solving specific optimization problems, and thus has gradually been applied to the research on mission replanning under power failures. However, when using direct methods represented by convex optimization technology to solve planning problems, the original trajectory planning problem needs to go through operations such as discretization, linearization, and convexification. Therefore, the solution trajectory corresponding to the processed optimization problem has a certain degree of conservativeness and is difficult to maximize the utility of the remaining vehicle carrying capacity. Summary of the Invention

[0004] The technical problem to be solved by the present invention is to overcome the deficiencies of the prior art and provide an online replanning method for low-orbit missions under power failures by integrating planning and correction. For a rocket performing a low-orbit transportation mission, in the hierarchical mission planning and solving framework, a trajectory correction strategy with key orbital parameters as iteration quantities is introduced to effectively compensate the planning and solving accuracy of direct methods represented by convex programming and weaken the conservativeness, maximize the utility of the vehicle carrying capacity, and improve the approximation degree of the original target orbit and the mission reconstruction success rate.

[0005] To solve the above technical problem, the present invention discloses an online replanning method for low-orbit missions under power failures by integrating planning and correction, including:

[0006] Step 1, taking the flight time and thrust vector as optimization variables, and solving the fuel-optimal trajectory planning problem of the original target orbit;

[0007] Step 2: Based on the solution result of the original target orbit fuel-optimal trajectory planning problem, determine whether it is possible to enter the target orbit of the original low-orbit transportation mission; if yes, keep the target orbit unchanged and proceed to Step 8; if not, proceed to Step 3;

[0008] Step 3: Take the thrust vector as the optimization variable and solve the problem of maximizing the reachable orbit within the current orbital plane;

[0009] Step 4: Based on the solution result of the problem of maximizing the reachable orbit within the current orbital plane, determine whether the highest reachable orbit is greater than the perigee altitude of the original target orbit; if yes, proceed to Step 5; if not, take the corrected orbit as the target orbit and proceed to Step 8;

[0010] Step 5: Take the thrust vector as the optimization variable and solve the problem of orbital plane similarity;

[0011] Step 6: Based on the solution result of the problem of orbital plane similarity, determine whether the orbital plane meets the mission constraints; if it meets, proceed to Step 7; if not, take the corrected orbit as the target orbit and proceed to Step 8;

[0012] Step 7: Take the thrust vector as the optimization variable and solve the problem of maximizing the orbital shape within the original target orbital plane for mission planning, take the corrected orbit as the target orbit, and proceed to Step 8;

[0013] Step 8: Implement iterative guidance according to the traditional iterative guidance solution method until the flight ends.

[0014] In the above online replanning method for low-orbit mission dynamic faults with integrated planning and correction, taking the flight time and thrust vector as optimization variables to solve the original target orbit fuel-optimal trajectory planning problem includes:

[0015] Taking the flight time and thrust vector as optimization variables, solve the following original target orbit fuel-optimal trajectory planning problem:

[0016] min - m(t0 + t f )

[0017]

[0018] where F represents the total thrust of the carrier, I sp represents the specific impulse of the engine, t0 represents the moment when the dynamic fault occurs, t fDenote the time of flight as \(t\), the thrust vector as \(F(t)\), the semi-major axis of the original target orbit as \(a_1\), the eccentricity of the original target orbit as \(e_1\), the Earth's gravitational constant as \(\mu\), the semi-focal length of the original target orbit as \(c_1\), the semi-minor axis of the original target orbit as \(b_1\), the flight time as \(t\), the position vector of the carrier at time \(t\) as \(r(t)\), the geocentric distance of the carrier at time \(t\) as \(r(t)\), the velocity vector of the carrier at time \(t\) as \(v(t)\), the mass of the carrier at time \(t\) as \(m(t)\), \(m(t_0 + t f ) represents the mass of the carrier at time \(t_0 + t f ; \(r(t_0 + t f ) represents the position vector of the carrier at time \(t_0 + t f ; \(v(t_0 + t f ) represents the velocity vector of the carrier at time \(t_0 + t f ; \(R go1 represents the transfer matrix between the launch inertial frame and the perifocal coordinate system of the original target orbit; \(r o (t_0 + t f ) represents the position vector of the carrier in the orbital coordinate system at time \(t_0 + t f ; \(r o,x (t_0 + t f ), \(r o,y (t_0 + t f ) and \(r o,z (t_0 + t f ) are the three-axis components of the vector \(r o (t_0 + t f ) respectively; \(v o (t_0 + t f ) represents the velocity vector of the carrier in the orbital coordinate system at time \(t_0 + t f ; \(v o,x (t_0 + t f ), \(v o,y (t_0 + t f ) and \(v o,z (t_0 + t f ) are the three-axis components of the vector \(v o (t_0 + t f ) respectively;

[0019] Obtain the optimized time of flight and the remaining mass

[0020] In the above online replanning method for low-orbit mission dynamic failures integrating planning and correction, the calculation formula for the transfer matrix \(R go1 between the launch inertial frame and the perifocal coordinate system of the original target orbit is as follows:

[0021] R go1 = R z (ω1)Rx (i1)R z (Ω1)R z (90° - λ0)R x (-B0)R y (90° + A0)

[0022]

[0023] Among them, ω1 represents the argument of perigee of the original target orbit, i1 represents the inclination of the original target orbit, Ω1 represents the right ascension of the ascending node of the original target orbit, A0 represents the azimuth of the launch point of the original low - earth orbit transportation mission, λ0 represents the longitude of the launch point of the original low - earth orbit transportation mission, B0 represents the latitude of the launch point of the original low - earth orbit transportation mission, R z (ω1) represents the matrix formed by rotating an angle ω1 around the z - axis, R x (i1) represents the matrix formed by rotating an angle i1 around the x - axis, R z (Ω1) represents the matrix formed by rotating an angle Ω1 around the z - axis, R z (90° - λ0) represents the matrix formed by rotating an angle (90° - λ0) around the z - axis, R x (-B0) represents the matrix formed by rotating an angle (-B0) around the x - axis, R y (90° + A0) represents the matrix formed by rotating an angle (90° + A0) around the y - axis.

[0024] In the above - mentioned online replanning method under the dynamic failure of the low - earth orbit mission with the integration of planning and correction, based on the solution result of the original target orbit fuel - optimal trajectory planning problem, it is judged whether it can enter the target orbit of the original low - earth orbit transportation mission, including:

[0025] If Satisfies Then it is determined that it can enter the target orbit of the original low - earth orbit transportation mission, keep the target orbit unchanged, and enter step 8;

[0026] If Does not satisfy Then it is determined that it cannot enter the target orbit of the original low - earth orbit transportation mission, and enter step 3;

[0027] Among them, m f Represents the mass of the carrier when the fuel is exhausted.

[0028] In the above - mentioned online replanning method under the dynamic failure of the low - earth orbit mission with the integration of planning and correction, with the thrust vector as the optimization variable, the problem of maximizing the reachable orbit within the current orbital plane is solved, including:

[0029] Calculate the parameters of the carrier's flight orbital plane at the current moment, including: the orbital inclination i2 and the right ascension of the ascending node Ω2 at the current moment:

[0030] Ω2 = arccos(N·u x )

[0031] h = R z (90° - λ0)R x (-B0)R y (90° + A0)(r(t0) × v(t0))

[0032]

[0033] u z =

[100] T ,u z =

[001] T

[0034] where h represents the instantaneous angular momentum vector, N represents the unit vector of the instantaneous orbital node line, u x represents the unit vector in the x - direction of the geocentric coordinate system, u z represents the unit vector in the z - direction of the geocentric coordinate system, r(t0) represents the position vector of the vehicle at the time of power failure, and v(t0) represents the velocity vector of the vehicle at the time of power failure;

[0035] The transfer matrix R between the launch inertial system and the near - focus coordinate system of the instantaneous orbital plane is calculated as go2 :

[0036] R go2 = R z (ω1)R x (i2)R z (Ω2)R z (90° - λ0)R x (-B0)R y (90° + A0)

[0037]

[0038] where R x (i2) represents the matrix formed by rotating by an angle i2 about the x - axis, and R z (Ω2) represents the matrix formed by rotating by an angle Ω2 about the z - axis;

[0039] Taking the thrust vector as the optimization variable, solve the following problem of maximizing the reachable orbit within the current orbital plane:

[0040]

[0041] where r(t0 + t f ) represents the geocentric distance of the vehicle at time t0 + t f and v(t0 + t f) represents the vehicle speed value at t0 + t f and v o (t0 + t f ) represents the magnitude of the velocity vector of the vehicle in the vehicle orbit system at t0 + t f and r o (t0 + t f ) represents the magnitude of the position vector of the vehicle in the vehicle orbit system at t0 + t f ;

[0042] The optimized terminal position vector r * (t0 + t f ) and the terminal velocity vector v * (t0 + t f ) are obtained.

[0043] In the above online replanning method for low - orbit mission power failures with the integration of planning and correction, based on the solution result of the maximum reachable orbit problem in the current orbital plane, it is judged whether the highest reachable orbit is greater than the perigee altitude of the original target orbit, including:

[0044] S41, taking as the first predicted target orbit parameters, and according to the large - step guidance period T1, the k - th large - step iterative guidance simulation calculation is completed to obtain the first predicted shutdown point mass where represents the semi - major axis of the first predicted target orbit, represents the eccentricity of the first predicted target orbit, represents the orbital inclination of the first predicted target orbit, represents the right ascension of the ascending node of the first predicted target orbit, represents the corrected orbit altitude of the (k - 1) - th time, R E represents the radius of the earth; for the first iteration, there is representing the initial value of the corrected orbit altitude;

[0045] S42, the corrected orbit altitude of the k - th time is obtained by solving according to the first predicted shutdown point mass

[0046]

[0047] where represents the partial derivative of the predicted orbit altitude ;

[0048] S43, it is judged whether the iteration is completed according to the following formula:

[0049]

[0050] Among them, if holds, the iteration terminates and step S44 is entered; if does not hold, then the value of k is incremented by 1 and the process returns to step S41; represents the corrected orbit altitude at the k-th time the corresponding predicted shutdown point mass, and ξ represents the given mass tolerance;

[0051] S44. Calculate the target orbit parameters after one update:

[0052]

[0053] Among them, represents the semi-major axis of the target orbit after one update, represents the eccentricity of the target orbit after one update, represents the orbital inclination of the target orbit after one update, represents the right ascension of the ascending node of the target orbit after one update;

[0054] S45. If satisfies then it is determined that the highest achievable orbit is greater than the perigee altitude of the original target orbit, and step 5 is entered; if does not satisfy then it is determined that the highest achievable orbit is not greater than the perigee altitude of the original target orbit, and the corrected orbit determined based on the target orbit parameters after one update is used as the target orbit, and step 8 is entered.

[0055] In the above online replanning method under the dynamic fault of low-orbit missions with the integration of planning and correction, taking the thrust vector as the optimization variable, the orbital plane similarity problem is solved, including:

[0056] Taking the thrust vector as the optimization variable, solve the following orbital plane similarity problem:

[0057] min|r o,z (t0 + t f )| + |v o,z (t0 + t f )

[0058]

[0059] Obtain the terminal position vector r ** (t0 + t f ) and the terminal velocity vector v ** (t0 + t f ).

[0060] In the above online replanning method under the dynamic fault of low-earth orbit missions with the integration of planning and correction, based on the solution results of the orbital plane similarity problem, it is judged whether the orbital plane meets the mission constraints, including:

[0061] S61, taking as the second predicted target orbit parameters, according to the large-step guidance period T1, complete the large-step iterative guidance simulation calculation for the k-th time, and obtain the second predicted shutdown point mass where, represents the semi-major axis of the second predicted target orbit, represents the eccentricity of the second predicted target orbit, represents the orbital inclination of the second predicted target orbit, represents the corrected orbital inclination at the (k - 1)-th time; represents the right ascension of the ascending node of the second predicted target orbit, represents the unit vector of the orbital node line planned, represents the planned angular momentum vector, For the first iteration, there is represents the initial value of the corrected orbital inclination;

[0062] S62, according to the second predicted shutdown point mass solve to obtain the corrected orbital inclination at the k-th time

[0063]

[0064] where, represents the partial derivative of the predicted orbital inclination ;

[0065] S63, judge whether to complete the iteration according to the following formula:

[0066]

[0067] where, if holds, the iteration terminates and enters step S64; if does not hold, increase the value of k by 1 and then return to step S61; represents the corrected orbital inclination at the k-th time corresponding to the predicted shutdown point mass;

[0068] S64, calculate the target orbit parameters after the second update:

[0069]

[0070] Among them, represents the semi-major axis of the target orbit after the second update, represents the eccentricity of the target orbit after the second update, represents the inclination of the target orbit after the second update, represents the right ascension of the ascending node of the target orbit after the second update;

[0071] S65, if satisfies then it is determined that the maximum achievable orbit inclination is greater than the original target orbit inclination, and proceed to step S66; if does not satisfy then it is determined that the maximum achievable orbit inclination is not greater than the original target orbit inclination, that is, the orbital plane does not meet the mission constraints. Take the corrected orbit determined based on the target orbit parameters after the second update as the target orbit, and proceed to step 8;

[0072] S66, use as the third predicted target orbit parameters, and according to the large-step guidance period T1, complete the k-th large-step iterative guidance simulation calculation to obtain the third predicted shutdown point mass Among them, represents the semi-major axis of the third predicted target orbit, represents the eccentricity of the third predicted target orbit, represents the inclination of the third predicted target orbit, represents the right ascension of the ascending node of the third predicted target orbit, represents the corrected right ascension of the ascending node at the (k - 1)-th time; for the first iteration, there is represents the initial value of the corrected right ascension of the ascending node;

[0073] S67, according to the third predicted shutdown point mass solve to obtain the corrected right ascension of the ascending node at the k-th time

[0074]

[0075] Among them, represents the partial derivative of the predicted right ascension of the ascending node ;

[0076] S68, judge whether the iteration is completed according to the following formula:

[0077]

[0078] Among them, if holds, the iteration terminates and step S69 is entered; if does not hold, then the value of k is incremented by 1 and the process returns to step S66; represents the corrected right ascension of the ascending node at the k-th time corresponding predicted shutdown point mass;

[0079] S69. Calculate the target orbit parameters after three updates:

[0080]

[0081] Among them, represents the semi-major axis of the target orbit after three updates, represents the eccentricity of the target orbit after three updates, represents the inclination of the target orbit after three updates, represents the right ascension of the ascending node of the target orbit after three updates;

[0082] S610. If satisfies then it is determined that the reachable maximum right ascension of the ascending node is less than the right ascension of the ascending node of the original target orbit, that is, the orbital plane meets the mission constraints, and step 7 is entered; if does not satisfy then it is determined that the reachable maximum right ascension of the ascending node is not less than the right ascension of the ascending node of the original target orbit, that is, the orbital plane does not meet the mission constraints. Take the corrected orbit determined based on the target orbit parameters after three updates as the target orbit, and enter step 8.

[0083] In the above online replanning method under the dynamic failure of low-orbit missions with the integration of planning and correction, taking the thrust vector as the optimization variable, solve the problem of maximizing the orbital shape within the original target orbital plane for mission planning, and conduct target orbit parameter planning, including:

[0084] S71. Taking the thrust vector as the optimization variable, solve the problem of maximizing the orbital shape within the original target orbital plane for mission planning:

[0085]

[0086] Obtain the terminal position vector r *** (t0 + t f ) and the terminal velocity vector v *** (t0 + t f ) after four optimizations;

[0087] S72. Taking as the fourth predicted target orbit parameters, according to the large-step guidance period T1, complete the k-th large-step iterative guidance simulation calculation to obtain the fourth predicted shutdown point mass Among them, represents the semi-major axis of the fourth predicted target orbit, represents the corrected semi-major axis of the (k - 1)-th orbit; represents the eccentricity of the fourth predicted target orbit, represents the orbital inclination of the fourth predicted target orbit, represents the right ascension of the ascending node of the fourth predicted target orbit, represents the argument of perigee of the fourth predicted target orbit, For the first iteration, there is represents the initial value of the corrected semi-major axis;

[0088] S73. According to the mass at the fourth predicted shutdown point solve to obtain the corrected semi-major axis of the k-th orbit

[0089]

[0090] Among them, represents the partial derivative of the predicted semi-major axis ;

[0091] S74. Judge whether the iteration is completed according to the following formula:

[0092]

[0093] Among them, if holds, the iteration terminates and proceeds to step S75; if does not hold, increment the value of k by 1 and return to step S72; represents the corrected semi-major axis of the k-th orbit corresponding to the mass at the predicted shutdown point;

[0094] S75. Calculate the target orbit parameters after four updates:

[0095]

[0096] Among them, represents the target orbit semi-major axis after four updates, represents the target orbit eccentricity after four updates, represents the target orbit inclination after four updates, represents the right ascension of the ascending node of the target orbit after four updates, represents the argument of perigee of the target orbit after four updates;

[0097] S76. Take the corrected orbit determined based on the target orbit parameters after four updates as the target orbit, and proceed to step 8.

[0098] In the above online replanning method for low-orbit mission power failures with integrated planning and correction, it further includes:

[0099] Obtain the overall vehicle parameters when a power failure occurs, including: F, I sp , t0, and m f ;

[0100] Obtain the flight state parameters of the vehicle when a power failure occurs, including: m(t0), r(t0), and v(t0); where m(t0) represents the mass of the vehicle when a power failure occurs.

[0101] Obtain the launch parameters of the original low-orbit transportation mission, including: A0, λ0, and B0;

[0102] Obtain the target orbit parameters of the original low-orbit transportation mission, including: a1, e1, i1, Ω1, and ω1.

[0103] The present invention has the following advantages:

[0104] (1) The present invention discloses an online replanning method for low-orbit mission power failures with integrated planning and correction, which adopts an online mission reconstruction strategy that integrates hierarchical planning and customized trajectory correction design. Based on the principle of maximizing target orbit similarity, a four-level mission planning problem is constructed according to different degrees of power failures. Each level of the mission planning problem is solved by a fusion design method of convex programming and customized iterative guidance. The constructed strategy can make full use of the finite-step convergence advantages of convex programming and iterative guidance technologies to ensure the online realizability of mission reconstruction.

[0105] (2) The present invention discloses an online replanning method for low-orbit mission power failures with integrated planning and correction, which adopts a trajectory correction strategy for iteratively key orbit parameters in the case of mission degradation. For two types of mission planning problems (maximum similarity of orbital planes and maximum similarity of orbital shapes) with mission degradation but sufficient vehicle carrying capacity, the orbital plane characteristic parameters and the apogee altitude are respectively used as the iteration quantities. Through large-step motion prediction, the planning solution accuracy of the original mission planning problem is gradually corrected, effectively improving the degree of target orbit similarity.

[0106] (3) The present invention discloses an online replanning method for low-orbit mission dynamic faults integrating planning and correction, which adopts a rescue trajectory planning strategy under near-minimum orbit injection capability. For the problem of maximizing the planning of accessible orbits with mission degradation and limited launch vehicle capacity, a prediction-correction technique with orbital altitude as the iteration quantity is adopted to reduce the conservativeness of direct method planning and solution, maximize the utility of launch vehicle capacity, and significantly improve the mission reconstruction success rate under the condition of near-minimum orbit injection capability. Description of the Drawings

[0107] Figure 1 is a flowchart of an online replanning method for low-orbit mission dynamic faults integrating planning and correction in an embodiment of the present invention. Detailed Embodiment

[0108] To make the objectives, technical solutions and advantages of the present invention clearer, the following will further describe in detail the disclosed embodiments of the present invention with reference to the drawings.

[0109] Refer to Figure 1 , in this embodiment, the online replanning method for low-orbit mission dynamic faults integrating planning and correction includes:

[0110] Step 0, data preparation.

[0111] In this embodiment, the data to be prepared includes but is not limited to:

[0112] Obtain the overall parameters of the launch vehicle when a dynamic fault occurs, including: the total thrust F of the launch vehicle, the specific impulse I sp of the engine, the time t0 when the dynamic fault occurs, and the mass m f of the launch vehicle at fuel exhaustion, etc.

[0113] Obtain the flight state parameters of the launch vehicle when a dynamic fault occurs, including: the mass m(t0) of the launch vehicle when the dynamic fault occurs, the position vector r(t0) of the launch vehicle when the dynamic fault occurs, and the velocity vector v(t0) of the launch vehicle when the dynamic fault occurs, etc.

[0114] Obtain the launch parameters of the original low-orbit transportation mission, including: the launch point azimuth A0 of the original low-orbit transportation mission, the longitude λ0 of the launch point of the original low-orbit transportation mission, and the latitude B0 of the launch point of the original low-orbit transportation mission, etc.

[0115] Obtain the target orbit parameters of the original low-orbit transportation mission, including: the semi-major axis a1 of the original target orbit, the eccentricity e1 of the original target orbit, the orbital inclination i1 of the original target orbit, the right ascension of the ascending node Ω1 of the original target orbit, and the argument of perigee ω1 of the original target orbit, etc.

[0116] Determine the transfer matrix R go1 between the launch inertial system and the near-focus coordinate system of the original target orbit:

[0117] R go1 = R z (ω1)R x (i1)R z (Ω1)R z (90° - λ0)R x (-B0)R y (90° + A0)

[0118]

[0119] where R z (ω1) represents the matrix formed by rotating an angle ω1 about the z-axis, R x (i1) represents the matrix formed by rotating an angle i1 about the x-axis, R z (Ω1) represents the matrix formed by rotating an angle Ω1 about the z-axis, R z (90° - λ0) represents the matrix formed by rotating an angle (90° - λ0) about the z-axis, R x (-B0) represents the matrix formed by rotating an angle (-B0) about the x-axis, R y (90° + A0) represents the matrix formed by rotating an angle (90° + A0) about the y-axis.

[0120] Step 1: Taking the flight time and thrust vector as optimization variables, solve the problem of the original target orbit fuel-optimal trajectory planning.

[0121] In this embodiment, taking the flight time and thrust vector as optimization variables, by solving the following original target orbit fuel-optimal trajectory planning problem:

[0122] min - m(t0 + t f )

[0123]

[0124] where t f represents the flight time, F(t) represents the thrust vector, μ represents the Earth's gravitational constant, c1 represents the semi-focal length of the original target orbit, b1 represents the semi-minor axis of the original target orbit, t represents the flight time, r(t) represents the position vector of the carrier at time t, r(t) represents the geocentric distance of the carrier at time t, v(t) represents the velocity vector of the carrier at time t, m(t) represents the mass of the carrier at time t, m(t0 + t f ) represents the mass of the carrier at time t0 + t f r(t0 + t f ) represents the position vector of the carrier at time t0 + t f v(t0 + t f ) represents the velocity vector of the carrier at time t0 + t f Rgo1 represents the transfer matrix between the launch inertial frame and the perifocal coordinate system of the original target orbit; r o (t0 + t f ) represents the position vector in the vehicle orbit system at time t0 + t f ; r o,x (t0 + t f ), r o,y (t0 + t f ) and r o,z (t0 + t f ) are the three-axis components of the vector r o (t0 + t f ) respectively; v o (t0 + t f ) represents the velocity vector in the vehicle orbit system at time t0 + t f ; v o,x (t0 + t f ), v o,y (t0 + t f ) and v o,z (t0 + t f ) are the three-axis components of the vector v o (t0 + t f ) respectively.

[0125] Then the optimized flight time and the remaining mass

[0126] Step 2: Based on the solution of the original target orbit fuel-optimal trajectory planning problem, determine whether it is possible to enter the target orbit of the original low-earth orbit transportation mission; if yes, keep the target orbit unchanged and go to Step 8; if not, go to Step 3.

[0127] In this embodiment, if satisfies then it is determined that it is possible to enter the target orbit of the original low-earth orbit transportation mission, keep the target orbit unchanged, and go to Step 8; if does not satisfy then it is determined that it is not possible to enter the target orbit of the original low-earth orbit transportation mission, and go to Step 3.

[0128] Step 3: Take the thrust vector as the optimization variable and solve the problem of maximizing the reachable orbit in the current orbital plane.

[0129] In this embodiment, first, calculate the parameters of the vehicle flight orbital plane at the current moment, including: the orbital inclination i2 and the right ascension of the ascending node Ω2 at the current moment:

[0130] Ω2 = arccos(N · u x )

[0131] h = R z (90° - λ0)R x (-B0)R y (90° + A0)(r(t0) × v(t0))

[0132]

[0133] u z =

[100] T ,u z =

[001] T

[0134] Among them, h represents the instantaneous angular momentum vector, N represents the unit vector of the instantaneous orbital node line, u x represents the unit vector in the x - direction of the geocentric coordinate system, u z represents the unit vector in the z - direction of the geocentric coordinate system, r(t0) represents the position vector of the vehicle at the time of power failure, and v(t0) represents the velocity vector of the vehicle at the time of power failure.

[0135] Furthermore, the transformation matrix R between the launch inertial system and the near - focus coordinate system of the instantaneous orbital plane is calculated as follows go2 :

[0136] R go2 = R z (ω1)R x (i2)R z (Ω2)R z (90° - λ0)R x (-B0)R y (90° + A0)

[0137]

[0138] Among them, R x (i2) represents the matrix formed by rotating by an angle i2 around the x - axis, R z (Ω2) represents the matrix formed by rotating by an angle Ω2 around the z - axis.

[0139] Taking the thrust vector as the optimization variable, by solving the following problem of maximizing the reachable orbit within the current orbital plane:

[0140]

[0141] Among them, r(t0 + t f ) represents the geocentric distance of the vehicle at time t0 + t f , v(t0 + t f ) represents the velocity value of the vehicle at time t0 + t f , v o (t0 + tf ) represents the magnitude of the velocity vector in the vehicle orbit system at time t0 + t f , and r o (t0 + t f ) represents the magnitude of the position vector in the vehicle orbit system at time t0 + t f .

[0142] Then the terminal position vector r * (t0 + t f ) and the terminal velocity vector v * (t0 + t f ) can be obtained after the second optimization.

[0143] Step 4: Based on the solution result of the maximum reachable orbit problem in the current orbital plane, determine whether the highest reachable orbit is greater than the perigee altitude of the original target orbit; if so, proceed to Step 5; if not, use the corrected orbit as the target orbit and proceed to Step 8.

[0144] In this embodiment, the following method can be used to determine whether the highest reachable orbit is greater than the perigee altitude of the original target orbit:

[0145] S41: Using as the first predicted target orbit parameters, according to the large-step guidance period T1, complete the k-th large-step iterative guidance simulation calculation to obtain the first predicted shutdown point mass where represents the semi-major axis of the first predicted target orbit, represents the eccentricity of the first predicted target orbit, represents the orbital inclination of the first predicted target orbit, represents the right ascension of the ascending node of the first predicted target orbit, represents the corrected orbit altitude at the (k - 1)-th time, and R E represents the radius of the Earth; for the first iteration, there is representing the initial value of the corrected orbit altitude.

[0146] S42: According to the first predicted shutdown point mass solve to obtain the corrected orbit altitude at the k-th time

[0147]

[0148] where represents the partial derivative of the predicted orbit altitude .

[0149] S43. Determine whether to complete the iteration according to the following formula:

[0150]

[0151] where, if holds, the iteration terminates and proceeds to step S44; if does not hold, increment the value of k by 1 and return to step S41; represents the corrected orbit altitude at the k-th time the corresponding predicted shutdown point mass, and ξ represents the given mass allowable error.

[0152] S44. Calculate the target orbit parameters after one update:

[0153]

[0154] where, represents the semi-major axis of the target orbit after one update, represents the eccentricity of the target orbit after one update, represents the orbital inclination of the target orbit after one update, represents the right ascension of the ascending node of the target orbit after one update.

[0155] S45. If satisfies then it is determined that the highest achievable orbit is greater than the perigee altitude of the original target orbit, and proceed to step 5; if does not satisfy then it is determined that the highest achievable orbit is not greater than the perigee altitude of the original target orbit, take the corrected orbit determined based on the target orbit parameters after one update as the target orbit, and proceed to step 8.

[0156] Step 5. Solve the orbital plane similarity problem with the thrust vector as the optimization variable.

[0157] In this embodiment, with the thrust vector as the optimization variable, by solving the following orbital plane similarity problem:

[0158] min|r o,z (t0 + t f )| + |v o,z (t0 + t f )|

[0159]

[0160] Then the terminal position vector r ** (t0 + t f ) and the terminal velocity vector v ** (t0 + t f ) can be obtained.

[0161] Step 6: Based on the solution result of the orbital plane similarity problem, determine whether the orbital plane meets the mission constraints; if it does, proceed to Step 7; if not, use the corrected orbit as the target orbit and proceed to Step 8.

[0162] In this embodiment, the following method can be used to determine whether the orbital plane meets the mission constraints:

[0163] S61: Take as the second predicted target orbit parameters, and according to the large-step guidance period T1, complete the k-th large-step iterative guidance simulation calculation to obtain the second predicted shutdown point mass where represents the semi-major axis of the second predicted target orbit, represents the eccentricity of the second predicted target orbit, represents the orbital inclination of the second predicted target orbit, represents the corrected orbital inclination at the (k - 1)-th time; represents the right ascension of the ascending node of the second predicted target orbit, represents the planned orbital node unit vector, represents the planned angular momentum vector, For the first iteration, there is represents the initial value of the corrected orbital inclination.

[0164] S62: According to the second predicted shutdown point mass solve to obtain the corrected orbital inclination at the k-th time

[0165]

[0166] where represents the partial derivative of the predicted orbital inclination with respect to.

[0167] S63: Determine whether the iteration is completed according to the following formula:

[0168]

[0169] where, if holds, the iteration terminates and proceeds to Step S64; if does not hold, increment the value of k by 1 and return to Step S61; represents the corrected orbital inclination at the k-th time The corresponding predicted shutdown point mass.

[0170] S64. Calculate the target orbit parameters after the second update:

[0171]

[0172] Among them, represents the semi-major axis of the target orbit after the second update, represents the eccentricity of the target orbit after the second update, represents the orbital inclination of the target orbit after the second update, represents the right ascension of the ascending node of the target orbit after the second update.

[0173] S65. If satisfies then it is determined that the reachable maximum orbital inclination is greater than the original target orbital inclination, and proceed to step S66; if does not satisfy then it is determined that the reachable maximum orbital inclination is not greater than the original target orbital inclination, that is, the orbital plane does not meet the mission constraints. Take the corrected orbit determined based on the target orbit parameters after the second update as the target orbit, and proceed to step 8.

[0174] S66. Use as the third predicted target orbit parameters. According to the large-step guidance period T1, complete the k-th large-step iterative guidance simulation calculation to obtain the third predicted shutdown point mass Among them, represents the semi-major axis of the third predicted target orbit, represents the eccentricity of the third predicted target orbit, represents the orbital inclination of the third predicted target orbit, represents the right ascension of the ascending node of the third predicted target orbit, represents the corrected right ascension of the ascending node at the (k - 1)-th time; for the first iteration, there is represents the initial value of the corrected right ascension of the ascending node.

[0175] S67. According to the third predicted shutdown point mass solve to obtain the corrected right ascension of the ascending node at the k-th time

[0176]

[0177] Among them, represents the partial derivative of the predicted right ascension of the ascending node with respect to.

[0178] S68, Determine whether to complete the iteration according to the following formula:

[0179]

[0180] where, if holds, the iteration terminates and proceeds to step S69; if does not hold, increment the value of k by 1 and return to step S66; represents the corrected right ascension of the ascending node at the k-th time corresponding predicted shutdown point mass.

[0181] S69, Calculate the target orbit parameters after three updates:

[0182]

[0183] where, represents the semi-major axis of the target orbit after three updates, represents the eccentricity of the target orbit after three updates, represents the inclination of the target orbit after three updates, represents the right ascension of the ascending node of the target orbit after three updates.

[0184] S610, If satisfies then it is determined that the reachable maximum right ascension of the ascending node is less than the right ascension of the ascending node of the original target orbit, that is, the orbital plane meets the mission constraints, and proceed to step 7; if does not satisfy then it is determined that the reachable maximum right ascension of the ascending node is not less than the right ascension of the ascending node of the original target orbit, that is, the orbital plane does not meet the mission constraints. Take the corrected orbit determined based on the target orbit parameters after three updates as the target orbit and proceed to step 8.

[0185] Step 7, Take the thrust vector as the optimization variable and solve the problem of maximizing the orbital shape within the original target orbital plane. Take the corrected orbit as the target orbit and proceed to step 8.

[0186] In this embodiment, the specific implementation manner of step 7 is as follows:

[0187] S71, Take the thrust vector as the optimization variable and solve the problem of maximizing the orbital shape within the original target orbital plane:

[0188]

[0189] Then the terminal position vector r *** (t0 + t f ) and the terminal velocity vector v *** (t0 + tf )。

[0190] S72, using as the fourth predicted target orbit parameter, and according to the large-step guidance period T1, complete the large-step iterative guidance simulation calculation for the k-th time to obtain the mass at the fourth predicted shutdown point wherein, represents the semi-major axis of the fourth predicted target orbit, represents the semi-major axis of the corrected orbit at the (k - 1)-th time; represents the eccentricity of the fourth predicted target orbit, represents the inclination of the fourth predicted target orbit, represents the right ascension of the ascending node of the fourth predicted target orbit, represents the argument of perigee of the fourth predicted target orbit, For the first iteration, there is represents the initial value of the semi-major axis of the corrected orbit.

[0191] S73, according to the mass at the fourth predicted shutdown point solve to obtain the semi-major axis of the corrected orbit at the k-th time

[0192]

[0193] wherein, represents the partial derivative of the predicted orbit semi-major axis with respect to.

[0194] S74, determine whether to complete the iteration according to the following formula:

[0195]

[0196] wherein, if holds, the iteration terminates and proceeds to step S75; if does not hold, increment the value of k by 1 and then return to step S72; represents the semi-major axis of the corrected orbit at the k-th time corresponding to the mass at the predicted shutdown point.

[0197] S75, calculate the target orbit parameters after four updates:

[0198]

[0199] wherein, represents the semi-major axis of the target orbit after four updates, denotes the target orbit eccentricity after four updates, denotes the target orbit inclination after four updates, denotes the right ascension of the ascending node of the target orbit after four updates, denotes the argument of perigee of the target orbit after four updates.

[0200] S76, take the corrected orbit determined based on the target orbit parameters after four updates as the target orbit, and enter step 8.

[0201] Step 8, implement iterative guidance according to the traditional iterative guidance solution method until the flight ends.

[0202] In summary, the present invention proposes an online replanning method for power failure during a low-earth orbit mission that integrates planning and correction for rockets performing low-earth orbit transportation missions. First, obtain the overall vehicle parameters, flight state parameters, launch parameters, and target orbit parameters of the original low-earth orbit transportation mission when a power failure occurs. Then, calculate the transfer matrix between the launch inertial system and the near-focus coordinate system of the original target orbit. Next, solve the problem of planning the trajectory with the least fuel consumption for the original target orbit and determine whether it is possible to enter the target orbit of the original high-earth orbit transportation mission. If not satisfied, calculate the parameters of the current vehicle flight orbital plane, obtain the transfer matrix between the launch inertial system and the near-focus coordinate system of the instantaneous orbital plane, and then solve the problem of maximizing the reachable orbit within the current orbital plane. Next, complete the large-step iterative guidance prediction and orbit height correction with the orbit height as a variable, and determine whether the maximum reachable orbit is greater than the perigee height of the original target orbit. If satisfied, solve the problem of orbital plane similarity, and then complete the large-step iterative guidance prediction and correction of the orbital inclination and the right ascension of the ascending node with the orbital inclination and the right ascension of the ascending node as variables. Next, determine whether the maximum reachable right ascension of the ascending node is less than the right ascension of the ascending node of the original target orbit. If satisfied, solve the problem of maximizing the orbital shape within the original target orbital plane, and then complete the large-step iterative guidance prediction and semi-major axis correction with the semi-major axis as a variable; finally, according to the planning situation, output the planned target orbit parameters and implement iterative guidance until the end. The method of the present invention introduces a trajectory correction technology with key orbital parameters as iteration quantities in the hierarchical task planning and solution framework, realizes effective compensation for the planning and solution accuracy of direct methods represented by convex programming and weakening of the conservatism, maximally exploits the utility of the launch capacity, and improves the approximation degree of the original target orbit and the success rate of mission reconstruction.

[0203] Although the present invention has been disclosed above with the preferred embodiments, it is not intended to limit the present invention. Any person skilled in the art can make possible changes and modifications to the technical solution of the present invention by using the methods and technical contents disclosed above without departing from the spirit and scope of the present invention. Therefore, all contents that do not depart from the technical solution of the present invention, any simple modifications, equivalent changes and decorations made to the above embodiments according to the technical essence of the present invention shall fall within the protection scope of the technical solution of the present invention.

[0204] The content not detailedly described in the specification of the present invention belongs to the well-known technology of those skilled in the art.

Claims

1. An online replanning method for low-earth orbit missions under power failures, which combines planning and correction, is characterized in that including: Step 1: Solve the problem of the most fuel-efficient trajectory planning for the original target orbit with the flight time and thrust vector as the optimization variables. Step 2: Based on the solution result of the problem of the most fuel-efficient trajectory planning for the original target orbit, determine whether it is possible to enter the target orbit of the original low-Earth orbit transportation mission. If yes, keep the target orbit unchanged and proceed to Step 8; if no, proceed to Step 3. Step 3: Solve the problem of maximizing the reachable orbit within the current orbital plane with the thrust vector as the optimization variable. Step 4: Based on the solution result of the problem of maximizing the reachable orbit within the current orbital plane, determine whether the highest reachable orbit is greater than the perigee altitude of the original target orbit. If yes, proceed to Step 5; if no, take the corrected orbit as the target orbit and proceed to Step 8. Step 5: Solve the problem of orbital plane similarity with the thrust vector as the optimization variable. Step 6: Based on the solution result of the problem of orbital plane similarity, determine whether the orbital plane meets the mission constraints; if it meets, proceed to Step 7; if it does not meet, take the corrected orbit as the target orbit and proceed to Step 8. Step 7: Solve the problem of maximizing the orbital shape within the original target orbital plane with the thrust vector as the optimization variable, take the corrected orbit as the target orbit, and proceed to Step 8. Step 8: Implement iterative guidance according to the traditional iterative guidance solution method until the flight ends.

2. The online replanning method under the dynamic fault of a low-earth orbit mission integrating planning and correction according to claim 1, characterized in that, Solving the problem of the most fuel-efficient trajectory planning for the original target orbit with the flight time and thrust vector as the optimization variables includes: Solving the following problem of the most fuel-efficient trajectory planning for the original target orbit with the flight time and thrust vector as the optimization variables: min-m(t0+t f ) Among them, F represents the total thrust of the vehicle, I sp represents the specific impulse of the engine, t0 represents the moment of power failure, t f represents the flight time, F(t) represents the thrust vector, a1 represents the semi-major axis of the original target orbit, e1 represents the eccentricity of the original target orbit, μ represents the gravitational constant of the Earth, c1 represents the semi-focal length of the original target orbit, b1 represents the semi-minor axis of the original target orbit, t represents the flight time, r(t) represents the position vector of the vehicle at time t, r(t) represents the geocentric distance of the vehicle at time t, v(t) represents the velocity vector of the vehicle at time t, m(t) represents the mass of the vehicle at time t, m(t0 + t f ) represents the mass of the vehicle at time t0 + t f , r(t0 + t f ) represents the position vector of the vehicle at time t0 + t f , v(t0 + t f ) represents the velocity vector of the vehicle at time t0 + t f , R go1 represents the transfer matrix between the launch inertial system and the near focus coordinate system of the original target orbit; r o (t0 + t f ) represents the position vector of the vehicle in the vehicle orbit system at time t0 + t f , r o,x (t0 + t f ), r o,y (t0 + t f ) and r o,z (t0 + t f ) are the three-axis components of the vector r o (t0 + t f ) respectively; v o (t0 + t f ) represents the velocity vector of the vehicle in the vehicle orbit system at time t0 + t f , v o,x (t0 + t f ), v o,y (t0 + t f ) and v o,z (t0 + t f ) are the three-axis components of the vector v o (t0 + t f ) respectively; Obtain the optimized time of flight once and the remaining mass 3. The online replanning method under the dynamic fault of a low-orbit mission with integrated planning and correction according to claim 2, characterized in that Transfer matrix R between the launch inertial system and the near focus coordinate system of the original target orbit go1 The calculation formula is as follows: R go1 = R z (ω1)R x (i1)R z (Ω1)R z (90° - λ0)R x (-B0)R y (90° + A0) Among them, ω1 represents the argument of perigee of the original target orbit, i1 represents the orbital inclination of the original target orbit, Ω1 represents the right ascension of the ascending node of the original target orbit, A0 represents the launch azimuth of the original low-Earth orbit transportation mission, λ0 represents the longitude of the launch point of the original low-Earth orbit transportation mission, B0 represents the latitude of the launch point of the original low-Earth orbit transportation mission, R z (ω1) represents the matrix formed by rotating an angle ω1 about the z-axis, R x (i1) represents the matrix formed by rotating an angle i1 about the x-axis, R z (Ω1) represents the matrix formed by rotating an angle Ω1 about the z-axis, R z (90° - λ0) represents the matrix formed by rotating an angle (90° - λ0) about the z-axis, R x (-B0) represents the matrix formed by rotating an angle (-B0) about the x-axis, R y (90° + A0) represents the matrix formed by rotating an angle (90° + A0) about the y-axis.

4. The online replanning method under power failure of low-orbit missions integrating planning and correction according to claim 3, characterized in that Based on the solution result of the problem of the most fuel-efficient trajectory planning for the original target orbit, determining whether it is possible to enter the target orbit of the original low-Earth orbit transportation mission includes: If Meet Then determine the target orbit that can enter the original low-orbit transportation mission, keep the target orbit unchanged, and enter step 8; If not satisfied it is determined that the target orbit of the original low-orbit transportation mission cannot be entered, and step 3 is entered; where m f represents the mass of the vehicle at fuel exhaustion.

5. The online replanning method under power failure of low-orbit missions integrating planning and correction according to claim 4, characterized in that Solving the problem of maximizing the reachable orbit within the current orbital plane with the thrust vector as the optimization variable includes: Calculating the flight orbital plane parameters of the carrier at the current moment, including: the orbital inclination i2 and the right ascension of the ascending node Ω2 at the current moment: Ω2 = arccos(N·u x ) h = R z (90° - λ0)R x (-B0)R y (90° + A0)(r(t0) × v(t0)) u z = [1 0 0] T ,u z = [0 0 1] T where h represents the instantaneous momentum vector, N represents the unit vector of the instantaneous orbital node line, u x represents the unit vector in the x direction of the geocentric coordinate system, u z represents the unit vector in the z direction of the geocentric coordinate system, r(t0) represents the position vector of the vehicle at the time of dynamic failure, and v(t0) represents the velocity vector of the vehicle at the time of dynamic failure; Calculate the transfer matrix R between the launch inertial frame and the near-focus coordinate system of the instantaneous orbital plane go2 : R go2 = R z (ω1)R x (i2)R z (Ω2)R z (90° - λ0)R x (-B0)R y (90° + A0) wherein, R x (i2) represents a matrix formed by rotating an angle i2 about the x-axis, R z (Ω2) represents a matrix formed by rotating an angle Ω2 about the z-axis; Solving the following problem of maximizing the reachable orbit within the current orbital plane with the thrust vector as the optimization variable: Among them, r(t0 + t f ) represents the geocentric distance of the vehicle at the moment of t0 + t f , v(t0 + t f ) represents the speed value of the vehicle at the moment of t0 + t f , v o (t0 + t f ) represents the magnitude of the velocity vector of the vehicle in the orbital system at the moment of t0 + t f , r o (t0 + t f ) represents the magnitude of the position vector of the vehicle in the orbital system at the moment of t0 + t f ; Obtain the terminal position vector r after secondary optimization * (t0 + t f ) and the terminal velocity vector v * (t0 + t f ).

6. The online replanning method under the power failure of a low-earth orbit mission integrating planning and correction according to claim 5, characterized in that Based on the solution result of the problem of maximizing the reachable orbit within the current orbital plane, determining whether the highest reachable orbit is greater than the perigee altitude of the original target orbit includes: S41, taking as the first predicted target orbit parameters, according to the large-step guidance period T1, complete the k-th large-step iterative guidance simulation calculation to obtain the first predicted shutdown point mass where represents the semi-major axis of the first predicted target orbit, represents the eccentricity of the first predicted target orbit, represents the orbital inclination of the first predicted target orbit, represents the right ascension of the ascending node of the first predicted target orbit, represents the corrected orbit altitude at the (k - 1)-th time, R E represents the radius of the Earth; for the first iteration, there is represents the initial value of the corrected orbit altitude; S42, according to the mass at the first predicted shutdown point Solve to obtain the corrected orbit altitude at the k-th time Among them, denotes the partial derivative with respect to the predicted orbital altitude S43: Determine whether the iteration is completed according to the following formula: Among them, if holds, the iteration terminates and step S44 is entered; if does not hold, then the value of k is incremented by 1 and the process returns to step S41; represents the corrected orbit altitude at the k-th time corresponding to the predicted shutdown point mass, and ξ represents the given mass tolerance; S44: Calculate the target orbit parameters after one update: Among them, represents the semi-major axis of the target orbit after one update, represents the eccentricity of the target orbit after one update, represents the orbital inclination of the target orbit after one update, represents the right ascension of the ascending node of the target orbit after one update; S45, if is satisfied then it is determined that the highest achievable orbit is greater than the perigee altitude of the original target orbit, and step 5 is entered; if is not satisfied then it is determined that the highest achievable orbit is not greater than the perigee altitude of the original target orbit, the corrected orbit determined based on the target orbit parameters after one update is taken as the target orbit, and step 8 is entered.

7. The online replanning method under the dynamic failure of a low-earth orbit mission with integrated planning and correction according to claim 6, characterized in that Solving the problem of orbital plane similarity with the thrust vector as the optimization variable includes: Solving the following problem of orbital plane similarity with the thrust vector as the optimization variable: min|r o,z (t0 + t f )| + |v o,z (t0 + t f )| Obtain the terminal position vector r optimized three times ** (t0 + t f ) and the terminal velocity vector v ** (t0 + t f ).

8. The online replanning method under power failure of low-orbit missions integrating planning and correction according to claim 7, characterized in that Based on the solution result of the problem of orbital plane similarity, determining whether the orbital plane meets the mission constraints includes: S61, taking as the second predicted target orbit parameters, according to the large-step guidance period T1, complete the k-th large-step iterative guidance simulation calculation to obtain the second predicted shutdown point mass where represents the semi-major axis of the second predicted target orbit, represents the eccentricity of the second predicted target orbit, represents the orbital inclination of the second predicted target orbit, represents the corrected orbital inclination at the (k - 1)-th time; represents the right ascension of the ascending node of the second predicted target orbit, represents the planned orbit nodal unit vector, represents the planned angular momentum vector, For the first iteration, there is represents the initial value of the corrected orbital inclination; S62, according to the mass at the second predicted shutdown point Solve to obtain the corrected orbital inclination at the k-th time Among them, denotes the partial derivative of the predicted orbital inclination S63: Determine whether the iteration is completed according to the following formula: Among them, if holds, the iteration terminates and proceeds to step S64; if does not hold, increment the value of k by 1 and return to step S61; represents the corrected orbit inclination at the k-th time corresponding predicted mass at the shutdown point; S64: Calculate the target orbit parameters after the second update: Among them, represents the semi-major axis of the target orbit after the second update, represents the eccentricity of the target orbit after the second update, represents the orbital inclination of the target orbit after the second update, represents the right ascension of the ascending node of the target orbit after the second update; S65, if is satisfied then it is determined that the reachable maximum orbit inclination is greater than the original target orbit inclination, and step S66 is entered; if is not satisfied then it is determined that the reachable maximum orbit inclination is not greater than the original target orbit inclination, that is, the orbital plane does not meet the mission constraints. The corrected orbit determined based on the target orbit parameters after the secondary update is used as the target orbit, and step 8 is entered; S66, taking as the third predicted target orbit parameters, according to the large-step guidance period T1, complete the large-step iterative guidance simulation calculation for the k-th time, and obtain the mass at the third predicted shutdown point where represents the semi-major axis of the third predicted target orbit, represents the eccentricity of the third predicted target orbit, represents the inclination of the third predicted target orbit, represents the right ascension of the ascending node of the third predicted target orbit, represents the corrected right ascension of the ascending node at the (k - 1)-th time; for the first iteration, there is represents the initial value of the corrected right ascension of the ascending node; S67, according to the mass at the third predicted shutdown point The right ascension of the ascending node at the k-th correction is obtained by solving Among them, denotes the partial derivative with respect to the predicted right ascension of the ascending node ; S68: Determine whether the iteration is completed according to the following formula: Among them, if holds, the iteration terminates and step S69 is entered; if does not hold, then the value of k is incremented by 1 and the process returns to step S66; represents the corrected right ascension of the ascending node at the k-th time corresponding predicted shutdown point mass; S69: Calculate the target orbit parameters after the third update: Among them, represents the semi-major axis of the target orbit after three updates, represents the eccentricity of the target orbit after three updates, represents the orbital inclination of the target orbit after three updates, represents the right ascension of the ascending node of the target orbit after three updates; S610, if is satisfied then it is determined that the reachable maximum right ascension of the ascending node is less than the right ascension of the ascending node of the original target orbit, that is, the orbital plane meets the mission constraints, and proceed to step 7; if is not satisfied then it is determined that the reachable maximum right ascension of the ascending node is not less than the right ascension of the ascending node of the original target orbit, that is, the orbital plane does not meet the mission constraints. Take the corrected orbit determined based on the target orbit parameters after three updates as the target orbit, and proceed to step 8.

9. The online replanning method under the dynamic failure of a low-earth orbit mission with integrated planning and correction according to claim 8, characterized in that, Solving the problem of maximizing the orbital shape within the original target orbital plane with the thrust vector as the optimization variable and performing target orbit parameter planning includes: S71: Solve the problem of maximizing the orbital shape within the original target orbital plane with the thrust vector as the optimization variable: Obtain the terminal position vector r after four optimizations *** (t0 + t f ) and the terminal velocity vector v *** (t0 + t f ); S72, taking as the fourth predicted target orbit parameter, and according to the large-step guidance period T1, completing the large-step iterative guidance simulation calculation for the k-th time to obtain the fourth predicted shutdown point mass where represents the semi-major axis of the fourth predicted target orbit, represents the semi-major axis of the corrected orbit for the (k - 1)-th time; represents the eccentricity of the fourth predicted target orbit, represents the orbital inclination of the fourth predicted target orbit, represents the right ascension of the ascending node of the fourth predicted target orbit, represents the argument of perigee of the fourth predicted target orbit, For the first iteration, there is representing the initial value of the semi-major axis of the corrected orbit; S73, according to the mass at the fourth predicted shutdown point Solve to obtain the corrected orbital semi-major axis at the k-th time Among them, denotes the partial derivative of the predicted orbital semi-major axis S74, determine whether the iteration is completed according to the following formula: Among them, if holds, the iteration terminates and step S75 is entered; if does not hold, the value of k is incremented by 1 and then step S72 is returned; represents the corrected orbital semi-major axis at the k-th time corresponding predicted mass at shutdown point; S75, calculate the target orbit parameters after four updates: Among them, represents the semi-major axis of the target orbit after four updates, represents the eccentricity of the target orbit after four updates, represents the orbital inclination of the target orbit after four updates, represents the right ascension of the ascending node of the target orbit after four updates, represents the argument of perigee of the target orbit after four updates; S76, take the corrected orbit determined based on the target orbit parameters after four updates as the target orbit, and proceed to step 8.

10. The online replanning method under the dynamic failure of a low-earth orbit mission integrating planning and correction according to claim 9, characterized in that It further includes: Obtain the overall parameters of the vehicle when a power failure occurs, including: F, I sp , t0 and m f ; Obtain the flight state parameters of the carrier vehicle when a power failure occurs, including: m(t0), r(t0), and v(t0); where m(t0) represents the mass of the carrier vehicle when a power failure occurs; Obtain the launch parameters of the original low-earth orbit transportation mission, including: A0, λ0, and B0; Obtain the target orbit parameters of the original low-earth orbit transportation mission, including: a1, e1, i1, Ω1, and ω1.

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