Mars atmospheric braking strategy design method

Abstract

The invention provides a Mars atmospheric braking strategy design method, which is characterized in that atmospheric braking is divided into at least three stages: a step-in stage, a main deceleration stage and a step-out stage in sequence, and comprises the following steps: designing an atmospheric braking corridor in each far fire point maneuvering cycle, and determining upper and lower boundaries of the corridor; the step-in section adopts an orbit control strategy that a near fire point is gradually reduced, and the near fire point is reduced to the upper boundary of the height of the current atmospheric braking corridor from a ring fire large elliptical orbit; the main deceleration section ensures that the near fire point height is within the current atmospheric braking corridor range, and when the far fire point height reaches the far fire point height of the target transition track, the step-out section is entered; the step-out section adopts an orbit control strategy of gradually lifting near a fire point; and when the near fire point speed reaches the near fire point speed corresponding to the task track after the last track lifting maneuver, rounding maneuver is implemented at the far fire point, so that the detector enters the task track.

Description

TECHNICAL FIELD

[0001] The application belongs to the technical field of deep space exploration, and relates to a Mars atmospheric braking strategy design and orbit correction method. BACKGROUND

[0002] Atmospheric braking is a type of aerodynamic auxiliary orbit transfer, and plays a very important role in future interplanetary exploration as a low-cost orbit transfer technology. Among the eight planets in the solar system, except Mercury, the remaining planets have an atmosphere, and using planetary atmosphere deceleration is an efficient and economical technical means to obtain speed increment.

[0003] In 1961, H. London first demonstrated the significance and feasibility of aerodynamic auxiliary orbit transfer in a paper at the American Astronautical Society. The concept of aerodynamic auxiliary orbit transfer is to combine pure impulse orbit transfer and aerodynamic orbit transfer, insert an atmospheric flight segment in the orbit transfer flight, change the orbit plane or altitude by aerodynamic force, and finally complete the entire orbit transfer requirements with minimum energy consumption and other technical indicators.

[0004] Compared with the traditional orbit transfer mode, the biggest advantage of atmospheric braking orbit transfer is that the probe can use atmospheric resistance to consume orbit energy and change the orbit plane, thereby greatly saving fuel and increasing the mass of the carried payload. For a Mars exploration mission, even if the most fuel-efficient scheme is used for orbit transfer, the required speed increment is still large, and the probe needs to carry a large amount of fuel using the traditional orbit transfer mode, which reduces the dry weight of the probe and increases the design difficulty of the probe.

[0005] The difficulty of atmospheric braking orbit transfer technology is that in order to obtain a certain speed increment, the flight altitude must be reduced, and at this time the flight trajectory needs to be strictly controlled to prevent the probe from being damaged due to excessive thermal environment and overload, and the orbit deviates too much due to excessive energy loss, and even cannot fly out of the atmosphere. At the same time, if the flight altitude is too high, the deceleration efficiency is too low, and the deceleration cannot be completed within the specified time, which affects the implementation of subsequent tasks. SUMMARY

[0006] The technical problem solved by the present application is to provide a Mars atmospheric braking strategy design and orbit correction method, which solves the problem of simultaneously meeting the task time constraint and safety during the atmospheric braking process in a Mars exploration mission.

[0007] The technical solution of the present application is: a Mars atmospheric braking strategy design method, the atmospheric braking is divided into at least three stages in turn: entry segment, main deceleration segment and exit segment, comprising:

[0008] Design an atmospheric braking corridor in each distant Mars point maneuver period, and determine the upper and lower boundaries of the corridor;

[0009] The step-in section adopts a track control strategy of gradually lowering the perihelion, and lowers the perihelion from a circular perihelion elliptical orbit to the upper boundary of the current atmospheric braking corridor;

[0010] The main deceleration section ensures that the perihelion height is within the current atmospheric braking corridor, and when the aphelion height reaches the aphelion height of the target transition orbit, the step-out section is entered;

[0011] The step-out section adopts a track control strategy of gradually raising the perihelion; after the last raising maneuver, the perihelion velocity reaches the perihelion velocity corresponding to the mission orbit, and a circularization maneuver is performed at the aphelion to make the probe enter the mission orbit.

[0012] Preferably, the upper and lower boundaries of the corridor are determined by:

[0013] According to the mission time constraint t max , the upper boundary h p,max of the corridor height is determined by a shooting method, to ensure that the predicted value of the current circle heat flux density is not greater than the heat flux density constraint;

[0014] According to the upper limit of the axial acceleration, the local atmospheric density is calculated, and the local atmospheric density is combined with the atmospheric prediction model to inversely solve the lower boundary h p,min of the corridor;

[0015] Considering the height error Δh p导航 caused by factors such as navigation error, orbit control error, atmospheric model error, atmospheric density daily fluctuation rate, and accelerometer measurement error, Δh p轨控 , Δh p大气模型 , Δh p日波动 , Δh p加计 , the corresponding perihelion height error ΔH is determined;

[0016] According to the perihelion height error ΔH, the upper and lower boundaries of the corridor considering the safety margin are determined.

[0017] Preferably,

[0018] Preferably, the lower boundary h p,min ' of the corridor considering the safety margin is h p,min + ΔH;

[0019] The upper boundary h p,max ' of the corridor considering the safety margin is max(h p,max , h p,min + 2ΔH).

[0020] Preferably, the track control strategy of gradually lowering the perihelion includes:

[0021] S1, the probe freely flies N1 circles, simultaneously completes orbit determination, and then implements a first step-in deorbiting maneuver at a far perigee point to reduce the perigee point height to the minimum resolution height h of the accelerometer p0 ;

[0022] S2, continues to freely fly N1 circles, simultaneously completes orbit determination, and then implements a second step-in deorbiting maneuver at a far perigee point to reduce the perigee point height to the upper boundary h of the atmospheric braking corridor p,max ’; N1≥2.

[0023] Preferably, the number of deorbiting maneuvers is determined according to orbit perturbation, considering the safety and time constraints of atmospheric braking flight. When the orbit perturbation is large (generally, the orbit inclination is less than 80° or greater than 100°), the perigee point height is reduced to the upper boundary of the atmospheric braking corridor through the S1 and S2 maneuvers, and then it is determined whether to increase an orbit correction maneuver to finely adjust the perigee point height according to the orbit control accuracy, so that the perigee point height requirement is met. When the orbit perturbation is small (generally, the near-polar orbit with an orbit inclination greater than or equal to 80° and less than or equal to 100°), the deorbiting maneuvers can be evenly divided into multiple times, and the high-altitude atmospheric model can be calibrated and corrected during this period, but the total number of step-in deorbiting maneuvers should not exceed 10.

[0024] Preferably, the step-out strategy reduces the orbit maneuver consumption as much as possible under the premise of ensuring the safe lifting of the perigee point height and not exceeding the longest atmospheric braking flight time constraint. Every N 自由 circle of free flight is followed by a perigee point lifting maneuver, and a total of N2 times are performed. After each maneuver, the perigee point height is lifted by h p步出 ; the above N 自由 ≥5, and the number of step-out lifting maneuvers N2 should not exceed 10;

[0025] h p步出 =(h j -h af ) / N2, h j is the perigee point height when the perigee point speed of the mission orbit is reached, and h af is the perigee point height of the target transfer orbit.

[0026] Preferably, the perigee point height of the target transfer orbit is recommended to be in the range of 800-2000 km.

[0027] Preferably, the main deceleration section is processed in the following manner:

[0028] It is determined whether the current perigee point has been freely flown for more than N3 circles since the last orbit control. If yes, the upper and lower boundaries of the corridor and the target perigee point height are calculated, and subsequent processing is performed. If not, no orbit change is performed.

[0029] The detector predicts the current perihelion height, judges whether the height is within the corridor upper and lower boundary range, and if so, does not perform orbit transfer; otherwise, performs a lift orbit maneuver or a drop orbit maneuver at the aphelion to make the current perihelion height to the target perihelion height;

[0030] N3 is the aphelion maneuver period.

[0031] Preferably, the aphelion maneuver period is not less than 2 revolutions; the target perihelion height of each maneuver of the main deceleration section at the aphelion is h p,目标 =h p,min +3(Δh p引力 +Δh p大气 ); h p,min ' is the current corridor lower boundary; Δh p引力 +Δh p大气 is the perihelion perturbation in one aphelion maneuver period.

[0032] The present application has the following beneficial effects compared with the prior art:

[0033] A Mars atmospheric braking strategy design and orbit correction method is proposed, which saves the fuel mass carried by the probe and increases the effective payload mass that can be carried. The following effects are achieved:

[0034] (1) The atmospheric braking process is divided into three stages, and an operable orbit control strategy is proposed for each stage.

[0035] (2) An atmospheric braking corridor design method is proposed to determine the upper and lower boundaries of the corridor, which is easy to implement.

[0036] (3) A method for determining the maneuver target perihelion height and period is proposed, so that the atmospheric braking orbit obtained by the method is close to the minimum total time and maintains a small number of orbit control maneuvers, thereby saving fuel. BRIEF DESCRIPTION OF DRAWINGS

[0037] Figure 1 It is a flowchart of the main deceleration section strategy. DETAILED DESCRIPTION

[0038] The present application will be further described below in conjunction with the embodiments.

[0039] The present application relates to a Mars atmospheric braking strategy design method, which needs to determine the maximum heat flux density constraint q max , the total mission time constraint t max and the atmospheric prediction model F 大气模型 of the mission period according to the overall scheme of the Mars exploration mission before the method:

[0040] ρ=F 大气模型 (dd,tt,h,λ,ψ) (1)

[0041] where p is the atmospheric density, dd is the date, tt is the local time, h is the altitude, and l is the longitude, is the latitude.

[0042] The probe is adjusted to a circular orbit with an orbit period of 1-3 days using a conventional propulsion braking method, and a de-orbiting maneuver is performed at the far apsis to reduce the altitude of the near apsis to the minimum resolution height h p0 of the accelerometer, and atmospheric braking is started. The atmospheric braking is divided into at least three stages: a step-in stage, a main deceleration stage, and a step-out stage. The method comprises:

[0043] 1. Designing an atmospheric braking corridor at each far apsis maneuver period, and determining the upper and lower boundaries of the corridor.

[0044] Far apsis maneuver period determination: Since the far apsis maneuver needs to be performed after the determination orbit is completed, and the need to simultaneously consider the need to save chemical fuel, the far apsis maneuver period should be determined in combination with the minimum period of the determination orbit and the minimum speed increment requirement of the orbit control. The maneuver period is generally not less than 2 orbits.

[0045] The criterion for determining the atmospheric braking corridor is to ensure the safety of the probe on the basis of ensuring the safety of the probe, and to make the total atmospheric braking time meet the mission time constraint.

[0046] Corridor altitude lower boundary design: The lower boundary of the corridor should be determined to ensure that the predicted value of the circumferential heat flux density is not greater than the heat flux density constraint q max . The theoretical calculation formula of the probe's windward surface heat flux density q is:

[0047] q = 0.5C H p v 3 (2)

[0048] where C H is the heat transfer coefficient at the location, p is the atmospheric density, and v is the flight speed of the probe relative to the atmosphere.

[0049] The axial acceleration a x measured using the accelerometer and the total velocity v of the Mars-fixed coordinate system obtained by integrating the three-axis acceleration can be used to estimate the local atmospheric density

[0050]

[0051] where m is the mass of the probe, S is the reference area, and C A is the axial force coefficient.

[0052] The probe's windward surface heat flux

[0053]

[0054] If the heat flux q is required to be less than the heat flux constraint q max , then

[0055]

[0056] Therefore, the acceleration must satisfy the constraint

[0057]

[0058] From the above equation, the upper limit of the axial acceleration a x,max is gradually increased with the decrease of the speed, that is, the lower boundary of the atmospheric braking corridor height decreases with time. Substitute a x,max into equation (3), and use the atmospheric prediction model F 大气模型 to inversely solve the lower boundary of the corridor height h p,min .

[0059] Corridor height lower boundary margin design: the height errors caused by the navigation error, orbit control error, atmospheric model error, atmospheric density daily fluctuation rate, accelerometer measurement error, etc. are Δh p导航 , Δh p轨控 , Δh p大气模型 , Δh p日波动 , Δh p加计 , respectively. According to this, the corresponding perihelion height error is determined

[0060]

[0061] Then the corridor height lower boundary h p,min ' considering the safety margin is h p,min + ΔH.

[0062] Wherein, the navigation error Δh p导航 , the orbit control error Δh p轨控 are determined according to the control system, the atmospheric model error Δh p大气模型 and the atmospheric density daily fluctuation rate Δh p日波动 are determined according to the atmospheric prediction model F 大气模型 and the flight measurement results, and the accelerometer measurement error Δh p加计 is determined according to the accelerometer index.

[0063] Corridor height upper boundary estimation: using the orbit mechanics simulation tool, according to the task time constraint t max , the corridor height upper boundary h p,max is determined by the shooting method. Then the corridor height upper boundary h p,max ' considering the safety margin is max(h p,max , h p,min + 2ΔH).

[0064] 2. Entry strategy

[0065] To mitigate risks arising from biases in the Martian upper atmosphere model and atmospheric changes, the approach phase employs a gradual descent strategy from near Mars. Specifically, the probe will perform N1 free-flight orbits (N1≥2) while completing its orbit determination. Then, it will execute its first approach-descent maneuver from the exoplanet to the altitude corresponding to the minimum resolution of the accelerometers. p0 Subsequently, it continued free flight for N1 orbits, completing the orbit determination, and then performed a second entry and descent maneuver at the far fire point, lowering its altitude to the upper boundary h of the current atmospheric braking corridor. p,max For atmospheric braking flight safety and time constraints, the number of descent maneuvers is determined based on the orbital perturbation. When the orbital perturbation is large (generally when the orbital inclination is less than 80° or greater than 100°), two maneuvers are required to lower the near-fire altitude to the upper boundary of the atmospheric braking corridor. Subsequently, depending on the orbital control accuracy, it is decided whether to add an orbital correction maneuver to fine-tune the near-fire altitude to meet the near-fire altitude requirements (see step 1 for details). When the orbital perturbation is small (generally when the orbital inclination is greater than or equal to 80° and less than or equal to 100° for near-polar orbits), the descent maneuvers can be divided into multiple equal parts. During this period, the upper atmosphere model can be calibrated and corrected, but the total number of descent maneuvers during the entry phase should not exceed 10.

[0066] 3. Main deceleration phase strategy

[0067] like Figure 1 As shown, the main deceleration section is processed according to the following steps:

[0068] Determine if the current distant fire point has been free-flying for more than N3 orbits since the last orbit control. If so, calculate the upper and lower boundaries of the corridor and the altitude of the target near fire point; then perform subsequent processing; otherwise, do not perform orbit change.

[0069] The detector predicts the current near-fire point height and determines whether the height is within the upper and lower boundaries of the corridor. If so, no orbit change is performed; otherwise, a raising or lowering maneuver is performed at the far-fire point to bring the current near-fire point height to the target near-fire point height.

[0070] N3 is the long-range fire control maneuver cycle.

[0071] Determining the near-fire height of a maneuvering target from a distant firing point: Taking the two-body problem as an example (similar relationships exist in more precise orbital mechanics models), the near-fire height h... p It can be represented as

[0072] h p =a(1-e)-R E (8)

[0073] In the formula, a is the semi-major axis, e is the eccentricity, and R... Eis the Mars reference radius. Therefore, the perihelion height rate of change can be expressed as

[0074]

[0075] i.e. the perihelion height h p affected by the semi-major axis a and the eccentricity e. The average effect of the gravitational perturbation on the semi-major axis a and the eccentricity e is 0, but it has a long-term effect on the longitude of the ascending node Ω and the perihelion argument ω, which is

[0076]

[0077] where J2 is the gravitational perturbation term, is the mean motion, and i is the inclination of the orbit. Although the gravitational perturbation does not directly affect the long-term changes of the semi-major axis a and the eccentricity e, the change of the perihelion argument ω causes the perihelion position to move within the orbital plane, thus the perihelion height h p may change due to the topography.

[0078] The effect of atmospheric drag on the orbital decay can be expressed as

[0079]

[0080]

[0081] where C D is the drag coefficient, S is the reference area, m is the mass, p is the atmospheric density, and m is the Mars gravitational constant, is the velocity, and r is the heliocentric distance. The atmospheric drag causes the semi-major axis a and the eccentricity e to decrease simultaneously, which further affects the perihelion height h p .

[0082] where the drag coefficient C D is calculated by aerodynamic methods. The atmospheric density p is obtained by atmospheric models, which is related to many factors such as altitude, season, day and night, solar activity, magnetic field, etc., and has large uncertainty. Commonly used Mars atmospheric models include MarsGRAM and MCD.

[0083] Through the above orbital perturbation model, the perihelion perturbation amount Dh p引力 + Dh p大气 can be predicted within a maneuver period, and the target perihelion height of each maneuver at the apohelion of the main deceleration phase is h p,目标 = h p,min ’ + 3(Dh p引 force + Dh p atmosphere).

[0084] Main deceleration phase trajectory correction method: To improve the accuracy and safety of trajectory maneuvers, during the free flight phase of the probe, atmospheric density can be retrieved in real time using measurement data from heat flow meters and / or accelerometers installed on the probe near the fire point. Based on the atmospheric retrieval data, the atmospheric model is corrected, and the target fire point altitude h is adjusted. p,目标 This leads to the correction of the maneuver speed increment of the far-field fire point. Among them, the atmospheric density inversion method based on the heat flow meter is derived from equation (2), and the atmospheric density inversion method based on the accelerometer is shown in equation (3).

[0085] The atmospheric model correction based on atmospheric inversion data refers to using k measurement data ρ 测量 Compared with atmospheric model output ρ 模型 By comparison, the systematic error of the atmospheric model is obtained through statistical methods. (Specific factors such as time, latitude and longitude, altitude, and other model parameters need to be considered), and then the atmospheric density is corrected to ρ'=ρ+Δρ.

[0086] 4. Step-out strategy

[0087] To ensure safe elevation of the near-fire point and to avoid exceeding the maximum atmospheric braking flight time constraint, while minimizing orbital maneuvering energy consumption, fully utilizing atmospheric braking efficiency, and seamlessly connecting with subsequent orbital maneuvers for the Mars circumduction mission, a gradual elevation of the near-fire point is employed during the walkout phase. Specifically, when the far-fire point altitude drops to the far-fire point altitude h of the target transition orbit... af (e.g., 1000km) per N 自由 After free flight (N_free ≥ 5), perform one high-arc maneuver, for a total of N2 times. After each maneuver, the near-fire altitude increases by h. p步出 After the final orbit-lifting maneuver, when the near-fire velocity reaches the corresponding near-fire velocity on the mission orbit, a circular maneuver is performed at the far-fire point to guide the probe into the mission orbit. The number of orbit-lifting maneuvers N2 during the step-out phase shall not exceed 10. p步出 =(h j -h af ) / N2,h j To achieve the proximity altitude corresponding to the proximity velocity on the mission trajectory, h af The altitude of the far-end of the target transition trajectory.

[0088] Example

[0089] The invention will be further illustrated using the Mars Reconnaissance Orbiter (MRO) detector as an example. The MRO detector has a mass of 1225 kg and a reference area of ​​36 m². 2 The drag coefficient is 2.2, the large-area heat transfer coefficient of the solar array is 0.8, the initial near-fire velocity is 4806 m / s, the final near-fire velocity is 3622 m / s, and the orbital inclination is 93°.

[0090] According to the overall scheme of Mars exploration mission, the maximum heat flux density constraint of the probe is determined to be 3 kW / m 2 , the total time constraint of the mission is 180 days, and the atmospheric prediction model is MarsGRAM 2000.

[0091] The probe is adjusted to a circum-Mars elliptical orbit with an orbital period of 1.5 days and a perimars point of 333 km by using a conventional propulsion braking mode, and a deorbiting maneuver is performed at the apomars point to reduce the perimars point height to 147 km.

[0092] Atmospheric braking corridor design: the criterion for determining the atmospheric braking corridor is that the total atmospheric braking time meets the mission time constraint on the basis of ensuring the safety of the probe.

[0093] Corridor height lower boundary design: the heat flux density constraint of 3 kW / m 2 is substituted into equation (6) to obtain the upper limit of the axial acceleration of 5×10 -5 m / s 2 . Substituting it into equation (3) and using the atmospheric prediction model MarsGRAM 2000 to inversely solve the corridor height lower boundary, the initial value is 100.8 km and the final value is 97.7 km.

[0094] Corridor height lower boundary margin design: the perimars point height error of 0.3 km is determined by factors such as navigation error, orbit control error, atmospheric model error, atmospheric density daily fluctuation rate, accelerometer measurement error, etc. Therefore, the initial value of the corridor height lower boundary considering the safety margin is 101.1 km and the final value is 98.0 km.

[0095] Corridor height upper boundary estimation: the corridor height upper boundary is determined by using the orbit mechanics simulation tool and the target shooting method according to the mission time constraint. Therefore, the corridor height upper boundary considering the safety margin is 113.5 km.

[0096] Apomars point maneuver period determination: when the (initial) orbital period is not less than 12 hours, the maneuver period is ≥2; when the orbital period is not greater than 4 hours, the maneuver period is ≥4.

[0097] Entry segment strategy design: the probe freely flies for 2 orbits while completing orbit determination, and then performs the first entry deorbiting maneuver at the apomars point to reduce the perimars point height to 130 km, which is the height corresponding to the minimum resolution of the accelerometer. Subsequently, the probe continues to freely fly for 2 orbits while completing orbit determination, and then performs the second entry deorbiting maneuver at the apomars point to reduce the perimars point height to 105 km, which is the upper boundary of the atmospheric braking corridor. The number of entry deorbiting maneuvers is 3.

[0098] Main deceleration segment strategy design: the processing is performed according to Figure 1 .

[0099] The far point maneuver target near point altitude determination: according to the gravitational perturbation model MGS85F2.GRV and the atmospheric perturbation model MarsGRAM 2000, the near point perturbation in a maneuver cycle is predicted to be 0.1km, then the initial height of the far point of the main deceleration section of the maneuver target is 101.4km, and the final height is 98.3km.

[0100] Orbit correction method: in order to improve the accuracy and safety of the orbit transfer maneuver, the atmospheric density near the near point is inversed in real time by using the heat flow meter and / or accelerometer installed on the probe, and the atmospheric density data obtained by the two inversion methods is fitted, and the far point maneuver velocity increment is corrected based on the fitting result.

[0101] Step-out section strategy design: when the far point altitude is reduced to the far point altitude 1000km of the target transition orbit, a far point orbit-raising maneuver is performed every 5 free flight circles, and the near point altitude is raised by about 40km after each maneuver. After the last orbit-raising maneuver, the near point velocity reaches the near point velocity corresponding to the mission orbit, and the probe enters the mission orbit by implementing a circularization maneuver at the far point. The number of orbit-raising maneuvers in the step-out section is 4 times.

[0102] 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.

Claims

1. A method for designing a Mars atmospheric braking strategy, said atmospheric braking being divided into at least three phases in succession: an entry phase, a main deceleration phase and an exit phase, characterized in that The method comprises the following steps: an atmospheric braking corridor is designed for each far point maneuver period, and upper and lower boundaries of the corridor are determined; a step-in section adopts a track control strategy of gradually reducing the near point, and the near point is reduced from the near point of the near-earth elliptical orbit to the upper boundary of the atmospheric braking corridor; the main deceleration section ensures that the near point height is within the range of the atmospheric braking corridor, and when the far point height reaches the far point height of the target transition orbit, a step-out section is entered; the step-out section adopts a track control strategy of gradually raising the near point, and after the last raising maneuver, the near point speed reaches the near point speed corresponding to the target orbit, and a circularization maneuver is performed at the far point to make the probe enter the target orbit.

2. The method of claim 1, wherein: The upper and lower boundaries of the corridor are determined in the following manner: According to the task time constraint t max , the corridor height upper boundary h p,max is determined by the targeting method, and it is ensured that the current cycle heat flux density prediction value is not greater than the heat flux density constraint; The local atmospheric density is calculated from the upper limit of the axial acceleration, and the lower boundary h under the corridor is inversely solved by combining the local atmospheric density with an atmospheric prediction model p,min ; Considering the altitude error Δh caused by navigation error, orbit control error, atmospheric model error, daily fluctuation rate of atmospheric density, and accelerometer measurement error. p导航 ,Δh p轨控 ,Δh p大气模型 ,Δh p日波动 ,Δh p加计 Based on this, the corresponding near-fire point height error ΔH is determined; The upper and lower boundaries of the corridor considering a safety margin are determined according to the near point height error ΔH.

3. The method of claim 2, wherein:

4. The method of claim 2 or 3, wherein: Corridor lower boundary h considering safety margin p,min ’ = h p,min + ΔH; Corridor upper boundary h considering safety margin p,max h = max(h p,max , h p,min + 2ΔH).

5. The method of claim 1, wherein: The track control strategy of gradually reducing the near point specifically comprises: S1, the probe free flight N1 circle, while completing the determination of the track, and then in the far fire point to implement the first step into the de-orbiting maneuver, the near fire point height to the accelerometer minimum resolution corresponding to the height h p0 ; S2, continue free flight for N1 orbits while completing the measurement orbit, then perform a second step-in de-orbit maneuver at the far point to reduce the far point altitude to the upper bound of the atmospheric braking corridor height h p,max N1≥ 2.

6. The method of claim 5, wherein: For the safety of atmospheric braking flight and time constraints, the number of lowering maneuvers is determined according to the orbit perturbation, when the orbit perturbation is large, that is, the orbit inclination is less than 80° or greater than 100°, the near point height is lowered to the upper boundary of the atmospheric braking corridor through two maneuvers S1 and S2, and then it is determined whether to increase an orbit correction maneuver to finely adjust the near point height to meet the near point height requirement; when the orbit perturbation is small, that is, the near polar orbit has an orbit inclination greater than or equal to 80° and less than or equal to 100°, the lowering maneuvers are evenly divided into multiple times, and the total number of lowering maneuvers in the step-in section should not exceed 10.

7. The method of claim 1, wherein: The step-out phase strategy, while ensuring a safe elevation to near-fire altitude and not exceeding the maximum atmospheric braking flight time constraint, minimizes orbital maneuvering energy consumption per N. 自由 After free flight, perform one high-point-of-fire maneuver, for a total of N2 maneuvers. After each maneuver, the near-fire altitude increases by h. p步出 The above N 自由 ≥5, the number of track lifting maneuvers N2 in the step-out section does not exceed 10 times; h p步出 = (h j -h af ) / N2, h j is the perigee height of the mission orbit corresponding to the perigee velocity, h af is the apogee height of the target transfer orbit.

8. The method of claim 1 or 7, wherein: The far point height of the target transition orbit is recommended to be in the range of 800-2000 km.

9. The method of claim 1, wherein: The main deceleration section is processed in the following manner: It is determined whether the current far point distance from the last track control is more than N3 circles of free flight, if yes, the upper and lower boundaries of the corridor and the target near point height are calculated, and subsequent processing is performed; otherwise, no orbit transfer is performed; The probe predicts the current near point height, and determines whether the height is within the range of the upper and lower boundaries of the corridor, if yes, no orbit transfer is performed; otherwise, a raising maneuver or a lowering maneuver is performed at the far point to make the current near point height reach the target near point height; N3 is the far point maneuver period.

10. The method of claim 9, wherein: The far fire point maneuver period is not less than 2 circles; the target near fire point height of each far fire point maneuver in the main deceleration section is h p,目标 = h p,min ’ + 3(Δh p引力 + Δh p大 air); h p,min ’ is the current corridor lower boundary; Δh p gravity + Δh p air is the near fire point perturbation in a far fire point maneuver period.

Images