CN113093776A

CN113093776A - Method and device for determining off-orbit parameters of spacecraft

Info

Publication number: CN113093776A
Application number: CN202110239649.9A
Authority: CN
Inventors: 李革非; 徐海涛; 马传令; 郝大功; 陈明; 盛庆轩; 欧阳琦; 李翠兰
Original assignee: Beijing Aerospace Control Center
Current assignee: Beijing Aerospace Control Center
Priority date: 2021-03-04
Filing date: 2021-03-04
Publication date: 2021-07-09
Anticipated expiration: 2041-03-04
Also published as: CN113093776B

Abstract

A method and a device for determining an off-orbit parameter of a spacecraft comprise the following steps: calculating an alternative derailment brake parameter according to an orbit parameter before the spacecraft leaves the orbit, an aiming reentry parameter after the spacecraft leaves the orbit and an aiming landing point, determining an actual reentry parameter for controlling the spacecraft after the spacecraft leaves the orbit according to the alternative derailment brake parameter and a preset guidance control algorithm adopted after the spacecraft leaves the orbit, correcting the aiming reentry parameter if the reentry deviation is determined not to meet the reentry precision requirement, executing the step of calculating the alternative derailment brake parameter until the reentry deviation is determined to meet the reentry precision requirement, calculating an actual landing point of the spacecraft according to the actual reentry parameter, correcting the aiming landing point if the landing deviation of the spacecraft is determined not to meet the landing precision requirement, executing the step of calculating the alternative derailment brake parameter until the landing deviation is determined to meet the landing precision requirement, obtaining the actual derailment brake parameter, initially determining the aiming reentry parameter according to the given reentry parameter, the targeted landing site is determined from the given landing site.

Description

Method and device for determining off-orbit parameters of spacecraft

Technical Field

The application relates to the technical field of spaceflight, in particular to a method and a device for determining an off-orbit parameter of a spacecraft.

Background

After the spacecraft flies on orbit for a period of time to complete a preset task, the braking engine can be utilized to generate thrust to reduce the flying speed, so that the spacecraft enters an elliptical orbit returning to the earth from an orbit surrounding the earth, and the off-orbit braking is realized.

The reentry section trajectory of the ballistic reentry spacecraft is determined by the reentry velocity and the reentry angle, and the position of the reentry point, the reentry angle and the reentry velocity are determined by the position of the braking point and the braking velocity, so that, for the ballistic reentry spacecraft, once the braking state is determined, the entire reentry trajectory is determined. That is, the deviation of the trajectory of the ballistic reentry spacecraft, the deviation of the position of the braking point, the deviation of the braking attitude, the deviation of the braking speed, the deviation of the atmospheric parameter, and the like are accumulated to the end point of the reentry trajectory, i.e., the landing point, which makes the landing point of the ballistic reentry spacecraft highly dispersive.

Because of the importance of reentry points, some ballistic reentry spacecraft braking control employ guidance control methods that target reentry parameters. However, because the real-time computing capability of the spacecraft is limited, an analytic guidance algorithm is generally adopted to calculate the control quantity in each control period and control the brake engine to work. Due to the simplification of the guidance algorithm and the setting of the control state, the guidance method aiming at the reentry parameter brings deviation to the reentry point. After the spacecraft reaches the reentry point, the spacecraft flies to the landing point according to the reentry section kinetic equation, so that the reentry point deviation is accumulated to the landing point, and the landing precision of the spacecraft is low.

Disclosure of Invention

The embodiment of the application provides a method and a device for determining an off-orbit parameter of a spacecraft, which are used for solving the problem that the spacecraft has low landing precision after off-orbit in the prior art.

In a first aspect, an embodiment of the present application provides a method for determining an off-orbit parameter of a spacecraft, including:

calculating alternative derailment braking parameters according to orbit parameters before the spacecraft leaves the orbit, aiming reentry parameters after the spacecraft leaves the orbit and aiming landing points, wherein the aiming reentry parameters are determined according to given reentry parameters at the beginning, and the aiming landing points are determined according to given landing points;

determining actual reentry parameters for controlling the spacecraft after the derailment according to the alternative derailment braking parameters and a preset guidance control algorithm adopted after the derailment;

if the reentry deviation of the spacecraft is determined to not meet the reentry precision requirement according to the actual reentry parameter and the given reentry parameter, correcting the aiming reentry parameter, and executing the step of calculating the alternative derailment braking parameter until the reentry deviation is determined to meet the reentry precision requirement, and calculating the actual landing point of the spacecraft according to the actual reentry parameter;

and if the landing deviation of the spacecraft is determined to not meet the landing precision requirement according to the actual landing point and the given landing point, correcting the aiming landing point, and executing the step of calculating the alternative off-orbit brake parameter until the landing deviation is determined to meet the landing precision requirement, and taking the current alternative off-orbit brake parameter as the actual off-orbit brake parameter.

In some possible embodiments, determining an actual reentry parameter for controlling the spacecraft after the derailment according to the candidate derailment braking parameter and a preset guidance control algorithm adopted after the derailment, includes:

calculating expected reentry parameters after controlling the spacecraft to be out of orbit according to the alternative off-orbit brake parameters;

taking the expected reentry parameter as an off-orbit brake control target, and gradually adjusting the motion parameter of the spacecraft after the off-orbit by adopting the preset guidance control algorithm;

and determining the actual reentry parameter of the spacecraft after the spacecraft is out of orbit based on the adjusted motion parameter of the spacecraft.

In some possible embodiments, calculating an actual landing site of the spacecraft from the actual reentry parameter comprises:

calculating an actual reentry point of the spacecraft according to the actual reentry parameter;

and calculating the actual landing point of the spacecraft according to the actual re-entry point and the preset position conversion relation between the re-entry point and the landing point.

In some possible embodiments, the aiming reentry parameters include an aiming reentry height and an aiming reentry angle, the reentry deviation including a reentry height deviation and a reentry angle deviation;

modifying the aiming reentry parameter, comprising:

the sum of the aiming re-entry height and the re-entry height deviation is taken as a new aiming re-entry height, and the sum of the aiming re-entry angle and the re-entry angle deviation is taken as a new aiming re-entry angle.

In a second aspect, an embodiment of the present application provides an apparatus for determining an off-orbit parameter of a spacecraft, including:

the off-orbit parameter determination module is used for calculating alternative off-orbit brake parameters according to orbit parameters before the spacecraft is off-orbit, aiming reentry parameters after the spacecraft is off-orbit and aiming landing points, and initially, the aiming reentry parameters are determined according to given reentry parameters, and the aiming landing points are determined according to given landing points;

the reentry parameter determination module is used for determining an actual reentry parameter for controlling the spacecraft after the derailment according to the alternative derailment braking parameter and a preset guidance control algorithm adopted after the derailment;

the reentry precision control module is used for correcting the aiming reentry parameter and executing the step of calculating an alternative derailment braking parameter if the reentry deviation of the spacecraft is determined to not meet the reentry precision requirement according to the actual reentry parameter and the given reentry parameter, and calculating the actual landing point of the spacecraft according to the actual reentry parameter until the reentry deviation is determined to meet the reentry precision requirement;

and the landing precision control module is used for correcting the aiming landing point and executing the step of calculating the alternative off-orbit brake parameter if the landing deviation of the spacecraft is determined to not meet the landing precision requirement according to the actual landing point and the given landing point, and taking the current alternative off-orbit brake parameter as the actual off-orbit brake parameter until the landing deviation is determined to meet the landing precision requirement.

In some possible embodiments, the reentry parameter determination module is specifically configured to:

In some possible embodiments, the reentry precision control module is specifically configured to:

the reentry precision control module is specifically configured to:

In a third aspect, an embodiment of the present application provides an electronic device, including: at least one processor, and a memory communicatively coupled to the at least one processor, wherein:

the memory stores instructions executable by the at least one processor to enable the at least one processor to perform the above-described method for determining off-orbit parameters of a spacecraft.

In a fourth aspect, embodiments of the present application provide a storage medium, where instructions in the storage medium are executed by a processor of an electronic device, and the electronic device is capable of executing the method for determining the off-orbit parameters of the spacecraft.

In the embodiment of the application, according to the orbit parameter before the spacecraft leaves the orbit, the aiming reentry parameter after the spacecraft leaves the orbit and the aiming landing point, the alternative derailment brake parameter is calculated, the actual reentry parameter after the spacecraft leaves the orbit is determined according to the alternative derailment brake parameter and the preset guidance control algorithm adopted after the spacecraft leaves the orbit, if the reentry deviation of the spacecraft is determined not to meet the requirement of the reentry precision, the aiming reentry parameter is corrected, the step of calculating the alternative derailment brake parameter is executed until the reentry deviation is determined to meet the requirement of the reentry precision, the actual landing point of the spacecraft is calculated according to the actual reentry parameter, if the landing deviation of the spacecraft is determined not to meet the requirement of the landing precision, the aiming landing point is corrected, the alternative derailment brake parameter is calculated until the landing deviation is determined to meet the requirement of the landing precision, the current alternative derailment brake parameter is taken as the actual derailment brake parameter, initially, the targeting reentry parameter is determined from the given reentry parameter and the targeting landing site is determined from the given landing site. Therefore, when the spacecraft is subjected to the derailment control, the reentry deviation of the spacecraft is corrected, the landing deviation of the spacecraft is corrected after the reentry deviation is determined to meet the reentry precision requirement, the situations that the reentry deviation is accumulated to a landing point and the landing deviation is increased can be well avoided, and the derailment parameter can be corrected reversely based on the landing deviation of the spacecraft, so that the finally determined actual derailment brake parameter can ensure the landing precision of the spacecraft after the derailment.

Drawings

The accompanying drawings, which are included to provide a further understanding of the application and are incorporated in and constitute a part of this application, illustrate embodiment(s) of the application and together with the description serve to explain the application and not to limit the application. In the drawings:

fig. 1 is a schematic view of a landing process of a spacecraft according to an embodiment of the present application;

fig. 2 is a schematic diagram illustrating a position relationship between an off-track point and a landing point according to an embodiment of the present disclosure;

fig. 3 is a flowchart of a method for determining an off-orbit parameter of a spacecraft according to an embodiment of the present application;

fig. 4 is a flowchart of a method for determining an off-orbit parameter of a spacecraft according to an embodiment of the present application;

fig. 5 is a schematic structural diagram of an apparatus for determining an off-orbit parameter of a spacecraft according to an embodiment of the present application;

fig. 6 is a schematic hardware structure diagram of an electronic device for implementing a method for determining an off-orbit parameter of a spacecraft according to an embodiment of the present application.

Detailed Description

In order to solve the problem that the landing precision of a spacecraft after the spacecraft leaves the orbit is low in the prior art, the embodiment of the application provides a method and a device for determining the off-orbit parameters of the spacecraft.

The preferred embodiments of the present application will be described below with reference to the accompanying drawings of the specification, it should be understood that the preferred embodiments described herein are merely for illustrating and explaining the present application, and are not intended to limit the present application, and that the embodiments and features of the embodiments in the present application may be combined with each other without conflict.

Fig. 1 is a schematic view of a spacecraft landing process according to an embodiment of the present disclosure, in which an orbit of the spacecraft around the earth before the spacecraft leaves the orbit is circular, the spacecraft enters an elliptical orbit returning to the earth after braking at an off-orbit point, a reentry point is located on the elliptical orbit, and then the spacecraft returns to a landing point on the earth surface from the reentry point.

Generally, the return trajectory of a spacecraft is divided into four segments:

(1) off-track section track

This segment is from the start of operation of the return brake or return transition engine to the end of its operation. Under the action of the return brake or return orbital transfer engine, the spacecraft leaves the original orbit and enters an orbit leading to the ground.

(2) Transition section track

This phase is from the end of the return braking or return orbital transfer engine operation to the passive phase prior to entering the earth's atmosphere. The earth atmosphere height is generally 80km-120km, and the orbit is not controlled in this stage.

(3) Reentry section rail

The section starts from the beginning of the spacecraft entering the atmosphere and ends at 10km from the ground. In this section, the spacecraft is subjected to severe aerodynamic heating, external pressure and large overload.

(4) Landing segment track

The segment is a final segment orbit which makes the spacecraft safely land on the earth surface by utilizing a parachute or other deceleration forms.

Because the dynamics of the spacecraft influenced by aerodynamic force outside and in the atmosphere are obviously different, the spacecraft motion is described by adopting a dynamic equation of an operation section in an off-orbit section and a transition section, and a dynamic equation of a reentry section is adopted in a reentry section and a landing section.

Equation of motion section dynamics

Under the assumption of an earth center gravitational field, a Cowell form is adopted to describe the spacecraft dynamics equation as follows:

in the formula,

spacecraft position and velocity vectors, respectively;

is the acceleration of the earth's non-spherical gravity;

is the atmospheric drag acceleration;

third body acceleration due to the sun, moon;

is the sunlight pressure acceleration;

is the thrust acceleration provided by the spacecraft engine.

Specific calculation models for each type of acceleration can be found in the relevant literature.

For a spacecraft under only limited thrust from the engine, the following kinetic equations apply:

wherein, F is the thrust of the engine,

(or write as

) For propellant consumption rate, I_spIs the effective exit velocity of the combustion products, also known as the specific impulse.

Equation of dynamics at reentry stage

And when the reentry flight trajectory and the landing point of the spacecraft are accurately calculated, an accurate reentry section kinetic equation is adopted.

The six-degree-of-freedom kinetic equation of the reentry section of the spacecraft is as follows:

wherein:

r is the ground-center distance of the spacecraft;

λ is longitude of the spacecraft;

the geocentric latitude of the spacecraft;

v is the spacecraft velocity;

gamma is a flight speed inclination angle;

psi is the airspeed azimuth;

d is pneumatic resistance;

l is aerodynamic lift;

m is the spacecraft mass;

g is the local gravitational acceleration;

ω_Eis the rotational angular velocity of the earth.

Method for calculating landing site based on reentry range matching reentry point

In reentry flight of the spacecraft, the ballistic flight plane is substantially unchanged, i.e. the reentry point and the landing point are substantially located in the same ballistic plane. And when the actual reentry point meets the requirement of the given reentry point, the reentry flight range is changed less compared with the given reentry trajectory. Therefore, according to the idea that the reentry flight is basically unchanged, the reentry trajectory flight is given by the spacecraft, and the reentry point and the landing point are matched in the flight trajectory plane.

Referring to fig. 2, let the re-entry point be B1, the landing point be B2, and the longitude of the re-entry point be λ_EThe geocentric latitude is

Track inclination angle is i, and given reentry range is

1) Formed by a spherical triangle AL₁B₁Calculate the arc AB₁Length d of₁：

2) Initially, aim and re-enter flight

Arc AB₂The length of (A) is as follows:

3) the geocentric latitude of the landing point located on the ballistic flight plane with the reentrant point is as follows:

4) arc AL₁Has a length of λ₁，AL₂Length of lambda₂Is calculated according to the following formula

The longitude of the landing site is:

λ_L＝λ_E+λ₁-λ₂ (8)

that is, the longitude relationship between the landing point and the re-entry point is:

5) calculating the actual voyage of the reentry point and the landing point as follows:

will actually make a voyage d_ELAnd given voyage

And comparing to obtain a range deviation:

6) d 'is corrected to d + delta d, d' is used as new d, and the operation returns to 2), the longitude and the latitude of the landing point are obtained through iterative calculation, and when | delta d | is less than or equal to epsilon, the longitude lambda of the landing point meeting the flight precision is obtained_LAnd latitude

Wherein epsilon is the convergence threshold.

In order to enable the spacecraft to land accurately finally, a given landing point can be regarded as a target for brake control, and in order to meet the requirement of reentry overload, the reentry angle of the spacecraft during reentry is also required. The reentry point and the landing point are two different characteristic points, and a certain range constraint requirement is required to be met between the two characteristic points.

Therefore, the brake control aiming target parameters of the spacecraft reentry landing comprise:

1) aiming at a landing site: longitude (G)

Latitude

Initially, the target landing site is a given landing site with a longitude

At a latitude of

2) Aiming at a reentry point: angle of re-entry for aiming

Initially, the aiming re-entry angle is a given re-entry angle

And (3) constraint:

1) given the reentry height:

spacecraft re-entry altitude is defined as the local altitude into the dense atmosphere, typically chosen to be 100 km;

2) and (3) giving a reentry section:

the spacecraft carries out braking control on a certain point on a flight orbit, and the reentry angle meets the reentry precision requirement and the spacecraft lands on a given landing point under the condition of an optimized index with the minimum braking speed increment value and under the condition of meeting a range constraint condition between the reentry point and the landing point by adjusting the braking time and the braking speed increment.

The braking control pulse along the opposite direction of the flight speed vector of the spacecraft is adopted, so that the optimization condition of the minimum braking speed increment value can be met.

The velocity coordinate system is defined as: the origin is located in the center of mass of the spacecraft, the X axis is along the velocity vector direction of the spacecraft, the Y axis is perpendicular to the orbit plane and points to the negative normal direction of the orbit plane, and the Z axis, the X axis and the Y axis form a right-hand system.

Setting the position vector of the inertial system of the spacecraft as

Velocity vector of

The three direction vectors of the speed coordinate system are respectively as follows:

as the longitude and latitude of the landing point need to be aimed and controlled, the brake control pulse comprises 2 parameters of the brake speed along the X-direction component and the Y-direction component of the speed system, and the corresponding spacecraft brake control attitude is the speed system directional brake attitude and is expressed by a yaw angle.

The spacecraft brake control planning variables are: braking time t_BThe component Deltav of the braking speed increment along the X direction of the speed system_BxAnd a Y-direction component Δ v_By. The target parameters are: longitude λ of landing Point_LLatitude and longitude

Reentry angle gamma_E. The solution relationship of 3 to 3.

The correspondence between the braking control plan variables and the given parameters is shown in the following table.

TABLE 1 brake control planning variables and target parameters

Calculation of initial values of brake control parameters (i.e. alternative brake control parameters)

The flight speed value v of the spacecraft at the braking point is as follows:

wherein a is a semi-long axis of the track, r is a geocentric distance, and mu is a gravitational force parameter.

Target velocity value v of the point^TComprises the following steps:

when the semi-major axis of the target

When, i.e. after brake control, the height of the target track in the vicinity thereof is equal to 0, wherein R_EThe equatorial radius of the earth.

The initial value of the X-direction component of the brake speed increment is obtained as follows:

Δv_Bx＝v^T-v (13)

the initial brake control position selects a position of orbital symmetry at a given landing point, and Δ v_By＝0。

Known landing site longitude λ_LAnd latitude

The track inclination angle is i, and the track amplitude u of the landing point can be obtained according to a spherical trigonometric formula_L：

Track amplitude u of brake control position_BComprises the following steps: u. of_B＝u_L-π。

Track amplitude u of brake control position_BDetermines the brake control time t_B。

Iterative calculation of brake control parameters

And obtaining an initial value of the braking control parameter by adopting a braking control initial value calculation method, and after the spacecraft applies the braking control speed increment, extrapolating the orbit to a reentry point to obtain a reentry point parameter. And obtaining the longitude and latitude of the landing point according to the reentry journey matching calculation method and the reentry point longitude and latitude. The calculated reentry angle, the longitude and the latitude of the landing point have certain deviation with the given reentry angle, the longitude and the latitude of the given landing point, and an iterative method is adopted to accurately calculate the brake control parameters and eliminate the deviation of the reentry angle, the longitude and the latitude of the landing point.

And performing iterative correction calculation by adopting a differential correction method.

And calculating the change of the target parameters according to the tiny deviation of the planning variables to obtain a sensitive matrix of the target parameters to the planning variables, which is also called a Jacobian matrix K.

The known planning variable is the braking time t_BIncrement of braking speed

Component Δ v in the direction of X, Y_Bx，ΔvBv

The target parameter being the reentry angle gamma_ELongitude of the landing site λ_LLatitude and longitude

The correction of the planned variable is calculated on the basis of the actual deviation of the target parameter.

The braking control plan variables are modified as follows: t'_B＝t_B+δt_B，Δv_Bx′＝Δv_Bx+δΔv_Bx，Δv_By′＝Δv_By+δΔv_By，

T 'is'_BAs new t_BWill Δ v_Bx' As a novel Δ v_BxWill Δ v_By' As a novel Δ v_By。

The relationship between the reentry point and the landing point of the reentry range matching constraint is an ideal model of the reentry section, the reentry point is used for calculating reentry parameters of the spacecraft orbit by adopting an on-orbit flight section dynamic model, and the landing point parameters can be obtained through the ideal model of the reentry range matching.

In actual flight, the spacecraft is guided and controlled by a Guidance Navigation and Control (GNC) system according to the position and the speed of the braking moment calculated on the ground and the position and the speed of a reentry point according to a Control cycle, the speed increment required by braking is calculated according to a Guidance algorithm in each Control cycle, the braking Control attitude is adjusted to be the vector direction of the braking speed increment, and the thruster is instructed to execute the braking speed increment Control.

The guidance algorithm adopts a two-body Lambert guidance algorithm.

According to the two-body orbit initial value theory, the position vector and the velocity vector of the initial time and the terminal time are connected through a Lagrange coefficient:

wherein F and G are Lagrange coefficients,

and delta f is an included angle between the two vectors, and P is the track half-diameter.

The boundary problem with the terminal reentry angle γ as a constraint can be derived:

in the formula,

let the current position vector of the spacecraft be

Velocity vector of

The location vector of the brake aiming re-entry point is

The targeted re-entry point velocity vector is

To satisfy the reentry angle of

The current-time target velocity vector of the guidance calculation can be solved according to equation (17):

the brake velocity increment vector of the guidance calculation is as follows:

applying the brake speed increment vector calculated by guidance to a spacecraft high-precision dynamic model, and extrapolating and calculating to the time of a reentry point to obtain parameters of the reentry point controlled by guidance

Wherein,

position vector and velocity vector, h, of the re-entry point obtained for guidance control, respectively_EG，γ_EGRespectively, the re-entry height and re-entry angle obtained by the guidance control.

In the related technology, after the spacecraft reaches the reentry point, the spacecraft flies to the landing point according to the reentry section kinetic equation, and the spacecraft ballistic reentry mode has no correction capability on the reentry ballistic, so that the reentry point deviation directly influences the landing point deviation. The guidance control adopts a two-body Lambert guidance algorithm with a simplified model, and the state of the reentry point of the spacecraft cannot reach the given reentry point state, so the guidance control deviation can influence the precision of the landing point of the spacecraft, and the actual landing point deviates from the given landing point.

Therefore, the double-layer iterative correction method for correcting the reentry point aiming parameters through inner layer guidance control and correcting the landing point aiming parameters through outer layer braking control is provided, and the landing precision of the ballistic reentry of the spacecraft is improved. Referring to fig. 3, the process is as follows.

1) Inputting orbit parameters before the spacecraft leaves the orbit and given reentry parameters after the spacecraft leaves the orbit: given reentrant height

Given reentrant angle

And given a landing site: longitude (G)

Latitude

2) Aiming reentry parameters are set according to given reentry parameters: aiming re-entry height

Angle of re-entry for aiming

And aiming the landing site according to the given landing site setting: longitude of the target landing Point

Latitude

3) According to the orbit parameters and aiming reentry parameters before the spacecraft leaves the orbit:

and aiming at the landing site:

and latitude

By taking voyage based on re-entry points and landing points

The matched brake control parameter calculation method is based on the kinetic equation of the operation section, and brake control parameters including the brake time t are obtained through iterative calculation_BPosition vector of braking time

Velocity vector

And velocity delta vector

And calculating expected reentry parameters after braking control based on the braking control parameters and the motion model of the operation segment of the spacecraft, including reentry time t_EReentry location vector

Velocity vector

Re-entry height h_EBAnd re-entry angle gamma_EBSo that when re-entering the height

When the re-entry angle, and the longitude and latitude of the landing point reach given values, i.e.

At the time of the first calculation,

4) using two-body Lambert guidanceAlgorithm according to braking time t_BPosition vector of spacecraft

Sum velocity vector

Re-entry point time t after brake control_EReentry location vector

Sum velocity vector

Calculating an actual reentry parameter after guidance control, including a reentry time t_EReentry location vector

Sum velocity vector

Re-entry height h_EGAnd re-entry angle gamma_EG。

5) Calculating the deviation of the actual reentry parameter after guidance control and the given reentry parameter:

6) modifying the targeting reentry parameters of the braking control, including the targeting reentry height and the targeting reentry angle:

will be provided with

As new

Will be provided with

As new

7) If the reentry height and reentry angle after guidance control meet the reentry precision requirement, namely delta h_G≤εh，Δγ_GTurning to 8 when the epsilon gamma is less than or equal to the epsilon gamma); otherwise, turning to 3), and continuing iterative calculation to eliminate the deviation of the re-entry height and the re-entry angle of the guidance control, wherein epsilon h and epsilon gamma are the precision requirements of the re-entry height and the re-entry angle.

7) According to reentry parameters after guidance control

Calculating the flight trajectory of the reentry section based on the reentry section kinetic equation to obtain the longitude lambda and the latitude of the actual landing point

8) Calculating the deviation of the actual landing point from the given landing point:

9) a targeted landing site for modified braking control, comprising longitude and latitude of the targeted landing site:

will be provided with

As new

Will be provided with

As new

10) If the longitude and latitude of the landing point both meet the landing accuracy requirement, that is

The calculation is finished, and the brake control parameters including the brake time and the brake speed increment vector are output; otherwise, turning to 3), continuing iterative computation, and finally eliminating the deviation of longitude and latitude of the landing point, wherein

Landing site longitude and latitude accuracy requirements.

The following describes simulation examples.

Assume the simulation input parameters are as follows:

1) spacecraft parameters

The mass of the spacecraft is 260kg, and the average area of the spacecraft is 0.95m²。

2) Spacecraft return given parameters

And (3) giving a landing point: 102.186 DEG E, 40.905 DEG N,

inertial velocity gives reentry angle: -1.9 °,

given the reentry height: the length of the beam is 100km,

and (3) giving a reentry section: 1100.0 km.

3) Spacecraft GNC control

And the spacecraft GNC performs guidance control according to the working cycle of 50 ms.

4) Number of orbits of spacecraft

Epoch time: 2020-05-06T 12: 32: 38,

number of orbits in J2000 inertial coordinate system:

semi-major axis 6641110m

The eccentricity is 0.0152,

the inclination angle is 41.011 degrees,

the right ascension point is 324.657 degrees,

the argument of the perigee is 171.139 degrees,

the mean anomaly angle is 103.425 °.

5) Dynamic model

The operation section dynamic model comprises: the gravity of the earth center, the gravity of the earth aspheric shape 32X32 order, the gravity of the mass in the daytime and the moon, and the atmospheric resistance. The atmosphere model adopts NRLMSISE 2000, the geomagnetic index AP is 10, the solar radiation flow F10.7 is 70, and the damping coefficient CD is 2.4.

The reentry section kinetic model comprises: gravity at the center of the earth, aerodynamic drag and aerodynamic lift.

6) Calculating convergence accuracy

Re-entering height: the thickness of the film is less than 100m,

reentry angle: less than 0.001 degree,

landing site longitude: less than 0.01 degree,

landing site latitude: is less than 0.01 degrees.

Simulation calculation result

Table 2 shows the results of the double-layer iterative calculation of guidance control and brake control for correcting the re-entry point deviation and the landing point deviation. And 3, aiming parameter correction results of reentry and landing of double-layer iteration of guidance control and braking control are shown in the table. And table 4 shows the results of the brake control parameters for the two-level iteration of guidance control and brake control.

TABLE 2 double-layer iterative computation results for brake control and guidance control

TABLE 3 aiming parameter modification results for reentry and landing

	Reentry height/km	Re-entry angle/°	Longitude/deg. of landing site	Landing Point latitude/°
					Given value	100.000	-1.900	102.186	40.905
Aiming value	100.528	-1.891	102.252	40.922
					Deviation of aim	0.528	+0.009	0.066	0.017
Actual value	100.075	-1.899	102.182	40.904
					Actual deviation	+0.075	+0.001	-0.004	-0.001

TABLE 4 braking control results

Moment of braking	Yaw angle/°	Braking speed/m/s
			2020-05-06T12：42：38	-175.536	77.058

Tables 2 and 3 show that:

1) during the first calculation, the reentry height of the guidance control is 99.459km, and the deviation of the reentry height is 541 m; the reentrant angle was-1.909 ° with a deviation of-0.009 °; the longitude of the landing site is 102.005 °, deviation from the given longitude is-0.181 °, and deviation in longitude direction distance is about 20.126 km; the latitude of the landing site was 40.877 °, with a deviation from the given latitude of-0.028 °, and a deviation in latitudinal distance of about 3.155 km.

2) And (3) inner-layer iterative computation of guidance control is performed for 2 times, the reentry height convergence precision is less than 100m, and the reentry angle convergence precision is less than 0.001 degrees. Subsequently, when the outer layer iteration calculation of the braking control is carried out, the reentry height and the reentry angle meet the requirement of convergence accuracy without iteration again.

3) The outer layer of the brake control is iterated and calculated for 3 times, and the convergence accuracy of the longitude and the latitude of the landing point is less than 0.01 degrees.

4) After double-layer iterative calculation of guidance control and brake control, the reentry height of the aiming parameters is offset by +528m, the reentry angle is offset by +0.009 degrees, the longitude of the landing point of the aiming parameters is offset by +0.066 degrees, the latitude is offset by +0.017 degrees, the reentry height 100.075km is realized by the final iterative calculation result, the reentry angle is-1.899 degrees, the longitude 101.182 degrees and the latitude 40.904 degrees of the landing point are realized, and the reentry point and the landing point reach given convergence accuracy. Compared with the method without iterative correction calculation, the method eliminates the deviation of about 20km in the longitude direction and about 3km in the latitude direction of the landing point.

Aiming at the problem that the landing deviation of a ballistic reentry spacecraft adopting a guidance method of aiming reentry parameters is large, the embodiment of the application provides a double-layer iterative correction method for correcting the reentry point aiming parameters by inner layer guidance control and correcting the landing point aiming parameters by outer layer brake control. A brake control parameter calculation method based on the route matching of the re-entry point and the landing point is designed, and a Lambert guidance control algorithm is established. During inner-layer iterative guidance control, correcting a reentry height and a reentry angle aimed by brake control according to the state parameter deviation of a reentry point; and during outer-layer iterative braking control, correcting the longitude and latitude of the landing point aimed by the braking control according to the landing point parameter deviation. Simulation results show that the method eliminates the landing deviation of the ballistic reentry spacecraft by adopting the aiming reentry parameter guidance method through double-layer iterative calculation of guidance control and braking control, meets the requirement that the reentry parameters and the landing parameters of the spacecraft meet the given precision, and solves the problem of accurate landing of the ballistic reentry spacecraft.

Fig. 4 is a flowchart of a method for determining an off-orbit parameter of a spacecraft, provided in an embodiment of the present application, including the following steps:

s401: and calculating alternative off-orbit brake parameters according to the orbit parameters before the spacecraft is off-orbit, the aiming reentry parameters after the spacecraft is off-orbit and the aiming landing points, wherein initially, the aiming reentry parameters are determined according to the given reentry parameters, and the aiming landing points are determined according to the given landing points.

In general, the given reentry parameters may include a given reentry height and a given reentry angle, and correspondingly, the aiming reentry parameters may include an aiming reentry height and an aiming reentry angle.

In specific implementation, the given reentry height may be directly used as the aiming reentry height, or the given reentry height may be offset according to a difference between the given reentry height and a preset height value and then used as the aiming reentry height, and the given reentry angle may be directly used as the aiming reentry angle, or the given reentry angle may be offset according to a difference between the given reentry angle and a preset reentry angle and then used as the aiming reentry angle.

The given landing site is generally a pair of longitude and latitude coordinates, and the given landing site can be directly used as a aiming landing site in specific implementation.

In addition, it should be noted that the process of calculating the candidate off-rail brake parameter is the process of calculating the initial value of the brake control parameter first and then iteratively calculating the brake control parameter, and the brake control parameter after iterative calculation is the candidate off-rail brake parameter.

S402: and determining actual reentry parameters after controlling the spacecraft to derail according to the alternative derailment braking parameters and a preset guidance control algorithm adopted after derailment.

During specific implementation, expected reentry parameters after the spacecraft is controlled to be out of orbit according to the alternative derailment brake parameters can be calculated, the expected reentry parameters serve as derailment brake control targets, the motion parameters after the spacecraft is out of orbit are gradually adjusted by adopting a preset guidance control algorithm, and the actual reentry parameters after the spacecraft is out of orbit are determined based on the adjusted motion parameters of the spacecraft, wherein the preset guidance control algorithm is such as a binary Lambert guidance algorithm.

The process can be implemented with reference to the above process of brake control of the spacecraft using a binary lambert guidance algorithm. Since the binary lambert guidance algorithm is a more ideal guidance algorithm, even if the desired reentry parameters are given, there will be reentry deviations after guidance.

S403: and calculating the reentry deviation of the spacecraft according to the actual reentry parameter and the given reentry parameter.

In a specific implementation, the difference between the given reentrant height and the actual reentrant height is calculated to obtain the reentrant height deviation, and the difference between the given reentrant angle and the actual reentrant angle is calculated to obtain the reentrant angle deviation. I.e. the re-entry deviation comprises a re-entry height deviation and a re-entry angle deviation.

S404: judging whether the reentry deviation of the spacecraft meets the reentry precision requirement, if not, entering S405; if yes, the process proceeds to S406.

Wherein, the reentry deviation meeting the reentry precision requirement means that: the reentry height deviation satisfies the reentry height accuracy requirement, and the reentry angle deviation satisfies the reentry angle accuracy requirement.

S405: and correcting the aiming reentry parameter to obtain a new aiming reentry parameter, and returning to S401.

For example, the sum of the aiming re-entry height and the re-entry height deviation is used as a new aiming re-entry height, and the sum of the aiming re-entry angle and the re-entry angle deviation is used as a new aiming re-entry angle, thereby obtaining a new aiming re-entry parameter.

S406: and calculating the actual landing point of the spacecraft according to the actual reentry parameter.

During specific implementation, the actual reentry point of the spacecraft can be calculated according to the actual reentry parameter, and then the actual landing point of the spacecraft is calculated according to the actual reentry point and the preset position conversion relation between the reentry point and the landing point.

S407: and determining the landing deviation of the spacecraft according to the actual landing point and the given landing point.

In specific implementation, the longitude difference between the given landing point and the actual landing point is calculated to obtain the longitude deviation during landing, and the latitude difference between the given landing point and the actual landing point is calculated to obtain the latitude deviation during landing. I.e. landing deviations include longitude deviations and latitude deviations upon landing.

S408: judging whether the landing deviation of the spacecraft meets the landing precision requirement, if not, entering S409; if yes, the process proceeds to S410.

Wherein, the landing deviation meeting the landing precision requirement means that: the longitude deviation upon landing satisfies the longitude accuracy requirement, and the latitude deviation upon landing satisfies the latitude accuracy requirement.

S409: and correcting the aiming landing point to obtain a new aiming landing point, and returning to S401.

For example, the sum of the longitude and longitude deviations of the aimed landing site is used as the longitude of the new aimed landing site, and the sum of the latitude and latitude deviations of the aimed landing site is used as the latitude of the new aimed landing site, so as to obtain the new aimed landing site.

S410: and taking the current alternative derailment braking parameter as an actual derailment braking parameter.

Subsequently, the actual off-orbit control parameters are adopted to control the off-orbit of the spacecraft, so that the reentry deviation and the landing deviation of the spacecraft after the off-orbit can meet the requirements.

When the method provided in the embodiments of the present application is implemented in software or hardware or a combination of software and hardware, a plurality of functional modules may be included in the electronic device, and each functional module may include software, hardware or a combination of software and hardware.

Fig. 5 is a schematic structural diagram of a device for determining an off-orbit parameter of a spacecraft, provided in an embodiment of the present application, and includes an off-orbit parameter determining module 501, a reentry parameter determining module 502, a reentry precision control module 503, and a landing precision control module 504.

An off-orbit parameter determination module 501, configured to calculate a candidate off-orbit brake parameter according to an orbit parameter before the spacecraft is off-orbit, an aiming reentry parameter after the spacecraft is off-orbit, and an aiming landing point, where initially, the aiming reentry parameter is determined according to a given reentry parameter, and the aiming landing point is determined according to a given landing point;

a reentry parameter determination module 502, configured to determine an actual reentry parameter for controlling the spacecraft after the derailment according to the candidate derailment braking parameter and a preset guidance control algorithm adopted after the derailment;

a reentry precision control module 503, configured to, if it is determined according to the actual reentry parameter and the given reentry parameter that the reentry deviation of the spacecraft does not meet the reentry precision requirement, correct the aiming reentry parameter, and perform a step of calculating an alternative off-orbit braking parameter until it is determined that the reentry deviation meets the reentry precision requirement, and calculate an actual landing point of the spacecraft according to the actual reentry parameter;

and a landing accuracy control module 504, configured to, if it is determined that the landing deviation of the spacecraft does not meet the landing accuracy requirement according to the actual landing point and the given landing point, correct the aimed landing point, and perform a step of calculating an alternative off-orbit braking parameter until it is determined that the landing deviation meets the landing accuracy requirement, and take the current alternative off-orbit braking parameter as the actual off-orbit braking parameter.

In some possible embodiments, the reentry parameter determination module 502 is specifically configured to:

In some possible embodiments, the reentry precision control module 502 is specifically configured to:

the reentry precision control module 503 is specifically configured to:

The division of the modules in the embodiments of the present application is schematic, and only one logic function division is provided, and in actual implementation, there may be another division manner, and in addition, each function module in each embodiment of the present application may be integrated in one processor, may also exist alone physically, or may also be integrated in one module by two or more modules. The coupling of the various modules to each other may be through interfaces that are typically electrical communication interfaces, but mechanical or other forms of interfaces are not excluded. Thus, modules described as separate components may or may not be physically separate, may be located in one place, or may be distributed in different locations on the same or different devices. The integrated module can be realized in a hardware mode, and can also be realized in a software functional module mode.

Fig. 6 is a schematic structural diagram of an electronic device according to an embodiment of the present disclosure, where the electronic device includes a transceiver 601 and a processor 602, and the processor 602 may be a Central Processing Unit (CPU), a microprocessor, an application specific integrated circuit, a programmable logic circuit, a large scale integrated circuit, or a digital Processing Unit. The transceiver 601 is used for data transmission and reception between the electronic device and other devices.

The electronic device may further comprise a memory 603 for storing software instructions executed by the processor 602, but may also store some other data required by the electronic device, such as identification information of the electronic device, encryption information of the electronic device, user data, etc. The Memory 603 may be a Volatile Memory (Volatile Memory), such as a Random-Access Memory (RAM); the Memory 603 may also be a Non-Volatile Memory (Non-Volatile Memory) such as a Read-Only Memory (ROM), a Flash Memory (Flash Memory), a Hard Disk (Hard Disk Drive, HDD) or a Solid-State Drive (SSD), or any other medium that can be used to carry or store desired program code in the form of instructions or data structures and that can be accessed by a computer, but is not limited thereto. The memory 603 may be a combination of the above memories.

The specific connection medium between the processor 602, the memory 603 and the transceiver 601 is not limited in the embodiment of the present application. In fig. 6, the embodiment of the present application is described by taking only the case where the memory 603, the processor 602, and the transceiver 601 are connected by the bus 604 as an example, the bus is shown by a thick line in fig. 6, and the connection manner between other components is merely illustrative and not limited. The bus may be divided into an address bus, a data bus, a control bus, etc. For ease of illustration, only one thick line is shown in FIG. 6, but this is not intended to represent only one bus or type of bus.

The processor 602 may be dedicated hardware or a processor running software, and when the processor 602 may run software, the processor 602 reads software instructions stored in the memory 603 and executes the off-orbit parameter determination method of the spacecraft involved in the foregoing embodiments under the driving of the software instructions.

The embodiment of the present application further provides a storage medium, and when instructions in the storage medium are executed by a processor of an electronic device, the electronic device is capable of executing the off-orbit parameter determination method for a spacecraft, which is referred to in the foregoing embodiments.

In some possible embodiments, the aspects of the method for determining an off-orbit parameter of a spacecraft provided by the present application may also be implemented in the form of a program product, which includes program code for causing an electronic device to perform the method for determining an off-orbit parameter of a spacecraft mentioned in the foregoing embodiments, when the program product is run on the electronic device.

The program product may employ any combination of one or more readable media. The readable medium may be a readable signal medium or a readable storage medium. A readable storage medium may be, for example, but not limited to, an electronic, magnetic, optical, electromagnetic, infrared, or semiconductor system, apparatus, or device, or any combination of the foregoing. More specific examples (a non-exhaustive list) of the readable storage medium include: an electrical connection having one or more wires, a portable Disk, a hard Disk, a RAM, a ROM, an Erasable Programmable Read-Only Memory (EPROM), a flash Memory, an optical fiber, a Compact Disk Read-Only Memory (CD-ROM), an optical storage device, a magnetic storage device, or any suitable combination of the foregoing.

The program product for determination of off-track parameters for a spacecraft in an embodiment of the present application may be a CD-ROM and comprise program code and may be run on a computing device. However, the program product of the present application is not limited thereto, and in this document, a readable storage medium may be any tangible medium that can contain, or store a program for use by or in connection with an instruction execution system, apparatus, or device.

A readable signal medium may include a propagated data signal with readable program code embodied therein, for example, in baseband or as part of a carrier wave. Such a propagated data signal may take many forms, including, but not limited to, electro-magnetic, optical, or any suitable combination thereof. A readable signal medium may also be any readable medium that is not a readable storage medium and that can communicate, propagate, or transport a program for use by or in connection with an instruction execution system, apparatus, or device.

Program code embodied on a readable medium may be transmitted using any appropriate medium, including but not limited to wireless, wireline, optical fiber cable, Radio Frequency (RF), etc., or any suitable combination of the foregoing.

Program code for carrying out operations of the present application may be written in any combination of one or more programming languages, including an object oriented programming language such as Java, C + + or the like and conventional procedural programming languages, such as the "C" programming language or similar programming languages. The program code may execute entirely on the user's computing device, partly on the user's device, as a stand-alone software package, partly on the user's computing device and partly on a remote computing device, or entirely on the remote computing device or server. In situations involving remote computing devices, the remote computing devices may be connected to the user computing device over any kind of Network, such as a Local Area Network (LAN) or Wide Area Network (WAN), or may be connected to external computing devices (e.g., over the internet using an internet service provider).

It should be noted that although several units or sub-units of the apparatus are mentioned in the above detailed description, such division is merely exemplary and not mandatory. Indeed, the features and functions of two or more units described above may be embodied in one unit, according to embodiments of the application. Conversely, the features and functions of one unit described above may be further divided into embodiments by a plurality of units.

Further, while the operations of the methods of the present application are depicted in the drawings in a particular order, this does not require or imply that these operations must be performed in this particular order, or that all of the illustrated operations must be performed, to achieve desirable results. Additionally or alternatively, certain steps may be omitted, multiple steps combined into one step execution, and/or one step broken down into multiple step executions.

As will be appreciated by one skilled in the art, embodiments of the present application may be provided as a method, system, or computer program product. Accordingly, the present application may take the form of an entirely hardware embodiment, an entirely software embodiment or an embodiment combining software and hardware aspects. Furthermore, the present application may take the form of a computer program product embodied on one or more computer-usable storage media (including, but not limited to, disk storage, CD-ROM, optical storage, and the like) having computer-usable program code embodied therein.

The present application is described with reference to flowchart illustrations and/or block diagrams of methods, apparatus (systems), and computer program products according to embodiments of the application. It will be understood that each flow and/or block of the flow diagrams and/or block diagrams, and combinations of flows and/or blocks in the flow diagrams and/or block diagrams, can be implemented by computer program instructions. These computer program instructions may be provided to a processor of a general purpose computer, special purpose computer, embedded processor, or other programmable data processing apparatus to produce a machine, such that the instructions, which execute via the processor of the computer or other programmable data processing apparatus, create means for implementing the functions specified in the flowchart flow or flows and/or block diagram block or blocks.

These computer program instructions may also be stored in a computer-readable memory that can direct a computer or other programmable data processing apparatus to function in a particular manner, such that the instructions stored in the computer-readable memory produce an article of manufacture including instruction means which implement the function specified in the flowchart flow or flows and/or block diagram block or blocks.

These computer program instructions may also be loaded onto a computer or other programmable data processing apparatus to cause a series of operational steps to be performed on the computer or other programmable apparatus to produce a computer implemented process such that the instructions which execute on the computer or other programmable apparatus provide steps for implementing the functions specified in the flowchart flow or flows and/or block diagram block or blocks.

While the preferred embodiments of the present application have been described, additional variations and modifications in those embodiments may occur to those skilled in the art once they learn of the basic inventive concepts. Therefore, it is intended that the appended claims be interpreted as including preferred embodiments and all alterations and modifications as fall within the scope of the application.

It will be apparent to those skilled in the art that various changes and modifications may be made in the present application without departing from the spirit and scope of the application. Thus, if such modifications and variations of the present application fall within the scope of the claims of the present application and their equivalents, the present application is intended to include such modifications and variations as well.

Claims

1. A method for determining an off-orbit parameter of a spacecraft is characterized by comprising the following steps:

2. The method of claim 1, wherein determining the actual reentry parameter for controlling the spacecraft after the derailment based on the candidate derailment braking parameter and a predetermined guidance control algorithm employed after the derailment comprises:

3. The method of claim 2, wherein calculating an actual landing site for the spacecraft from the actual reentry parameter comprises:

4. The method of claim 1, wherein the aiming re-entry parameters include an aiming re-entry height and an aiming re-entry angle, and the re-entry bias includes a re-entry height bias and a re-entry angle bias;

modifying the aiming reentry parameter, comprising:

5. An off-orbit parameter determination apparatus for a spacecraft, comprising:

6. The apparatus of claim 5, wherein the reentry parameter determination module is specifically configured to:

7. The apparatus of claim 6, wherein the reentry precision control module is specifically configured to:

8. The apparatus of claim 5, wherein the aiming re-entry parameters include an aiming re-entry height and an aiming re-entry angle, the re-entry bias including a re-entry height bias and a re-entry angle bias;

the reentry precision control module is specifically configured to:

9. An electronic device, comprising: at least one processor, and a memory communicatively coupled to the at least one processor, wherein:

the memory stores instructions executable by the at least one processor to enable the at least one processor to perform the method of any one of claims 1-4.

10. A storage medium, wherein instructions in the storage medium, when executed by a processor of an electronic device, enable the electronic device to perform the method of any of claims 1-4.