Mars capture braking strategy calculation method and system

Abstract

The present invention provides a Mars capture braking strategy calculation method and system, which relates to the field of spacecraft overall design technology. The method comprises: step S1: collecting relevant data of the hyperbolic asymptotic vector, the earth direction and the sun direction of entering Mars, and establishing an auxiliary coordinate system; step S2: calculating the semi-major axis of the capture orbit through the established auxiliary coordinate system; step S3: calculating the eccentricity of the capture orbit through the established auxiliary coordinate system; step S4: calculating the capture orbit inclination, the right ascension of the ascending node and the argument of the perigee through the established auxiliary coordinate system; step S5: calculating the angles related to the earth and the sun through the established auxiliary coordinate system, and comparing and analyzing the conditions including energy and communication in the capture stage according to the calculated relevant data. The present invention can improve the calculation speed of the capture orbit parameters and the design efficiency of the Mars capture strategy.

Description

Technical Field

[0001] The present invention relates to the technical field of overall spacecraft design, and in particular to a method for quickly calculating a Mars capture braking strategy, and in particular to a method and system for calculating a Mars capture braking strategy. Background Art

[0002] Mars capture braking refers to the process in which the probe forms a Mars orbit after decelerating from a hyperbolic orbit through the perigee. Due to the large velocity increment and long deceleration time of capture braking, the capture effect directly determines the success or failure of the entire mission. The design of the near-Mars capture strategy needs to consider factors such as the orbit determination accuracy of the near-Mars segment, the visible arc segment of the tracking and control station, the duration of the Mars star occultation for ground communication, the shadow time of Mars, the altitude of the Martian atmosphere, the orbit control error, the selection of the capture inclination, and the subsequent relay communication mission and remote sensing mission requirements of the Mars orbit. In the capture strategy design and comparison stage, it is necessary to quickly calculate the capture orbit parameters. In view of the problem that the numerical integration method takes a lot of time in state setting and numerical integration calculation, a calculation method needs to be proposed to improve the calculation speed of the capture orbit parameters and the design efficiency of the Mars capture strategy.

[0003] The invention patent with publication number CN102168981B discloses an autonomous astronomical navigation method for the Mars capture phase of a deep space probe. The state model of the deep space probe is established based on the circular restricted four-body orbital dynamics model; the pixel image line information of Mars, Mars satellites and background stars is obtained by using sensors, and the obtained pixel image line information is converted into the angle information of Mars, Mars satellites and background stars to establish the angle measurement model of Mars, Phobos and Deimos; and the q-method method is used to estimate the probe attitude information, and the Unscented Kalman filter is combined to estimate the position and speed of the deep space probe. The invention mainly focuses on establishing a state model based on the orbital dynamics model, using the angle measurement information of Mars, Phobos and Deimos, and performing autonomous navigation for Mars capture.

[0004] The invention patent with the publication number CN105509750B discloses a Mars capture segment navigation method combining astronomical velocity measurement and ground radio, including the following steps: the ground station obtains the distance between the probe and the ground through radio ranging; the ground station obtains the line-of-sight velocity between the probe and the ground station through radio Doppler velocity measurement; the probe obtains the line-of-sight velocity between the probe and a certain star through an autonomous velocity measurement navigation sensor; and the position and velocity estimation of the earth-sky combined navigation is obtained through the extended Kalman filter. This method mainly obtains the position and velocity of the earth-sky combined navigation based on the radio ranging and Doppler velocity measurement of the ground station, combined with the line-of-sight velocity between the probe and a certain star measured by the autonomous velocity measurement navigation sensor.

[0005] The invention patent with publication number CN103412569B discloses a method for autonomous management of the Mars capture process, including: a management sequence before the Mars capture of the Mars probe, a management sequence during the Mars capture of the Mars probe, and a management sequence after the Mars capture of the Mars probe. The three management sequences are used to manage the attitude and orbit control subsystem, the power supply subsystem, the integrated electronic subsystem, and the measurement and control data transmission subsystem. Through reasonable planning of probe management events before, during and after capture, and through multidisciplinary optimization design of probe energy, communication, attitude, lighting, etc., a complete autonomous management sequence of the capture process is established. This method mainly realizes autonomous management of the probe capture process by reasonably planning probe management events before, during and after capture, and through multidisciplinary optimization design of probe energy, communication, attitude, lighting, etc.

[0006] The invention patent with publication number CN104252546A discloses a method for optimizing the parameters of Mars capture brake. First, the dynamics of Mars capture brake are modeled and simplified, and then the optimization parameters are determined by the objective function of fuel optimization. The fuel optimization is converted into the shortest ignition duration and used as the optimization parameter for optimization. Then, given the search variable and its range, the particle swarm algorithm is used for search to obtain a capture strategy with a relatively small fitness value. Finally, if a solution that satisfies the constraints can be found within a given time range, the time range setting value can be reduced and the search can be repeated until the boundary of the time range limit value is found.

[0007] The invention patent with the publication number CN111414002A discloses a Mars probe secondary braking capture control method based on Newton iteration. After the Mars probe fails during the capture braking, the on-board computer is reset. After that, the Mars probe is subjected to secondary braking by using the capture braking optimization method based on Newton iteration to ensure that the probe can smoothly enter the target orbit during the Mars capture braking process. The secondary capture braking strategy can be independently designed for the failure during the capture braking. In addition, a capture braking safety interval is set, including the shortest ignition duration and the longest ignition duration for braking capture. When the on-board time is within the shortest ignition duration and the longest ignition duration, the speed increment reaches the nominal speed increment, or the on-board time reaches the maximum ignition duration, the engine shutdown operation is performed, so that the Mars probe can form a Mars orbit during the braking capture and will not collide with Mars due to the long ignition time. Summary of the invention

[0008] In view of the defects in the prior art, the present invention provides a Mars capture braking strategy calculation method and system.

[0009] According to a Mars capture braking strategy calculation method and system provided by the present invention, the scheme is as follows:

[0010] In a first aspect, a Mars capture braking strategy calculation method is provided, the method comprising:

[0011] Step S1: Collect relevant data of the hyperbolic asymptotic vector entering Mars, the Earth's direction and the Sun's direction, and establish an auxiliary coordinate system;

[0012] Step S2: calculating the semi-major axis of the capture orbit through the established auxiliary coordinate system;

[0013] Step S3: calculating the capture orbit eccentricity through the established auxiliary coordinate system;

[0014] Step S4: calculating the capture orbit inclination, the right ascension of the ascending node and the argument of perigee through the established auxiliary coordinate system;

[0015] Step S5: By establishing the auxiliary coordinate system, according to the orbital parameters obtained in steps S2, S3 and S4, the angles related to the earth and the sun are calculated, and based on the relevant data obtained by calculation, the conditions including energy and communication in the capture phase are compared and analyzed, so as to select the capture orbit corresponding to the optimal energy and communication conditions.

[0016] Preferably, the step S1 specifically includes:

[0017] Hyperbolic asymptotic vectors for known entry into Mars The Earth points to The sun points to Near fire point

[0018] First establish an auxiliary coordinate system P-OX j Y j Z j , with its origin being the center of Mars, z j Axis The unit vector, X j The axis is located In-plane, Y j Axis and X j , z j The axis satisfies the right-hand rule;

[0019] Then the axes of the coordinate system are calculated as follows:

[0020] Then the transfer matrix between the auxiliary coordinate system and the inertial coordinate system is:

[0021] Preferably, step S2 comprises:

[0022] Known hyperbolic asymptotic vector Expressed as (V inf,δ,β), where (δ,β) is the azimuth and elevation angle of the asymptotic unit vector in the inertial coordinate system, V inf represents the hyperbolic residual velocity;

[0023] The same asymptotic vector can capture different inclinations, and the asymptotic vector is located in the capture orbit plane;

[0024] The semi-major axis of the capture orbit a = -U m / C3, where U m is the Martian gravitational constant, which is 42828.28596; C3 represents the escape energy.

[0025] Preferably, the step S3 of calculating the capture orbit eccentricity specifically includes:

[0026] Assume that the capture perigee height is H P , then r P =H P +r m , where r m is the average radius of Mars, which is 3397 km;

[0027] H P Indicates the height of the captured near-fire point;

[0028] r P Indicates the distance between the captured perigee and the center of Mars;

[0029] Capture orbit eccentricity e = 1 + r p C3 / U m .

[0030] Preferably, the step S4 specifically includes:

[0031] Assume that the capture orbit plane normal is And there is definition In the auxiliary coordinate system, j The angle between the axes is u, for The azimuth and elevation angles in the inertial coordinate system, Serving size:

[0032]

[0033] Capture orbit inclination

[0034] The right ascension of the ascending node of the capture orbit is Ω: Ω=π / 2+α;

[0035] Find the nearest fire point vector

[0036] Ascending node right ascension vector:

[0037] get:

[0038]

[0039] in,

[0040] Thus, the near-fire angle is obtained

[0041] Preferably, the step S5 specifically includes:

[0042] Calculate the near fire vector Pointing to the Earth Angle:

[0043] Calculate the near fire vector Pointing with the sun Angle:

[0044] According to the above calculation process, by changing different u values, different There are different orbital planes and numbers of captured orbital elements;

[0045] The selection of reasonable capture orbit parameters needs to be determined based on relevant factors including the velocity increment, ∠POS and ∠POE required to transform the captured orbit parameters to the target orbit parameters.

[0046] In a second aspect, a Mars capture braking strategy calculation system is provided, the system comprising:

[0047] Module M1: Establish an auxiliary coordinate system based on the known hyperbolic approach vector into Mars, the Earth's orientation, and the Sun's orientation;

[0048] Module M2: Calculate the semi-major axis of the capture orbit according to the auxiliary coordinate system;

[0049] Module M3: Calculating the capture orbit eccentricity according to the auxiliary coordinate system;

[0050] Module M4: Calculate the capture orbit inclination, right ascension of ascending node and argument of perigee according to the auxiliary coordinate system;

[0051] Module M5: By establishing the auxiliary coordinate system, according to the orbital parameters obtained by modules M2, M3 and M4, the angles related to the earth and the sun are calculated. According to the relevant data obtained by calculation, the conditions including energy and communication in the capture phase are compared and analyzed, so as to select the capture orbit corresponding to the optimal energy and communication conditions.

[0052] Preferably, the module M1 specifically includes:

[0053] Hyperbolic asymptotic vectors for known entry into Mars The Earth points to The sun points to Near fire point

[0054] First establish an auxiliary coordinate system P-OX j Y j Z j , with its origin being the center of Mars, z j Axis The unit vector, X j The axis is located In-plane, Y j Axis and X j , z j The axis satisfies the right-hand rule;

[0055] Then the axes of the coordinate system are calculated as follows:

[0056] Then the transfer matrix between the auxiliary coordinate system and the inertial coordinate system is:

[0057] Preferably, the module M2 comprises:

[0058] Known hyperbolic asymptotic vector Expressed as (V inf ,δ,β), where (δ,β) is the azimuth and elevation angle of the asymptotic unit vector in the inertial coordinate system, V inf represents the hyperbolic residual velocity;

[0059] The same asymptotic vector can capture different inclinations, and the asymptotic vector is located in the capture orbit plane;

[0060] The semi-major axis of the capture orbit a = -U m / C3, where U m is the Martian gravitational constant, which is 42828.28596; C3 represents the escape energy.

[0061] Preferably, the module M3 calculates the capture orbit eccentricity specifically including:

[0062] Assume that the capture perigee height is H P , then r P =H P +r m , where r m is the average radius of Mars, which is 3397 km;

[0063] H PIndicates the height of the captured near-fire point;

[0064] r P Indicates the distance between the captured perigee and the center of Mars;

[0065] Capture orbit eccentricity e = 1 + r p C3 / U m .

[0066] Compared with the prior art, the present invention has the following beneficial effects:

[0067] 1. The present invention focuses on the rapid calculation of Mars capture braking strategy, which can enter the hyperbolic asymptotic vector The direction and size of the spacecraft are calculated, and the capture orbit parameters of different inclinations, as well as the angles between the perigee vector and the Earth and the Sun, are used to achieve comparative analysis of energy, communication and other conditions during the capture phase.

[0068] 2. The present invention solves the problem that the numerical integration method consumes a lot of time in state setting and numerical integration calculation;

[0069] 3. The application of the present invention has a good effect on the design of Mars capture braking strategy and the improvement of orbit calculation and analysis efficiency, and significantly improves the calculation speed of capture orbit parameters and the design efficiency of Mars capture strategy. BRIEF DESCRIPTION OF THE DRAWINGS

[0070] Other features, objects and advantages of the present invention will become more apparent from the detailed description of non-limiting embodiments made with reference to the following drawings:

[0071] Figure 1 The auxiliary coordinate system P-OX j Y j Z j .

[0072] Figure 2 Schematic diagram of the capture orbit. DETAILED DESCRIPTION

[0073] The present invention is described in detail below in conjunction with specific embodiments. The following embodiments will help those skilled in the art to further understand the present invention, but are not intended to limit the present invention in any form. It should be noted that, for those of ordinary skill in the art, several changes and improvements can also be made without departing from the concept of the present invention. These all belong to the protection scope of the present invention.

[0074] The embodiment of the present invention provides a Mars capture braking strategy calculation method, referring to Figure 1 As shown, the specific steps of this method are as follows:

[0075] Step S1: Collect relevant data on the hyperbolic asymptotic vector entering Mars, the Earth’s orientation, and the Sun’s orientation, and establish an auxiliary coordinate system.

[0076] Hyperbolic asymptotic vectors for known entry into Mars The Earth pointing (unit vector from Mars to Earth) is The sun pointing (unit vector of Mars pointing to the sun) is Near fire point

[0077] First establish an auxiliary coordinate system P-OX j Y j Z j , with its origin being the center of Mars, z j Axis The unit vector, X j The axis is located In-plane, Y j Axis and X j , z j The axis satisfies the right-hand rule;

[0078] Then the axes of the coordinate system are calculated as follows:

[0079] Then the transfer matrix between the auxiliary coordinate system and the inertial coordinate system is:

[0080] Step S2: Calculate the semi-major axis of the capture orbit:

[0081] Known hyperbolic asymptotic vector It can be expressed as (V inf ,δ,β), where (δ,β) is the azimuth and elevation angle of the asymptotic unit vector in the inertial coordinate system, V inf Represents the hyperbolic residual velocity; the same asymptotic vector can be captured to obtain different inclinations, but the asymptotic vector should lie in the capture orbit plane.

[0082] The semi-major axis of the capture orbit a = -U m / C3, where U m is the Martian gravitational constant, which is 42828.28596; C3 represents the escape energy.

[0083] Step S3: Calculate the capture orbit eccentricity:

[0084] Assume that the capture perigee height is H P , then r P =H P +r m , where r m is the average radius of Mars, which is 3397 km;

[0085] H P Indicates the height of the captured near-fire point;

[0086] r P Indicates the distance between the captured perigee and the center of Mars;

[0087] Capture orbit eccentricity e = 1 + r p C3 / U m .

[0088] Step S4: Calculate the capture orbit inclination, right ascension of the ascending node, and the argument of perigee:

[0089] Assume that the capture orbit plane normal is And there is definition In the auxiliary coordinate system, j The angle between the axes is u, for The azimuth and elevation angles in the inertial coordinate system, Serving size:

[0090]

[0091] Capture orbit inclination

[0092] The right ascension of the ascending node of the capture orbit is Ω: Ω=π / 2+α;

[0093] Find the nearest fire point vector

[0094] Ascending node right ascension vector:

[0095] get:

[0096]

[0097] in,

[0098] Thus, the near-fire angle is obtained

[0099] Step S5: Refer to Figure 1 and Figure 2 As shown, calculate the angles related to the earth and the sun:

[0100] Calculate the near fire vector Pointing with the Earth Angle:

[0101] Calculate the near fire vector Pointing with the sun Angle:

[0102] According to the above calculation process, different values ​​of u can be obtained by changing them. There are different orbital planes and numbers of captured orbital elements.

[0103] The selection of reasonable capture orbit parameters needs to be determined based on relevant factors including the velocity increment, ∠POS and ∠POE required to transform the captured orbit parameters to the target orbit parameters.

[0104] The embodiment of the present invention provides a Mars capture braking strategy calculation method and system, which can enter the hyperbolic asymptotic vector The direction and size of the capture orbit parameters at different inclinations, as well as the angles between the perigee vector and the Earth and the Sun, are calculated to achieve comparative analysis of energy, communication and other conditions during the capture phase. The problem of the numerical integration method being time-consuming in state setting and numerical integration calculation is solved. The application of the present invention has a good effect on the design of Mars capture braking strategy and improving the efficiency of orbit calculation and analysis.

[0105] Those skilled in the art know that, in addition to realizing the system and its various devices, modules, and units provided by the present invention in a purely computer-readable program code, it is entirely possible to realize the same functions in the form of logic gates, switches, application-specific integrated circuits, programmable logic controllers, and embedded microcontrollers by logically programming the method steps. Therefore, the system and its various devices, modules, and units provided by the present invention can be considered as a hardware component, and the devices, modules, and units included therein for realizing various functions can also be regarded as structures within the hardware component; the devices, modules, and units for realizing various functions can also be regarded as both software modules for realizing the method and structures within the hardware component.

[0106] The above describes the specific embodiments of the present invention. It should be understood that the present invention is not limited to the above specific embodiments, and those skilled in the art can make various changes or modifications within the scope of the claims, which does not affect the essence of the present invention. In the absence of conflict, the embodiments of the present application and the features in the embodiments can be combined with each other arbitrarily.

Claims

1. A Mars capture braking strategy calculation method, characterized in that: include: Step S1: Collect relevant data of the hyperbolic asymptotic vector entering Mars, the Earth's direction and the Sun's direction, and establish an auxiliary coordinate system; Step S2: calculating the semi-major axis of the capture orbit through the established auxiliary coordinate system; Step S3: calculating the capture orbit eccentricity through the established auxiliary coordinate system; Step S4: calculating the capture orbit inclination, the right ascension of the ascending node and the argument of perigee through the established auxiliary coordinate system; Step S5: By establishing the auxiliary coordinate system, according to the orbital parameters obtained in steps S2, S3 and S4, the angles relative to the earth and the sun are calculated, and the conditions including energy and communication in the capture phase are compared and analyzed, so as to select the capture orbit corresponding to the optimal energy and communication conditions.

2. The Mars capture braking strategy calculation method according to claim 1, characterized in that: The step S1 specifically includes: Hyperbolic asymptotic vectors for known entry into Mars The Earth points to The sun points to Near fire point First establish an auxiliary coordinate system P-OX j Y j Z j , with its origin at the center of Mars, Z j Axis The unit vector, X j The axis is located In-plane, Y j Axis and X j , Z j The axis satisfies the right-hand rule; Then the axes of the coordinate system are calculated as follows: Then the transfer matrix between the auxiliary coordinate system and the inertial coordinate system is:

3. The Mars capture braking strategy calculation method according to claim 1, characterized in that: The step S2 comprises: Known hyperbolic asymptotic vector Expressed as (V inf ,δ,β), where (δ,β) is the azimuth and elevation angle of the asymptotic unit vector in the inertial coordinate system, V inf represents the hyperbolic residual velocity; The same asymptotic vector can capture different inclinations, and the asymptotic vector is located in the capture orbit plane; The semi-major axis of the capture orbit a = -U m / C3, where U m is the Martian gravitational constant, which is 42828.28596; C3 represents the escape energy.

4. The Mars capture braking strategy calculation method according to claim 1, characterized in that: The step S3 of calculating the capture orbit eccentricity specifically includes: Assume that the capture height of the perigee is H P , then r P =H P +r m , where r m is the average radius of Mars, which is 3397 km; H P Indicates the height of the captured near-fire point; r P Indicates the distance between the captured perigee and the center of Mars; Capture orbit eccentricity e = 1 + r p C3 / U m .

5. The Mars capture braking strategy calculation method according to claim 1, characterized in that: The step S4 specifically includes: Assume that the capture orbit plane normal is And there is definition In the auxiliary coordinate system, j The angle between the axes is u, for The azimuth and elevation angles in the inertial coordinate system, Serving size: Capture orbit inclination i: The right ascension of the ascending node of the capture orbit is Ω: Ω=π / 2+α; Find the nearest fire point vector Ascending node right ascension vector: get: in, Thus, the near-fire angle is obtained 6. The Mars capture braking strategy calculation method according to claim 1, characterized in that: The step S5 specifically includes: Calculate the near fire vector Pointing with the Earth Angle: Calculate the near fire vector Pointing with the sun Angle: According to the above calculation process, by changing different u values, different There are different orbital planes and numbers of captured orbital elements; The selection of reasonable capture orbit parameters needs to be determined based on relevant factors including the velocity increment, ∠POS and ∠POE required to transform the captured orbit parameters to the target orbit parameters.

7. A Mars capture braking strategy calculation system, characterized in that: include: Module M1: Establish an auxiliary coordinate system based on the known hyperbolic approach vector into Mars, the Earth's orientation, and the Sun's orientation; Module M2: Calculate the semi-major axis of the capture orbit according to the auxiliary coordinate system; Module M3: Calculating the capture orbit eccentricity according to the auxiliary coordinate system; Module M4: Calculate the capture orbit inclination, right ascension of ascending node and argument of perigee according to the auxiliary coordinate system; Module M5: By establishing the auxiliary coordinate system, according to the orbital parameters obtained by modules M2, M3 and M4, the angles related to the earth and the sun are calculated, and the conditions including energy and communication in the capture phase are compared and analyzed, so as to select the capture orbit corresponding to the optimal energy and communication conditions.

8. The Mars capture braking strategy calculation system according to claim 7, characterized in that: The module M1 specifically includes: Hyperbolic asymptotic vectors for known entry into Mars The Earth points to The sun points to Near fire point First establish an auxiliary coordinate system P-OX j Y j Z j , with its origin being the center of Mars, z j Axis The unit vector, X j The axis is located In-plane, Y j Axis and X j , z j The axis satisfies the right-hand rule; Then the axes of the coordinate system are calculated as follows: Then the transfer matrix between the auxiliary coordinate system and the inertial coordinate system is:

9. The Mars capture braking strategy calculation system according to claim 7, characterized in that: The module M2 comprises: Known hyperbolic asymptotic vector Expressed as (V inf ,δ,β), where (δ,β) is the azimuth and elevation angle of the asymptotic unit vector in the inertial coordinate system, V inf represents the hyperbolic residual velocity; The same asymptotic vector can capture different inclinations, and the asymptotic vector is located in the capture orbit plane; The semi-major axis of the capture orbit a = -U m / C3, where U m is the Martian gravitational constant, which is 42828.28596; C3 represents the escape energy.

10. The Mars capture braking strategy calculation system according to claim 7, characterized in that: The module M3 calculates the capture orbit eccentricity specifically including: Assume that the capture height of the perigee is H P , then r P =H P +r m , where r m is the average radius of Mars, which is 3397 km; H P Indicates the height of the captured near-fire point; r P Indicates the distance between the captured perigee and the center of Mars; Capture orbit eccentricity e = 1 + r p C3 / U m .

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