A method for multi-stage joint planning of lunar sampling return fixed-point soft landing orbit

Abstract

The application discloses a kind of lunar sampling return fixed point soft landing orbit multi-stage joint planning methods, belong to aerospace field.The application implementation method is: according to task requirement calculation orbit key parameter, determine nominal lunar orbit inclination, nominal lunar orbit inclination latitude argument and nominal lunar orbit inclination size of nominal lunar orbit inclination latitude argument;According to lunar orbit inclination, carry out earth-moon transfer orbit planning, calculate earth-moon transfer orbit parameter;Using current earth-moon transfer orbit parameter carries out high-precision earth-moon transfer orbit recursion to obtain perigee parameter;According to perigee parameter calculation near moon braking parameter, recursion to four-device separation point, calculate four-device separation point parameter;Calculate lunar orbit inclination parameter, recursion to power descent point, calculate power descent point orbit parameter and moon landing point parameter;According to four-device separation point parameter and moon landing point parameter, adjust earth-moon transfer orbit parameter and braking and maneuver parameter, realize lunar sampling return fixed point soft landing orbit multi-stage joint planning.

Description

Technical Field

[0001] This invention relates to a multi-stage joint planning method for lunar sample return fixed-point soft landing orbit, belonging to the field of aerospace. Background Technology

[0002] Lunar exploration is a general term for human activities involving the exploration and observation of the Moon. Among these, a soft landing on the Moon allows for close contact with the Moon and is one of the most effective methods of lunar exploration. Building upon a soft landing, lunar sample collection, lunar ascent, and Moon-Earth return operations can be performed to bring samples back to Earth for further research. The significant advantages of lunar sample return missions in lunar science research have led to widespread attention and the extensive development of lunar sample return technology. Among these, orbit design technology is a prerequisite and key to sending a probe to the lunar surface. The lunar sample return soft landing flight process is complex, including stages such as launch, Earth-Moon transfer, lunar capture, lunar orbit flight, descent maneuver, and powered descent. The lunar soft landing orbit design must meet various system constraints, such as launch capacity constraints, maneuverability constraints, telemetry and control coverage constraints, illumination constraints, landing area constraints, and ascent and rendezvous constraints. The complex flight stages and engineering constraints together make it difficult to jointly plan a fixed-point soft landing orbit on the Moon. Segmented planning introduces a huge workload and makes it impossible to obtain a globally optimal orbit. To improve the computational efficiency of lunar fixed-point soft landing orbits and enhance the optimality of planning results, a multi-stage joint planning method is urgently needed.

[0003] Among the developed lunar stationary soft landing orbit design and planning methods, the advanced technology [1] (Bai Yuzhu, Xi Xiaoning, Gao Yudong, et al. Research on lunar soft landing mission window and orbit design method [J]. Journal of Astronautics, 2009(6):2092-2098.) established a set of window calculation and orbit planning methods for lunar soft landing missions under multiple constraints. The method uses the reverse calculation method of lunar landing point, near-lunar point, orbit insertion point and launch point to carry out the through-planning of lunar soft landing mission. The advantage of this method is that it can plan the entire orbit before lunar soft landing. The disadvantage is that this method cannot consider the constraint that the orbiter and ascender need to rendezvous and dock in the sample return mission.

[0004] The prior art [2] (Wang Zhongsheng, Meng Zhanfeng, Gao Shan. Research on lunar fixed-point landing orbit change strategy [J]. Spacecraft Engineering, 2017, 26(2):29-37.) compared and analyzed different fixed-point landing orbit change strategies such as orbit adjustment, phasing and orbital plane adjustment, and adopted 2-to-2 and 3-to-3 aiming maneuvers to achieve target altitude and landing point latitude. The advantage of this method is that it can perform orbit planning from insertion to lunar landing for sampling return missions. The disadvantage is that the single parameter adjustment capability is poor, the nested iterative planning calculation is large, and it is difficult to obtain the global optimal planning result. Summary of the Invention

[0005] The technical problem addressed by the multi-stage joint planning method for lunar sample return and fixed-point soft landing orbits disclosed in this invention is: to achieve lunar fixed-point soft landing orbit planning under the complex dynamic environment of the Earth-Moon system and the engineering constraints of lunar exploration missions, considering the lunar surface ascent, rendezvous, and docking requirements of lunar sample return missions, and to achieve joint planning of multi-stage orbits throughout the entire process from launch to powered descent. Furthermore, this invention is compatible and can solve the multi-stage orbit joint planning problem for lunar exploration missions with other soft landing requirements.

[0006] The objective of this invention is achieved through the following technical solution:

[0007] This invention discloses a multi-stage joint planning method for a lunar sample return and fixed-point soft landing trajectory. The method calculates key trajectory parameters based on overall mission requirements, determining the nominal lunar orbit inclination, the nominal lunar descent maneuver point latitude argument, and the nominal lunar descent maneuver magnitude. Based on the lunar orbit inclination, a rapid Earth-Moon transfer trajectory is planned, and Earth-Moon transfer trajectory parameters are calculated. Using the current Earth-Moon transfer trajectory parameters, a high-precision Earth-Moon transfer trajectory recursion is performed to obtain the perilune parameters. Based on the perilune parameters, lunar braking parameters are calculated and recursively extended to the separation point, calculating the separation point parameters. Based on the separation point parameters, lunar descent maneuver parameters are calculated and recursively extended to the powered descent point, calculating the powered descent point trajectory parameters and the lunar landing point parameters. Based on the separation point parameters and the lunar landing point parameters, the Earth-Moon transfer trajectory parameters and braking and maneuver parameters are adjusted to achieve multi-stage joint planning of the lunar sample return and fixed-point soft landing trajectory.

[0008] This invention discloses a multi-stage joint planning method for lunar sample return fixed-point soft landing orbits, comprising the following steps:

[0009] Step 1: Calculate the key orbital parameters according to the overall mission requirements, and determine the nominal lunar orbit inclination, the nominal lunar descent maneuver point latitude argument, and the nominal lunar descent maneuver magnitude;

[0010] Based on the mission objectives, the nominal lunar longitude and latitude of the landing site are obtained; based on the overall lander design, the nominal powered descent range and the altitude of the powered descent initiation point are obtained; based on the lunar surface work mission plan, the lunar surface work time is obtained; and based on the overall lunar exploration requirements, the nominal lunar orbit altitude of the probe is obtained. The lunar orbit inclination is calculated using the nominal lunar latitude of the landing site, the lunar surface work time, and the lunar rotation angular velocity; the latitude argument of the lunar descent maneuver point is calculated using the nominal lunar latitude of the landing site and the lunar orbit inclination; and the magnitude of the nominal lunar descent maneuver is calculated using the altitude of the powered descent initiation point and the nominal lunar orbit altitude. The first and second lunar descent maneuvers are calculated according to the overall mission requirements; the level is defined as level zero.

[0011] Step 2: Based on the lunar orbit inclination, perform rapid planning of the Earth-Moon transfer orbit and calculate the Earth-Moon transfer orbit parameters;

[0012] Using the Earth-Moon transfer conic section splicing model, the direction of the lunar velocity at the lunar entry point is analytically calculated based on the nominal Earth-Moon transfer time, nominal lunar orbit altitude, and lunar orbit inclination. The direction of the Earth velocity at the entry point is analytically calculated based on the lunar ephemeris data at the entry point time. The initial guesses of the semi-major axis, right ascension of the ascending node, and argument of the perigee of the Earth-Moon transfer orbit are calculated using the Earth-Moon transfer velocity directions and the lunar velocity directions, respectively. These are called the Earth-Moon transfer orbit parameters.

[0013] The level is increased by one level, the magnitude of the second lunar braking normal is set to zero, and the current Earth-Moon transfer orbit parameters are calculated using the geocentric orbit parameters of the Earth-Moon transfer orbit.

[0014] Step 3: Use the current Earth-Moon transfer orbit parameters to perform high-precision Earth-Moon transfer orbit recursion to obtain the perigee parameters, and execute step 3 or step 4 based on the recursion results;

[0015] The initial state of the Earth-Moon transfer orbit is calculated using the current Earth-Moon transfer orbit parameters and the altitude, inclination, and true anomaly of the insertion point. The Earth-Moon transfer orbit parameters are then recursively derived to the lunar perigee using the insertion time as the initial time under a high-precision dynamic model. The lunar perigee orbit altitude, inclination, and true anomaly of the lunar perigee are obtained and are referred to as the lunar perigee parameters.

[0016] If the level is Level 1 and the perilune parameters do not meet the accuracy requirements, adjust the current Earth-Moon transfer orbit parameters and repeat step 3; if the level is Level 1 and the perilune parameters meet the accuracy requirements, upgrade the level by one level and proceed to step 4; if the level is not Level 1, proceed to step 4.

[0017] Step 4: Calculate the lunar braking parameters based on the near-lunar point parameters, and then extrapolate to the separation point of the four instruments to calculate the separation point parameters. Execute either Step 3 or Step 5 based on the separation point parameters.

[0018] Based on the lunar orbit period after the first lunar braking, the first lunar braking is calculated analytically. Based on the lunar point parameters obtained in step three, the first lunar braking is applied to obtain the lunar orbit parameters after the first lunar braking. These parameters are then recursively applied to the second lunar braking point. Based on the lunar orbit period after the second lunar braking, the tangential magnitude of the second lunar braking is calculated analytically. The tangential magnitude and normal magnitude of the second lunar braking are combined to form the second lunar braking. The second lunar braking is applied to obtain the lunar orbit parameters after the second lunar braking. These parameters are then recursively applied to the third lunar braking point. Based on the nominal lunar orbit height, the third lunar braking is calculated analytically. The third lunar braking is applied to obtain the lunar orbit parameters after the third lunar braking. These parameters are then recursively applied to the separation point of the four spacecraft to obtain the orbit height and inclination at the separation point, which are called the orbit parameters at the separation point.

[0019] If the level is level 2 and the orbital parameters of the true perigee angle and the separation point of the four spacecraft do not meet the accuracy requirements, then adjust the current Earth-Moon transfer orbit parameters and repeat step 3; if the level is level 2 and the orbital parameters of the true perigee angle and the separation point of the four spacecraft meet the accuracy requirements, then upgrade the level by one level and proceed to step 5; if the level is not level 2, then proceed to step 5.

[0020] Step 5: Calculate the lunar descent maneuver parameters based on the separation point parameters of the four spacecraft, and extrapolate to the powered descent point. Calculate the orbital parameters of the powered descent point and the lunar landing point parameters. Execute either Step 5 or Step 6 based on the results.

[0021] Based on the number of lunar orbit rotations from the separation point of the four spacecraft to the first lunar descent maneuver point, and the latitude argument of the lunar descent maneuver point, the parameters of the separation point are recursively derived to the first lunar descent maneuver point. The first lunar descent maneuver is then applied, yielding the lunar orbit parameters after the maneuver. Based on the number of lunar orbit rotations from the first to the second lunar descent maneuver, the lunar orbit parameters after the first maneuver are recursively derived to the second lunar descent maneuver point. The second lunar descent maneuver is then applied, yielding the lunar orbit parameters after the second maneuver. Based on the number of lunar orbit rotations after the second maneuver, the lunar orbit parameters after the second maneuver are recursively derived to the powered descent point, and the altitude of the powered descent point is calculated. The landing time and the latitude and longitude of the landing point are calculated using the nominal powered descent range and nominal powered descent time.

[0022] If the level is level three and the altitude of the powered descent point and the latitude of the lunar landing point do not meet the accuracy requirements, then adjust the latitude argument of the current lunar descent maneuver point, perform a second lunar descent maneuver, and repeat step five; if the level is level three and the altitude of the powered descent point and the latitude of the lunar landing point meet the accuracy requirements, then upgrade the level by one level and execute step six; if the level is not level three, then execute step six.

[0023] Step Six: Based on the separation point parameters of the four spacecraft and the lunar landing point parameters, adjust the Earth-Moon transfer orbit parameters and braking and maneuvering parameters to achieve multi-stage joint planning of the lunar sample return fixed-point soft landing orbit. Execute Step Three or complete the multi-stage joint planning of the lunar sample return fixed-point soft landing orbit based on the planning results to obtain the multi-stage joint planning results of the lunar sample return fixed-point soft landing orbit. Achieve the lunar sample return fixed-point soft landing based on the multi-stage joint planning results.

[0024] If the level is level four and the true anomaly of the perihelion, the orbital altitude of the separation point of the four spacecraft, the inclination, the altitude of the powered descent point, the landing time, and the latitude of the landing point do not meet the accuracy requirements, then adjust the semi-major axis of the current Earth-Moon transfer orbit, the right ascension of the ascending node, the argument of the perigee, the magnitude of the second lunar braking tangent, the latitude argument of the lunar descent maneuver point, and the magnitude of the second lunar descent maneuver, and repeat step three; if the level is level four and the true anomaly of the perihelion, the orbital altitude of the separation point of the four spacecraft, the inclination, the altitude of the powered descent point, the landing time, and the latitude of the landing point meet the accuracy requirements, then upgrade the level by one level; if the level is level five and the true anomaly of the perihelion, the orbital altitude of the separation point of the four spacecraft, the inclination, the altitude of the powered descent point, and the landing time... If the latitude and longitude of the landing point do not meet the accuracy requirements, adjust the semi-major axis of the current Earth-Moon transfer orbit, the right ascension of the ascending node, the argument of the perigee, the magnitude of the second lunar braking tangent, the magnitude of the second lunar braking normal, the latitude and argument of the lunar descent maneuver point, and the magnitude of the second lunar descent maneuver, and repeat step three. If the level is five and the true perigee angle of the lunar point, the orbital altitude of the separation point of the four spacecraft, the inclination, the altitude of the powered descent point, the landing time, and the latitude and longitude of the landing point meet the accuracy requirements, then complete the multi-stage joint planning of the lunar sample return fixed-point soft landing orbit, obtain the multi-stage joint planning result of the lunar sample return fixed-point soft landing orbit, and realize the lunar sample return fixed-point soft landing according to the multi-stage joint planning result.

[0025] Beneficial effects:

[0026] 1. The present invention discloses a multi-stage joint planning method for lunar sample return and fixed-point soft landing orbits. The method decomposes the complete lunar sample return and fixed-point soft landing orbit into multiple sub-processes with progressive relationships. This breaks down the high-dimensional problem of complex engineering orbit design into smaller parts. Compared with the complete process planning method, it reduces the low-dimensional problem of optimization design and improves the convergence of the multi-stage joint solution process.

[0027] 2. To overcome the difficulty of traditional design methods in considering complex engineering constraints, this invention discloses a multi-stage joint planning method for lunar sample return and fixed-point soft landing orbits. Based on the basic constraints of the lunar sample return mission, this method transforms complex control and illumination constraints that are sensitive to changes in design variables into easily considered process parameters by setting detailed process parameters, such as the lunar orbit period after each lunar braking maneuver, the number of lunar orbits, the number of lunar orbits after the separation point of the four spacecraft, and the landing time parameters. This solves the joint planning of lunar sample return and fixed-point soft landing orbits under the constraints of entry point altitude, inclination, true perigee angle, and lunar point latitude and longitude.

[0028] 3. The present invention discloses a multi-stage joint planning method for lunar sample return and fixed-point soft landing orbits. It establishes a design framework based on design levels, dynamically adjusts the design levels according to the current design results, captures changes in design levels and adjusts the design process in a timely manner accordingly, so as to flatten and streamline the multi-stage joint design problem with complex flight processes and strong stage coupling relationships, and improve the flexibility, globality and adaptability of the multi-stage joint planning method for lunar sample return and fixed-point soft landing orbits.

[0029] 4. The multi-stage joint planning method for lunar sample return fixed-point soft landing orbit disclosed in this invention, on the basis of achieving the above-mentioned beneficial effects 1 and 2, has good convergence, can be applied to the orbit planning of lunar sample return missions with different windows, different scenarios and different payloads, requires low computational load, and can improve the planning speed of lunar sample return orbit. Attached Figure Description

[0030] Figure 1 This is a flowchart of a multi-stage joint planning method for a lunar sample return fixed-point soft landing orbit disclosed in this invention.

[0031] Figure 2 The geocentric orbit diagram is a design result of an example of this invention.

[0032] Figure 3 The lunar orbit diagram is designed as an example of the present invention. Detailed Implementation

[0033] To better illustrate the purpose and advantages of the present invention, the invention will be further described below in conjunction with the accompanying drawings and examples.

[0034] In this example, the Earth-Moon transfer window is selected as the Julian Day 2460433 window, with an estimated landing time of MJD60462.9.

[0035] This embodiment discloses a method for reconstructing a hybrid satellite constellation failure after satellite failure, the specific implementation method of which is as follows:

[0036] Step 1: Calculate the key orbital parameters based on the overall mission plan, and determine the nominal lunar orbit inclination, the nominal lunar descent maneuver point latitude argument, and the nominal lunar descent maneuver magnitude.

[0037] In this example, the nominal lunar landing site coordinates are 154°W and 43.2°S, the nominal powered descent range is 600 km, the powered descent initiation altitude is 15 km, the lunar surface working time is 2 days, and the nominal lunar orbit altitude is 200 km. The lunar orbit inclination, the nominal lunar descent maneuver argument at the latitude of the maneuver point, and the nominal lunar descent maneuver magnitude are calculated using the following methods.

[0038]

[0039]

[0040]

[0041] Where: i MoonStd φ is the lunar orbital inclination. LandStd ω represents the lunar latitude of the nominal landing point. Moon Let t be the angular velocity of the moon's rotation. Work For lunar working time, S DecentStd For the nominal power descent range, R Moon Let μ be the radius of the moon. Moon h is the lunar gravitational constant. MoonStd h is the nominal lunar orbital altitude. DecentStd This refers to the nominal starting height of the descent.

[0042] In this example, the first and second lunar orbit descent maneuvers are calculated using the following method.

[0043]

[0044] Wherein: F DEC1 The thrust of the engine used in the first lunar orbit descent maneuver, m DEC1 For the mass of the probe during its first lunar orbit descent maneuver, Δt DEC1 For the first lunar orbit descent maneuver test spray time, △v DEC1 With △v DEC2 These refer to the first and second lunar orbit descent maneuvers, respectively. The above equation indicates the case where the first lunar orbit descent maneuver is used as the calibration of the thrust for the landing complex's descent.

[0045] In this example, the calculated lunar orbit inclination is 136.037 degrees, the nominal lunar descent maneuver point latitude argument is 79.785 degrees, and the nominal lunar descent maneuver magnitude is 40.372 m / s.

[0046] Step 2: Based on the lunar orbit inclination, quickly design the Earth-Moon transfer orbit and calculate its parameters.

[0047] In this example, the following method is used for rapid design of the Earth-Moon transfer orbit.

[0048]

[0049]

[0050]

[0051]

[0052]

[0053]

[0054] r EAF =p rr r BF +p vr v BF ,v EAF =p rv r BF +p vv v BF

[0055] Where: μ Earth r is the gravitational constant of the Earth's core. EAF For the rapid design of the Earth-Moon transfer incident point location, β EAF For the rapid design of the Earth-Moon transfer incident point trajectory angle, r BF For the rapid design of the Earth-Moon transfer entry point location, v BF For rapid design of the Earth-Moon transfer entry point velocity, e EF For the eccentricity of the Earth-Moon transfer orbit in rapid design, f EAF For the true anomaly angle of the Earth-Moon transfer orbit insertion point in rapid design, f BF For the true anterior angle of the Earth-Moon transfer entry point in rapid design, p rr ,p vr ,p rv ,p vv For state transition parameters in rapid design, r BF With v BF For the position and velocity vectors of the Earth-Moon transfer entry point in rapid design, r EAF With v EAF The position and velocity vectors of the Earth-Moon transfer incident point in rapid design.

[0056] After completing the rapid design of the Earth-Moon transfer orbit, the parameters of the Earth-Moon transfer orbit are calculated in this example using the following method.

[0057]

[0058]

[0059]

[0060] Ω EtoLOI1 =arg(Ω) EtoLOI1 )

[0061]

[0062] Where: a EtoLOI1 L is the semi-major axis of the Earth-Moon transfer orbit. EtoLOI1 Ω is the Laplace vector of the Earth-Moon transfer orbit. EtoLOI1 Ω is the right ascension vector of the ascending node of the Earth-Moon transfer orbit. EtoLOI1 Let ω be the right ascension of the ascending node of the Earth-Moon transfer orbit, arg(·) denote the principal argument of the vector in the complex plane, and ω be the right ascension of the ascending node of the Earth-Moon transfer orbit. EtoLOI1 The perigee argument of the Earth-Moon transfer orbit.

[0063] In this example, the semi-major axis of the rapidly designed Earth-Moon transfer orbit is 191,531 km, the right ascension of the ascending node is 349 degrees, and the argument of the perigee is 239 degrees.

[0064] Step 3: Use the current Earth-Moon transfer orbit parameters to perform a high-precision Earth-Moon transfer orbit recursion, and execute either Step 3 or Step 4 based on the results.

[0065] In this example, the initial state of the Earth-Moon transfer orbit is calculated using the following method.

[0066]

[0067] Where: h Inject R is the altitude of the entry point. Earth The radius is the Earth's radius.

[0068] The dynamic model used in this example for the Earth-Moon transfer recursion is:

[0069] a EtoM =a EarthCenter +a EarthNonSphere +a MoonThirdBody +a MoonNonSphere +a SRP

[0070] Where: a EtoM Let a be the total acceleration experienced by the probe during its Earth-Moon transfer. EarthCenter a is the gravitational acceleration at the Earth's center.EarthNonSphere For the non-spherical acceleration of Earth, a MoonThirdBody a is the gravitational acceleration due to the third body of the moon. MoonNonSphere For the lunar third-body acceleration, a SRP This refers to solar radiation pressure acceleration.

[0071] This example uses the Newton-Raphson method to adjust the current Earth-Moon transfer orbit parameters.

[0072]

[0073] Where: h LOI1 i represents the lunar orbital altitude at the perigee. LOI1 f is the lunar orbital inclination at the perigee. LOI1 The true anterior angle of the lunar orbit around the perigee. This is a function of the Earth-Moon transfer orbit parameters to the perigee parameters. This is the Jacobian matrix representing the Earth-Moon transfer orbit parameters to the perigee parameters.

[0074] In this example, the orbital parameters selected are an insertion point altitude of 200km, an inclination angle of 23.3 degrees, and a true anomaly angle of 22.5 degrees. The Earth-Moon transfer time is set to 112 hours, and the Earth-Moon transfer orbit parameters calculated in step two are used.

[0075] The initial state of the Earth-Moon transfer orbit was calculated using the semi-major axis, right ascension of the ascending node, argument of perigee, altitude of the entry point, inclination, and true anomaly. The Earth-Moon transfer time was recursively calculated using the entry point as the initial time under a high-precision dynamic model. The obtained lunar orbital altitude at the perigee was 66.608 km, the inclination was 133.768 degrees, and the true anomaly was -11.927 degrees.

[0076] In this example, the required lunar orbit altitude at the perigee is 200 km, and the required true anomaly angle is 0 degrees. At this stage, the design level is Level 1, and the lunar orbit altitude, inclination, and true anomaly angle do not meet the accuracy requirements. The semi-major axis of the Earth-Moon transfer orbit is adjusted to 191664.729 km, the right ascension of the ascending node to 349.240 degrees, and the argument of perigee to 239.175 degrees. Step three is then repeated, resulting in a lunar orbit altitude of 204.287 km, an inclination of 135.943 degrees, and a true anomaly angle of -1.131 degrees. It can be seen that the deviation of the target parameters is decreasing. After repeatedly adjusting the parameters of the Earth-Moon transfer orbit and repeating step three, the lunar orbital altitude, inclination, and true perigee angle that meet the accuracy requirements are obtained. The corresponding semi-major axis of the Earth-Moon transfer orbit is 191665.036 km, the right ascension of the ascending node is 349.233 degrees, and the argument of perigee is 239.179 degrees.

[0077] Step 4: Calculate the lunar braking parameters based on the near-lunar point parameters, and then extrapolate to the separation point of the four spacecraft. Calculate the orbital parameters at the separation point of the four spacecraft, and proceed to Step 3 or Step 5 based on the results.

[0078] In this example, the first lunar braking is calculated using the following method.

[0079]

[0080] Wherein: T LOI1toLOI2 For the lunar orbital period following the first lunar braking, a LOI1 Let Δv be the semi-major axis of the lunar orbit at the perigee. LOI1 This was the first near-lunar braking maneuver.

[0081] In this example, the lunar orbit parameters after the first lunar braking are calculated using the following method.

[0082]

[0083]

[0084] in: This represents the lunar orbital position vector prior to the first lunar braking maneuver. This represents the lunar orbital velocity vector prior to the first lunar braking maneuver. This represents the lunar orbit position vector after the first lunar braking maneuver. This represents the lunar orbital velocity vector following the first lunar braking maneuver.

[0085] The dynamic model used in this example for lunar recursion is:

[0086] a MoonOrbit =a MoonCenter +a MoonNonSphere +a EarthThirdBody +a SRP

[0087] Where: a MoonOrbit Let a be the total acceleration experienced by the probe during its lunar orbit. MoonCenter a is the gravitational acceleration at the center of the moon. MoonNonSphere For the non-spherical acceleration of the moon, a EarthThirdBody For the gravitational acceleration due to Earth's third body, a SRP This refers to solar radiation pressure acceleration.

[0088] In this example, the following method is used to calculate the magnitude of the second lunar braking tangential.

[0089]

[0090] Where: h LOI2 The orbital altitude of the second lunar braking point, T LOI2toLOI3For the lunar orbital period after the second lunar braking, a LOI2 For the semi-major axis before the second near-lunar braking, △v LOI2,T This refers to the magnitude of the second near-lunar braking tangential direction.

[0091] In this example, the lunar orbit parameters after the second lunar braking are calculated using the following method.

[0092]

[0093]

[0094] in: This represents the lunar orbital position vector prior to the second lunar braking maneuver. This represents the lunar orbital velocity vector prior to the second lunar braking maneuver. This represents the lunar orbit position vector after the second lunar braking maneuver. Let Δv be the lunar orbital velocity vector after the second lunar braking maneuver. LOI2,N This represents the magnitude of the normal direction of the second near-lunar braking.

[0095] In this example, the third lunar braking is calculated using the following method.

[0096]

[0097] Where: h LOI3 The orbital altitude of the third lunar braking point, a LOI3 For the semi-major axis before the third near-lunar braking, △v LOI3 This is the third near-lunar braking.

[0098] In this example, the lunar orbit parameters after the third lunar braking are calculated using the following method.

[0099]

[0100]

[0101] in: This represents the lunar orbital position vector prior to the third lunar braking maneuver. This represents the lunar orbital velocity vector prior to the third lunar braking maneuver. This is the lunar orbit position vector after the third lunar braking maneuver. This is the lunar orbital velocity vector after the third lunar braking maneuver.

[0102] This example uses the Newton-Raphson method to adjust the current Earth-Moon transfer orbit parameters.

[0103]

[0104] Where: hSeparation The track height at the separation point of the four instruments, i Separation The track inclination angle at the separation point of the four instruments. It is a function of the Earth-Moon transfer orbit parameters to the true anomaly angle at the perigee and the orbit parameters at the separation point of the four spacecraft. The Jacobian matrix represents the parameters of the Earth-Moon transfer orbit to the true anomaly angle at the perigee and the orbit parameters at the separation point of the four spacecraft.

[0105] In this example, the lunar orbital period after the first lunar braking is 24 hours, with one lunar orbit. The lunar orbital period after the second lunar braking is 4 hours, with four lunar orbits. The separation time of the four spacecraft is MJD60460.4. Based on the lunar center orbital parameters obtained in step three, the true anomaly angle of the lunar point is calculated to be 0 degrees. The first lunar braking speed is 274.649 m / s, the second lunar braking speed is 285.647 m / s, and the third lunar braking speed is 266.103 m / s. The orbital altitude at the separation point is 161.666 km, and the inclination angle is 136.507 degrees.

[0106] In this example, the required orbital altitude for the separation point of the four spacecraft is 200 km in lunar orbit. At this point, the design level is Level II, and the true anomaly angle at the perigee meets the requirements. However, the orbital altitude and inclination at the separation point do not meet the accuracy requirements. The semi-major axis of the Earth-Moon transfer orbit is adjusted to 191681.836 km, the right ascension of the ascending node to 349.281 degrees, and the argument of perigee to 239.132 degrees. Step three is repeated, resulting in an orbital altitude of 200.009 km, an orbital inclination of 136.235 degrees, and a true anomaly angle at the perigee of -0.004 degrees, meeting the accuracy requirements. Step five is then executed.

[0107] Step 5: Based on the separation point parameters of the four spacecraft, calculate the lunar descent maneuver parameters to the lunar descent point, and then calculate the orbital parameters of the powered descent point and the lunar landing point parameters. Execute either Step 5 or Step 6 based on the results.

[0108] In this example, the lunar orbit parameters after the first lunar orbit descent maneuver are calculated using the following method.

[0109]

[0110]

[0111] in: This represents the lunar orbital position vector prior to the first lunar orbital descent maneuver. This represents the lunar orbital velocity vector prior to the first lunar orbital descent maneuver. This represents the lunar orbital position vector after the first lunar orbital descent maneuver. This represents the lunar orbital velocity vector after the first lunar orbital descent maneuver.

[0112] In this example, the lunar orbit parameters after the second lunar orbit descent maneuver are calculated using the following method.

[0113]

[0114]

[0115] in: This represents the lunar orbital position vector before the second lunar orbital descent maneuver. This represents the lunar orbital velocity vector prior to the second lunar orbital descent maneuver. This represents the lunar orbit position vector after the second lunar orbit descent maneuver. This represents the lunar orbital velocity vector after the second lunar orbital descent maneuver.

[0116] In this example, the Newton-Raphson method is used to adjust the latitude argument of the current lunar descent maneuver point and the second lunar descent maneuver.

[0117]

[0118] Where: h Decent φ is the height of the descent point. Land For the lunar latitude, Let be the latitude argument of the lunar descent maneuvering point, the altitude from the second lunar descent to the dynamic descent point, and the latitude of the landing point. The Jacobian matrix represents the latitude argument of the lunar descent maneuvering point and the altitude of the second lunar descent to the dynamic descent point and the latitude of the landing point.

[0119] In this example, the number of lunar orbits from the separation point of the four spacecraft to the first lunar descent maneuver is 12, and the number of lunar orbits from the first lunar descent maneuver to the second lunar descent maneuver is 12. The nominal powered descent range is 600 km, the nominal powered descent time is 850 s, the mass of the landing and overland assembly is 3750 kg, the thrust of the landing and overland assembly engine is 7500 N, and the test firing time of the landing and overland assembly engine is 15 s. The corresponding speeds for the first lunar descent maneuver are 30 m / s and the second lunar descent maneuver is 10.372 m / s. Based on the separation point parameters obtained in step four, the calculated altitude of the powered descent point is 31.112 km, and the latitude of the lunar landing point is 29.744 degrees south latitude.

[0120] In this example, the required altitude for the powered descent point is 15 km. At this stage, the design level is three, and the altitude of the powered descent point and the latitude of the lunar landing point do not meet the accuracy requirements. The latitude argument of the lunar descent maneuver point is adjusted to 69.785 degrees, and the second lunar descent maneuver is performed at 12.328 m / s. Step five is repeated, resulting in a powered descent point altitude of 22.821 km and a lunar landing point latitude of 36.365 degrees south. It can be seen that the deviation of the target parameters is decreasing. Multiple lunar descent maneuver points with the same latitude argument and the second lunar descent maneuver, followed by repeating step five, yield a powered descent point altitude and lunar landing point latitude that meet the accuracy requirements. The corresponding lunar descent maneuver point latitude argument is 54.463 degrees, and the second lunar descent maneuver is performed at 13.913 m / s.

[0121] Step Six: Based on the separation point parameters of the four spacecraft and the lunar landing point parameters, adjust the Earth-Moon transfer orbit parameters and braking and maneuvering parameters to achieve a multi-stage joint design of the lunar sample return fixed-point soft landing orbit. Execute Step Three or obtain the design results based on the design results.

[0122] In this example, the Broyden method is used to adjust the true anomaly angle of the current lunar point, the orbital altitude and inclination of the separation point of the four spacecraft, the altitude of the powered descent point, the landing time, and the latitude of the lunar point.

[0123]

[0124] Where: t Land For the sake of the month, The Jacobian matrix △X represents the semi-major axis of the Earth-Moon transfer orbit, the right ascension of the ascending node, the argument of the perigee, the magnitude of the second lunar braking tangent, the latitude argument of the lunar descent maneuver point, the magnitude of the second lunar descent maneuver to the true perigee angle, the orbital altitude of the separation point, the inclination, the altitude of the powered descent point, the landing time, and the latitude of the landing point. VI The current adjustments to the semi-major axis of the Earth-Moon transfer orbit, the right ascension of the ascending node, the argument of the perigee, the magnitude of the second lunar braking tangent, the latitude argument of the lunar descent maneuver point, and the magnitude of the second lunar descent maneuver, ΔF. VI The current deviations of the true perigee angle, orbital altitude and inclination at the separation point of the four spacecraft, altitude at the dynamic descent point, landing time, and latitude of the landing point are as follows:

[0125] In this example, the Broyden method is used to adjust the true perigee angle of the current lunar point, the orbital altitude and inclination of the separation point of the four spacecraft, the altitude of the powered descent point, the landing time, and the latitude and longitude of the landing point.

[0126]

[0127] Where: λ Land For the longitude of the moon, The Jacobian matrix (ΔX) represents the semi-major axis of the Earth-Moon transfer orbit, the right ascension of the ascending node, the argument of the perigee, the magnitude of the second lunar braking tangential, the magnitude of the second lunar braking normal, the argument of the lunar descent maneuver point in latitude, the magnitude of the second lunar descent maneuver to the true anomaly of the perigee, the altitude of the separation point of the four spacecraft, the inclination, the altitude of the powered descent point, the landing time, and the latitude and longitude of the landing point. VII The current adjustments ΔF represent: the semi-major axis of the Earth-Moon transfer orbit, the right ascension of the ascending node, the argument of the perigee, the magnitude of the second lunar braking tangential, the magnitude of the second lunar braking normal, the latitude argument of the lunar descent maneuver point, and the magnitude of the second lunar descent maneuver. VII This represents the current deviations of the true perigee angle, orbital altitude and inclination at the separation point of the four spacecraft, altitude at the dynamic descent point, landing time, and latitude and longitude of the landing point.

[0128] In this example, the required lunar landing time is MJD60462.924. At this time, the calculated true anomaly angle of the lunar perihelion is 0 degrees, the orbital altitude of the separation point of the four spacecraft is 200.009 km, the inclination is 136.235 degrees, the altitude of the powered descent point is 14.946 km, the lunar landing time is MJD60462.907, and the latitude of the lunar landing point is 42.981 degrees south latitude. At this time, the design level is level four. The true anomaly angle of the lunar perihelion, the orbital altitude of the separation point of the four spacecraft, the inclination, the altitude of the powered descent point, and the latitude of the lunar landing point meet the accuracy requirements, but the lunar landing time does not meet the accuracy requirements. The semi-major axis of the Earth-Moon transfer orbit was adjusted to 191679.791 km, the right ascension of the ascending node to 349.284 degrees, the argument of perigee to 239.130 degrees, the magnitude of the second lunar braking tangential motion to 285.487 m / s, the argument of latitude at the lunar descent maneuver point to 53.463 degrees, and the magnitude of the second lunar descent maneuver to 14.232 m / s. Step three was repeated, and the obtained true perigee angle at the lunar point was 0 degrees, the orbital altitude at the separation point of the four spacecraft was 200.023 km, the inclination was 136.176 degrees, the altitude at the powered descent point was 13.412 km, the landing time was MJD60462.912, and the latitude of the landing point was 43.282 degrees south latitude. It can be seen that the deviation of the target parameters is decreasing. After multiple adjustments to the semi-major axis of the Earth-Moon transfer orbit, the right ascension of the ascending node, the argument of perigee, the magnitude of the second lunar braking tangent, the latitude argument of the lunar descent maneuver point, and the magnitude of the second lunar descent maneuver, and by repeating step three, the true anomaly of the perigee, the orbital altitude of the separation point of the four spacecraft, the inclination, the altitude of the powered descent point, the landing time, and the latitude of the landing point were obtained, all meeting the accuracy requirements. The corresponding semi-major axis of the Earth-Moon transfer orbit is 191674.890 km, the right ascension of the ascending node is 349.290 degrees, the argument of perigee is 239.124 degrees, the magnitude of the second lunar braking tangent is 285.158 m / s, the latitude argument of the lunar descent maneuver point is 54.537 degrees, and the magnitude of the second lunar descent maneuver is 13.860 m / s.

[0129] In this example, the calculated true anomaly angle of the lunar perihelion is 0 degrees, the orbital altitude of the separation point of the four spacecraft is 200.001 km, the inclination is 136.036 degrees, the altitude of the powered descent point is 15 km, the landing time is MJD60462.924, the longitude of the landing point is 156.946 degrees west, and the latitude of the landing point is 43.2 degrees south. At this time, the design level is level five. The true anomaly angle of the lunar perihelion, the orbital altitude of the separation point of the four spacecraft, the inclination, the altitude of the powered descent point, the landing time, and the latitude of the landing point meet the accuracy requirements, but the longitude of the landing point does not meet the accuracy requirements. The semi-major axis of the Earth-Moon transfer orbit was adjusted to 191680.377 km, the right ascension of the ascending node to 349.283 degrees, the argument of perigee to 239.131 degrees, the tangential magnitude of the second lunar braking maneuver to 285.199 m / s, the normal magnitude of the second lunar braking maneuver to 8.019 m / s, the latitude argument of the lunar descent maneuver point to 53.537 degrees, and the magnitude of the second lunar descent maneuver to 14.060 m / s. Step three was repeated, yielding a true anomaly of 0 degrees, an orbital altitude of 200.035 km at the separation point of the four spacecraft, an inclination of 136.036 degrees, an altitude of 14.069 km at the powered descent point, a lunar landing time of MJD60462.925, a lunar landing longitude of 154.944 degrees West, and a lunar landing latitude of 43.433 degrees South. It can be seen that the deviation of the target parameters is decreasing. Multiple adjustments were made to the semi-major axis of the Earth-Moon transfer orbit, the right ascension of the ascending node, the argument of perigee, the magnitude of the tangential and normal axes of the second lunar braking maneuver, the latitude argument of the lunar descent maneuver point, and the magnitude of the second lunar descent maneuver. Step three was then repeated to obtain the true perigee angle, orbital altitude and inclination of the separation point of the four spacecraft, the altitude of the powered descent point, the landing time, and the latitude and longitude of the landing point, all meeting the accuracy requirements. This completed the multi-stage joint design of the lunar sample return and stationary soft landing orbit. The corresponding Earth-Moon transfer orbit has a semi-major axis of 191,744.088 km, an ascending node right ascension of 349.195 degrees, a perigee argument of 239.212 degrees, a second lunar braking tangential magnitude of 284.329 m / s, a second lunar braking normal magnitude of 100.096 m / s, a lunar descent maneuver point latitude argument of 54.775 degrees, and a second lunar descent maneuver magnitude of 13.880 m / s. The corresponding geocentric orbit based on the design results is as follows: Figure 2 The design results correspond to the lunar orbit as follows: Figure 3 Show.

[0130] The above detailed description further illustrates the purpose, technical solution, and beneficial effects of the invention. It should be understood that the above description is only a specific embodiment of the present invention and is not intended to limit the scope of protection of the present invention. Any modifications, equivalent substitutions, improvements, etc., made within the spirit and principles of the present invention should be included within the scope of protection of the present invention.

Claims

1. A multi-stage joint planning method for lunar sample return fixed-point soft landing orbit, characterized in that: Includes the following steps, Step 1: Calculate the key orbital parameters according to the overall mission requirements, and determine the nominal lunar orbit inclination, the nominal lunar descent maneuver point latitude argument, and the nominal lunar descent maneuver magnitude; Based on the mission objectives, the nominal lunar longitude and latitude of the landing site are obtained; based on the overall lander design, the nominal powered descent range and the altitude of the powered descent initiation point are obtained; based on the lunar surface work mission plan, the lunar surface work time is obtained; and based on the overall lunar exploration requirements, the nominal lunar orbit altitude of the probe is obtained. The lunar orbit inclination is calculated using the nominal lunar latitude of the landing site, the lunar surface work time, and the lunar rotation angular velocity; the latitude argument of the lunar descent maneuver point is calculated using the nominal lunar latitude of the landing site and the lunar orbit inclination; and the magnitude of the nominal lunar descent maneuver is calculated using the altitude of the powered descent initiation point and the nominal lunar orbit altitude. The first and second lunar descent maneuvers are calculated according to the overall mission requirements; the level is defined as level zero. Step 2: Based on the lunar orbit inclination, perform rapid planning of the Earth-Moon transfer orbit and calculate the Earth-Moon transfer orbit parameters; Using the Earth-Moon transfer conic section splicing model, the direction of the lunar velocity at the lunar entry point is analytically calculated based on the nominal Earth-Moon transfer time, nominal lunar orbit altitude, and lunar orbit inclination. The direction of the Earth velocity at the entry point is analytically calculated based on the lunar ephemeris data at the entry point time. The initial guesses of the semi-major axis, right ascension of the ascending node, and argument of the perigee of the Earth-Moon transfer orbit are calculated using the Earth-Moon transfer velocity directions and the lunar velocity directions, respectively. These are called the Earth-Moon transfer orbit parameters. The level is increased by one level, the magnitude of the second lunar braking normal is set to zero, and the current Earth-Moon transfer orbit parameters are calculated using the geocentric orbit parameters of the Earth-Moon transfer orbit. Step 3: Use the current Earth-Moon transfer orbit parameters to perform high-precision Earth-Moon transfer orbit recursion to obtain the perigee parameters, and execute step 3 or step 4 based on the recursion results; The initial state of the Earth-Moon transfer orbit is calculated using the current Earth-Moon transfer orbit parameters and the altitude, inclination, and true anomaly of the insertion point. The Earth-Moon transfer orbit parameters are then recursively derived to the lunar perigee using the insertion time as the initial time under a high-precision dynamic model. The lunar perigee orbit altitude, inclination, and true anomaly of the lunar perigee are obtained and are referred to as the lunar perigee parameters. If the level is Level 1 and the perilune parameters do not meet the accuracy requirements, adjust the current Earth-Moon transfer orbit parameters and repeat step 3; if the level is Level 1 and the perilune parameters meet the accuracy requirements, upgrade the level by one level and proceed to step 4; if the level is not Level 1, proceed to step 4. Step 4: Calculate the lunar braking parameters based on the near-lunar point parameters, and then extrapolate to the separation point of the four instruments to calculate the separation point parameters. Execute either Step 3 or Step 5 based on the separation point parameters. Based on the lunar orbit period after the first lunar braking, the first lunar braking is calculated analytically; based on the lunar point parameters obtained in step three, the first lunar braking is applied to obtain the lunar orbit parameters after the first lunar braking, and then the lunar orbit parameters after the first lunar braking are recursively extrapolated to the second lunar braking point; based on the lunar orbit period after the second lunar braking, the tangential magnitude of the second lunar braking is calculated analytically, and the tangential magnitude of the second lunar braking is combined with the normal magnitude to form the second lunar braking; A second lunar braking maneuver is performed to obtain the lunar orbit parameters after the second lunar braking maneuver. These parameters are then used to extrapolate to the third lunar braking point. The third lunar braking maneuver is then calculated analytically based on the nominal lunar orbit altitude. A third lunar braking was applied to obtain the lunar orbit parameters after the third lunar braking. The lunar orbit parameters after the third lunar braking were then extrapolated to the separation point of the four spacecraft to obtain the orbital height and inclination of the separation point of the four spacecraft, which are called the orbital parameters of the separation point of the four spacecraft. If the level is level 2 and the orbital parameters of the true perigee angle and the separation point of the four spacecraft do not meet the accuracy requirements, then adjust the current Earth-Moon transfer orbit parameters and repeat step 3; if the level is level 2 and the orbital parameters of the true perigee angle and the separation point of the four spacecraft meet the accuracy requirements, then upgrade the level by one level and proceed to step 5; if the level is not level 2, then proceed to step 5. Step 5: Calculate the lunar descent maneuver parameters based on the separation point parameters of the four spacecraft, and extrapolate to the powered descent point. Calculate the orbital parameters of the powered descent point and the lunar landing point parameters. Execute either Step 5 or Step 6 based on the results. Based on the number of lunar orbits from the separation point of the four spacecraft to the first lunar orbit descent maneuver point and the latitude argument of the lunar orbit descent maneuver point, the parameters of the separation point of the four spacecraft are recursively derived to the first lunar orbit descent maneuver point; the first lunar orbit descent maneuver is applied to obtain the lunar orbit parameters after the first lunar orbit descent maneuver. Based on the number of lunar orbit rotations from the first to the second lunar orbit reduction maneuver, the lunar orbit parameters after the first lunar orbit reduction maneuver are recursively extrapolated to the point of the second lunar orbit reduction maneuver. A second lunar orbit descent maneuver was performed, and the lunar orbit parameters after the second lunar orbit descent maneuver were obtained. Based on the number of lunar orbits after the second lunar orbit descent maneuver, the lunar orbit parameters after the second lunar orbit descent maneuver are recursively extrapolated to the powered descent point, and the height of the powered descent point is calculated. The landing time and landing point latitude and longitude are calculated using the nominal powered descent range and nominal powered descent time. If the level is level three and the altitude of the powered descent point and the latitude of the lunar landing point do not meet the accuracy requirements, then adjust the latitude argument of the current lunar descent maneuver point, perform a second lunar descent maneuver, and repeat step five; if the level is level three and the altitude of the powered descent point and the latitude of the lunar landing point meet the accuracy requirements, then upgrade the level by one level and execute step six; if the level is not level three, then execute step six. Step Six: Based on the separation point parameters of the four spacecraft and the lunar landing point parameters, adjust the Earth-Moon transfer orbit parameters and braking and maneuvering parameters to achieve multi-stage joint planning of the lunar sample return fixed-point soft landing orbit. Execute Step Three or complete the multi-stage joint planning of the lunar sample return fixed-point soft landing orbit based on the planning results, obtain the multi-stage joint planning results of the lunar sample return fixed-point soft landing orbit, and achieve the lunar sample return fixed-point soft landing based on the multi-stage joint planning results.

2. The multi-stage joint planning method for lunar sample return fixed-point soft landing orbit as described in claim 1, characterized in that: Step six is ​​implemented as follows: If the level is level four and the true anomaly of the perihelion, the orbital altitude of the separation point of the four spacecraft, the inclination, the altitude of the powered descent point, the landing time, and the latitude of the landing point do not meet the accuracy requirements, then adjust the semi-major axis of the current Earth-Moon transfer orbit, the right ascension of the ascending node, the argument of the perigee, the magnitude of the second lunar braking tangent, the latitude argument of the lunar descent maneuver point, and the magnitude of the second lunar descent maneuver, and repeat step three; if the level is level four and the true anomaly of the perihelion, the orbital altitude of the separation point of the four spacecraft, the inclination, the altitude of the powered descent point, the landing time, and the latitude of the landing point meet the accuracy requirements, then upgrade the level by one level; if the level is level five and the true anomaly of the perihelion, the orbital altitude of the separation point of the four spacecraft, the inclination, the altitude of the powered descent point, and the landing time... If the latitude and longitude of the landing point do not meet the accuracy requirements, adjust the semi-major axis of the current Earth-Moon transfer orbit, the right ascension of the ascending node, the argument of the perigee, the magnitude of the second lunar braking tangent, the magnitude of the second lunar braking normal, the latitude and argument of the lunar descent maneuver point, and the magnitude of the second lunar descent maneuver, and repeat step three. If the level is five and the true perigee angle of the lunar point, the orbital altitude of the separation point of the four spacecraft, the inclination, the altitude of the powered descent point, the landing time, and the latitude and longitude of the landing point meet the accuracy requirements, then complete the multi-stage joint planning of the lunar sample return fixed-point soft landing orbit, obtain the multi-stage joint planning result of the lunar sample return fixed-point soft landing orbit, and realize the lunar sample return fixed-point soft landing according to the multi-stage joint planning result.

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