Satellite-rocket separation design method of inclined layout satellite

Abstract

The invention relates to a satellite-rocket separation design method for an inclined layout satellite, and belongs to the technical field of satellite separation design. A multi-satellite inclined structure is designed and comprises a control cabin, a satellite support, n satellites and n satellite-rocket separation mechanisms. The satellite-rocket separation mechanism adopts an ejector rod-spring structure as an energy storage mechanism for satellite-rocket separation; a locking device is arranged on the satellite-rocket separation mechanism, after the locking device is unlocked, elastic potential energy stored in a spring is released, the satellite is pushed to be separated from the satellite support through an ejector rod, and satellite-rocket separation is achieved; the friction coefficient of the ejector rod and the corresponding satellite contact surface is optimized; the spring force is optimized; the spring moment is optimized; after the satellite and the rocket are separated, designing an attitude control time sequence and an attitude adjustment angle of the satellite bracket to realize satellite and rocket separation safety control; according to the method, the problems of satellite attitude angular velocity deviation control and satellite-rocket separation safety after large-dip-angle satellite support layout satellite separation are solved, and the satellite-rocket separation safety is guaranteed.

Description

Technical Field

[0001] The present invention belongs to the technical field of satellite separation design, and relates to a method for designing the separation of a satellite and a rocket for a satellite with an inclined layout. Background Art

[0002] With the development of space technology, the demand for satellite launches is strong, and the rocket carrying capacity has been improved, which has promoted the development of satellite layout methods and separation technologies. In view of the limited available envelope of the fairing and the case of using a large-inclination satellite support to layout satellites, problems such as centroid offset, serious imbalance of torque during the separation process, and unclear friction coefficient have emerged, resulting in a large angular velocity of the satellite and the satellite support after separation, leading to problems such as the satellite being unable to enter the orbit and separation interference. Summary of the Invention

[0003] The technical problem solved by the present invention is: overcoming the deficiencies of the prior art, and proposing a method for designing the separation of a satellite and a rocket for a satellite with an inclined layout, solving the problems of controlling the attitude angular velocity deviation of the satellite after separation and the safety of satellite-rocket separation for a satellite layout with a large-inclination satellite support, and ensuring the safety of satellite-rocket separation.

[0004] The solution for the present invention to solve the technology is as follows:

[0005] A method for designing the separation of a satellite and a rocket for a satellite with an inclined layout, including:

[0006] Designing a multi-satellite inclined structure, including a control cabin, a satellite support, n satellites, and n satellite-rocket separation mechanisms; the satellite support is installed on the top of the control cabin; the n satellites are installed on the satellite support; each satellite is connected to the satellite support through 1 satellite-rocket separation mechanism;

[0007] The satellite-rocket separation mechanism uses a push rod-spring structure as the energy storage mechanism for satellite-rocket separation; a locking device is provided on the satellite-rocket separation mechanism. After the locking device is unlocked, the elastic potential energy stored in the spring is released, and the satellite is pushed away from the satellite support through the push rod to achieve satellite-rocket separation;

[0008] Optimizing the friction coefficient of the contact surface between the push rod and the corresponding satellite;

[0009] Optimizing the spring force;

[0010] Optimizing the spring torque;

[0011] After satellite-rocket separation, designing the attitude control timing and attitude adjustment angle of the satellite support to achieve safe control of satellite-rocket separation.

[0012] In the above method for designing the separation of a satellite and a rocket for a satellite with an inclined layout, the bottom plane of the satellite support is docked with the top of the control cabin; the top of the satellite support is a regular n-sided pyramid structure; each satellite is correspondingly installed on 1 side surface of the satellite support; the included angle between each side surface of the satellite support and the axis is θ.

[0013] In the above method for designing the separation of the satellite and the launch vehicle with an inclined layout, the design method for the friction coefficient of the contact surface between the push rod and the corresponding satellite is as follows:

[0014] S31. Fabricate the contact surface of the same material as the satellite at the end of the push rod for experiments;

[0015] S32. According to the data of the friction coefficient obtained from the experiments with respect to the contact pressure and the tangential relative velocity, use the B-spline interpolation method to obtain the functional relationship between the friction coefficient and different contact pressures and tangential relative velocities;

[0016] S33. Establish a dynamic model for the separation of the satellite and the launch vehicle, and substitute the functional relationship between the friction coefficient and different contact pressures and tangential relative velocities into the dynamic model for the separation of the satellite and the launch vehicle; Set different friction coefficients μ in the dynamic model for the separation of the satellite and the launch vehicle d Conduct simulations and select the optimized friction coefficient value.

[0017] In the above method for designing the separation of the satellite and the launch vehicle with an inclined layout, in S31, the specific content of the experiment is as follows:

[0018] Set 30 groups of experiments with the pressure ranging from 100 N to 1600 N and different relative sliding velocities ranging from 10 mm / s to 200 mm / s.

[0019] In the above method for designing the separation of the satellite and the launch vehicle with an inclined layout, in S33, the method for selecting the optimized friction coefficient value is as follows:

[0020] Set different friction coefficients μ in the dynamic model for the separation of the satellite and the launch vehicle d Conduct simulations to obtain the final angular velocity values under different friction coefficients μ d Select the final angular velocity value that is closest to the actual angular velocity, and the corresponding friction coefficient μ d is the optimized friction coefficient value.

[0021] In the above method for designing the separation of the satellite and the launch vehicle with an inclined layout, the spring includes an upper spring and a lower spring; the upper spring is arranged at the top end of the corresponding satellite; the lower spring is arranged at the bottom end of the corresponding satellite.

[0022] In the above method for designing the separation of the satellite and the launch vehicle with an inclined layout, the optimization method for the spring force is as follows:

[0023] Set the acting force of the lower spring as F1, the reaction force of the lower spring as F1', the acting force of the upper spring as F2, and the reaction force of the upper spring as F2'; the optimized spring force satisfies:

[0024] 1550 N ≤ F1 ≤ 1700 N, 1950 N ≤ F2 ≤ 2050 N, F2 > F1.

[0025] In the above satellite-rocket separation design method for an inclined layout satellite, the optimization method for the spring moment is as follows:

[0026] Set the moment of F1 acting on the satellite's center of mass as M1, and the moment of F2 acting on the satellite's center of mass as M2; after optimization, it satisfies:

[0027] M1 = M2, and the lever arm length is 1 - 4 m.

[0028] In the above satellite-rocket separation design method for an inclined layout satellite, set the moment of F1' acting on the center of mass of the satellite support as M1', and the moment of F2' acting on the center of mass of the satellite support as M2'; after optimization, it satisfies:

[0029] The total moment acting on the center of mass of the satellite support is M1' + M2'.

[0030] In the above satellite-rocket separation design method for an inclined layout satellite, the design method for the attitude control timing sequence and attitude adjustment angle of the satellite support is as follows:

[0031] Set the satellite-rocket separation duration as t 1 ; after satellite-rocket separation, simultaneously enter the attitude adjustment of the satellite support and the preparation for orbit departure attitude adjustment; among them, the duration of the attitude adjustment of the satellite support is t 2 , the adjustment angle is X; the duration of the preparation for orbit departure attitude adjustment is t 3 ;

[0032] 30 s ≤ t 1 ≤ 50 s, 3 s ≤ t 2 ≤ 8 s, 150 s ≤ t 3 ≤ 300 s, 30° ≤ X ≤ 50°.

[0033] The beneficial effects of the present invention compared with the prior art are as follows:

[0034] (1) The present invention first proposes a method for arranging satellites on a satellite support with a large inclination angle, solves the problems of separation safety and layout limitations, and maximizes the available envelope of the fairing;

[0035] (2) The present invention proposes an angular velocity control method for an inclined layout satellite, solves the problem of uncertain friction coefficient between the separation spring ejector rod and the satellite contact surface, and improves the inherent reliability of the separation design and the overall performance of the rocket;

[0036] (3) Through the collaborative joint design of attitude control, the present invention solves the problems of collision interference and rotation caused by excessive angular velocity of the last stage, and the problem that the main engine or the final correction jet of the rocket approaches and contaminates the satellite, and improves the satellite-rocket separation safety. Description of the Drawings

[0037] Figure 1Schematic diagram of the multi-satellite inclined structure of the present invention;

[0038] Figure 2 Schematic diagram of the friction coefficient substitution simulation of the present invention;

[0039] Figure 3 Schematic diagram of the attitude control timing of the satellite bracket of the present invention. Specific implementation manners

[0040] The present invention will be further described below in conjunction with embodiments.

[0041] The present invention provides a star-ship separation design method for inclined-layout satellites. By combining research means such as numerical simulation and experiments, the research problem of the contact dynamics of the star-ship separation ejector is focused on. The separation scheme considers the combined design method of separation attitude control, solves the problems of controlling the satellite attitude angular velocity deviation after the separation of the satellites arranged on the inclined satellite bracket and the safety of star-ship separation, and ensures the safety of star-ship separation.

[0042] The star-ship separation design method for inclined-layout satellites specifically includes the following steps:

[0043] Design a multi-satellite inclined structure, as Figure 1 shown, including a control cabin, a satellite bracket, n satellites and n star-ship separation mechanisms; the satellite bracket is installed on the top of the control cabin; the n satellites are installed on the satellite bracket; each satellite is connected to the satellite bracket through 1 star-ship separation mechanism.

[0044] The bottom plane of the satellite bracket is docked with the top of the control cabin; the top of the satellite bracket is a regular n-sided pyramid structure; each satellite is correspondingly installed on 1 side surface of the satellite bracket; the angle between each side surface of the satellite bracket and the axis is θ.

[0045] When a single satellite is in an inclined layout, the satellite bracket is located above the control cabin. The satellite bracket and the control cabin are connected by bolts on the docking surface to form a control cabin + satellite bracket combination (also abbreviated as the last stage). The installation surface between the satellite bracket and the satellite is designed to be inclined, and the angle θ is the angle between the star-ship docking surface and the axis. θ can be adjusted according to the available envelope and centroid balance of the satellite layout in the fairing. Among them, the installation surface of the satellite bracket is in an inclined angle θ state, and the satellite is above the installation surface of the satellite bracket. In this state, the satellite and the satellite bracket are connected through a separation device. Similarly, for a multi-satellite inclined layout, multiple satellite bracket installation surfaces with different inclined angles can be set on the satellite bracket, and satellites can be arranged on the corresponding inclined installation surfaces, as Figure 1 shown.

[0046] The satellite-rocket separation mechanism uses a push rod-spring structure as the energy storage mechanism for satellite-rocket separation; a locking device is provided on the satellite-rocket separation mechanism. After the locking device is unlocked, the elastic potential energy stored in the spring is released, and the satellite is pushed away from the satellite bracket by the push rod to achieve satellite-rocket separation.

[0047] Optimize the friction coefficient of the contact surface between the push rod and the corresponding satellite.

[0048] In the satellite-rocket separation mechanism, the push rod-spring device is the energy storage mechanism for satellite-rocket separation. When the separation device is unlocked, the elastic potential energy stored in the spring separates the satellite from the rocket. The separation state of the satellite is directly related to the interaction between the push rod and the bottom surface of the satellite, and the friction coefficient of the contact interface will change with the relative speed and contact pressure between the end of the push rod and the ground contact surface of the satellite.

[0049] The design method of the friction coefficient of the contact surface between the push rod and the corresponding satellite is as follows:

[0050] S31. Make a contact surface of the same material as the satellite at the end of the push rod for experiments. In order to accurately describe the interaction between the push rod and the satellite, first use a friction testing machine to obtain the dynamic friction coefficient under different contact pressures and different relative sliding speeds. The specific method is: use the Figure 2 friction test piece in it to make a contact surface of the same material as the bottom surface of the satellite at the end of the push rod for experiments. According to the working conditions during the actual satellite separation process, 30 groups of experiments are formulated with the contact pressure ranging from 100 N to 1600 N and different relative sliding speeds ranging from 10 mm / s to 200 mm / s.

[0051] S32. According to the data of the friction coefficient obtained from the experiments with respect to the contact pressure and the tangential relative speed, use the B-spline interpolation method to obtain the functional relationship between the friction coefficient and different contact pressures and tangential relative speeds, as shown in Figure 2 shown.

[0052] S33. Establish a satellite-rocket separation dynamics model, substitute the functional relationship between the friction coefficient and different contact pressures and tangential relative speeds into the satellite-rocket separation dynamics model, and set different friction coefficients μ d in the satellite-rocket separation dynamics model for simulation, and select the optimized friction coefficient value.

[0053] The method for selecting the optimized friction coefficient value is as follows:

[0054] Set different friction coefficients μ d in the satellite-rocket separation dynamics model for simulation, obtain the final angular velocity values under different friction coefficients μ d , select the final angular velocity value closest to the actual angular velocity, and the corresponding friction coefficient μ d is the optimized friction coefficient value.

[0055] Establish a separation ejector model for the satellite and the rocket. The dynamic friction coefficients in the Coulomb friction model are respectively selected as μ d = 0.16 / 0.18 / 0.20 / 0.22 / 0.24, the spring stiffness k = 41 N / mm, and calculations are respectively carried out using the Coulomb friction model and the experimental friction model. The satellite angular velocity values are shown in the following table:

[0056] <![CDATA[Coefficient of kinetic friction μ d > 0.16 0.18 0.20 0.22 0.24 Experimental friction model Final angular velocity (° / s) -0.4032 -0.2754 -0.1471 -0.0184 0.1108 -0.1772

[0057] It can be seen from the calculation results that: when using the Coulomb friction and the dynamic friction coefficient μ d = 0.2, the final values of the satellite angular velocity ω s are respectively -0.1471° / s and -0.1772° / s. At this time, the gap is the smallest, only 0.03° / s. Thus, it can be seen that choosing the Coulomb friction model with a dynamic friction coefficient μ d = 0.20 is closer to the experimental friction model.

[0058] Optimize the spring force.

[0059] Adopt a large-inclination satellite bracket layout for the satellite separation design. The torque acting on the upper stage is M upper stage = M1'+M2'. The upper stage will rotate rapidly, resulting in a shorter acting time of the upper spring compared to the lower spring, and the torque M satellite = M1 - M2 acting on the satellite is unbalanced, thus affecting the angular velocity during satellite separation. At the same time, it is necessary to consider that the relative separation speed between the satellite and the upper stage is not less than 500 mm / s as the baseline to optimize the separation spring.

[0060] To ensure the angular velocity requirement after satellite separation and make M satellite = M1 - M2 balanced, the initial force exerted by the upper spring should be greater than that of the lower spring. The force exerted by the spring on the satellite during this separation process is non-linearly changing, and it is necessary to analyze through dynamic simulation modeling whether the separation angular velocity and separation speed meet the requirements.

[0061] 1. When optimizing using springs of the same specification, it is necessary to increase the compression amount of the upper spring and at the same time decrease the compression amount of the lower spring.

[0062] 2. When using springs of different specifications, it is necessary to obtain the appropriate spring force through simulation analysis, and then carry out spring selection or design work.

[0063] The spring includes an upper spring and a lower spring; the upper spring is arranged at the top corresponding to the satellite; the lower spring is arranged at the bottom corresponding to the satellite. The optimization method of the spring force is as follows:

[0064] Set the acting force of the lower spring as F1, the reaction force of the lower spring as F1', the acting force of the upper spring as F2, and the reaction force of the upper spring as F2'; the optimized spring force satisfies:

[0065] 1550N ≤ F1 ≤ 1700N, 1950N ≤ F2 ≤ 2050N, F2 > F1.

[0066] Optimize the spring torque. The optimization method of the spring torque is as follows:

[0067] Set the torque of F1 acting on the satellite's center of mass as M1, and the torque of F2 acting on the satellite's center of mass as M2; after optimization, it satisfies:

[0068] M1 = M2, and the lever arm length is 1 - 4m.

[0069] Set the torque of F1' acting on the center of mass of the satellite bracket as M1', and the torque of F2' acting on the center of mass of the satellite bracket as M2'; after optimization, it satisfies:

[0070] The total torque acting on the center of mass of the satellite bracket is M1' + M2'.

[0071] After the satellite and rocket separate, design the attitude control timing and attitude adjustment angle of the satellite bracket to achieve safe control of the satellite and rocket separation.

[0072] Since the satellite bracket + control cabin combination has a large angular velocity after separation, in order to eliminate the interference caused by attitude control, avoid collisions caused by random angular velocity, and ensure the near-field safety after separation, the attitude system stops controlling at the moment of satellite separation, and the attitude system starts controlling after the satellite has separated for t 1 duration, at this time the separated body has reached a certain safe distance.

[0073] When the attitude control starts again, the direction with a large change in attitude angle requires the attitude control to be adjusted to X within t 2 duration and remain stable until t 3 moment for orbit departure attitude adjustment.

[0074] As Figure 3 shown, the design method of the attitude control timing and attitude adjustment angle of the satellite bracket is as follows:

[0075] Set the satellite and rocket separation duration as t 1 ; after the satellite and rocket separate, they simultaneously enter the attitude adjustment of the satellite bracket and prepare for orbit departure attitude adjustment; among them, the duration of the attitude adjustment of the satellite bracket is t 2 , the adjustment angle is X; the duration of preparing for orbit departure attitude adjustment is t 3 ;

[0076] 30s ≤ t 1 ≤ 50s, 3s ≤ t 2 ≤ 8s, 150s ≤ t 3 ≤ 300s, 30° ≤ X ≤ 50°.

[0077] The present invention first proposes a method for arranging satellites on a large-inclination satellite support, solves the problems of separation safety and limited layout, and maximizes the available envelope of the fairing.

[0078] The present invention proposes an angular velocity control method for inclined layout satellites, solves the problem of uncertain friction coefficient between the separation spring ejector and the satellite contact surface, and improves the inherent reliability of the separation design and the overall performance of the rocket.

[0079] Through collaborative attitude control design, the present invention solves the problems of collision interference caused by excessive angular velocity of the upper stage and the problem that the main engine of the rocket or the final correction jet flow approaches and contaminates the satellite due to rotation, and improves the separation safety of the satellite and the rocket.

[0080] Although the present invention has been disclosed above with preferred embodiments, it is not intended to limit the present invention. Any person skilled in the art can make possible changes and modifications to the technical solution of the present invention by using the methods and technical contents disclosed above without departing from the spirit and scope of the present invention. Therefore, any simple modifications, equivalent changes and decorations made to the above embodiments according to the technical essence of the present invention without departing from the technical solution of the present invention all fall within the protection scope of the technical solution of the present invention.

Claims

1. A satellite-rocket separation design method for an inclined satellite, characterized in that: include: Design of multi-satellite tilting structure, including control cabin, satellite bracket, n satellites and n satellite-rocket separation mechanisms; The satellite bracket is installed on top of the control cabin; n satellites are installed on a satellite bracket; each satellite is connected to the satellite bracket via a satellite-rocket separation mechanism; The satellite-rocket separation mechanism adopts a push rod-spring structure as an energy storage mechanism for satellite-rocket separation; a locking device is provided on the satellite-rocket separation mechanism. After the locking device is unlocked, the elastic potential energy stored in the spring is released, and the push rod is used to push the satellite out of the satellite bracket to achieve satellite-rocket separation; Optimize the friction coefficient between the push rod and the corresponding satellite contact surface; Optimize spring force; Optimize spring torque; After the separation of the satellite and the rocket, the attitude control timing and attitude adjustment angle of the satellite bracket are designed to achieve safe control of the separation of the satellite and the rocket.

2. The satellite-rocket separation design method for an inclined satellite according to claim 1, characterized in that: The bottom plane of the satellite bracket is butted against the top of the control cabin; the top of the satellite bracket is a positive n-face cone structure; each satellite is correspondingly mounted on one side of the satellite bracket; and the angle between each side of the satellite bracket and the axis is θ.

3. The satellite-rocket separation design method for an inclined satellite according to claim 1, characterized in that: The design method of the friction coefficient between the push rod and the corresponding satellite contact surface is: S31, make the contact surface between the end of the push rod and the satellite with the same material for experiment; S32, according to the experimental data of the friction coefficient with the contact pressure and the tangential relative speed, the B-spline interpolation method is used to obtain the functional relationship between the friction coefficient and different contact pressures and tangential relative speeds; S33, establishing a satellite-rocket separation dynamics model, and substituting the functional relationship between the friction coefficient and different contact pressures and the tangential pair velocity into the satellite-rocket separation dynamics model; Setting different friction coefficients μ in the satellite-rocket separation dynamics model d Perform simulation and select the optimized friction coefficient value.

4. The satellite-rocket separation design method for an inclined satellite according to claim 3, characterized in that: In S31, the specific contents of the experiment are: 30 sets of experiments were conducted with pressure ranging from 100N to 1600N and relative sliding speeds ranging from 10mm / s to 200mm / s.

5. The satellite-rocket separation design method for an inclined satellite according to claim 3, characterized in that: In S33, the method for selecting the optimized friction coefficient value is: Setting different friction coefficients μ in the satellite-rocket separation dynamics model d Simulate and obtain different friction coefficients μ d The final angular velocity value under the condition of , select the final angular velocity value that is closest to the actual angular velocity, and its corresponding friction coefficient μ d This is the optimized friction coefficient value.

6. The satellite-rocket separation design method for an inclined satellite according to claim 1, characterized in that: The spring comprises an upper spring and a lower spring; the upper spring is arranged at the top end of the corresponding satellite; and the lower spring is arranged at the bottom end of the corresponding satellite.

7. The satellite-rocket separation design method for an inclined satellite according to claim 6, characterized in that: The optimization method of the spring force is: Set the action force of the lower spring to F1, the reaction force of the lower spring to F1', the action force of the upper spring to F2, and the reaction force of the upper spring to F2'; after optimization, the spring force satisfies: 1550N≤F1≤1700N, 1950N≤F2≤2050N, F2>F1.

8. The satellite-rocket separation design method for an inclined satellite according to claim 7, characterized in that: The optimization method of the spring torque is: Assume that the moment of F1 acting on the satellite's mass center is M1, and the moment of F2 acting on the satellite's mass center is M2; after optimization, the following conditions are satisfied: M1=M2 lever arm length 1~4m.

9. The satellite-rocket separation design method for an inclined satellite according to claim 8, characterized in that: Assume that the moment of F1' acting on the mass center of the satellite bracket is M1', and the moment of F2' acting on the mass center of the satellite bracket is M2'. After optimization, the following conditions are satisfied: The total moment acting on the center of mass of the satellite bracket is M1'+M2'.

10. The satellite-rocket separation design method for an inclined satellite according to claim 1, characterized in that: The design method of the attitude control timing and attitude adjustment angle of the satellite bracket is: Set the satellite-rocket separation duration to t1; after the satellite-rocket separation, the satellite bracket will simultaneously enter the attitude adjustment and deorbit preparation attitude adjustment; the satellite bracket attitude adjustment duration is t2, the adjustment angle is X; the deorbit preparation attitude adjustment duration is t3; 30s≤t1≤50s、3s≤t2≤8s、150s≤t3≤300s、30°≤X≤50°。

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