Aerodynamic capture maneuver guidance method based on sideslip correction

Abstract

The application discloses a kind of aerodynamic capture machine dynamic guidance methods based on sideslip correction, it is related to the field of orbit braking energy dissipation optimization, including establishing the maneuvering dynamics model of spacecraft in the atmospheric flight section during atmospheric capture maneuver;S2, profile analysis is carried out to maneuvering dynamics model based on saturation function and target aerodynamic capture tilt angle reference trajectory is designed;The basic parameters of the saturation function are set by the spacecraft roll maneuvering capability, and the guidance parameters modulated by the tilt angle in the aerodynamic capture process are determined.The application can significantly reduce the sensitivity of the guidance parameters by approximating the step-type tilt angle time profile of the saturation function, and enhance the robustness of the guidance system.

Description

TECHNICAL FIELD

[0001] The present application relates to the field of orbit braking energy consumption optimization, and particularly relates to a method for aerocapture maneuvering guidance based on sideslip correction. BACKGROUND

[0002] Low-energy orbit braking in space missions such as planetary exploration and lunar return is a key link in the implementation of the mission, and how to reduce the fuel consumption of the spacecraft itself is a major issue in the overall design of the mission. Considering that the capture braking process belongs to the process of spacecraft kinetic energy decay, applying atmospheric ablation deceleration is a potential low-fuel consumption capture method. This aerocapture method not only replaces the engine thrust with aerodynamic overload to slow down, but also has stronger maneuvering capability, which can enhance the reachable range of the spacecraft terminal state, and is considered to be a widely used high-speed spacecraft capture method in the future.

[0003] For this reason, many scholars have studied this problem, especially focusing on the corresponding guidance algorithm research. At present, in the research of aerocapture guidance algorithm, the optimal aerocapture guidance algorithm proposed by the American scholar Lu Ping has the best performance in robustness and optimality. He converts the guidance trajectory design problem into a shooting problem of individual parameters, and combines the optimal control profile to propose a performance optimal aerocapture algorithm.

[0004] However, the significant disadvantage of this method is that the algorithm performance is heavily dependent on the roll maneuvering capability of the spacecraft during atmospheric flight, which leads to the easy occurrence of control parameter saturation during atmospheric flight. This control saturation can significantly reduce the maneuvering fuel consumption of the spacecraft after capture, and further affect the aerocapture performance.

[0005] Therefore, there is an urgent need for an aerocapture maneuvering guidance method based on sideslip correction to improve the above problems. SUMMARY

[0006] To solve the above problems, the present application proposes an aerocapture maneuvering guidance method based on sideslip correction, comprising the following steps:

[0007] S1, establishing a maneuvering dynamics model of the spacecraft in the atmospheric flight segment during the atmospheric capture maneuvering process based on the polar coordinate system of the spacecraft during atmospheric flight;

[0008] S2, performing profile analysis on the maneuvering dynamics model based on a saturation function and designing a target aerocapture roll angle reference trajectory;

[0009] The basic parameters of the saturation function are set by the roll maneuvering capability of the spacecraft, and the guidance parameters modulated by the roll angle during the aerocapture process are determined according to the mapping relationship between the roll angle bang-bang optimal profile and the saturation function. The basic parameters include switching time and amplitude.

[0010] Preferably, the maneuvering dynamics model of the atmospheric flight segment in the spacecraft aerodynamic capture process in S1 is:

[0011]

[0012] In the above formula, respectively represent the position vector rate of change, the latitude rate of change, the speed rate of change, the longitude rate of change, the track angle rate of change and the heading angle rate of change, g r , g φ respectively represent the radial and tangential gravitational acceleration of the spacecraft, g r , g φ The expression is:

[0013]

[0014] Wherein, r is the distance between the spacecraft mass center and the center of the central celestial body; V is the speed; γ is the track angle of the spacecraft, ψ is the heading angle of the spacecraft; σ is the flight roll angle, which is a control variable; θ, φ are respectively the longitude and latitude; ω is the planetary rotation angular velocity; μ is the planetary gravitational constant, R is the planetary radius; J2 is the second order spherical harmonic coefficient of the planet; the lift acceleration L, the drag acceleration D and the lateral force acceleration Q are respectively:

[0015]

[0016] Wherein, S is the reference area of the spacecraft; C L , C D and C Q are respectively the lift coefficient, the drag coefficient and the side slip coefficient; ρ is the atmospheric density; m is the mass of the spacecraft.

[0017] Preferably, the specific content of the profile analysis of the maneuvering dynamics model and the design of the target aerodynamic capture roll angle reference trajectory based on the saturation function in S2 is:

[0018] S201, on the basis of roll angle modulation, a side slip correction component is added to the saturation function with the side slip acceleration coefficient as the design parameter, and the aerodynamic capture guidance process is integrated into a first serial parameter homing execution stage and a second serial parameter homing execution stage through the side slip correction component;

[0019] S202, the mapping relationship between the guidance parameters and the aerodynamic capture efficiency characterization of the spacecraft in the atmospheric capture maneuvering process is constructed, and the homing parameters of the first serial parameter homing execution stage and the second serial parameter homing execution stage are determined according to the mapping relationship;

[0020] S203, the basic parameters of the saturation function with the roll angle as the reference and the guidance parameters are determined through single target optimization, and the entry speed pulse representing the aerodynamic capture efficiency calculated by the terminal state is given.

[0021] S204, according to the guidance parameters and terminal state obtained in S203, the trajectory of the atmospheric flight process of the spacecraft in the guidance period, the time sequence profile of the roll angle, and the time sequence profile of the sideslip force coefficient can be obtained.

[0022] Preferably, the specific content of the roll angle modulation in S201 is:

[0023] The time sequence profile of the roll angle is the time-varying trajectory of the roll angle, which is a reference profile based on the bang-bang structure of a single jump. The reference profile expression is:

[0024]

[0025] Where t is the current time, ts is the jump time of the roll angle without considering the change rate limit, σ min is the lower limit of the roll angle, σ max is the upper limit of the roll angle, and k is the smooth jump coefficient considering the roll angle change rate. The expression of k is:

[0026]

[0027] The upper limit of the roll angle change rate is s max , that is,

[0028] Where is the roll angle change rate, s max is the boundary value of the roll angle change rate;

[0029] The first step of the open-loop guidance process of the guidance parameters is to determine ts by targeting.

[0030] Preferably, in S201, the sideslip correction component is added to the saturation function with the sideslip acceleration coefficient as the design parameter. The specific content is:

[0031] The sideslip coefficient is unbounded, and C Q is set to:

[0032]

[0033] Where C Qmin and C Qmax are the lower and upper bounds of the sideslip force coefficient C Q , respectively.

[0034] Through the above expression, the boundary-limited sideslip force coefficient C Q is mapped to the unconstrained variable C b .

[0035] Preferably, in S201, the first serial parameter targeting execution stage takes CQ the targeting parameters as the guidance;

[0036] the second serial parameter targeting execution stage, i.e. the sideslip correction stage, is C b the targeting parameters as the guidance.

[0037] Preferably, the basic parameters of the saturation function based on the bank angle and the guidance parameters are determined by single-target optimization in S203, and the entry speed pulse representing the aerocapture efficiency calculated based on the terminal state is given, and the specific content is:

[0038] The efficiency of the aerocapture process is represented by the entry maneuver pulse after exiting the atmosphere;

[0039] After the spacecraft exits the atmosphere, it enters the target orbit through the first pulse and the second pulse maneuver;

[0040] The first pulse ΔV1 is collinear with the speed at the apogee, which increases the perigee to the target orbit radius, and the second pulse ΔV2 raises or lowers the new apogee at the perigee to the target orbit apogee;

[0041] The sum of the sizes of the first pulse and the second pulse is:

[0042]

[0043] Where, r atgt , r ptgt are the apogee radius and perigee radius of the target orbit, r a0 and r p0 are the apogee and perigee radii of the orbit after aerocapture;

[0044]

[0045] Where, a is the semi-major axis of the orbit after aerocapture, r EI , V exit and γ exit are the position vector magnitude, speed and flight path angle under the exit condition of the atmosphere;

[0046] The semi-major axis of the orbit after aerocapture is:

[0047]

[0048] Preferably, in the bank angle modulation guidance part, the guidance parameter is ts, and the guidance parameter is solved as:

[0049] ts→minΔV(ts);

[0050] In the sideslip correction guidance part, the guidance parameter is C b , i.e. the guidance parameter is solved as

[0051] C b →minΔV(C b );

[0052] The target's guidance parameters ts and C are obtained quickly through numerical Newton's method or gradient descent method. b .

[0053] Given the target's guidance parameters ts and C b Then, the optimal aerodynamic capture orbital velocity pulse size minΔV is obtained.

[0054] Preferably, based on the guidance parameters and terminal status obtained in S203, the specific contents of the spacecraft's atmospheric flight trajectory, tilt angle time-series profile, and sideslip force coefficient time-series profile during the guidance cycle can be obtained as follows:

[0055] Given the target's guidance parameters ts and C b Then, the roll angle σ and sideslip force coefficient C from the current moment to the final moment are obtained. Q The tilt angle σ forms a tilt angle time-series profile, and the sideslip force coefficient C Q Constructing a time-series profile of the sideslip force coefficient;

[0056] Given the tilt angle time profile and the sideslip force coefficient time profile, the trajectory of the spacecraft's atmospheric flight process under one guidance cycle is obtained by integrating the dynamic equations.

[0057] The orbital dynamics equation is:

[0058] Starting from the current time t, the integral to the atmospheric inlet is represented as:

[0059]

[0060] Where, x EI , t exit ,x,g(), These are the state vector at the atmospheric exit, the time at the atmospheric exit, the time-varying state vector, the differential expression of the state quantity, and the differential vector of the state quantity, respectively. The cutoff condition for the orbital integral is Γ0 = r. EI -r atm =0, and the integral trajectory, tilt angle time-varying profile, and sideslip acceleration coefficient time-varying profile under one guidance cycle are obtained.

[0061] In summary, the aerodynamic capture maneuver guidance method based on sideslip correction of the present invention has the following advantages compared with traditional technologies:

[0062] (1) The aerodynamic capture maneuvering guidance method based on sideslip correction provided by the present application can significantly reduce the sensitivity of guidance parameters and enhance the robustness of the guidance system by approximating the step-type sideslip angle time profile with a saturation function;

[0063] (2) The sideslip correction provided by the present application can significantly enhance the problem of limited control ability caused by sideslip angle saturation and fully utilize the effect of sideslip force to enhance the efficiency of the guidance system;

[0064] (3) The guidance method provided by the present application converts the guidance parameters into unconstrained single-target optimization problems during the calculation of the guidance parameters, so the calculation efficiency is high and the convergence is strong.

[0065] The technical method of the present application will be further described below through the accompanying drawings and examples. BRIEF DESCRIPTION OF DRAWINGS

[0066] Figure 1 The figure is a schematic diagram of the overall process of the embodiment of the present application;

[0067] Figure 2 The figure is a schematic diagram of the trajectory of the aerodynamic capture maneuvering process of the embodiment of the present application;

[0068] Figure 3 The figure is a sideslip angle time profile of the embodiment of the present application;

[0069] Figure 4 The figure is a sideslip force coefficient time profile of the embodiment of the present application;

[0070] Figure 5 The figure is a trajectory of the aerodynamic capture atmospheric flight process of the embodiment of the present application. DETAILED DESCRIPTION

[0071] The technical method of the present application will be further described below through the accompanying drawings and examples. It should be noted that the relative arrangement, numerical expression and values of the components and steps set forth in these examples do not limit the scope of the present application unless otherwise specifically stated.

[0072] The following description of at least one exemplary embodiment is merely exemplary in nature and is in no way intended to limit the present application or its application or uses.

[0073] Techniques, systems, and devices known to those of ordinary skill in the relevant art can not be discussed in detail herein, but should be considered as part of the specification, where appropriate.

[0074] In all examples shown and discussed herein, any specific values should be interpreted as merely exemplary, and not as a limitation. Thus, other examples of the exemplary embodiments can have different values.

[0075] Unless otherwise defined, technical terms or scientific terms used in the present application shall have the ordinary meaning as understood by a person of ordinary skill in the art to which the present application pertains.

[0076] The present application provides a method for aerocapture maneuvering based on sideslip correction, comprising the following steps:

[0077] S1, establishing a maneuvering dynamics model of the spacecraft in the atmospheric flight segment during the aerocapture maneuvering process based on the polar coordinate system of the spacecraft in the atmospheric flight process;

[0078] Preferably, the maneuvering dynamics model of the spacecraft in the atmospheric flight segment during the aerocapture process in S1 is:

[0079]

[0080] In the above formula, respectively represent the position vector rate of change, the latitude rate of change, the velocity rate of change, the longitude rate of change, the track angle rate of change and the heading angle rate of change, g r , g φ respectively represent the radial and tangential gravitational acceleration of the spacecraft, g r , g φ The expression is:

[0081]

[0082] Wherein, r is the distance between the spacecraft mass center and the center of the central celestial body; V is the speed; γ is the track angle of the spacecraft, ψ is the heading angle of the spacecraft; σ is the flight roll angle, which is a control variable; θ, φ are respectively the longitude and latitude; ω is the planetary rotation angular velocity; μ is the planetary gravitational constant, R is the planetary radius; J2 is the second order spherical harmonic coefficient of the planet; the lift acceleration L, the drag acceleration D and the lateral force acceleration Q are respectively:

[0083]

[0084] Wherein, S is the reference area of the spacecraft; C L , C D and C Q are respectively the lift coefficient, the drag coefficient and the sideslip coefficient; ρ is the atmospheric density; m is the mass of the spacecraft.

[0085] S2, profile analysis of the maneuvering dynamics model is performed based on the saturation function, and a target aerocapture roll angle reference trajectory is designed;

[0086] The basic parameters of the saturation function are set by the spacecraft roll maneuvering capability, and the guidance parameters modulated by the roll angle in the aerocapture process are determined according to the mapping relationship between the bang-bang optimal profile and the saturation function, and the basic parameters include switching time and amplitude.

[0087] The prior art has proved that the optimal roll angle time profile is a bang-bang structure with a single jump, so the designed roll angle time profile will be based on this structure, but since the spacecraft has limited ability to perform attitude maneuvering, the roll angle change rate of the spacecraft has an upper limit, and the upper and lower limits of the roll angle itself are also limited, that is, there is a lower limit σ min and an upper limit σ max .

[0088] The roll angle time profile is the time-varying trajectory of the roll angle, which is given based on the bang-bang structure of the optimal aerocapture atmospheric flight characteristics.

[0089] Preferably, the specific content of the roll angle modulation in S201 is:

[0090] The roll angle time profile, which is the time-varying trajectory of the roll angle, is a reference profile based on the bang-bang structure with a single jump, and the expression of the reference profile is:

[0091]

[0092] Where t is the current time, ts is the jump time of the roll angle without considering the change rate limit, σ min is the lower limit of the roll angle, σ max is the upper limit of the roll angle, and k is the smooth jump coefficient considering the roll angle change rate, and the expression of k is:

[0093]

[0094] The upper limit of the roll angle change rate is s max , that is

[0095] Where is the roll angle change rate, and s max is the boundary value of the roll angle change rate;

[0096] The determination of the guidance profile only leaves the jump time ts, so the first step of the open-loop guidance process with the roll angle as the guidance parameter is to determine ts.

[0097] Preferably, the specific content of S2 based on the saturation function to analyze the maneuvering dynamics model and design the target aerocapture roll angle reference trajectory is:

[0098] In order to overcome the boundary saturation problem, here the sideslip correction is added, S201, on the basis of the roll angle modulation, the sideslip correction component is added in the saturation function with the sideslip acceleration coefficient as the design parameter, and the aerodynamic capture guidance process is integrated into the first serial parameter homing execution stage and the second serial parameter homing execution stage through the sideslip correction component.

[0099] The sideslip correction component is added in the dynamic sideslip acceleration part, and if the sideslip correction component is added in the dynamics, the corresponding sideslip acceleration is no longer 0.

[0100] Preferably, in S201, the sideslip correction component is added in the saturation function with the sideslip acceleration coefficient as the design parameter, and the specific content is:

[0101] The size of the sideslip acceleration is completely controlled by the sideslip coefficient C Q Therefore, in the sideslip correction guidance link, it is used as the homing parameter. Considering that the sideslip control ability is limited, and the control ability is reflected in the sideslip force coefficient C Q Therefore, the boundary of the sideslip correction amount needs to be limited. Assuming that the sideslip force coefficient C Q There is a lower bound and an upper bound, that is, C Qmin And C Qmax At this time, in order to make the guidance parameter easy to converge as a single variable in the homing process, the sideslip coefficient is unbounded, and C Q is set to:

[0102]

[0103] Wherein, C Qmin And C Qmax are the lower bound and the upper bound of the sideslip force coefficient C Q ;

[0104] Through the above expression, the boundary-limited sideslip force coefficient C Q is mapped to the unbounded variable C b .

[0105] The boundary-limited sideslip force coefficient C Q is mapped to the unbounded variable C b , and the sideslip correction link uses C b as the homing parameter of the guidance.

[0106] S202, the mapping relationship between the guidance parameter and the aerodynamic capture efficiency of the spacecraft in the atmospheric capture maneuver process is constructed, and the homing parameters of the first serial parameter homing execution stage and the second serial parameter homing execution stage are determined according to the mapping relationship;

[0107] Preferably, in S201, the first serial parameter homing execution stage uses C QThe target parameters as guidance;

[0108] The second serial parameter target execution stage, i.e. the sideslip correction link, is executed by C b The target parameters as guidance.

[0109] S203, determining the basic parameters of the saturation function based on the roll angle and the guidance parameters by single-target optimization, and giving the entry orbit speed pulse representing the aerodynamic capture efficiency calculated based on the terminal state;

[0110] The terminal state is obtained by the prediction open loop of each guidance, and the termination condition of the prediction open loop is that the spacecraft flies out of the atmosphere, and the state at this time is the terminal state.

[0111] Preferably, in S203, the basic parameters of the saturation function based on the roll angle and the guidance parameters are determined by single-target optimization, and the entry orbit speed pulse representing the aerodynamic capture efficiency calculated based on the terminal state is given, and the specific content is:

[0112] The efficiency of the aerodynamic capture process is represented by the post-atmospheric entry orbit maneuver pulse;

[0113] When the spacecraft flies out of the atmosphere, it enters the target orbit by the first pulse and the second pulse maneuver;

[0114] The first pulse ΔV1 is collinear with the speed at the apogee, which increases the perigee to the target orbit radius, and the second pulse ΔV2 at the perigee raises or lowers the new apogee to the target orbit apogee;

[0115] As shown in Figure 2 For the aerodynamic capture process, its efficiency is represented by the post-atmospheric entry orbit maneuver pulse, when the spacecraft flies out of the atmosphere, a double-pulse maneuver is needed to finally enter the target orbit, the first pulse ΔV1 (collinear with the speed at the apogee) increases the perigee to the target orbit radius, and the second pulse ΔV2 (at the perigee) raises or lowers the new apogee to the target orbit apogee.

[0116] The sum of the sizes of the first pulse and the second pulse is:

[0117]

[0118] Where, r atgt , r ptgt are the apogee radius and perigee radius of the target orbit respectively, r a0 and r p0 are the apogee and perigee radii of the post-aerodynamic capture orbit respectively;

[0119] Where,

[0120] where a is the semi-major axis of the post-aerocapture orbit, r EI , V exit and γ exit are the position vector magnitude, velocity and flight path angle at the atmospheric exit conditions;

[0121] The semi-major axis of the post-aerocapture orbit is:

[0122]

[0123] Preferably, in the bank angle modulation guidance segment, the guidance parameter is ts, the guidance objective is to minimize ΔV, and the guidance parameter is solved as:

[0124] ts→minΔV(ts);

[0125] In the sideslip correction guidance segment, the guidance parameter is C b , the guidance objective is to minimize ΔV, and the mapping relationship between the guidance parameter and the aerocapture effectiveness characterization quantity is also a single-objective optimization problem, i.e., the guidance parameter is solved as

[0126] C b →minΔV(C b );

[0127] No matter whether it is the bank angle modulation guidance segment or the sideslip correction guidance segment, the solving process of the above single objective is similar, and the guidance parameters tsand C b are quickly searched by the numerical Newton method or gradient descent method.

[0128] After the guidance parameters tsand C b are given, the optimal aerocapture entry velocity impulse size minΔV is obtained.

[0129] S204, according to the guidance parameters and the terminal state obtained in S203, the trajectory of the spacecraft in the atmospheric flight process, the bank angle time sequence profile and the sideslip force coefficient time sequence profile in the guidance period are obtained.

[0130] Preferably, according to the guidance parameters and the terminal state obtained in S203, the specific content of the trajectory of the spacecraft in the atmospheric flight process, the bank angle time sequence profile and the sideslip force coefficient time sequence profile in the guidance period is:

[0131] After the guidance parameters tsand C b are given, the bank angle σ and the sideslip force coefficient C Q from the current time to the final time are obtained, the bank angle σ forms the bank angle time sequence profile, and the sideslip force coefficient C Q forms the sideslip force coefficient time sequence profile.

[0132] Given the bank angle time profile and the sideslip force coefficient time profile, the trajectory of the spacecraft in the atmospheric flight process in a guidance period is obtained by integrating the dynamic equation.

[0133] wherein the orbit dynamic equation is

[0134] From the current time t, the integration to the atmospheric entrance is expressed as:

[0135]

[0136] wherein x EI , t exit , x, g(), and are respectively the state vector at the atmospheric exit, the time at the atmospheric exit, the time-varying state vector, the differential expression of the state quantity, and the differential vector of the state quantity, the cutoff condition of the orbit integration is Γ0=r EI -r atm =0, the integral trajectory in a guidance period and the bank angle time-varying profile and the sideslip acceleration coefficient time-varying profile are obtained.

[0137] The specific implementation is as follows:

[0138] The spacecraft in the aerocapture task embodiment adopts the NASA "Orion" spacecraft, the mass of which is 10387kg, and the nominal lift-drag ratio is 0.27. The earth gravity constant μ=398600km 3 / s 2 , the earth radius is R=6378km, and the atmospheric height is 121.9km. The task scenario is an aerocapture of a lunar-earthen return spacecraft, and the corresponding state quantity initial value is shown in the following table.

[0139] Table 1 nominal inertial entry condition

[0140]

[0141] The target orbit is a h aim 200km high circular orbit. The calculated whole-process bank angle time profile, sideslip force coefficient time profile, and trajectory of the aerocapture atmospheric flight process are shown in FIGS. Figure 3 、 4 、 and 5, and the obtained orbital speed increment ΔV is 47.29m / s;

[0142] Therefore, the application adopts the above-mentioned aerodynamic capture maneuvering guidance method based on sideslip correction, can significantly reduce the sensitivity of the guidance parameters by approximating the step-type roll angle time profile with a saturation function, enhance the robustness of the guidance system, and can significantly enhance the problem of limited control ability caused by roll angle saturation, fully utilize the effect of sideslip force to enhance the efficiency of the guidance system. In addition, in the calculation of the guidance parameters, they are all converted into unconstrained single-objective optimization problems, so the calculation efficiency is high and the convergence is strong.

[0143] Finally, it should be noted that the above examples are only used to illustrate the technical method of the present application and not to limit it. Although the present application has been described in detail with reference to the preferred embodiments, those skilled in the art should understand that the technical method of the present application can still be modified or equivalently replaced, and these modifications or equivalent replacements cannot make the modified technical method deviate from the spirit and scope of the technical method of the present application.

Claims

1. A method of aerodynamically capturing a maneuvering target based on sideslip correction, characterized by, The method comprises the following steps: S1, establishing a maneuver dynamics model of the spacecraft in the atmospheric flight segment in the atmospheric capture maneuver process based on a polar coordinate system of the spacecraft in the atmospheric flight process; S2, profile analysis of the maneuver dynamics model based on a saturation function and design of a target aerodynamic capture roll angle reference trajectory; basic parameters of the saturation function are set by a roll maneuvering capability of the spacecraft, and a guidance parameter modulated by the roll angle in the aerodynamic capture process is determined according to a mapping relationship between a bang-bang optimal profile of the roll angle and the saturation function, the basic parameters including a switching time and an amplitude; the maneuver dynamics model of the spacecraft in the atmospheric flight segment in the aerodynamic capture process in S1 is: ； In the above formulae, , , , , , respectively represent the position radial rate of change, the latitude rate of change, the speed rate of change, the longitude rate of change, the track angle rate of change and the heading angle rate of change, , respectively represent the radial and tangential gravitational acceleration acting on the spacecraft, , The expression is: , ; wherein, is the distance of the spacecraft's center of mass from the center of the central celestial body; is the speed magnitude; is the flight path angle of the spacecraft, is the heading angle of the spacecraft; is the flight bank angle, which is the control variable; are the longitude and latitude, respectively; is the planet's angular rotation speed; is the planet's gravitational constant, is the planet's radius; is the planet's second order spherical harmonic coefficient; lift acceleration , drag acceleration , and lateral force acceleration , respectively. ； wherein, Arefis a reference area for the spacecraft; , and are lift, drag, and side-slip force coefficients, respectively; is atmospheric density; is the mass of the spacecraft; the specific content of the profile analysis of the maneuver dynamics model based on the saturation function and the design of the target aerodynamic capture roll angle reference trajectory in S2 is: S201, adding a sideslip correction component in the saturation function by taking a sideslip acceleration coefficient as a design parameter on the basis of roll angle modulation, and integrating the aerodynamic capture guidance process into a first serial parameter homing execution stage and a second serial parameter homing execution stage through the sideslip correction component; S202, constructing a mapping relationship between the guidance parameter and an aerodynamic capture efficiency representation of the spacecraft in the atmospheric capture maneuver process, and determining homing parameters of the first serial parameter homing execution stage and the second serial parameter homing execution stage according to the mapping relationship; S203, determining the basic parameters of the saturation function with the roll angle as the reference and the guidance parameter through single-objective optimization, and giving an entry speed pulse representing the aerodynamic capture efficiency calculated according to the terminal state; S204, obtaining the trajectory of the spacecraft in the atmospheric flight process, the roll angle time sequence profile and the sideslip force coefficient time sequence profile under the guidance period according to the guidance parameter and the terminal state obtained in S203.

2. A method of aerodynamically capturing a vehicle according to claim 1, wherein the specific content of the roll angle modulation in S201 is: the roll angle time sequence profile, that is, the time-varying trajectory of the roll angle, is a reference profile with a single jump bang-bang structure reference, and the reference profile expression is: ； wherein, is the current time, is the step time of the roll angle without considering the rate limit, is the lower limit of the roll angle, is the upper limit of the roll angle, is the smooth jump coefficient after considering the roll angle rate, the expression of is: ； The upper limit of the roll angle change rate is i.e. ; wherein is a rate of change of the roll angle, is a boundary value for the magnitude of the rate of change of the roll angle; The first step in the open loop guidance process of the guidance parameters is the target determination .

3. A method of aerodynamically capturing a vehicle according to claim 2, wherein, the specific content of adding the sideslip correction component in the saturation function by taking the sideslip acceleration coefficient as the design parameter in S201 is: The lateral slip coefficient is unbounded, and the lateral slip coefficient is set as: is set as: ； wherein and are the lower and upper bounds, respectively, of the side slip coefficient of the vehicle. By the above expression, the side slip force coefficient of the boundary restriction is mapped as an unconstrained variable .

4. A method of aerodynamically capturing a vehicle according to claim 3, wherein In S201, the first serial parameter targeting execution stage is executed with as the guided targeting parameter; The second serial parameter targeting execution phase is a side slip correction link to as a guided targeting parameter.

5. A method of aerodynamically capturing a vehicle according to claim 4, wherein, the specific content of determining the basic parameters of the saturation function with the roll angle as the reference and the guidance parameter through single-objective optimization and giving the entry speed pulse representing the aerodynamic capture efficiency calculated according to the terminal state in S203 is: the efficiency of the aerodynamic capture process is represented by an entry orbit maneuver pulse after the spacecraft flies out of the atmosphere; after the spacecraft flies out of the atmosphere, the spacecraft enters a target orbit through a first pulse and a second pulse; the first impulse at the apogee co-linear with the velocity, raising the apogee to the target orbit radius, the second impulse at the perigee raising or lowering the new apogee to the target orbit apogee; the sum of the first pulse and the second pulse is: ； wherein, , are the apogee radius and the perigee radius of the target orbit, respectively, and are the apogee and perigee radii of the post-aerocapture orbit, respectively; wherein , ; wherein, a is the semi-major axis of the post-capture orbit, , and are the position vector magnitude, velocity and track angle under atmospheric exit conditions; the semi-major axis of the orbit after the aerodynamic capture is: 。 6. A method of aerodynamically capturing a vehicle according to claim 5, wherein, In the bank angle modulation guidance section, the guidance parameter is , and the guidance parameter is solved as ； In the guidance part of sideslip correction, the guidance parameter is i.e. the guidance parameter is solved as ； The guidance parameters of the target are quickly searched by numerical Newton method or gradient descent method and ; Guidance parameters for a given target and After which the optimal aerocapture insertion velocity pulse size is obtained .

7. A method of aerodynamically capturing a vehicle according to claim 6, wherein, the specific content of obtaining the trajectory of the spacecraft in the atmospheric flight process, the roll angle time sequence profile and the sideslip force coefficient time sequence profile under the guidance period according to the guidance parameter and the terminal state obtained in S203 is: Guidance parameters for a given target and a roll angle from the current time to the final time and a side slip force coefficient the roll angle forms a roll angle time profile, the side slip force coefficient constitutes a side slip force coefficient time profile; given the roll angle time sequence profile and the sideslip force coefficient time sequence profile, the trajectory of the spacecraft in the atmospheric flight process under a guidance period is obtained by integrating the dynamics equation; where the orbit dynamics equation is ; From the current time Start, integrate to the atmosphere entrance, expressed as: ； wherein, , , , , , are the state vector of the atmosphere exit, the time of the atmosphere exit, the time-varying state vector, the differential expression of the state quantity, the differential vector of the state quantity, respectively, the cutoff condition of the orbit integral is , the integral trajectory and the time-varying profile of the roll angle, the time-varying profile of the sideslip acceleration coefficient in a guidance cycle are obtained.

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