Rapid earth-moon emergency return orbit design method for manned lunar exploration

Abstract

The invention discloses a manned lunar exploration-oriented earth-moon emergency return orbit rapid design method, which comprises the following steps of: establishing a high-precision orbit kinetic equation in a geocentric inertial coordinate system by taking the earth as a central celestial body, and describing the motion state of an aircraft in an earth-moon space; determining a general transfer orbit and a free return orbit of the manned lunar exploration task by adopting a B-plane design method in combination with a high-precision orbit kinetic equation of the aircraft; globally searching an initial value range of a design variable under the condition of comprehensively considering reentry point height and reentry angle constraints by taking a general transfer orbit or a free return orbit as a baseline; and on the basis of the initial value range of the design variable, further considering the reentry range, the reentry inclination angle and the coplanar constraint with the drop point, and determining a final design variable and an emergency return orbit by adopting a differential correction optimization algorithm. The method has an important reference value for rapid design and parameter selection of the emergency return orbit in a manned lunar exploration task.

Description

Technical Field

[0001] The present invention relates to a rapid design method for an Earth-Moon emergency return orbit for manned lunar exploration, and in particular to a rapid design method for an emergency return orbit and a key parameter initial value search method under a complex high-precision dynamic model taking into account multiple engineering constraints, and belongs to the field of spacecraft orbit design and optimization. Background Art

[0002] Lunar and deep space exploration are important exploration areas for human space activities, and they play an incomparable role in future scientific research and resource development. As an important part of the overall mission design, the orbit design and optimization technology of the exploration mission is one of the key technologies in the field of deep space exploration. The optimal transfer orbit is the result of the combined gravitational effect in a multi-celestial body system, and the rapid orbit design faces significant influences and challenges from a complex dynamic environment. At the same time, in the execution of manned lunar exploration missions, once a dangerous situation that endangers the life safety of astronauts occurs during the flight, how to make full use of the spacecraft resources, quickly design an emergency return orbit to implement emergency rescue for astronauts, and ensure the safe return of astronauts to the ground. In this context, it is necessary and essential to study the rapid design strategy of emergency return orbit for manned lunar exploration.

[0003] There are two main types of manned lunar exploration mission orbits: general transfer and free return. How to effectively search and calculate feasible key parameters to design initial values ​​to reduce sensitivity after a spacecraft failure. How to simultaneously meet multiple engineering constraints such as reentry point height, reentry angle, reentry range, and coplanarity with the landing area aiming point, and consider the impact of emergency return on mission fuel consumption and return flight time under general transfer orbit and free return orbit, and quickly design emergency return orbit under high-precision dynamic model is worthy of in-depth research.

[0004] In response to the emergency return problem in manned lunar missions, Lu Lin used the MultuStart algorithm to globally optimize the lunar emergency three-pulse return orbit based on a circular restricted three-body model, but did not verify it in a high-precision model. (Lu Lin, Li Haiyang, Liu Jianghui, Yang Luyi. Optimization design of lunar emergency return orbit [J]. Systems Engineering and Electronics, 2020, 42(2): 420-426). Liu proposed a rapid optimization method for lunar escape energy-saving transfer orbit based on the low energy consumption characteristics of the Earth-Moon space, but there are problems such as long return time, which is not suitable for manned lunar exploration missions (Liu Y, Jing WX, Xiao YY. Finit thrust transfer strategy designing for low energy moon returnbased on GPM / VSM [J]. International Journal of Automation and Computing, 2015, 12(5): 490-496). 。 Based on the manned Earth-Moon hybrid orbit, Lv Pengwei proposed three modes: direct return, immediate maneuverable return around the moon, and maneuverable return after orbiting the moon. He studied the characteristics of emergency return transfer time and velocity increment after failure at different times, but did not fully consider the constraints such as the return landing area (Lv Pengwei, Shang Zhi, Wang Ping. Analysis of one-pulse emergency return orbit of manned Earth-Moon hybrid orbit [J]. Chinese Space Science and Technology, 2020, 40(1): 1-6). Li Jiuren analyzed the adaptability of manned spacecraft to the autonomous emergency return landing area, constructed mathematical models such as sub-satellite point trajectory and emergency landing area boundary, and formed an adaptability analysis criterion. However, compared with near-Earth space missions, the flight trajectory of manned lunar exploration missions is more complicated due to the influence of the gravity of multiple celestial bodies, and more design constraints need to be considered (Li Jiuren, Lin Xiqiang, Li Yingliang, Zhang Hailian. Analysis of adaptability of manned spacecraft to the autonomous emergency return landing area [J]. Manned Spaceflight, 2015, 21(5): 469-471). Summary of the invention

[0005] The technical problem solved by the present invention is: to overcome the shortcomings of the prior art and provide a rapid design method for an Earth-Moon emergency return orbit for manned lunar exploration, which has important reference value for the rapid design and parameter selection of emergency return orbits in manned lunar exploration missions.

[0006] The technical solution of the present invention is to provide a method for rapid design of an Earth-Moon emergency return orbit for manned lunar exploration, comprising:

[0007] With the Earth as the celestial body at the center, a high-precision orbital dynamics equation is established in the geocentric inertial coordinate system to describe the motion state of the spacecraft in the Earth-Moon space;

[0008] Combined with the high-precision orbital dynamics equations of the spacecraft, the B-plane design method is used to determine the general transfer orbit and free return orbit of the manned lunar exploration mission;

[0009] Taking the general transfer orbit or free return orbit as the baseline, the initial value range of the design variables is globally searched under the conditions of comprehensive consideration of the reentry point altitude and reentry angle constraints;

[0010] On the basis of the initial value range of the design variables, the reentry range, reentry inclination angle and coplanarity constraints with the landing point are further considered, and the differential correction optimization algorithm is used to determine the final design variables and emergency return orbit.

[0011] Preferably, in the geocentric inertial coordinate system, the high-precision orbital dynamics equation of the aircraft is:

[0012]

[0013] in, represents the vehicle acceleration vector, μ E represents the Earth's gravitational constant, μ M is the lunar gravitational constant, μ S is the solar gravitational constant, r EP represents the distance between the earth and the spacecraft, r MP represents the distance between the moon and the spacecraft, r EM represents the distance between the Earth and the Moon, r SP represents the distance between the sun and the spacecraft, r ES represents the distance between the sun and the earth, represents the position vector of the aircraft relative to the earth, represents the position vector of the spacecraft relative to the moon, represents the moon position vector, represents the position vector of the aircraft relative to the sun, represents the sun position vector, It represents the influence of other perturbation accelerations, K is the shadow coefficient, which is 0 when the detector is blocked and 1 otherwise, and the solar light pressure coefficient S p satisfy:

[0014]

[0015] Among them, C r is the reflection coefficient, A and m are the area and mass of the aircraft respectively, is the light pressure flux at 1AU, c is the speed of light, r AU is 1AU.

[0016] Preferably, in the B-plane design method, the origin of the B-plane coordinate system is located at the center of the moon, and the unit vector in the direction of the asymptote of the spacecraft orbit is the S axis, that is:

[0017]

[0018] Among them, v ∞ represents the velocity vector; N represents the normal unit vector of the lunar orbital plane, T and R represent the mutually perpendicular unit vectors in the B plane, then:

[0019]

[0020] Vector B is the vector from the center of the moon to the intersection of the asymptote of the spacecraft orbit and the B plane. T represents the projection of vector B on the T axis; B R Represents the projection of vector B on the R axis.

[0021] Preferably, the design variable D is:

[0022] D=[τ,ΔV,t F ] T

[0023] Where: τ is the time for emergency return to the track change point, ΔV is the track change speed increment, t F To achieve emergency return flight time by limiting the emergency return flight time.

[0024] Preferably, the constraint equation F(D) formed by considering the reentry point height and the reentry angle constraints is expressed as:

[0025]

[0026] Where: h alt is the height of the reentry point, γ is the reentry angle; the superscript “*” indicates the expected value of the constraint, that is, represents the expected lunar orbit altitude, γ * is the expected reentry angle.

[0027] Preferably, when searching globally for the initial value range of the design variables, a genetic algorithm is used to comprehensively consider the orbital height and reentry angle constraints, that is:

[0028]

[0029] In the formula, J GA is the performance index function of the genetic algorithm, and the coefficient κ ranges from [0,10];

[0030] Combine the Lambert algorithm with the genetic algorithm and calculate 10 times in a loop, and record the J GA Results and obtained design variables, screen out the minimum J GA , and determine the initial value range of the final design variables accordingly.

[0031] Preferably, the method for determining the final design variables and emergency return trajectory is:

[0032] Based on the initial value range of the design variables, the constraint equation F(D) considering the reentry point altitude and reentry angle constraints is further expanded to the constraint equation F considering the reentry range, reentry inclination angle and the coplanarity constraint with the landing point. N (D);

[0033] Derive the constraint equation F N (D) Regarding the partial derivative relationship of the design variable D, the differential correction optimization algorithm is used to determine the final design variables and emergency return orbit.

[0034] Preferably, the constraint equation F(D) considering the reentry point altitude and reentry angle constraints is further expanded to a constraint equation F(D) considering the reentry range, reentry inclination angle, and coplanarity with the landing point. N (D):

[0035]

[0036] Where: h alt is the height of the reentry point, γ is the reentry angle; represents the expected lunar orbit altitude, γ * is the expected reentry angle; L is the reentry range, i is the reentry inclination angle, ε eta For the points to be coplanar, L * In anticipation of reentry, i * is the desired reentry inclination angle, is the coplanar constraint of the desired landing point, that is, the coplanar constraint of the aircraft position vector and the emergency return aiming point. Here

[0037] Preferably, the constraint equation F is derived N (D) Regarding the partial derivative relationship of the design variable D, when the differential correction optimization algorithm is used to determine the final design variables and emergency return trajectory, the constraint equation F N (D) The partial derivative relationship about the design variable D is:

[0038]

[0039] After iteratively solving the design variable D using the differential correction algorithm, an emergency return orbit is finally designed that meets the constraints of re-entry point altitude, re-entry angle, re-entry range, re-entry inclination, and landing point coplanarity.

[0040] Compared with the prior art, the present invention has the following advantages:

[0041] (1) The global and local optimization design strategies proposed in the present invention can autonomously search for the track change point position and the emergency return speed increment under the constrained return time through genetic algorithm and differential correction algorithm under a high-precision dynamic model.

[0042] (2) According to the stage of spacecraft failure and the requirements of the emergency return mission, the present invention can effectively complete the return orbit design task of emergency return to the designated landing area while controlling the return time, and can provide a reference for emergency rescue of future manned lunar exploration missions.

[0043] (3) The present invention comprehensively considers the re-entry index constraint, the gravity of celestial bodies in the Earth-Moon space, and the spacecraft velocity increment constraints, and quickly optimizes and searches for design variables without introducing too many constraints, and can effectively determine the appropriate initial value range of the design variables. BRIEF DESCRIPTION OF THE DRAWINGS

[0044] Figure 1 A geometrical schematic diagram of the geocentric high-precision orbital dynamics model in step 1 of the present invention;

[0045] Figure 2 This is a schematic diagram of plane B in step 2 of the present invention;

[0046] Figure 3 is a typical general transfer orbit in step 2 of the present invention;

[0047] Figure 4 It is a typical free return orbit in step 2 of the present invention;

[0048] Figure 5 The schematic diagram of the emergency return track and the design variables in step 3 of the present invention;

[0049] Figure 6 It is an emergency return trajectory based on a general transfer orbit according to an embodiment of the present invention;

[0050] Figure 7 The embodiment of the present invention generally transfers the emergency return parameter change curve during different flight periods;

[0051] Figure 8 The emergency return trajectory based on the free return orbit of the embodiment of the present invention;

[0052] Fig. 9 This is a curve showing the change of emergency return parameters during free return in different flight periods according to an embodiment of the present invention. DETAILED DESCRIPTION

[0053] The present invention discloses a method for rapid design of an Earth-Moon emergency return orbit for manned lunar exploration, and the specific steps are as follows:

[0054] Step 1: With the Earth as the celestial body, establish a high-precision orbital dynamics equation in the geocentric inertial coordinate system to describe the motion state of the spacecraft in the Earth-Moon space, as shown in the attached figure. Figure 1 shown.

[0055] In the geocentric inertial coordinate system, the high-precision orbital dynamics equation of the aircraft is:

[0056]

[0057] in, represents the vehicle acceleration vector, μ E represents the Earth's gravitational constant, μ M is the lunar gravitational constant, μ S is the solar gravitational constant, r EP represents the distance between the earth and the spacecraft, r MP represents the distance between the moon and the spacecraft, r EM represents the distance between the Earth and the Moon, r SP represents the distance between the sun and the spacecraft, r ES represents the distance between the sun and the earth, represents the position vector of the aircraft relative to the earth, represents the position vector of the spacecraft relative to the moon, represents the moon position vector, represents the position vector of the aircraft relative to the sun, represents the sun position vector, It represents the influence of other perturbation accelerations, K is the shadow coefficient, which is 0 when the detector is blocked and 1 otherwise, and the solar light pressure coefficient S p satisfy:

[0058]

[0059] Among them, C r is the reflection coefficient, A and m are the area and mass of the aircraft respectively, is the light pressure flux at 1AU, c is the speed of light, r AU is 1AU.

[0060] Step 2: Determine the general transfer orbit and free return orbit of the manned lunar exploration mission by combining the high-precision orbital dynamics equations of the spacecraft.

[0061] The Earth-Moon transfer orbits for manned lunar exploration missions mainly include two types: general transfer orbits and free return orbits. In the high-precision orbital dynamics equations, the general transfer orbit and free return orbit are designed using the B-plane design method, as shown in the attached Figure 2 As shown. The origin of the B plane coordinate system is located at the center of the moon, and the unit vector in the direction of the asymptote of the spacecraft orbit is the S axis, that is,

[0062]

[0063] Among them, v ∞ represents the velocity vector; N represents the normal unit vector of the lunar orbital plane, T and R represent the mutually perpendicular unit vectors in the B plane, then:

[0064]

[0065] Vector B is the vector from the center of the moon to the intersection of the asymptote of the spacecraft orbit and the B plane. T represents the projection of vector B on the T axis; B R Represents the projection of vector B on the R axis.

[0066] The B-plane design method can effectively design a general transfer orbit and a free return orbit that meet the target lunar orbit altitude, inclination or ascending node right ascension, such as Figure 3 and Figure 4 shown.

[0067] Step 3: Using the general transfer orbit or free return orbit as a baseline, globally search for the approximate range of the initial values ​​of the design variables.

[0068] The design variables mainly include the emergency return time τ and the orbit change speed increment ΔV, as well as the emergency return flight time t F When solving the approximate range of the initial values ​​of the design variables, the general transfer orbit or free return orbit is used as the basis, and the height h of the reentry point at the end of the emergency return orbit is considered. alt , reentry angle γ constraint, combined with genetic algorithm and Lambert algorithm to perform global search of design variables. The schematic diagram of emergency return orbit transfer is shown in the attached figure. Figure 5 shown.

[0069] The design variable D and constraint equation F(D) are expressed as:

[0070]

[0071] The superscript “*” in F(D) represents the expected value of the constraint, that is, represents the expected lunar orbit altitude, γ * is the expected reentry angle.

[0072] At the same time, the performance index of the genetic algorithm comprehensively considers the orbit altitude and reentry angle constraints, that is,

[0073]

[0074] In the formula, J GA is the performance index function of the genetic algorithm, and the coefficient κ takes values ​​in the range of [0,10].

[0075] In order to avoid the randomness of the single optimization result of the genetic algorithm, the Lambert algorithm was combined with the genetic algorithm to calculate 10 times in a cycle, and the J GA Results and obtained design variables, screen out the minimum J GA , and based on this, determine the approximate range of the initial values ​​of the final design variables.

[0076] Step 4: Based on the approximate range of the initial values ​​of the design variables in step 3, the differential correction optimization algorithm is further used to determine the final design variables and emergency return orbit.

[0077] Through step 3, the approximate range of the initial values ​​of the design variables can be determined, but it cannot guarantee that the emergency return orbit end under the high-precision dynamic equation will also meet the emergency return orbit end reentry point height h alt , reentry angle γ, reentry range L, reentry inclination angle i and coplanarity with the landing point ε eta Constraints require the use of a differential correction algorithm, combined with the approximate range of the initial values ​​in step three, to further iterate and solve the design variables.

[0078] The constraint equation F(D) in step 3 is expanded to F N (D):

[0079]

[0080] Where, L * In anticipation of reentry, i * is the expected emergency return inclination, i.e., the reentry inclination, is the coplanar constraint between the aircraft position vector and the desired emergency return aiming point, i.e., the landing point is coplanar ε eta The expected value of

[0081] When calculating the differential correction algorithm, it is necessary to derive the constraint equation F N (D) Partial derivative relationship about design variable D After iteratively solving the design variable D using the differential correction algorithm, we can eventually design an emergency return orbit that meets the constraints of re-entry point altitude, re-entry angle, re-entry range, re-entry inclination, and landing point coplanarity.

[0082] The present invention is based on the general transfer orbit and free return orbit under the high-precision orbital dynamics equation of the Earth-Moon space, uses Lambert and genetic algorithms to perform global search, quickly determines the approximate range of the initial values ​​of the design variables of the emergency return orbit, and then optimizes and solves the emergency return orbit at the end of the emergency return orbit to meet the constraints of reentry point height, reentry angle, reentry range, reentry inclination, and landing point coplanarity through a differential correction algorithm, so as to realize the emergency orbit design for returning to the Earth under the condition of a spacecraft failure. This method can effectively solve the problem of initial value conjecture of the emergency return orbit, has a wide range of application, and can be applied to emergency orbit designs such as Earth-Moon transfer and Moon-Earth return.

[0083] Example:

[0084] In order to verify the feasibility of the method, an emergency return orbit is designed for a general transfer orbit. Assuming that the separation time of the spacecraft and the rocket is 2030-1-202:17:54.017 Beijing time, the spacecraft needs to return in about 4 hours. The design parameters of each step in the typical emergency return orbit design are as follows. The flight trajectory and key parameter change curves are shown in the attached figure. Figures 6-7 shown.

[0085] Step 1: With the Earth as the celestial body, establish a high-precision orbital dynamics equation in the geocentric inertial coordinate system. E The value is 398600km 3 / s 2 , μ M Value: 4903km 3 / s 2 , μ S The value is 1.32712e+11km 3 / s 2 , It is determined based on the positional relationship between the sun, the earth, the moon and the spacecraft at a given epoch. When calculating the solar light pressure coefficient, C r Take the value 1, the area A and mass m of the aircraft are 20m 2 and 20t.

[0086] Step 2: Combine the high-precision orbital dynamics equations of the spacecraft to determine the general transfer orbit and free return orbit of the manned lunar exploration mission. When using the B-plane design, the high-precision orbital dynamics equations of step 1 can be integrated to determine B T =-4000km and B R =-500km.

[0087] Step 3: Taking the general transfer orbit or free return orbit as the baseline, consider the reentry point height at the end of the emergency return orbit is 120 km, and the reentry angle is γ * is -6.5°. After global search of design variables by genetic algorithm and Lambert algorithm, the approximate range of the initial value of the design variable satisfies: emergency return time τ∈(3.2,4.2) hours, orbit speed increment ΔV∈(1400,1600) m / s, emergency return flight time t F ∈(12,36) hours.

[0088] Step 4: Based on the approximate range of the initial values ​​of the design variables in step 3, consider the reentry point height is 120km, reentry angle γ * =-6.5°, reentry range L≈7000km, reentry inclination i≈50° and coplanar with the landing point ε eta=90° constraint, and further use the differential correction optimization algorithm to determine the final design variables and emergency return orbit. It is finally determined that the spacecraft can apply an emergency orbit change velocity increment ΔV = 1596.7m / s after τ = 3.7 hours after the separation of the rocket and the spacecraft, and then continue to fly for about t F ≈24 hours later, it reaches the 120km reentry point. Secondly, based on the same steps and constraints mentioned above, emergency return at different fault times is designed. The simulation results are shown in the attached figure. Figure 7 As shown. Figure 7 From the parameters, it can be seen that the emergency return time and the emergency orbit change speed increment have opposite changing trends. Under this condition, the spacecraft can return to the designated landing site in an emergency within 48 hours after the separation of the spacecraft and the rocket.

[0089] In order to further verify the effectiveness of the proposed design strategy, the emergency return orbit of the free return orbit is designed and analyzed. The design process and the values ​​of the design variables in each step are consistent with the general transfer orbit. The changes of the key parameters of the emergency return orbit and the speed increment and flight time under different working conditions are shown in the attached figure. Figures 8-9 As shown. If a fault occurs after the separation point of the spacecraft, it is required to complete the emergency orbit change within 3 hours and return to the designated Dongfeng landing site. Under this condition, the spacecraft can apply an emergency orbit change pulse of 1426.2m / s 2.19 hours after the separation of the ship and the rocket, and continue to sail for about 17.6 hours after completing the orbit change to reach the re-entry point, and the re-entry range is about 6667km. Similar to the general transfer orbit characteristics, as the emergency return orbit change speed increment increases, the return time will gradually shorten. Therefore, under the premise of meeting the re-entry index requirements, the spacecraft's orbit change capability and the urgency of fault handling can be combined to increase the orbit change speed increment within the spacecraft's capabilities and shorten the emergency return time.

[0090] Secondly, according to statistics, for general transfer orbits and free return orbits, the time for initial value search and complete design of an emergency return orbit is less than 10 minutes, and the computer configuration is Intel(R)Core(TM)i7-6700 CPU@3.40GHz.

[0091] The contents not described in detail in the specification of the present invention belong to the prior art known to the professional and technical personnel in this field.

Claims

1. A rapid design method for Earth-Moon emergency return orbit for manned lunar exploration, characterized in that include: With the Earth as the celestial body at the center, a high-precision orbital dynamics equation is established in the geocentric inertial coordinate system to describe the motion state of the spacecraft in the Earth-Moon space; Combined with the high-precision orbital dynamics equations of the spacecraft, the B-plane design method is used to determine the general transfer orbit and free return orbit of the manned lunar exploration mission; Taking the general transfer orbit or free return orbit as the baseline, the initial value range of the design variables is globally searched under the conditions of comprehensive consideration of the reentry point altitude and reentry angle constraints; On the basis of the initial value range of the design variables, the reentry range, reentry inclination angle and the coplanarity constraint with the landing point are further considered, and the differential correction optimization algorithm is used to determine the final design variables and emergency return orbit.

2. The method for rapid design of an Earth-Moon emergency return orbit for manned lunar exploration according to claim 1, characterized in that: In the geocentric inertial coordinate system, the high-precision orbital dynamics equation of the aircraft is: in, represents the vehicle acceleration vector, μ E represents the Earth's gravitational constant, μ M is the lunar gravitational constant, μ S is the solar gravitational constant, r EP represents the distance between the earth and the spacecraft, r MP represents the distance between the moon and the spacecraft, r EM represents the distance between the Earth and the Moon, r SP represents the distance between the sun and the spacecraft, r ES represents the distance between the sun and the earth, represents the position vector of the aircraft relative to the earth, represents the position vector of the spacecraft relative to the moon, represents the moon position vector, represents the position vector of the aircraft relative to the sun, represents the sun position vector, It represents the influence of other perturbation accelerations, K is the shadow coefficient, which is 0 when the detector is blocked and 1 otherwise, and the solar light pressure coefficient S p satisfy: Among them, C r is the reflection coefficient, A and m are the area and mass of the aircraft respectively, is the light pressure flux at 1AU, c is the speed of light, r AU is 1AU.

3. The rapid design method of an Earth-Moon emergency return orbit for manned lunar exploration according to claim 1 is characterized in that: In the B-plane design method, the origin of the B-plane coordinate system is located at the center of the moon, and the unit vector in the direction of the asymptote of the spacecraft orbit is the S axis, that is: Among them, v ∞ represents the velocity vector; N represents the normal unit vector of the lunar orbital plane, T and R represent the mutually perpendicular unit vectors in the B plane, then: Vector B is the vector from the center of the moon to the intersection of the asymptote of the spacecraft orbit and the B plane. T represents the projection of vector B on the T axis; B R Represents the projection of vector B on the R axis.

4. The method for rapid design of an Earth-Moon emergency return orbit for manned lunar exploration according to claim 1, characterized in that: The design variable D is: D=[τ,ΔV,t F ] T Where: τ is the time for emergency return to the track change point, ΔV is the track change speed increment, t F To achieve emergency return flight time by limiting the emergency return flight time.

5. The rapid design method of an Earth-Moon emergency return orbit for manned lunar exploration according to claim 4 is characterized in that: The constraint equation F(D) formed by considering the reentry point height and reentry angle constraints is expressed as: Where: h alt is the height of the reentry point, γ is the reentry angle; the superscript "*" indicates the expected value of the constraint, that is, represents the expected lunar orbit altitude, γ * is the expected reentry angle.

6. The method for rapid design of an Earth-Moon emergency return orbit for manned lunar exploration according to claim 5, characterized in that: When searching the initial value range of the design variables globally, a genetic algorithm is used to comprehensively consider the orbital height and reentry angle constraints, namely: In the formula, J GA is the performance index function of the genetic algorithm, and the coefficient κ ranges from [0,10]; Combine the Lambert algorithm with the genetic algorithm and calculate 10 times in a loop, and record the J GA Results and obtained design variables, screen out the minimum J GA , and determine the initial value range of the final design variables accordingly.

7. The method for rapid design of an Earth-Moon emergency return orbit for manned lunar exploration according to claim 1, characterized in that: The method to determine the final design variables and emergency return trajectory is: Based on the initial value range of the design variables, the constraint equation F(D) considering the reentry point altitude and reentry angle constraints is further expanded to the constraint equation F considering the reentry range, reentry inclination angle and the coplanarity constraint with the landing point. N (D); Derive the constraint equation F N (D) Regarding the partial derivative relationship of the design variable D, the differential correction optimization algorithm is used to determine the final design variables and emergency return orbit.

8. The method for rapid design of an Earth-Moon emergency return orbit for manned lunar exploration according to claim 7, characterized in that: The constraint equation F(D) considering the reentry point altitude and reentry angle constraints is further expanded to consider the reentry range, reentry inclination angle and the coplanarity constraint with the landing point. N (D): Where: h alt is the height of the reentry point, γ is the reentry angle; represents the expected lunar orbit altitude, γ * is the expected reentry angle; L is the reentry range, i is the reentry inclination angle, ε eta For the points to be coplanar, L * In anticipation of reentry, i * is the desired reentry inclination angle, is the coplanar constraint of the desired landing point, that is, the coplanar constraint of the aircraft position vector and the emergency return aiming point. Here 9. The method for rapid design of an Earth-Moon emergency return orbit for manned lunar exploration according to claim 7, characterized in that: Derive the constraint equation F N (D) Regarding the partial derivative relationship of the design variable D, when the differential correction optimization algorithm is used to determine the final design variables and emergency return trajectory, the constraint equation F N (D) The partial derivative relationship about the design variable D is: After iteratively solving the design variable D using the differential correction algorithm, an emergency return orbit is finally designed that meets the constraints of re-entry point altitude, re-entry angle, re-entry range, re-entry inclination, and landing point coplanarity.

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