Three-pulse near-moon braking fixed-time and fixed-point landing aiming method

Abstract

The invention discloses a three-pulse near-moon braking fixed-time and fixed-point landing aiming method. The method comprises the steps that input parameters are obtained; determining a power descending point moment initial value, a second-time near-moon braking moment, a four-device separation point moment initial value, a third-time near-moon braking moment and a lunar surface take-off moment; a finite thrust model and an accurate forecasting model are adopted for orbit control strategy solving, and orbit parameters at the four-device separation point moment are obtained through calculation; inner and outer ring strategy solving is carried out according to four-device separation point moment orbit parameters; and outputting an orbit control strategy calculation result. According to the method, the characteristic of perturbation rate difference of lunar elliptical orbit planes of different periods is utilized, a 4h lunar parking elliptical orbit is designed, the parking orbit flight time corresponding to the orbit plane adjustment amount is determined by establishing the drift rate difference between the 4h lunar parking elliptical orbit and a 2h target lunar orbit plane, the flight time of the nearly one month from capturing to descending is fully utilized, and the flight speed of the target lunar orbit plane is determined. On the premise of not additionally consuming a propellant, high-precision aiming of a landing point is realized.

Description

TECHNICAL FIELD

[0001] The application belongs to the technical field of deep space exploration orbit design, and particularly relates to a three-impulse near-moon braking timing and fixed-point landing aiming method. BACKGROUND

[0002] In the unmanned sampling return mission on the back of the moon, the probe reaches the near-moon point via the earth-moon transfer orbit, and three near-moon braking controls are implemented at the near-moon point in sequence to enter the moon orbit with a period of about 12h, 4h and 2h respectively. After the probe enters the moon orbit with a height of about 200km and a period of about 2h, the separation operation is implemented at the predetermined four-vehicle separation time, the orbit-return combination body flies in orbit, and the landing-ascending combination body performs two de-orbit controls on the moon orbit and reaches the powered descent point at the predetermined time for implementation.

[0003] Due to the rugged terrain on the back of the moon, the longitude allowable range of the sampling area on the back of the moon is very small, and the accuracy of the powered descent point is high. Meanwhile, due to the launch window requirement of the moon-back sampling return mission, which is two months in a year and two days in a month, the orbital planes of different windows when reaching the moon have certain differences, which leads to the existence of longitude differences of the moon orbit when reaching the powered descent point, and the orbit control is needed to realize the aiming at the target landing point. In the past moon exploration missions, the adjustment of the orbit inclination is generally realized by increasing the normal velocity increment or increasing the normal correction of the moon orbit in the near-moon braking orbit strategy, so as to achieve the purpose of timing and fixed-point landing aiming, but this will lead to the increase of the velocity increment and propellant consumption of the probe.

[0004] Since the unmanned sampling return mission on the back of the moon does not have additional propellant for adjusting the orbital plane parameters, the traditional timing and fixed-point landing aiming method is not applicable, and therefore a new near-moon braking strategy is needed to realize the task requirement of landing in the specified sampling area for all launch windows without consuming additional propellant. SUMMARY

[0005] The technical problem of the application is to overcome the shortcomings of the prior art, provide a three-impulse near-moon braking timing and fixed-point landing aiming method, utilize the characteristics of the drift rate difference of different period moon orbit planes, carefully design a 4h moon parking elliptical orbit, determine the parking orbit flight time corresponding to the orbit plane adjustment amount by establishing the drift rate difference between the 4h moon parking elliptical orbit and the 2h target moon orbit plane, and fully utilize the flight time from capture to the month before descent to realize the high-precision aiming at the landing point without additional consumption of propellant.

[0006] In order to solve the above technical problems, the application discloses a three-impulse near-moon braking timing and fixed-point landing aiming method, which comprises the following steps:

[0007] acquiring input parameters;

[0008] determining a powered descent point time initial value, a second lunar approach braking time, a four-device separation point time initial value, a third lunar approach braking time, and a lunar surface takeoff time based on the input parameters;

[0009] Trajectory control strategy solving is performed using a finite thrust model and an accurate prediction model, and the four-device separation point time orbit parameters are calculated;

[0010] Inner loop strategy solving is performed according to the four-device separation point time orbit parameters;

[0011] Outer layer strategy solving is performed according to the inner loop strategy solving result;

[0012] Output the trajectory control strategy calculation result.

[0013] In the above three-impulse lunar approach braking timing and point landing aiming method, the input parameters include: geolunar transfer orbit parameters, lunar surface target sampling point parameters, powered descent point parameters, a first lunar approach braking post-control orbit period T1, a second lunar approach braking post-control orbit period T2, and a third lunar approach braking post-control orbit period T3; wherein the geolunar transfer orbit parameters include: an initial time t0, an initial time orbit element E0, a flight time ΔT E2M , a departure and arrival mode; the lunar surface target sampling point parameters include: sampling point longitude and latitude lunar surface working time ΔT LS ; the powered descent point parameters include: powered descent point altitude h Des , powered descent point latitude amplitude angle ω Des .

[0014] In the above three-impulse lunar approach braking timing and point landing aiming method, the powered descent point time initial value is determined by the following method:

[0015] According to and ΔT LS , the powered descent point time fixed inclination initial value I0 and the ascending node longitude initial value λ Des0 that satisfy the lunar surface working time requirement are calculated:

[0016]

[0017]

[0018] wherein ω Moon represents the lunar rotation angular velocity;

[0019] According to E0 and ΔT E2M , the orbit element E f, and then convert it to obtain the orbital position velocity X in the lunar center inertial system LOI1 =(r LOI1 ,v LOI1 );According to the perigee time t f =t0+ΔT E2M and X LOI1 , coordinate transformation is performed to obtain the orbital position velocity of the lunar center fixed connection at the time of perilunar point pass The conversion of the orbital elements yields the fixed system inclination i at the perigee LOI1 and the longitude of the ascending node λ LOI1 ;

[0020] Solve to obtain the flight time Δλ from the braking to the powered descent near the moon LOI1toDes :

[0021] Δλ LOI1toDes =λ Des0 -λ LOI1

[0022] According to T2 and T3, the precession rate of the ascending node of the lunar orbit after the second near-moon braking is calculated respectively. and the precession rate of the ascending node of the lunar orbit after the third near-moon braking

[0023] According to ω Moon and Solve to obtain the flight time ΔT from near-month braking to powered descent LOI1toDes :

[0024]

[0025] Among them, k is the forward and backward flag, the forward track k is -1, and the backward track k is 1;

[0026] Calculate the initial value of the powered descent point according to different lunar orbit types:

[0027] For a prograde lunar orbit, ΔT LOI1toDes The repetition period T of the ground station tracking arc VD Round down and take the remainder to get the integer part and the remainder The initial value of the power drop point at this time as follows:

[0028]

[0029] For a retrograde lunar orbit, ΔT LOI1toDes T VD Round up and take the remainder to get the integer part and remainder The initial value of the power drop point at this time as follows:

[0030]

[0031] Among them, t LOI1 Indicates the first near-month braking moment.

[0032] In the above three-pulse near-moon braking timing and fixed-point landing aiming method, the second near-moon braking time is: one day after the first near-moon braking and the moment when it enters the common view arc of the ground tracking and control station; where t LOI2 =t LOI1 +T VD , t LOI2 Indicates the second near-month braking moment.

[0033] In the above three-pulse near-moon braking timing and fixed-point landing aiming method, the initial value of the four-device separation point is: day, and is within the common view arc of the ground tracking and control station; among them, Indicates the initial value of the four-device separation point, N Sep The value range is 1 to 3 days.

[0034] In the above three-pulse near-moon braking timing and fixed-point landing aiming method, the third near-moon braking time is determined by the following method:

[0035] according to Calculate the residual deviation of the longitude of the orbital ascending node at the time of power descent

[0036]

[0037] According to the difference in the precession rate of the ascending node of the lunar orbit after the second and third near-moon braking, The maximum adjustable amount of the orbital ascending node longitude drift is calculated

[0038]

[0039] in, t Sep Indicates the moment of power drop point;

[0040] like According to The third near-month braking time t is calculated LOI3 : At this time, the power drops at the moment t Des and the separation point of the four devices t Sep Remain unchanged as: Where Δλ1day Indicates the angle of the moon's rotation in one day;

[0041] like but:

[0042] For prograde orbits around the Moon:

[0043] Calculated Corresponding longitude of the ascending node of the lunar orbit

[0044] One day after the power drop point is calculated Corresponding longitude of the ascending node of the lunar orbit

[0045] according to and Calculate the eastern boundary λ of the adjustable range of the ascending node longitude east and the western boundary of the ascending node longitude adjustable λ west :

[0046] Calculate the target value of the ascending node longitude adjustment

[0047] Determine whether the following conditions are met:

[0048] If satisfied The third near-month braking time t LOI3 for: At this time, the power drops at the moment t Des and the separation point of the four devices t Sep Remain unchanged as:

[0049] If not satisfied but:

[0050] like The third near-month braking time t LOI3 for: At this time, the power drops at the moment t Des and the separation point of the four devices t Sep Remain unchanged as: The ascending node longitude δλ that needs to be adjusted for the first near-moon braking LOI1 for: in, Indicates the minimum flight time between the third near-moon braking and the separation point of the four vehicles;

[0051] like Then postpone the power drop point by one day: The third near-month braking time t LOI3 Set to: The first perilune braking needs to adjust the ascending node longitude amount δλ LOI1 is: The four-vehicle separation point time t Sep The corresponding update of the power descent point time is: t Sep = t Des -N Sep T VD ; wherein, The minimum flight time between the second perigee braking and the third perigee braking is represented by T;

[0052] For a circumlunar retrograde orbit:

[0053] The power descent point time t The corresponding ascending node longitude of the circumlunar orbit

[0054] The day before the power descent point t The corresponding ascending node longitude of the circumlunar orbit

[0055] According to and The west boundary of the adjustable range of the ascending node longitude λ west , the east boundary of the adjustable ascending node longitude λ east are respectively:

[0056] The target value of the ascending node longitude adjustment λ

[0057] Determine whether the following conditions are met:

[0058] If The third perigee braking time t LOI3 is: At this time, the power descent time t Des and the four-vehicle separation point time t Sep remain unchanged, which are:

[0059] If Then:

[0060] If The third perigee braking time t LOI3 is: At this time, the power descent time t Des and the four-vehicle separation point time t Sep remain unchanged, which are: The first perigee braking needs to adjust the ascending node longitude amount δλ LOI1 is: wherein, denotes the minimum flight duration between the third lunar proximity maneuver and the four-impulse separation point;

[0061] If the powered descent point time is advanced one day: the third lunar proximity maneuver time t LOI3 is: the first lunar proximity maneuver requires adjustment of the ascending node longitude amount δλ LOI1 is: the four-impulse separation point time t Sep is updated accordingly with the powered descent point time: Sep = t Des -N Sep T VD ; wherein, denotes the minimum flight duration between the second lunar proximity maneuver and the third lunar proximity maneuver.

[0062] In the above three-impulse lunar proximity maneuver timing and point landing targeting method, the lunar surface takeoff time t Des is calculated based on the powered descent point time t Asc : t Asc = t Des + ΔT LS .

[0063] In the above three-impulse lunar proximity maneuver timing and point landing targeting method, the finite thrust model and the precise prediction model are used to solve the orbit control strategy, and the four-impulse separation point time orbit parameters are calculated, including:

[0064] According to the finite thrust model, the first lunar proximity maneuver is applied, the design variable is the on-time duration ΔT LOI1 , the target variable is T1, and after iterative convergence using the differential correction method, the first lunar proximity maneuver orbit control parameters ΔV LOI1 and the controlled orbit parameters E LOI1 are obtained;

[0065] According to E LOI1 , the second lunar proximity maneuver time t LOI2 is predicted to stop, the finite thrust model is applied to the second lunar proximity maneuver, the design variable is the on-time duration ΔT LOI2 , the target variable is T2, and after iterative convergence using the differential correction method, the second lunar proximity maneuver orbit control parameters ΔV LOI2 and the controlled orbit parameters E LOI2 are obtained;

[0066] According to E LOI2 , the third lunar proximity maneuver time t LOI3 is predicted, the finite thrust model is applied to the third lunar proximity maneuver, the design variable is the on-time duration ΔT LOI3The target variable is T3. After iterative convergence using the differential correction method, the orbit control parameter ΔV for the third near-moon braking is obtained. LOI3 and the orbital parameter E after control LOI3 ;

[0067] According to E LOI3 Forecast to the time of separation of the four devices t LOI3 , obtain the orbital parameters E at the separation point of the four devices Sep =(a Sep ,e Sep ,i Sep ,Ω Sep ,ω Sep ,f Sep ); where a Sep represents the semi-major axis of the orbit at the moment of separation, e Sep represents the eccentricity at the separation moment, i Sep represents the orbital inclination at the separation moment, Ω Sep represents the right ascension of the ascending node at the moment of separation, ω Sep represents the argument of perigee at the moment of separation, f Sep Indicates the true anomaly at the moment of separation.

[0068] In the above three-pulse near-moon braking timing and fixed-point landing aiming method, the inner loop strategy is solved according to the orbit parameters at the separation point of the four devices, including:

[0069] According to E Sep =(a Sep ,e Sep ,i Sep ,Ω Sep ,ω Sep ,f Sep ), aiming at the target parameters of the powered descent point according to the lunar orbit descent strategy, the design variable is the latitude argument u of the four-device separation point Sep =ω Sep +f Sep , the start-up time of two lunar orbit reduction control operations ΔT DM1 and ΔT DM2 , the target variable is the moment t at which the power drops Des , height h Des and the argument of perigee ω Des After iterative convergence using the differential correction method, the control parameters ΔV for the two lunar orbit reductions are obtained. DM1 and ΔV DM2 , orbital parameter E corresponding to the moment of dynamic descent Des ;

[0070] According to E Sep =(a Sep ,e Sep ,i Sep ,ΩSep ,ω Sep ,f Sep ), according to the phase modulation strategy aiming at the target parameters of the shift point, the design variable is the start-up time length of the four-phase modulation control, the target variable is the orbit semi-major axis, eccentricity and latitude amplitude angle at the shift point, and the control parameters of the four-phase modulation and the orbit parameters E Asc at the lunar surface take-off moment are obtained after iterative convergence by using the differential correction method.

[0071] According to E Des and E Asc , the angles LPA Des and LPA Asc between the position vectors of the sampling points at the powered descent point and the lunar surface take-off moment and the orbital plane are calculated.

[0072] E Des is converted into position and velocity X Des =(r Des ,v Des ), further coordinate conversion is performed to obtain the position and velocity of the powered descent point in the fixed connection with the moon center X , and the angle LPA Des between the position vector of the sampling point at the powered descent point and the orbital plane is calculated.

[0073]

[0074] wherein, represents the orbital angular momentum of the powered descent point, represents the position vector of the sampling point,

[0075] Similarly, the angle LPA Asc between the position vector of the sampling point at the lunar surface take-off moment and the orbital plane is calculated.

[0076] In the above three-impulse near-moon braking timing and fixed-point landing aiming method, the outer strategy solving is performed according to the inner loop strategy solving results, including:

[0077] According to the task requirements of timing and fixed-point landing and timely take-off, the orbital inclination i Sep and the ascending node right ascension Ω Sep of the four-separation point are selected as the design variables p1, and the angles between the position vectors of the sampling points at the powered descent point t Des and the lunar surface take-off moment t Asc and the orbital plane are 0° as the aiming target q1:

[0078]

[0079] The deviation Δq1 between the target variable and the target value is calculated.

[0080]

[0081] According to the function relation between Δq1, the design variable p1 and the aiming target variable q1: p1=f(q1), the orbit inclination i Sep and the ascending node right ascension Ω Sep of the four-vehicle separation point orbit parameter E

[0082]

[0083] The orbit inclination i Sep and the ascending node right ascension Ω Sep of the four-vehicle separation point orbit parameter E Sep are updated, and iterative calculation is carried out until Δq1 is less than the allowed error threshold, the calculation is ended, and the four-vehicle separation point target orbit parameter satisfying the requirements of the timing and point landing and the on-time take-off task is obtained

[0084] According to the lunar perigee orbit inclination of the earth-moon transfer and the lunar perigee braking strategy are solved, the design variable p2 is the lunar perigee orbit inclination i LOI1 , the thrust azimuth angle A of the first lunar perigee braking LOI1 , the control orbit period T3 of the third lunar perigee braking, the target variable q2 is the four-vehicle separation point latitude amplitude orbit inclination and the ascending node right ascension

[0085]

[0086] The differential correction method is adopted, the deviation of the variable p2 and the target value q2 is calculated, and then the correction amount of the design variable is obtained, the lunar perigee inclination target of the earth-moon transfer orbit is updated, and the calculation is ended until Δq2 is less than the allowed error threshold.

[0087] The present application has the following advantages:

[0088] The application discloses a three-impulse near-moon braking timing and fixed-point landing aiming method, which realizes aiming at target parameters out of the orbital plane through adjustment of in-plane parameters; can take into account the task requirements of timing and fixed-point landing and coplanar take-off on time, utilizes the perturbation drift of different period lunar orbits, greatly reduces the normal velocity increment consumption of orbital plane correction, can finely adjust the semi-major axis of the lunar orbit, has the ability of phase adjustment, and meets the requirements of timing landing. In addition, the measurement and control requirements of orbit transfer events are considered, the method is applicable to a lunar unmanned sampling return task with a long lunar time (more than 20 days), is not limited by the lunar surface working time and sampling point position, and can accurately aim at different landing point positions by adjusting the inclination of the near-moon point of the earth-moon transfer orbit and the drift time of the elliptical orbit. BRIEF DESCRIPTION OF DRAWINGS

[0089] Figure 1 is a flowchart of a three-impulse near-moon braking timing and fixed-point landing aiming method in the embodiment of the application;

[0090] Figure 2 is a schematic diagram of the ascending node adjustment capability of an elliptical orbit in a lunar orbit prograde orbit state, which meets the adjustment requirements;

[0091] Figure 3 is a schematic diagram of the ascending node adjustment capability of an elliptical orbit in a lunar orbit prograde orbit state, which does not meet the requirements;

[0092] Figure 4 is a schematic diagram of the ascending node adjustment capability of an elliptical orbit in a lunar orbit retrograde orbit state, which meets the requirements;

[0093] Figure 5 is a schematic diagram of the ascending node adjustment capability of an elliptical orbit in a lunar orbit retrograde orbit state, which does not meet the requirements. DETAILED DESCRIPTION

[0094] To make the objectives, technical solutions and advantages of the application clearer, the following further describes the disclosed embodiments of the application with reference to the drawings.

[0095] Reference Figure 1 In the embodiment, the three-impulse near-moon braking timing and fixed-point landing aiming method comprises the following steps.

[0096] S1, acquiring input parameters.

[0097] In the embodiment, the input parameters mainly include: the geolunar transfer orbit parameters, the lunar surface target sampling point parameters, the powered descent point parameters, the control post-orbit period T1 of the first near-moon braking, the control post-orbit period T2 of the second near-moon braking, and the control post-orbit period T3 of the third near-moon braking. Further, the geolunar transfer orbit parameters mainly include: the initial time t0, the initial time orbit element E0, the flight time ΔT E2M , the departure and arrival mode (elevated track); the lunar surface target sampling point parameters mainly include: the sampling point longitude and latitude , the lunar surface working time ΔT LS ; the powered descent point parameters mainly include: the powered descent point height h Des , the powered descent point latitude amplitude angle ω Des .

[0098] S2, determine the powered descent point time initial value.

[0099] In the embodiment, the powered descent point time initial value can be determined by the following method:

[0100] 2.1) according to and ΔT LS , the powered descent point time fixed connection inclination initial value I0 and the ascending node longitude initial value λ Des0 satisfying the lunar surface working time requirement are calculated:

[0101]

[0102] wherein, ω Moon represents the lunar rotation angular velocity.

[0103] 2.2) according to E0 and ΔT E2M , the orbit element E f at the geolunar transfer arrival near-moon point time is predicted and calculated, and then the orbit position and velocity X LOI1 =(r LOI1 , v LOI1 ) in the lunar center inertia system is converted; according to the near-moon point time t f =t0+ΔT E2M and X LOI1 , the coordinate conversion is performed to obtain the lunar center fixed connection system orbit position and velocity at the near-moon point time Through the conversion between and the orbit element, the fixed connection inclination i LOI1 and the ascending node longitude λ LOI1 at the near-moon point time are obtained.

[0104] 2.3) the flight time Δλ LOI1toDes from the near-moon braking to the powered descent is solved:

[0105] Δλ LOI1toDes=λ Des0 -λ LOI1

[0106] 2.4) Based on T2 and T3, calculate the precession rate of the ascending node of the lunar orbit after the second near-moon braking. and the precession rate of the ascending node of the lunar orbit after the third near-moon braking

[0107] 2.5) According to ω Moon and Solve to obtain the flight time ΔT from near-month braking to powered descent LOI1toDes :

[0108]

[0109] Among them, k is the forward and reverse flag, the forward track k is -1, and the reverse track k is 1.

[0110] 2.6) Calculate the initial value of the powered descent point according to different lunar orbit types:

[0111] (2.6.1) For the prograde lunar orbit, replace ΔT LOI1toDes The repetition period T of the ground station tracking arc VD Round down and take the remainder to get the integer part and the remainder The initial value of the power drop point at this time as follows:

[0112]

[0113] Among them, T VD =24h50min.

[0114] This part needs to be calculated by multiples of T VD Adjust the landing time to achieve the remainder Adjustments must be made using the ascending node drift caused by the elliptical orbit perturbation.

[0115] (2.6.2) For a retrograde lunar orbit, replace ΔT LOI1toDes T VD Round up and take the remainder to get the integer part and the remainder The initial value of the power drop point at this time as follows:

[0116]

[0117] Among them, t LOI1 Indicates the first near-month braking moment.

[0118] 2.7) The final step S2 outputs the initial value of the power drop point

[0119] S3, determine the initial values ​​of the second near-month braking time and the four-device separation point time.

[0120] In this embodiment, the second near-moon braking time is: 1 day after the first near-moon braking and the time when it enters the common viewing arc of the ground measurement and control station; the initial value of the four-device separation point time is: N before the power drops Sep day, and is within the common view arc of the ground tracking and control stations; then:

[0121] t LOI2 =t LOI1 +T VD

[0122]

[0123] Among them, t LOI2 Indicates the second near-month braking moment; Indicates the initial value of the four-device separation point, N Sep In engineering, the general value range is 1 to 3 days.

[0124] S4, determining the third near-month braking time.

[0125] In this embodiment, based on the initial values ​​of the powered descent point, the second near-moon braking time, and the four-device separation point, the flight time ΔT from near-moon braking to powered descent is calculated. LOI1toDes and the repetition period T of the ground station tracking arc VD The remainder of Determine the third near-month braking time t LOI3 Specifically:

[0126] 4.1) According to Calculate the residual deviation of the longitude of the orbital ascending node at the time of power descent

[0127]

[0128] 4.2) Based on the difference in the precession rate of the ascending node of the lunar orbit after the second and third near-lunar braking The maximum adjustable amount of the orbital ascending node longitude drift is calculated

[0129]

[0130] in, t Sep Indicates the moment of power drop. The target period of the second and third lunar proximity maneuvers, and the flight time from the second lunar proximity maneuver to the four-vehicle separation point are related. The second lunar proximity maneuver enters a 4h period elliptical orbit, the third lunar proximity maneuver enters a 2h period elliptical orbit, and the flight time from the second lunar proximity maneuver to the four-vehicle separation point is 20 days, for example.

[0131] 4.3) If then according to the third lunar proximity maneuver time t LOI3 is calculated.

[0132]

[0133] where Δλ 1day represents the angle of the moon rotation in one day, Δλ 1day = ω Moon · T VD , 13.2°, which means that the adjustment ability of the elliptical orbit ascending node can completely satisfy the adjustment of , and t LOI3 can be directly calculated according to .

[0134] At this time, the dynamic descent time t Des and the four-vehicle separation point time t Sep remain unchanged, and are as follows:

[0135]

[0136] 4.4) If This means that the maximum adjustable amount of the orbit ascending node longitude drift may not completely cover The third lunar proximity maneuver time needs to be determined according to the size relationship between the ascending node adjustment ability of the elliptical orbit and , and the specific process is as follows:

[0137] 4.4.1) For a lunar orbiting prograde orbit:

[0138] a) The corresponding lunar orbiting orbit ascending node longitude is calculated.

[0139] b) The corresponding lunar orbiting orbit ascending node longitude after one day of the dynamic descent point is calculated.

[0140] c) According to and , the east side boundary λ east of the adjustable range of the ascending node longitude and the west side boundary λ west of the adjustable ascending node longitude are calculated:

[0141] d) Calculate the target value of the ascending node longitude adjustment

[0142] e) Determine whether the following condition is met:

[0143] As shown in Figure 2 , if is met, it means that the ascending node adjustment capability of the elliptical orbit meets the adjustment requirement of , and the third perigee braking time t LOI3 is:

[0144]

[0145] At this time, the power descent time t Des and the four-device separation point time t Sep remain unchanged, and are:

[0146]

[0147] As shown in Figure 3 , if is not met, it means that exceeds the ascending node adjustment capability of the elliptical orbit, and the ascending node residual deviation needs to be adjusted in combination with the ascending node drift capability and the perigee braking normal pulse. The specific process is as follows:

[0148] If , the third perigee braking time t LOI3 is:

[0149]

[0150] wherein represents the minimum flight time from the third perigee braking to the four-device separation point, which is 1 measurement and control repetition period, i.e. This represents that the maximum ascending node longitude drift capability of the elliptical orbit is used to meet the adjustment requirement of , and the remaining unadjustable part is adjusted by the normal pulse of the first perigee braking. At this time, the power descent time t Des and the four-device separation point time t Sep remain unchanged, and are:

[0151]

[0152] The ascending node longitude δλ LOI1 that needs to be adjusted by the first perigee braking is:

[0153]

[0154] If Then the powered descent point is delayed by one day:

[0155]

[0156] Third lunar proximity braking time t LOI3 Set to:

[0157]

[0158] Where, The minimum flight time between the second lunar proximity braking and the third lunar proximity braking, which is 1 TT&C repetition period, is denoted as This means that the elliptical orbit drift is not used to adjust the longitude deviation of the ascending node, The adjustment requirement of is completely adjusted by adjusting the powered descent time combined with the normal pulse of the first lunar proximity braking.

[0159] The ascending node longitude amount δλ that needs to be adjusted by the first lunar proximity braking LOI1 Is:

[0160]

[0161] Four-vehicle separation point time t Sep The powered descent point time is updated accordingly:

[0162] t Sep = t Des -N Sep T VD

[0163] 4.4.2) For a lunar retrograde orbit:

[0164] a) Calculate the powered descent point time The corresponding lunar orbit ascending node longitude

[0165] b) Calculate one day before the powered descent point The corresponding lunar orbit ascending node longitude

[0166] c) According to And Calculate the west boundary of the adjustable range of the ascending node longitude λ west The east boundary of the adjustable range of the ascending node longitude λ east Respectively:

[0167] d) Calculate the target value of the ascending node longitude adjustment

[0168] e) Determine whether the following conditions are met:

[0169] As shown in Figure 4 , if , it means that the ascending node adjustment capability of the elliptical orbit fully meets the adjustment requirement of , and the third perigee braking time t LOI3 is:

[0170]

[0171] At this time, the power descent time t Des and the four-device separation point time t Sep remain unchanged, and are:

[0172]

[0173] As shown in Figure 5 , if , it means that exceeds the ascending node adjustment capability of the elliptical orbit, and the ascending node residual deviation needs to be adjusted in combination with the ascending node drift capability and the perigee braking normal pulse. The specific process is as follows:

[0174] If , the third perigee braking time t LOI3 is:

[0175]

[0176] Among them, represents the minimum flight time from the third perigee braking to the four-device separation point, which is 1 measurement and control repetition period, that is, This represents that the maximum drift capability of the ascending node longitude of the elliptical orbit is used to meet the adjustment requirement of , and the remaining part that cannot be adjusted is adjusted by the normal pulse of the first perigee braking. At this time, the power descent time t Des and the four-device separation point time t Sep remain unchanged, and are:

[0177]

[0178] The ascending node longitude δλ LOI1 that needs to be adjusted by the first perigee braking is:

[0179]

[0180] If , the power descent point time is advanced by one day:

[0181]

[0182] The third perigee braking time tLOI3 is:

[0183]

[0184] wherein, represents the minimum flight duration between the second perigee braking and the third perigee braking, which is 1 TT&C repetition period, i.e. This represents that the elliptical orbit drift is not used to adjust the longitude deviation of the ascending node, The adjustment requirement of is completely adjusted by adjusting the power descent time combined with the normal pulse of the first perigee braking. The first perigee braking needs to adjust the longitude amount of the ascending node LOI1 is:

[0185]

[0186] The four-vehicle separation time t Sep The power descent time t Sep is updated accordingly: t Des -N Sep T VD .

[0187] 4.5) The final output results of this step S4 are: the third perigee braking time t LOI3 , the four-vehicle separation time t Sep , the power descent time t Des , and the longitude adjustment amount of the ascending node of the first perigee braking LOI1 .

[0188] S5, determine the lunar take-off time.

[0189] In this embodiment, the lunar take-off time t Des can be calculated according to the power descent time t Asc : t Asc = t Des + ΔT LS .

[0190] S6, use the finite thrust model and the accurate prediction model to solve the orbit control strategy, and calculate the orbit parameters at the four-vehicle separation time.

[0191] In this embodiment, the key event times and the orbit parameters obtained in the above steps are used as inputs to solve the orbit control strategy by using the finite thrust model and the accurate prediction model, and the orbit parameters at the four-vehicle separation time are calculated. Specifically:

[0192] 6.1) According to the finite thrust model, the first perigee braking is applied, and the design variable is the on time ΔT LOI1, the target variable is T1, and the orbit control parameter AV of the first lunar proximity braking is obtained after iterative convergence by using the differential correction method LOI1 and the controlled orbit parameter E LOI1 .

[0193] 6.2) According to E LOI1 , the time t LOI2 of the second lunar proximity braking is predicted, the second lunar proximity braking is applied according to the finite thrust model, the design variable is the starting time length AT LOI2 , the target variable is T2, and the orbit control parameter AV of the second lunar proximity braking is obtained after iterative convergence by using the differential correction method LOI2 and the controlled orbit parameter E LOI2 .

[0194] 6.3) According to E LOI2 , the time t LOI3 of the third lunar proximity braking is predicted, the third lunar proximity braking is applied according to the finite thrust model, the design variable is the starting time length AT LOI3 , the target variable is T3, and the orbit control parameter AV of the third lunar proximity braking is obtained after iterative convergence by using the differential correction method LOI3 and the controlled orbit parameter E LOI3 .

[0195] 6.4) According to E LOI3 , the time t LOI3 of the fourth device separation is predicted, and the orbit parameter E Sep of the fourth device separation point is obtained = (a Sep , e Sep , i Sep , Ω Sep , ω Sep , f Sep ). Wherein, a Sep represents the orbit semi-major axis at the separation time, e Sep represents the eccentricity at the separation time, i Sep represents the orbit inclination at the separation time, Ω Sep represents the ascending node right ascension at the separation time, ω Sep represents the near point amplitude angle at the separation time, and f Sep represents the true near point angle at the separation time.

[0196] S7, according to the orbit parameter of the fourth device separation point, the inner loop strategy is solved.

[0197] In this embodiment, the inner loop strategy solving process is as follows:

[0198] 7.1) According to E Sep = (a Sep , e Sep , i Sep , ΩSep ,ω Sep ,f Sep ), the design variable is the latitude amplitude u Sep =ω Sep +f Sep , the on-time of the two circumlunar deorbit controls ΔT DM1 and ΔT DM2 , the target variable is the time t Des , the altitude h Des and the perigee amplitude ω Des , the control parameters ΔV DM1 and ΔV DM2 of the two circumlunar deorbit controls and the orbit parameter E Des corresponding to the time of the powered landing point are obtained by using the differential correction method to iteratively converge.

[0199] 7.2) According to E Sep = (a Sep , e Sep , i Sep , Ω Sep , ω Sep , f Sep ), the target parameters of the shift point are aimed at according to the phase modulation strategy, the design variable is the on-time of the four-phase modulation control, the target variable is the orbit semi-major axis, eccentricity and latitude amplitude at the time of the shift point, and the control parameters of the four-phase modulation and the orbit parameter E Asc corresponding to the time of the lunar surface take-off are obtained by using the differential correction method to iteratively converge.

[0200] 7.3) According to E Des and E Asc , the position vector and the orbit plane angle LPA Des and LPA Asc of the sampling point at the time of the powered landing point and the time of the lunar surface take-off are calculated:

[0201] E Des is converted into position and velocity X Des = (r Des , v Des ), further coordinate conversion is performed to obtain the position and velocity of the lunar center fixed system at the time of the powered landing point and the position vector and the orbit plane angle LPA Des of the sampling point at the time of the powered landing point are calculated:

[0202]

[0203] wherein, represents the orbit plane angular momentum of the powered landing point, represents the position vector of the sampling point,

[0204] Similarly, the angle LPA between the position vector of the sampling point at the moment of lunar takeoff and the orbital plane is calculated Asc LPA Asc The calculation method is similar to LPA Des The calculation method is similar to LPA and will not be described here.

[0205] S8, according to the result of the inner loop strategy, the outer loop strategy is solved.

[0206] In this embodiment, the outer loop strategy solving process is as follows:

[0207] 8.1) According to the task requirement of timing and fixed point landing, on-time takeoff, the orbital inclination i of the four-device separation point is selected Sep and the ascending node right ascension Ω Sep as the design variable p1, and the angle between the position vector of the sampling point at the moment of t Des and the lunar takeoff moment t Asc is 0° as the target quantity q1:

[0208]

[0209] Then, the deviation Δq1 of the target variable and the target value is calculated:

[0210]

[0211] 8.2) According to the function relationship between Δq1, design variable p1 and target variable q1: p1=f(q1); further obtain the correction amount Δp1 of the orbital inclination i Sep and the ascending node right ascension Ω Sep

[0212]

[0213] 8.3) Update the orbital inclination i Sep and the ascending node right ascension Ω Sep of the four-device separation point orbital parameter E Sep in step S6, and perform iterative calculation according to steps S6-S8 until Δq1 is less than the allowable error threshold, end the calculation, and obtain the target orbital parameter of the four-device separation point that meets the task requirement of timing and fixed point landing, on-time takeoff

[0214] 8.4) According to the lunar orbit inclination of the earth-moon transfer and the near-lunar braking strategy are solved, and the design variable p2 is the near-lunar orbit inclination i LOI1 , the thrust azimuth angle A LOI1 ​, the target value q2 of the third time near-moon braking control post-orbit period T3, the target value q2 of the latitude amplitude of the four-device separation point orbit inclination and the ascending node right ascension

[0215]

[0216] The differential correction method is used to calculate the deviation of the variable p2 from the target value q2, and then the correction amount of the design variable is obtained, the near-moon point inclination target of the earth-moon transfer orbit is updated, and the calculation is ended until the Δq2 is less than the allowable error threshold.

[0217] S9, outputting the orbit control strategy calculation result.

[0218] In the embodiment, the finally output orbit control strategy calculation result includes:

[0219] Key point orbit parameters: orbit parameters at the initial and ending moments of the earth-moon transfer, orbit parameters at the four-device separation point, orbit parameters at the power descent point, and orbit parameters at the handover point.

[0220] Orbit control parameters: near-moon braking, moon orbit descent, and orbit phasing of the orbit control on-off time, velocity increment, propellant consumption, and pre-control and post-control orbit parameters.

[0221] Although the present application has been disclosed with the above preferred embodiments, it is not intended to limit the present application, and any person skilled in the art can make possible changes and modifications to the technical solutions of the present application by using the disclosed methods and technical contents without departing from the spirit and scope of the present application. Therefore, any simple modification, equivalent change and modification made to the above embodiments according to the technical essence of the present application, which does not deviate from the technical solutions of the present application, belongs to the protection scope of the technical solutions of the present application.

[0222] The contents not described in detail in the specification of the present application are known to those skilled in the art.

Claims

1. A three-pulse near-moon braking timing and fixed-point landing aiming method, characterized in that: include: Get input parameters; Based on the input parameters, the initial values ​​of the powered descent point, the second near-moon braking time, the initial values ​​of the four-vehicle separation point, the third near-moon braking time, and the lunar surface takeoff time are determined; The finite thrust model and the precise prediction model are used to solve the orbit control strategy and calculate the orbit parameters at the separation point of the four vehicles. Solve the inner loop strategy based on the orbital parameters at the separation point of the four devices; According to the solution of the inner loop strategy, the outer layer strategy is solved; Output orbit control strategy calculation results.

2. The three-pulse near-moon braking timing and fixed-point landing aiming method according to claim 1 is characterized in that: Input parameters include: Earth-Moon transfer orbit parameters, lunar surface target sampling point parameters, powered descent point parameters, orbital period T1 after one near-moon braking, orbital period T2 after two near-moon braking, and orbital period T3 after three near-moon braking. Among them, Earth-Moon transfer orbit parameters include: initial time t0, number of orbital elements E0 at the initial time, flight time ΔT E2M , departure and arrival methods; lunar target sampling point parameters, including: sampling point latitude and longitude Working time on the moon ΔT LS ;Powered descent point parameters, including: powered descent point height h Des 、Latitude angle of dynamic descent point ω Des .

3. The three-pulse near-moon braking timing and fixed-point landing aiming method according to claim 2 is characterized in that: The initial value of the power drop point is determined as follows: according to and ΔT LS , calculate the initial value of the fixed connection inclination I0 and the initial value of the ascending node longitude λ at the time of the powered descent point to meet the lunar surface working time requirements Des0 : Among them, ω Moon represents the angular velocity of the moon's rotation; According to E0 and ΔT E2M , predict and calculate the orbital element E when the Earth-Moon transfer reaches the perigee f , and then convert it to obtain the orbital position velocity X in the lunar center inertial system LOI1 =(r LOI1 ,v LOI1 );According to the perigee time t f =t0+ΔT E2M and X LOI1 , coordinate transformation is performed to obtain the orbital position velocity of the lunar center fixed connection at the time of perigee pass The conversion of the orbital elements yields the fixed system inclination i at the perigee LOI1 and the longitude of the ascending node λ LOI1 ; Solve to obtain the flight time Δλ from the near-month braking to the powered descent LOI1toDes : Dl LOI1toDes =λ Des0 -l LOI1 According to T2 and T3, the precession rate of the ascending node of the lunar orbit after the second near-moon braking is calculated respectively. and the precession rate of the ascending node of the lunar orbit after the third near-moon braking According to ω Moon and Solve to obtain the flight time ΔT from near-month braking to powered descent LOI1toDes : Among them, k is the forward and backward flag, the forward track k is -1, and the backward track k is 1; Calculate the initial value of the powered descent point according to different lunar orbit types: For a prograde lunar orbit, ΔT LOI1toDes The repetition period T of the ground station tracking arc VD Round down and take the remainder to get the integer part and the remainder The initial value of the power drop point at this time as follows: For a retrograde lunar orbit, ΔT LOI1toDes T VD Round up and take the remainder to get the integer part and the remainder The initial value of the power drop point at this time as follows: Among them, t LOI1 Indicates the first near-month braking moment.

4. The three-pulse near-moon braking timing and fixed-point landing aiming method according to claim 3 is characterized in that: The second near-moon braking time is: one day after the first near-moon braking and the moment when the vehicle enters the common viewing arc of the ground tracking and control station; where t LOI2 =t LOI1 +T VD , t LOI2 Indicates the second near-month braking moment.

5. The three-pulse near-moon braking timing and fixed-point landing aiming method according to claim 4 is characterized in that: The initial value of the four-device separation point is: N before power drops Sep day, and is within the common view arc of the ground tracking and control station; among them, Indicates the initial value of the four-device separation point, N Sep The value range is 1 to 3 days.

6. The three-pulse near-moon braking timing and fixed-point landing aiming method according to claim 5 is characterized in that: The third near-month braking time is determined as follows: according to Calculate the residual deviation of the longitude of the orbital ascending node at the time of power descent According to the difference in the precession rate of the ascending node of the lunar orbit after the second and third near-moon braking, The maximum adjustable amount of the orbital ascending node longitude drift is calculated in, t Sep Indicates the moment of power drop point; like According to The third near-month braking time t is calculated LOI3 : At this time, the power drops at the moment t Des and the separation point of the four devices t Sep Remain unchanged as: Where Δλ 1day Indicates the angle of the moon's rotation in one day; like but: For prograde orbits around the Moon: Calculated Corresponding longitude of the ascending node of the lunar orbit One day after the power drop point is calculated Corresponding longitude of the ascending node of the lunar orbit according to and Calculate the eastern boundary λ of the adjustable range of the ascending node longitude east and the western boundary of the ascending node longitude adjustable λ west : Calculate the target value of the ascending node longitude adjustment Determine whether the following conditions are met: If satisfied The third near-month braking time t LOI3 for: At this time, the power drops at the moment t Des and the separation point of the four devices t Sep Remain unchanged as: If not satisfied but: like The third near-month braking time t LOI3 for: At this time, the power drops at the moment t Des and the separation point of the four devices t Sep Remain unchanged as: The ascending node longitude δλ that needs to be adjusted for the first near-moon braking LOI1 for: in, Indicates the minimum flight time between the third near-moon braking and the separation point of the four vehicles; like Then postpone the power drop point by one day: The third near-month braking time t LOI3 Set to: The ascending node longitude δλ that needs to be adjusted for the first near-moon braking LOI1 for: The separation point of the four devices is t Sep The corresponding update of the power drop point is: t Sep =t Des -N Sep T VD ;in, Indicates the minimum flight time between the second near-moon braking and the third near-moon braking; For retrograde orbits around the Moon: Calculate the moment of power descent point Corresponding longitude of the ascending node of the lunar orbit Calculate the day before the power drop point Corresponding longitude of the ascending node of the lunar orbit according to and Calculate the western boundary λ of the adjustable range of the ascending node longitude west , the eastern boundary of the adjustable longitude of the ascending node λ east They are: Calculate the target value of the ascending node longitude adjustment Determine whether the following conditions are met: If satisfied The third near-month braking time t LOI3 for: At this time, the power drops at the moment t Des and the separation point of the four devices t Sep Remain unchanged as: If not satisfied but: like The third near-month braking time t LOI3 for: At this time, the power drops at the moment t Des and the separation point of the four devices t Sep Remain unchanged as: The ascending node longitude δλ that needs to be adjusted for the first near-moon braking LOI1 for: in, Indicates the minimum flight time between the third near-moon braking and the separation point of the four vehicles; like Then the power drop point time is advanced by one day: The third near-month braking time t LOI3 for: The ascending node longitude δλ that needs to be adjusted for the first near-moon braking LOI1 for: The separation point of the four devices is t Sep The corresponding update of the power drop point is: t Sep =t Des -N Sep T VD ;in, Indicates the minimum flight time between the second near-moon braking and the third near-moon braking.

7. The three-pulse near-moon braking timing and fixed-point landing aiming method according to claim 6 is characterized in that: According to the dynamic descent point time t Des , calculate the lunar takeoff time t Asc :t Asc =t Des +ΔT LS .

8. The three-pulse near-moon braking timing and fixed-point landing aiming method according to claim 7 is characterized in that: The finite thrust model and the precise prediction model are used to solve the orbit control strategy and calculate the orbit parameters at the separation point of the four vehicles, including: According to the finite thrust model, the first near-month braking is applied, and the design variable is the startup time ΔT LOI1 The target quantity is T1. After iterative convergence using the differential correction method, the orbit control parameter ΔV for the first near-moon braking is obtained. LOI1 and the orbital parameter E after control LOI1 ; According to E LOI1 Forecast to the second near-month braking time t LOI2 Stop, apply the second near-month braking according to the finite thrust model, and the design variable is the startup time ΔT LOI2 , the target variable is T2, and after iterative convergence using the differential correction method, the orbit control parameter ΔV of the second near-moon braking is obtained LOI2 and the orbital parameter E after control LOI2 ; According to E LOI2 Forecast to the third near-month braking time t LOI3 , according to the limited thrust model, the third near-month braking is applied, and the design variable is the startup time ΔT LOI3 The target variable is T3. After iterative convergence using the differential correction method, the orbit control parameter ΔV for the third near-moon braking is obtained. LOI3 and the orbital parameter E after control LOI3 ; According to E LOI3 Forecast to the time of separation of the four devices t LOI3 , obtain the orbital parameters E at the separation point of the four devices Sep =(a Sep ,e Sep ,i Sep ,Ω Sep ,ω Sep ,f Sep ); where a Sep represents the semi-major axis of the orbit at the moment of separation, e Sep represents the eccentricity at the separation moment, i Sep represents the orbital inclination at the separation moment, Ω Sep represents the right ascension of the ascending node at the moment of separation, ω Sep represents the argument of perigee at the moment of separation, f Sep Indicates the true anomaly at the moment of separation.

9. The three-pulse near-moon braking timing and fixed-point landing aiming method according to claim 8 is characterized in that: According to the orbital parameters at the separation point of the four devices, the inner loop strategy is solved, including: According to E Sep =(a Sep ,e Sep ,i Sep ,Ω Sep ,ω Sep ,f Sep ), aiming at the target parameters of the powered descent point according to the lunar orbit descent strategy, the design variable is the latitude argument u of the four-device separation point Sep =ω Sep +f Sep , the start-up time of two lunar orbit reduction control operations ΔT DM1 and ΔT DM2 , the target variable is the moment t at which the power drops Des , height h Des and the argument of perigee ω Des After iterative convergence using the differential correction method, the control parameters ΔV for the two lunar orbit reductions are obtained. DM1 and ΔV DM2 , orbital parameter E corresponding to the moment of dynamic descent Des ; According to E Sep =(a Sep ,e Sep ,i Sep ,Ω Sep ,ω Sep ,f Sep ), aiming at the target parameters of the handover point according to the phase adjustment strategy, the design variable is the start-up time of the four-phase adjustment control, the target variable is the orbit semi-major axis, eccentricity and latitude angle at the handover point, and the differential correction method is used to iteratively converge to obtain the control parameters of the four-phase adjustment and the orbit parameter E corresponding to the lunar takeoff time. Asc ; According to E Des and E Asc Calculate the angle LPA between the position vector of the sampling point and the orbital plane at the time of powered descent and lunar takeoff Des and LPA Asc : E Des Convert to position speed X Des =(r Des ,v Des ), and further coordinate transformation is performed to obtain the position velocity of the lunar center fixed connection at the time of the dynamic descent point And calculate the angle LPA between the position vector of the sampling point and the orbital plane at the time of power descent Des : in, represents the angular momentum of the orbital plane at the power descent point, represents the sampling point position vector, Similarly, the angle LPA between the position vector of the sampling point and the orbital plane at the time of takeoff from the lunar surface is calculated Asc .

10. The three-pulse near-moon braking timing and fixed-point landing aiming method according to claim 9 is characterized in that: According to the results of the inner loop strategy solution, the outer layer strategy solution is performed, including: According to the mission requirements of landing at a fixed time and taking off on time, the orbital inclination angle i of the separation point of the four vehicles is selected. Sep and the right ascension of the ascending node Ω Sep As the design variable p1, the power drop point time t Des and the lunar takeoff time t Asc The angle between the sampling point position vector and the orbital plane is 0°, which is the aiming target quantity q1: The deviation Δq1 between the target variable and the target value is calculated: According to the functional relationship between Δq1, design variable p1 and target variable q1: p1 = f(q1); further obtain the orbit inclination i Sep and the right ascension of the ascending node Ω Sep The correction value Δp1: Update the orbital parameters E of the four-device separation point Sep The orbital inclination i Sep and the right ascension of the ascending node Ω Sep , iterative calculation, until Δq1 is less than the allowed error threshold, end the calculation, and obtain the target orbit parameters of the four-vehicle separation point that meet the requirements of landing at a fixed time and taking off on time. according to The perilunar orbit inclination and perilunar braking strategy of the Earth-Moon transfer are solved. The design variable p2 is the perilunar orbit inclination i LOI1 , thrust azimuth angle A of the first near-moon braking LOI1 The post-control orbit period T3 of the third near-moon braking, the target quantity q2 is the latitude argument of the four-device separation point orbital inclination and right ascension of the ascending node The differential correction method is used to calculate the deviation between the variable p2 and the target value q2, and then the correction amount of the design variable is obtained. The target perigee inclination of the Earth-Moon transfer orbit is updated until Δq2 is less than the allowable error threshold and the calculation is terminated.

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