A Design Method for Mars Atmospheric Braking Strategy

Abstract

This invention provides a Mars atmospheric braking strategy design method, wherein the atmospheric braking is divided into at least three stages: an approach stage, a main deceleration stage, and an exit stage. The method includes: designing an atmospheric braking corridor for each apogee maneuver cycle and determining the upper and lower boundaries of the corridor; the approach stage employs a trajectory control strategy of gradually lowering the perigee from the highly elliptical orbit around Mars to the upper boundary of the current atmospheric braking corridor height; the main deceleration stage ensures that the perigee height remains within the current atmospheric braking corridor range; when the apogee height reaches the apogee height of the target transition orbit, the method enters the exit stage; the exit stage employs a trajectory control strategy of gradually raising the perigee; and after the final orbit raising maneuver, when the perigee velocity reaches the perigee velocity corresponding to the mission orbit, a circular maneuver is performed at the apogee to allow the probe to enter the mission orbit.

Description

Technical Field

[0001] This invention belongs to the field of deep space exploration technology and relates to a design of a Martian atmospheric braking strategy and a method for orbit correction. Background Technology

[0002] Atmospheric braking is a type of aerodynamic-assisted orbital transfer, and as a low-cost orbital transfer technology, it plays a very important role in future interplanetary exploration. Of the eight planets in our solar system, all except Mercury have atmospheres, and using planetary atmospheres for deceleration is an efficient and economical technique for obtaining velocity increments.

[0003] In a 1961 paper presented to the American Institute of Astronautics, H. London first demonstrated the significance and feasibility of aerodynamic-assisted orbit change. The concept of aerodynamic-assisted orbit change combines pure impulse orbit change with aerodynamic orbit change, inserting an atmospheric flight segment into the orbital flight to change the orbital plane or altitude with the help of aerodynamic forces, and ultimately completing all orbit change requirements with minimal energy consumption and other technical indicators.

[0004] Compared to traditional orbit transfer methods, the biggest advantage of atmospheric braking orbit change is that the probe can utilize atmospheric drag to consume orbital energy and change the orbital plane, thereby significantly saving fuel and increasing the mass of the payload. For Mars exploration missions, even with the most fuel-efficient orbit change method, the required velocity increment is still large. Using traditional orbit transfer methods requires the probe to carry a large amount of fuel, reducing the probe's dry weight and increasing the design complexity.

[0005] The challenge of atmospheric braking orbit change technology lies in the fact that to achieve a certain speed increment, the flight altitude must be reduced. This requires strict control of the flight trajectory to prevent damage to the probe due to excessive thermal stress and G-forces, as well as excessive energy loss leading to significant orbital deviations or even failure to escape the atmosphere. Furthermore, if the flight altitude is too high, the deceleration efficiency will be too low, making it impossible to complete deceleration within the specified time, thus affecting subsequent mission execution. Summary of the Invention

[0006] The technical problem solved by this invention is to propose a Mars atmospheric braking strategy design and orbit correction method, which will solve the problem of simultaneously meeting mission time constraints and safety in the atmospheric braking process of Mars exploration missions.

[0007] The solution of this invention is: a Mars atmospheric braking strategy design method, wherein the atmospheric braking is divided into at least three stages: an approach stage, a main deceleration stage, and an exit stage, including:

[0008] Design an atmospheric braking corridor for each long-range firing point maneuver cycle and determine the upper and lower boundaries of the corridor;

[0009] The approach phase employs a trajectory control strategy that gradually lowers the near-fire point, reducing the near-fire point from the highly elliptical orbit around Mars to the upper boundary of the current atmospheric braking corridor.

[0010] The main deceleration phase ensures that the near-fire point altitude is within the current atmospheric braking corridor. When the far-fire point altitude reaches the far-fire point altitude of the target transition orbit, it enters the step-out phase.

[0011] The step-out phase employs a trajectory control strategy of gradually raising the near-fire point; when the near-fire point speed reaches the speed corresponding to the near-fire point on the mission trajectory after the last trajectory raising maneuver, a circular maneuver is performed at the far-fire point to bring the probe into the mission trajectory.

[0012] Preferably, the upper and lower boundaries of the corridor are determined in the following manner:

[0013] Based on the task time constraint t max The upper boundary h of the corridor height was determined by the target shooting method. p,max To ensure that the current predicted value of heat flux density does not exceed the heat flux density constraint;

[0014] The local atmospheric density is calculated based on the upper limit of axial acceleration. This local atmospheric density, combined with an atmospheric prediction model, is then used to inversely solve for the lower boundary h of the corridor. p,min ;

[0015] 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;

[0016] The upper and lower boundaries of the corridor, taking into account the safety margin, are determined based on the near-fire point height error ΔH.

[0017] Preferably,

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

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

[0020] Preferably, the orbit control strategy of gradually lowering the proximity point is specifically as follows:

[0021] S1. The probe flies freely for N1 orbits, completing its orbit determination. Then, it performs its first entry and descent maneuver at the far end of the orbit, lowering its altitude to the minimum resolution altitude h corresponding to the accelerometer. p0 ;

[0022] S2. Continue free flight for N1 orbits while completing the orbit determination, then perform a second entry and descent maneuver at the far fire point, lowering the altitude to the upper boundary of the atmospheric braking corridor (h). p,max ';N1≥2.

[0023] Preferably, considering 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, S1 and S2, 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. 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.

[0024] Preferably, 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;

[0025] h 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.

[0026] Preferably, the range of the far-point altitude of the target transition orbit is 800-2000 km.

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

[0028] 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.

[0029] 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.

[0030] N3 is the long-range fire control maneuver cycle.

[0031] Preferably, the maneuver cycle of the far-end fire point is not less than 2 revolutions; the target near-end fire point height for each maneuver of the far-end fire point during the main deceleration phase is h. p,目标 =h p,min '+3(Δh p引力 +Δh p大气 );h p,min ' represents the lower boundary of the current corridor; Δh p引力 +Δh p大气 It represents the perturbation of the near-fire point within one far-fire point maneuver cycle.

[0032] The beneficial effects of this invention compared to the prior art are:

[0033] A Martian atmospheric braking strategy and orbit correction method are proposed to save fuel mass on the probe and increase the mass of the payload. The results are as follows:

[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 for track control.

[0036] (3) A method for determining the near-fire point altitude and period of the maneuvering target is proposed, which makes the atmospheric braking trajectory obtained by this method close to the minimum total time and maintains fewer orbit control maneuvers, thus saving fuel. Attached Figure Description

[0037] Figure 1 The flowchart for the main deceleration phase strategy. Detailed Implementation

[0038] The present invention will be further described below with reference to the embodiments.

[0039] This invention relates to a design method for a Mars atmospheric braking strategy. Prior to this method, it is necessary to determine the maximum heat flux density constraint q that the probe can withstand, based on the overall Mars exploration mission plan. max Total task time constraint t max and atmospheric prediction model F during the mission period 大气模型 :

[0040] ρ=F 大气模型 (dd,tt,h,λ,ψ) (1)

[0041] In the formula, ρ is the atmospheric density, dd is the date, tt is the local time, h is the altitude, and λ is the longitude. Latitude.

[0042] Using conventional propulsion and braking methods, the probe was adjusted to a highly elliptical orbit around Mars with an orbital period of 1 to 3 days. A descent maneuver was then performed at the far fire point to lower the altitude to the minimum resolution height h corresponding to the accelerometer. p0 Atmospheric braking is then initiated. Atmospheric braking is divided into at least three phases: the approach phase, the main deceleration phase, and the exit phase. The method includes:

[0043] 1. Design an atmospheric braking corridor for each long-range fire point maneuver cycle and determine the upper and lower boundaries of the corridor;

[0044] Determining the maneuver cycle of the distant firing point: Since the maneuver of the distant firing point needs to be carried out after the orbit determination is completed, and the need to fully conserve chemical fuel must also be considered, the maneuver cycle of the distant firing point should be determined by comprehensively considering the minimum orbit determination period and the requirement to minimize the increment of orbit control speed. The maneuver cycle is generally not less than 2 revolutions.

[0045] The criterion for determining the atmospheric braking corridor is: to ensure the safety of the detector while meeting the mission time constraints for the total atmospheric braking time.

[0046] Corridor height lower boundary design: The lower boundary of the corridor should be determined to ensure that the predicted heat flux density of the current loop does not exceed the heat flux density constraint q. max The theoretical formula for calculating the heat flux density q on the windward surface of the detector is:

[0047] q = 0.5C H ρv 3 (2)

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

[0049] The axial acceleration a measured by the accelerometer x The total velocity v in the Martian fixed coordinate system obtained by integrating the triaxial accelerations can be used to estimate the local atmospheric density.

[0050]

[0051] In the formula, m is the detector mass, S is the reference area, and C A This is the axial force coefficient.

[0052] The heat flow on the windward surface of the detector

[0053]

[0054] If the heat flux q is required to not exceed the heat flux constraint q specified in (1) max ,but

[0055]

[0056] Therefore, the acceleration must satisfy the constraints.

[0057]

[0058] From the above equation, we can see that the upper limit of axial acceleration a x,max It gradually increases as speed decreases, meaning the lower boundary of the atmospheric braking corridor height decreases over time. Let a... x,max Substitute into equation (3) and use the atmospheric prediction model F 大气模型 Inverse solution for the lower boundary h of the corridor height p,min .

[0059] The lower boundary margin design of the corridor height: The height error caused by factors such as navigation error, track control error, atmospheric model error, daily fluctuation rate of atmospheric density, and accelerometer measurement error is Δh. 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 is determined.

[0060]

[0061] Then consider the lower boundary h of the corridor height for safety margin. p,min '=h p,min +ΔH.

[0062] Among them, the navigation error Δh p导航 Track control error Δh p轨控 Based on the control system, the atmospheric model error Δh is determined. p大气模型 and the daily fluctuation rate of atmospheric density Δh p日波动 According to atmospheric prediction model F 大气模型 Together with the flight measurement results, the accelerator error Δh is determined. p加计 Determined based on accelerometer specifications.

[0063] Corridor height upper boundary estimation: Using orbital mechanics simulation tools, based on the mission time constraint t max The upper boundary h of the corridor height was determined by the target shooting method. p,max Then, considering the safety margin, the upper boundary h of the corridor height... p,max =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 while completing the orbit determination. Then, at the far fire point, it performed a second entry and descent maneuver, lowering its altitude from near fire point to the upper boundary of the current atmospheric braking corridor (h). 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... EThis is the Martian reference radius. Therefore, the rate of change of perihelion altitude can be expressed as...

[0074]

[0075] That is, the height of the fire point h p Influenced by the semi-major axis *a* and the eccentricity *e*, gravitational perturbations have an average effect of 0 on the semi-major axis *a* and the eccentricity *e*, but they have a long-term effect on the right ascension of the ascending node *Ω* and the argument ω of the perihelion.

[0076]

[0077] In the formula, J2 is the gravitational perturbation term. Let i represent the average motion, and i be the orbital inclination. Although gravitational perturbations do not directly affect the long-term changes in the semi-major axis a and eccentricity e, the change in the periapsis argument ω causes the periapsis position to shift within the orbital plane; therefore, the periapsis height h... p It may change due to terrain variations.

[0078] The effect of atmospheric drag on orbital attenuation can be expressed as:

[0079]

[0080]

[0081] In the formula, C D Let S be the drag coefficient, S be the reference area, m be the mass, ρ be the atmospheric density, and μ be the Martian gravitational constant. Let be the velocity, and r be the distance from the fire center. Atmospheric drag reduces both the semi-major axis 'a' and the eccentricity 'e', ​​thus affecting the near-fire height 'h'. p .

[0082] Wherein, the drag coefficient C D The atmospheric density ρ is calculated using aerodynamic methods. It is obtained through atmospheric models and is related to various factors such as latitude and longitude, season, day and night, solar activity, and magnetic field, resulting in significant uncertainty. Commonly used Martian atmospheric models include MarsGRAM and MCD.

[0083] The above orbital perturbation model can be used to predict the near-fire perturbation Δh within one maneuver cycle. p引力 +Δh p大气 Then, the target near-fire height for each maneuver during the main deceleration phase is h. p,目标 =h p,min '+3(Δh p引 Force + Δh 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] Based on the overall plan for the Mars exploration mission, the maximum heat flux density constraint that the probe can withstand is determined to be 3 kW / m³. 2 The total mission time is constrained to 180 days, and the atmospheric prediction model is MarsGRAM 2000.

[0091] Using conventional propulsion and braking methods, the probe was adjusted to a highly elliptical orbit around the fire with an orbital period of 1.5 days and a periapsis of 333 km. At the far fire point, a descent maneuver was performed to lower the periapsis altitude to 147 km.

[0092] Atmospheric braking corridor design: The criterion for determining the atmospheric braking corridor is to ensure the safety of the detector while meeting the mission time constraints.

[0093] Corridor height lower boundary design: constrain heat flux density to 3kW / m 2 Substituting into equation (6), we obtain the upper limit of axial acceleration as 5 × 10⁻⁶. -5 m / s 2 Substituting it into equation (3), and using the atmospheric prediction model MarsGRAM 2000 to reverse-engineer the lower boundary of the corridor height, we find that it is 100.8 km in the initial stage and 97.7 km in the final stage.

[0094] The lower boundary margin design of the corridor height: A near-fire height error of 0.3 km is determined based on factors such as navigation error, orbit control error, atmospheric model error, daily fluctuation rate of atmospheric density, and accelerometer measurement error. Therefore, considering the safety margin, the initial lower boundary of the corridor height is 101.1 km, and the final lower boundary is 98.0 km.

[0095] Estimation of the upper boundary of the corridor height: Using orbital mechanics simulation tools and based on mission time constraints, the upper boundary of the corridor height is determined by a target-shooting method. Considering a safety margin, the upper boundary of the corridor height is 113.5 km.

[0096] Determining the maneuver cycle of the distant firing point: when the (initial) orbital period is not less than 12 hours, the maneuver cycle is ≥2; when the orbital period is not greater than 4 hours, the maneuver cycle is ≥4.

[0097] The entry phase strategy is as follows: The probe will fly freely for two orbits, completing its orbit determination. Then, it will perform its first entry-and-descend maneuver at the far end of the orbit, lowering its altitude to the minimum resolution altitude of the accelerometers at 130 km. Afterward, it will continue free flight for two more orbits, completing its orbit determination again. Finally, it will perform a second entry-and-descend maneuver at the far end of the orbit, lowering its altitude to the upper boundary of the atmospheric braking corridor at 105 km. The entry phase will involve three entry-and-descend maneuvers.

[0098] Main deceleration phase strategy design: according to Figure 1 Process it.

[0099] Determination of the near-fire point height of the maneuvering target at the far-fire point: Based on the gravitational perturbation model MGS85F2.GRV and the atmospheric perturbation model MarsGRAM 2000, the near-fire point perturbation is predicted to be 0.1 km in one maneuver cycle. Therefore, the height of the far-fire point in each maneuver during the main deceleration phase is initially 101.4 km and later 98.3 km.

[0100] Orbit correction method: To improve the accuracy and safety of orbit change maneuvers, the atmospheric density is inverted in real time using the measurement data of the heat flow meter and / or accelerometer installed on the probe near the fire point during the free flight phase. The atmospheric density data obtained by the two inversion methods are then fitted, and the maneuver speed increment at the far fire point is corrected based on the fitting results.

[0101] The walkout phase strategy is as follows: When the far-point altitude drops to 1000km, the near-point altitude is raised after every 5 free-flight orbits. Each maneuver raises the near-point altitude by approximately 40km. After the final maneuver, when the near-point velocity reaches the velocity corresponding to the mission orbit, a circular maneuver is performed at the far-point to guide the probe into the mission orbit. The walkout phase involves 4 maneuvers.

[0102] Although the present invention has been disclosed above with reference to preferred embodiments, it is not intended to limit the present invention. Any person skilled in the art can make possible changes and modifications to the technical solutions of the present invention by utilizing the methods and techniques disclosed above without departing from the spirit and scope of the present invention. Therefore, any simple modifications, equivalent changes and alterations made to the above embodiments based on the technical essence of the present invention without departing from the content of the technical solutions of the present invention shall fall within the protection scope of the technical solutions of the present invention.

Claims

1. A method for designing a Mars atmospheric braking strategy, wherein the atmospheric braking is divided into at least three stages: an approach stage, a main deceleration stage, and an exit stage, characterized in that... include: Design an atmospheric braking corridor for each long-range firing point maneuver cycle and determine the upper and lower boundaries of the corridor; The approach phase employs a trajectory control strategy that gradually lowers the near-fire point, reducing the near-fire point from the highly elliptical orbit around Mars to the upper boundary of the current atmospheric braking corridor. The main deceleration phase ensures that the near-fire point altitude is within the current atmospheric braking corridor. When the far-fire point altitude reaches the far-fire point altitude of the target transition orbit, it enters the step-out phase. The step-out phase adopts a track control strategy of gradually raising the near-fire point; when the near-fire point speed reaches the near-fire point speed corresponding to the mission orbit after the last track-raising maneuver, a circular maneuver is performed at the far-fire point to allow the probe to enter the mission orbit. The upper and lower boundaries of the corridor shall be determined in the following manner: Based on task time constraints t max The upper boundary of the corridor height was determined by target practice. h p,max To ensure that the current predicted value of heat flux density does not exceed the heat flux density constraint; The local atmospheric density is calculated based on the upper limit of axial acceleration, and the lower boundary of the corridor is solved by inverse calculation using the local atmospheric density and an atmospheric prediction model. h p,min ; Considering the altitude error Δ caused by navigation error, orbit control error, atmospheric model error, daily fluctuation rate of atmospheric density, and accelerometer measurement error. h p导航 , Δ h p轨控 , Δ h p大气模型 , Δ h p日波动 , Δ h p加计 Based on this, the corresponding near-fire point height error is determined. ; Based on the error of the near-fire point height Determine the upper and lower boundaries of the corridor, taking into account safety margins; Lower boundary of the corridor considering safety margin h p,min '= h p,min +Δ H ; Upper boundary of the corridor considering safety margin h p,max =max( h p,max , h p,min +2Δ H ).

2. The method according to claim 1, characterized in that: The orbit control strategy of gradually lowering the proximity point is as follows: S1, the probe is flying freely. N One orbit is completed, and the orbit determination is completed simultaneously. Then, the first entry and descent maneuver is performed at the far end of the orbit, lowering the altitude of the near end of the orbit to the altitude corresponding to the minimum resolution of the accelerometer. h p0 ; S2, Continue Free Flight N One orbit was completed, and the orbit determination was also completed. Then, a second entry and descent maneuver was performed at the far end of the orbit, lowering the altitude of the near end of the orbit to the upper boundary of the atmospheric braking corridor. h p,max '; N 1≥2.

3. The method according to claim 2, characterized in that: 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, i.e., the orbital inclination is less than 80° or greater than 100°, two maneuvers, S1 and S2, 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. When the orbital perturbation is small, i.e., the orbital inclination is greater than or equal to 80° and less than or equal to 100° in near-polar orbit, the descent maneuvers are divided into multiple equal parts, and the total number of descent maneuvers in the step-up phase should not exceed 10.

4. The method according to claim 1, characterized in that: 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. N 自由 After free flight, perform one long-range fire point lifting maneuver, for a total of N Twice, the near-fire height increased after each maneuver. h p步出 The above N 自由 ≥5, number of track lifting maneuvers in the step-out section N 2. No more than 10 times; h p步出 =(h j - h af ) / N 2,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.

5. The method according to claim 1 or 4, characterized in that: The recommended range for the far-point altitude of the target transition orbit is 800-2000 km.

6. The method according to claim 1, characterized in that: The main deceleration section is handled in the following way: 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. 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. N3 is the long-range fire control maneuver cycle.

7. The method according to claim 6, characterized in that: The maneuver cycle of the far-field fire point is no less than 2 revolutions; the target near-fire point height for each maneuver of the far-field fire point during the main deceleration phase is... h p,目标 = h p,min '+3(Δ h p引力 +Δ h p大气 ); h p,min 'This represents the lower boundary of the current corridor; Δ h p引力 +Δ h p大气 It represents the perturbation of the near-fire point within one far-fire point maneuver cycle.

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