A Reentry Near-Space Aerodynamic Large-Skew Orbit Reconfiguration Method

Abstract

A method for reentry near-space aerodynamic large off-plane orbit reconstruction, which solves the problems of high energy consumption of current orbit control engines for orbit change and difficulty in large-range orbit change by aerodynamic assist, belongs to the field of aerospace technology. The present invention includes: designing the braking reentry section orbit of the aircraft from the initial orbit into the near space, designing the aerodynamic orbit change section orbit of the aircraft in the near space, designing the ascending orbit insertion section orbit of the aircraft from the near space into the target orbit, taking the designed quantities as integral performance indicators, setting the index function J, and seeking the state and control (x(t), u(t)) that make the selected index function J optimal, that is: #imgabs0# represents the state variable of the aircraft in the aerodynamic orbit change section, u(t) represents the control variable, t is the time, X represents the process index, φ[x(t f ),t f represents the performance index at time t f ; #imgabs1# t0 ≤ t ≤ t f , f[x(t), u(t), t] represents the motion dynamics equation of the aircraft; and setting the constraint conditions and boundary conditions, and performing optimization solution to obtain the optimal control variable u(t).

Description

Technical Field

[0001] The present invention relates to a method for re - entry near - space aerodynamic large - skew - plane orbit reconstruction, belonging to the field of aerospace technology. Background Technique

[0002] In future outer - space confrontations, information confrontation is the primary aspect, and reconnaissance and anti - reconnaissance are the main means. For the reconnaissance, interference, or even destruction of spacecraft, it is necessary to approach the target spacecraft closely or even accompany it, which requires the reconnaissance spacecraft to have the ability of large - skew - plane orbit transfer and be able to quickly approach the target spacecraft. Anti - reconnaissance is to quickly maneuver when the enemy's reconnaissance spacecraft approaches and monitors our spacecraft, get rid of the enemy's tracking, and jump out of the enemy's field of vision, that is, achieve stealth through maneuvering. Therefore, the rapid maneuverability of spacecraft is an important technical approach to achieve close - in reconnaissance and counter - close - in reconnaissance.

[0003] Using an orbit - control engine for orbit transfer is the main way to achieve rapid orbit maneuver. However, the large consumption of fuel by the orbit - control engine severely limits the orbit - maneuvering range and number of times. For example, the US X - 37B vehicle needs to consume 1660.8 kg of propellant (1 / 3 of its own take - off mass) to change the orbit inclination by 9.1482°, and the Zhongxing series satellites need to consume nearly half of their own take - off mass of energy to change the orbit inclination by nearly 10°.

[0004] Aerodynamic - assisted orbit transfer is an effective orbit - transfer method and has been studied and applied to a certain extent. However, these applications mainly use the weak aerodynamic force of the spacecraft during low - orbit flight to achieve fine - tuning of the orbit and are difficult to achieve large - range orbit maneuvers. Some scholars have also proposed using the "slingshot effect", where the spacecraft realizes large - range orbit maneuvers for interstellar flight by "skipping" through re - entry into the atmosphere. However, there is currently no complete technical solution for "re - entry - aerodynamic orbit transfer - repeated re - entry into orbit", and the proposed aerospace vehicles can only achieve assisted landing through aerodynamic force and do not have the ability to achieve large - range lateral maneuvers through aerodynamic force. For example, after re - entry, the US X - 37B uses a lifting - body shape to achieve lifting re - entry and horizontal landing, but its orbit maneuver still uses an orbit - control engine, the time required for the orbit - transfer process is long, the change in orbit inclination is small, and large - skew - plane orbit transfer cannot be achieved. Summary of the Invention

[0005] Aiming at the problems of high energy consumption of orbit - control engine orbit transfer and difficulty in large - range orbit transfer by aerodynamic assistance, the present invention provides a method for re - entry near - space aerodynamic large - skew - plane orbit reconstruction.

[0006] A method for re - entry near - space aerodynamic large - skew - plane orbit reconstruction according to the present invention includes:

[0007] S1: Design the braking re - entry section orbit of the aircraft from the initial orbit into near - space:

[0008] The aircraft performs a braking ignition by applying a velocity increment ΔV1 at point A.

[0009] μ is the Earth's gravitational constant, r1 is the geocentric distance of the initial orbit, r a is the geocentric distance of the atmosphere, V0 is the velocity entering the atmosphere, and γ0 is the reentry angle;

[0010] S2: Design the orbit of the aircraft in the aerodynamic orbit change section in the near space:

[0011] The starting point of the aerodynamic orbit change section is point B. The aircraft enters the atmosphere at point B with a velocity V0 and a reentry angle γ0. When flying in the atmosphere, it uses aerodynamic lift to change the orbit inclination and flies out of the atmosphere at point D with a velocity V f and a trajectory inclination γ f The velocity change amount ΔV2 and fuel consumption Δm of the aerodynamic orbit change section;

[0012] ΔV2 = V f - V0, Δm = m1(t0) - m1(t f );

[0013] m1(t0) represents the fuel mass at point B, t0 represents the time at point B, and m1(t f ) represents the fuel mass at point D, and t f represents the time at point D;

[0014] S3: Design the orbit of the aircraft in the ascending orbit insertion section from the near space to the target orbit:

[0015] The starting point of the ascending orbit insertion section is point D. From point D, with a velocity V f and a track inclination γ f After jumping out of the atmosphere boundary, the orbit inclination becomes i f , and ΔV3 is applied at point E to enter the target orbit. ΔV3 is the ascending velocity increment;

[0016]

[0017] r2 represents the radius of the target orbit;

[0018] S4: Set the index function J and find the state and control (x(t), u(t)) that make the selected index function J optimal, that is:

[0019]

[0020] x(t) represents the state variable of the aircraft in the aerodynamic orbit change section, u(t) represents the control variable, t is the time, X represents the process index, and φ[x(t f ), tf represents the performance index at time t f ; f[x(t), u(t), t] represents the kinematic equation of the aircraft;

[0021] And set the constraint conditions and boundary conditions, conduct optimization to solve, and obtain the optimal control variable u(t).

[0022] Preferably, the constraint conditions include heat flux density constraint, overload constraint and dynamic pressure constraint;

[0023] The heat flux density constraint is:

[0024]

[0025] In the formula, Q(t) is the heat flux density at the stagnation point, R is the radius of the aircraft head, h w is the wall enthalpy, h0 is the total enthalpy, and ρ is the atmospheric density;

[0026] The overload constraint is:

[0027] Among them, n(t) is the three-axis overload of the aircraft in the velocity coordinate system, g0 is the gravitational acceleration at sea level, C D and C L are the drag coefficient and lift coefficient respectively, and n max represents the maximum value of the three-axis overload of the aircraft in the velocity coordinate system;

[0028] The dynamic pressure constraint:

[0029] In the formula, q(t) is the dynamic pressure, and q max represents the maximum value of the dynamic pressure.

[0030] Preferably, the constraint conditions further include control constraints;

[0031] The control constraint is:

[0032] {u ∈ R m : u L ≤ u ≤ u R}

[0033] u L represents the lower boundary of the control variable;

[0034] u R represents the upper boundary of the control variable.

[0035] Preferably, the boundary conditions include:

[0036] The initial orbit where the aircraft is located and the transfer orbit of the braking reentry section are tangent at the apogee, and there are boundary conditions at the initial time:

[0037]

[0038] After the aircraft jumps out of the atmosphere, it enters the transfer orbit of the ascending and orbital insertion section. The apogee of the transfer orbit is tangent to the target orbit, and there are terminal time boundary conditions:

[0039]

[0040]

[0041] i f represents the orbital inclination angle at point D, and ψ f represents the yaw angle at point D, represents the latitude at point D.

[0042] The beneficial effects of the present invention are as follows: The present invention proposes an orbital reconstruction scheme of "re-entry into the near space - aerodynamic large off-plane orbit change - repeated orbital insertion". By re-entering the near space and adjusting the attitude of the aircraft and applying a certain thrust during the aerodynamic orbit change process, rapid large off-plane maneuvering orbit change can be achieved; by using aerodynamic force-assisted orbit change, the fuel consumption for strong maneuvering orbit change can be reduced, and multiple orbit changes can be realized; the present invention describes the orbital optimization problem as an optimal control problem of the aircraft from the initial state to the final state and the problem of obtaining the extreme value of the objective function, and sets constraint conditions and boundary conditions, and can be solved by different methods according to the mission requirements. Different system parameters can be set according to different initial conditions, and it can be applied to orbital transformations of various configurations. Therefore, the applicable range of the target orbit is wide. BRIEF DESCRIPTION OF THE DRAWINGS

[0043] Figure 1 is a schematic diagram of the scheme of the re-entry into the near space aerodynamic large off-plane orbital reconstruction method;

[0044] Figure 2 is a schematic diagram of the orbit change of "retrograde re-entry - aerodynamic orbit change - ascending and orbital insertion";

[0045] Figure 3 is a schematic diagram of a common reference coordinate system;

[0046] Figure 4 is a schematic diagram of the trajectory of the aerodynamic orbit change section;

[0047] Figure 5 is a schematic diagram of the aircraft-fixed coordinate system. DETAILED DESCRIPTION OF THE INVENTION

[0048] Next, the technical solutions in the embodiments of the present invention will be clearly and completely described in conjunction with the accompanying drawings in the embodiments of the present invention. Obviously, the described embodiments are only a part of the embodiments of the present invention, rather than all the embodiments. All other embodiments obtained by those of ordinary skill in the art based on the embodiments of the present invention without creative efforts shall fall within the protection scope of the present invention.

[0049] It should be noted that, without conflict, the embodiments in the present invention and the features in the embodiments may be combined with each other.

[0050] The present invention will be further described below in conjunction with the accompanying drawings and specific embodiments, but it is not limited to the present invention.

[0051] A reentry near-space aerodynamic large-off-plane orbit reconstruction method in this embodiment includes:

[0052] Step 1: Define the reference coordinate systems required in the orbit design and calculate the conversion relationships between the coordinate systems;

[0053] Refer to Figure 1 and Figure 3 , and define the following common reference coordinate systems:

[0054] 1) Geocentric inertial coordinate system O i X i Y i Z i

[0055] The origin O of the coordinate system i is the center of the earth, the basic plane is the earth's equatorial plane, O i X i axis points to the vernal equinox point in the basic plane, O i Z i axis is perpendicular to the celestial equatorial plane and points to the north pole, O i Y i axis is in the equatorial plane and perpendicular to O i X i axis, and forms a right-handed rectangular coordinate system with O i X i axis and O i Z i axis.

[0056] 2) Geocentric fixed coordinate system OXYZ

[0057] The origin O of the coordinate system is the center of the earth, the OX axis points to the starting sub-meridian at a certain moment t0 in the equatorial plane, the OZ axis is perpendicular to the equatorial plane and points to the north pole, and the direction of the OY axis is determined by the right-hand rule. This coordinate system rotates with the earth's angular velocity ω e relative to the geocentric inertial coordinate system.

[0058] 3) Ground coordinate system O1X1Y1Z1

[0059] The origin O1 of the coordinate system is the launch point on the Earth's sea level, that is, the intersection point of the line connecting the center of mass of the aircraft and the center of the Earth with the sea level. The O1X1 axis is along the sea level and points due east. The O1Y1 axis is the line connecting the center of mass of the aircraft and the center of the Earth, with the upward direction being positive. The O1Z1 axis is a line on the sea level, and its direction is determined by the right-hand rule.

[0060] 4) Orbital coordinate system O o X o Y o Z o

[0061] The coordinate origin O o is the center of mass of the aircraft. The aircraft's orbital plane is the coordinate plane. The O o Z o axis points from the center of mass of the aircraft to the center of the Earth. The O o X o axis is in the orbital plane and is perpendicular to the O o Z o axis and points in the direction of the aircraft's velocity. The O o Y o axis is right-handed orthogonal to the O o X o and O o Z o axes and is parallel to the normal of the orbital plane.

[0062] 5) Body coordinate system O b X b Y b Z b

[0063] The coordinate origin O b is the center of mass of the aircraft. The O b X b axis is the axis of symmetry of the aircraft body, pointing towards the head of the aircraft. The O b Y b axis is in the main symmetry plane of the aircraft and is perpendicular to the O b X b axis. The O b Z b axis is right-handed orthogonal to the O b X b and O b Y b axes.

[0064] 6) Flight path coordinate system O2X2Y2Z2

[0065] The coordinate origin O2 is the instantaneous center of mass of the aircraft. The O2X2 axis coincides with the velocity vector v of the spacecraft. The O2Y2 axis is located in the vertical plane containing the velocity vector v and is positive upward perpendicular to the O2X2 axis. The O2Z2 axis is perpendicular to the other two axes and forms a right-handed coordinate system.

[0066] 7) Velocity coordinate system O v X v Y v Z v

[0067] The coordinate origin O v is the center of mass of the aircraft. The O v X v axis is the velocity direction of the aircraft. The O v Y v axis is in the main symmetry plane of the aircraft and perpendicular to the O v X v axis. The O v Z v axis is perpendicular to the O v X v Y v plane and points to the right.

[0068] From the definitions of the ground coordinate system O1X1Y1Z1 and the track coordinate system O2X2Y2Z2, it can be seen that the mutual relationship between the ground coordinate system and the track coordinate system is usually determined by the angle of track inclination and the yaw angle, and their definitions are as follows:

[0069] Track inclination γ: The angle between the velocity vector v of the spacecraft and the horizontal plane. When the velocity vector points above the horizontal plane, γ is positive.

[0070] Yaw angle ψ: The angle between the projection of the velocity vector v of the spacecraft in the horizontal plane and the O1X1 axis of the ground coordinate system. If the rotation from the O1X1 axis to the O2X2 axis is counterclockwise, the ψ angle is positive; otherwise, it is negative.

[0071] The transformation matrix relationship from the ground coordinate system [x1, y1, z1] to the track coordinate system [x2, y2, z2] is:

[0072]

[0073] where γ is the track inclination and ψ is the yaw angle.

[0074] The transformation matrix relationship from the velocity coordinate system [x v , y v , z v to the body coordinate system [x b , y b , z b is:

[0075]

[0076] In the formula, α is the angle of attack and β is the sideslip angle.

[0077] From the velocity coordinate system [x v , y v , z v to the ground coordinate system [x1, y1, z1], the conversion matrix relationship is:

[0078]

[0079] In the formula, γ is the flight path angle, ψ is the yaw angle, and σ is the roll angle.

[0080] Step 2: Design the braking re-entry section of the aircraft from the initial orbit into the near space. Consider the aircraft and the Earth as mass points. Before braking, it flies along the initial circular orbit. The braking re-entry section is as shown by the arc AB in Figure 1 and Figure 2 . Among them, i0 is the initial orbit inclination. The aircraft applies a velocity increment ΔV1 at point A to ignite and brake, and enters an elliptical transfer orbit with the perigee located within the atmosphere; ΔV1 is the braking velocity increment. Referring to Figure 2 , according to the conservation of angular momentum, we can obtain:

[0081]

[0082]

[0083]

[0084] In the formula, μ is the Earth's gravitational constant, r1 is the geocentric distance of the initial orbit, a1 is the semi-major axis of the braking re-entry section orbit, r a is the geocentric distance of the atmosphere, V A is the velocity at r1 after braking, γ0 is the re-entry angle, and V0 is the velocity entering the atmosphere.

[0085] From the formula, the expression of a1 is derived as:

[0086] All the variables on the right side of the formula are known, and the value of a1 can be calculated. Furthermore, the braking velocity V A can be obtained.

[0087] The initial orbit circular velocity V r1 is:

[0088]

[0089] Furthermore, the braking velocity increment ΔV1 can be obtained as:

[0090]

[0091] Step 3: Design the orbit of the aircraft during the aerodynamic orbit transfer section in the near space. The aerodynamic orbit transfer section is shown as the BD arc segment in Figure 1 and Figure 2 . The aircraft enters the atmosphere at point B with a speed of V0 and a reentry angle of γ0. When flying in the atmosphere, the orbit inclination is changed by using the aerodynamic lift, and the aircraft flies out of the atmosphere at point D with a speed of V f and a trajectory inclination of γ f .

[0092] Referring to Figure 2 and Figure 4 , point C on the transfer orbit is the lowest point of the orbit altitude. The aircraft enters the near space at point B (corresponding to time t0) with V0 on the transfer orbit, and changes the flight state of the aircraft in the atmosphere through optimal control. At point D (corresponding to time t f ), it flies out of the near space with V f , completing the aerodynamic assisted orbit transfer process. A certain thrust is applied during the flight of the aircraft, and the thrust direction coincides with the O b X b axis of the aircraft body coordinate system.

[0093] Assume that the atmosphere is relatively stationary with respect to the Earth, the aircraft control is in an ideal state, and due to the small Earth rotation speed ω e , the atmospheric rotation speed is much smaller than the flight speed of the aircraft, and the influence of the Earth's rotation can be ignored.

[0094] The velocity change ΔV2 and fuel consumption Δm during the aerodynamic orbit transfer section;

[0095] ΔV2 = V f - V0 (8)

[0096] Δm = m1(t0) - m1(t f ) (9);

[0097] m1(t0) represents the fuel mass at point B, t0 represents the time at point B, m1(t f ) represents the fuel mass at point D, and t f represents the time at point D;

[0098] Step 4: Design the orbit of the aircraft during the ascending orbit insertion section from the near space to the target orbit. The ascending orbit insertion section is shown as the DE arc segment in Figure 1 , Figure 2 . After jumping out of the atmosphere boundary from point D with a speed of V f and a track inclination of γ f , the orbit inclination becomes i f . At point E, ΔV3 is applied to enter the target orbit, and ΔV3 is the ascending velocity increment;

[0099] According to the conservation of angular momentum and the conservation of energy, it can be obtained that:

[0100]

[0101]

[0102]

[0103] The semi-major axis a3 of the transfer orbit in the ascending section is expressed as:

[0104]

[0105] The circumferential velocity V of the target orbit r2 is:

[0106]

[0107] From equations (10), (11) and (12), the ascending velocity increment ΔV3 can be obtained:

[0108]

[0109] r2 represents the radius of the target orbit;

[0110] Step 5: Optimize and solve the designed transfer orbit. As can be seen from the above steps, when the parameters of the initial orbit and the target orbit are given, by setting reasonable initial conditions (V0 and γ0) and optimizing and solving the aerodynamic orbit transfer section (BD section), the entire transfer orbit can be calculated.

[0111] The orbit optimization problem can be described as solving an optimal control problem: Set the index function J, and find the state and control (x(t), u(t)) that make the selected index function J optimal, that is:

[0112]

[0113] x(t) represents the state variable of the aircraft in the aerodynamic orbit transfer section, u(t) represents the control variable, t is the time, X represents the process index, and its specific form can be designed according to the mission requirements; φ[x(t f ),t f represents the performance index at time t f ; f[x(t), u(t), t] represents the motion dynamics equation of the aircraft; represents the integral performance index, reflecting the requirements for the entire process during the transfer process of the system from one point to another under the action of the control input u(t);

[0114] Set the constraint conditions and boundary conditions, and perform optimization to obtain the optimal control variable u(t).

[0115] Given the kinematic equation of the aircraft Given the initial time t0, the state x(t0) at the initial time, and the boundary conditions of the terminal state x(t f ), given the constraint conditions, find an admissible control u(t) that, within the specified time interval t0 ≤ t ≤ t f , starting from the initial state, transfers to the terminal state under the action of u(t), and the transfer process satisfies the constraint conditions, and minimizes the performance index J. The pseudospectral method or optimization can be carried out according to indicators such as fuel efficiency and orbit transfer time.

[0116] The orbit reconstruction scheme of this embodiment fully combines the atmospheric environment characteristics of the near space and the flight state of the aircraft, designs the braking reentry section, the aerodynamic orbit transfer section, and the ascent and orbit insertion section, and optimizes the designed orbit to form the orbit reconstruction scheme of this embodiment. This embodiment can achieve multiple and large off-plane orbit maneuvers of the aircraft.

[0117] In this embodiment, referring to Figure 5 , from the definition and transformation relationship of the reference coordinate system, the position coordinates (O i X i Y i Z i ) and velocity coordinates (O v x v y v z v ) of the aircraft can be expressed in the form of polar coordinates and (v, γ, ψ). Based on the above assumptions, the kinematic equation of the aircraft is established as follows:

[0118]

[0119] where x(t) = [r, φ, θ, v, γ, ψ, m], referring to Figure 5 , in the formula are the coordinates of the aircraft in the aircraft-fixed coordinate system (flight path coordinate system), m is the total mass of the aircraft, r and v are the scalar of the geocentric distance and flight speed respectively, and θ are the local geographical latitude and longitude of the aircraft respectively, γ and ψ are the flight path inclination angle and yaw angle of the aircraft respectively; D and L are the aerodynamic drag and aerodynamic lift respectively; c is the thrust coefficient; u(t) = [T, α, σ], T is the thrust, α is the angle of attack, and σ is the roll angle;

[0120] x(t) ∈ R n , R n is a column vector of dimension n, u(t) ∈ Rm , R m is a column vector with dimension m.

[0121] f(x, u, t) ∈ R n , is a continuous function of x(t), u(t), and t, and is differentiable with respect to x(t) and t.

[0122] The constraint conditions in this embodiment include heat flux density constraint, overload constraint, and dynamic pressure constraint;

[0123] The heat flux density constraint is:

[0124]

[0125] In the formula, Q(t) is the heat flux density at the stagnation point, R is the radius of the aircraft nose, h w is the wall enthalpy, h0 is the total enthalpy, and ρ is the atmospheric density;

[0126] The overload constraint is:

[0127] where n(t) is the three-axis overload of the aircraft in the velocity coordinate system, g0 is the acceleration due to gravity at sea level, C D and C L are the drag coefficient and lift coefficient respectively, and n max represents the maximum value of the three-axis overload of the aircraft in the velocity coordinate system;

[0128] The dynamic pressure constraint:

[0129] In the formula, q(t) is the dynamic pressure, and q max represents the maximum value of the dynamic pressure.

[0130] Equations (16) to (18) can be expressed as the constraint conditions that the system needs to satisfy:

[0131] C(x(t), u(t), t) ≤ 0, t0 ≤ t ≤ t f

[0132] Near space refers to the airspace with a height between 20 and 100 km, and the atmospheric density can be described by the atmospheric exponential model:

[0133] ρ = ρ S exp(-Hβ) (19)

[0134] In the formula, ρ is the atmospheric density, and ρ s is the atmospheric density at the Earth's sea level, H represents the flight altitude, and β = 1 / 6900m -1 The formula is H = r - r e , r e is the radius of the Earth.

[0135] The aerodynamic drag D and aerodynamic lift L of this embodiment are respectively:

[0136]

[0137] In the formula, S is the effective windward area of the aircraft.

[0138] The drag coefficient C D and lift coefficient C L of this embodiment are respectively:

[0139]

[0140] In the formula, C D0 , C D1 , C D2 , C L0 , C L1 are aerodynamic model parameters, which are C D0 = 0.06304, C D1 = -0.0061592, C D2 = 0.006214, C L0 = -0.20704, C L1 = 0.029244.

[0141] The constraint conditions of this embodiment also include control constraints;

[0142] The control constraint is:

[0143] {u ∈ R m : u L ≤ u ≤ u R} (22)

[0144] u L represents the lower boundary of the control variable;

[0145] u R represents the upper boundary of the control variable.

[0146] The boundary conditions of this embodiment include:

[0147] The initial orbit where the aircraft is located and the transfer orbit of the braking reentry section are tangent at the apogee, and there are boundary conditions at the initial moment:

[0148]

[0149] After the aircraft jumps out of the atmosphere, it enters the transfer orbit of the ascending orbit insertion section, and its apogee is tangent to the target orbit. There are boundary conditions at the terminal moment:

[0150]

[0151] if Denotes the orbital inclination angle of point D, ψ f Denotes the yaw angle of point D, Denotes the latitude of point D.

[0152] The constraint conditions at the initial and terminal times can be expressed as boundary conditions:

[0153] φ[x(t0), t0, x(t f ), t f = 0

[0154] φ[x(t f ), t f represents the final state performance index (at time t f ), reflecting the requirements for the final state performance. φ[x(t), t] represents the state performance index of the system at time t. φ[x(t0), t0, x(t f ), t f = 0 represents the constraint conditions (boundary conditions) at the initial and terminal times.

[0155] This embodiment proposes an orbital reconstruction scheme of "re - entry into near - space - aerodynamic large - plane - change orbit - repeated orbit - entry", including the orbital design of the braking re - entry section, the orbital design of the aerodynamic orbit - change section, the orbital design of the ascending orbit - entry section, and orbit optimization; the orbital design of the braking re - entry section establishes a motion model with reference to the two - body problem and calculates the braking velocity increment ΔV1; the orbital design of the aerodynamic orbit - change section combines the characteristics of the near - space atmosphere environment and the flight state of the aircraft to establish a near - space atmosphere model and a re - entry vehicle motion model; the orbital design of the ascending orbit - entry section establishes a transfer - orbit vehicle motion model based on the two - body motion model and target orbit parameters and calculates the ascending velocity increment ΔV3; the orbit optimization describes the orbit optimization problem as the optimal control of the aircraft from the initial state to the final state and the acquisition of the extreme value of the objective function, selects thrust, angle of attack, and roll angle as control variables, and sets constraint conditions and boundary conditions; this embodiment can reduce the fuel consumption of strong - maneuver orbit - change, achieve large - plane strong - maneuver orbit - change, enhance the orbit - change ability and increase the number of orbit - changes, and improve the space confrontation ability of future spacecraft.

[0156] Although the present invention has been described herein with reference to specific embodiments, it should be understood that these embodiments are merely examples of the principles and applications of the present invention. Therefore, it should be understood that many modifications can be made to the exemplary embodiments, and other arrangements can be designed, as long as they do not deviate from the spirit and scope of the present invention as defined by the appended claims. It should be understood that the different dependent claims and the features described herein can be combined in a manner different from that described in the original claims. It should also be understood that the features described in connection with a single embodiment can be used in other described embodiments.

Claims

1. A method for re - entry near - space aerodynamic large - different - plane orbit reconstruction, characterized in that, The method includes: S1: Design the braking reentry section orbit for the aircraft to enter the near space from the initial orbit: The aircraft performs a braking ignition by applying an acceleration increment ΔV1 at point A. μ is the Earth's gravitational constant, r1 is the geocentric distance of the initial orbit, r a is the geocentric distance of the atmosphere, V0 is the velocity entering the atmosphere, and γ0 is the reentry angle; S2: Design the aerodynamic orbit-changing section orbit for the aircraft in the near space: The starting point of the aerodynamic orbit transfer section is point B. The aircraft enters the atmosphere at point B with a velocity V0 and a reentry angle γ0. When flying in the atmosphere, the aerodynamic lift is used to change the orbit inclination, and at point D, it flies out of the atmosphere with a velocity V f , trajectory inclination γ f , the velocity change ΔV2 and fuel consumption Δm in the aerodynamic orbit transfer section; ΔV2 = V f - V0, Δm = m1(t0) - m1(t f ) m1(t0) represents the fuel mass at point B, t0 represents the time at point B, m1(t f ) represents the fuel mass at point D, t f represents the time at point D; S3: Design the ascending orbit-insertion section orbit for the aircraft to enter the target orbit from the near space: The starting point of the ascending orbit insertion segment is point D, and from point D at a velocity V f , the flight path inclination γ f After jumping out of the atmosphere boundary, the orbit inclination becomes i f , and at point E, ΔV3 is applied to enter the target orbit. ΔV3 is the ascending velocity increment; r2 represents the radius of the target orbit; S4: Set the index function J, and seek the state and control (x(t), u(t)) that make the selected index function J optimal, that is: $x(t)$ represents the state variables of the aircraft during the aerodynamic orbit transfer phase, $u(t)$ represents the control variables, $t$ is the time, $X$ represents the process index, and $\varphi[x(t f ),t f $ represents the performance index at time $t$; f ; $f[x(t),u(t),t]$ represents the kinematic equation of the aircraft; And set the constraint conditions and boundary conditions, perform optimization and solution to obtain the optimal control variable u(t).

2. The reentry near-space aerodynamic large different-plane orbit reconstruction method according to claim 1, wherein The where \(u(t)=[T,\alpha,\sigma]\), \(T\) is the thrust, \(\alpha\) is the angle of attack, and \(\sigma\) is the roll angle. \(m\) is the total mass of the aircraft, \(r\) and \(v\) are the geocentric distance and the scalar of the flight speed respectively. \(\lambda\) and \(\theta\) are the local geographic latitude and longitude of the aircraft respectively, \(\gamma\) and \(\psi\) are the flight path angle and the yaw angle of the aircraft respectively; \(D\) and \(L\) are the aerodynamic drag and the aerodynamic lift respectively; \(c\) is the thrust coefficient. x(t) ∈ R n , R n is a column vector of dimension n, u(t) ∈ R m , R m is a column vector of dimension m.

3. The reentry near-space aerodynamic large different-plane orbit reconstruction method according to claim 2, characterized in that The constraint conditions include heat flux density constraint, overload constraint and dynamic pressure constraint; The heat flux density constraint is: where \(Q(t)\) is the heat flux density at the stagnation point, \(R\) is the radius of the aircraft nose, \(h\) w is the wall enthalpy, \(h_0\) is the total enthalpy, and \(\rho\) is the atmospheric density; The overload constraint is: Among them, n(t) is the three-axis overload of the aircraft in the velocity coordinate system, g0 is the acceleration due to gravity at sea level, C D and C L are the drag coefficient and the lift coefficient respectively, and n max represents the maximum value of the three-axis overload of the aircraft in the velocity coordinate system; The dynamic pressure constraint: where q(t) is the dynamic pressure, and q max represents the maximum value of the dynamic pressure.

4. The reentry near-space aerodynamic large different-plane orbit reconstruction method according to claim 3, characterized in that, The atmospheric density is: ρ = ρ S exp(-Hβ) where ρ s is the atmospheric density at the Earth's sea level, H represents the flight altitude, and β = 1 / 6900 m -1 .

5. The reentry near-space aerodynamic large different-plane orbit reconstruction method according to claim 4, characterized in that The aerodynamic drag D and aerodynamic lift L are respectively: In the formula, S is the effective windward area of the aircraft.

6. The reentry near-space aerodynamic large different-plane orbit reconstruction method according to claim 5, wherein C D = C D0 + C D1 α + C D2 α 2 C L = C L0 + C L1 α where C D0 , C D1 , C D2 , C L0 , C L1 are pneumatic model parameters, specifically C D0 = 0.06304, C D1 = -0.0061592, C D2 = 0.006214, C L0 = -0.20704, C L1 = 0.029244.

7. The reentry near-space aerodynamic large different-plane orbit reconstruction method according to claim 3, wherein The constraint conditions also include control constraints; The control constraint is: {u ∈ R m : u L ≤ u ≤ u R} u L represents the lower boundary of the control variable; u R represents the upper bound of the control variable.

8. The reentry near-space aerodynamic large-aspect-ratio orbit reconstruction method according to claim 3, characterized in that The boundary conditions include: The initial orbit where the aircraft is located is tangent to the transfer orbit of the braking reentry section at the apogee, and there are boundary conditions at the initial moment: After the aircraft jumps out of the atmosphere, it enters the transfer orbit of the ascending orbit-insertion section, and its apogee is tangent to the target orbit, and there are boundary conditions at the terminal moment: i f represents the orbital inclination of point D, ψ f represents the yaw angle of point D, represents the latitude of point D.

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