Active electric propulsion and passive air resistance deorbit composite spacecraft deorbit method

Abstract

The invention relates to a deorbit method for an active electric propulsion and passive air resistance deorbit composite spacecraft, and belongs to the technical field of space debris slowing. Under the condition of different deorbit shift heights H1, the duration and deorbit mass of an electric propulsion deorbit process and an air resistance deorbit process are calculated respectively, and the total duration T of a fixed task is calculated; and calculating the minimum value of the total mass MDEO of the deorbit device, traversing the deorbit shift change height H1 and comparing the MDEO to obtain the optimal deorbit shift change height H1 and the total mass MDEO of the deorbit device. The advantages of the space electric propulsion technology and the passive air resistance deorbit technology are combined, the deorbit process of large space debris is divided into an active electric propulsion deorbit stage and an air resistance passive deorbit stage, and the high specific impulse of the active electric propulsion deorbit technology on a high orbit is fully exerted; and due to the advantages of long service life and large-area long-term air resistance speed reduction of the unfolding structure on a low orbit, a deorbit scheme with better overall cost, deorbit capability and quality cost is formed.

Description

TECHNICAL FIELD

[0001] The present application belongs to the field of space debris mitigation technology, and particularly relates to a method for deorbiting a composite space vehicle by active electric propulsion and passive air resistance deorbiting. BACKGROUND

[0002] Space debris refers to all in-orbit, abandoned man-made objects in space, as large as satellite disintegration caused debris, and also includes rocket final stage, abandoned satellites and other abandoned spacecraft; as small as the coating dropped by the spacecraft, unburned solid particles of solid rocket engine, etc. Although most of the space debris is small in size, but due to the average speed of space debris and spacecraft collision is as high as 10km / s, debris larger than 1cm will cause serious damage to spacecraft. Therefore, relevant international organizations have made a 25-year deorbiting regulation, and it is necessary to carry out research on related spacecraft autonomous deorbiting technology.

[0003] In view of the increasingly serious safety problem of on-orbit operation of spacecraft caused by space debris orbit space occupation, the deorbiting is carried out by using the resistance of thin atmosphere molecules on the space orbit, which has the advantages of simple structure, low cost, easy realization and no fuel consumption, and does not need attitude control, and has great technical advantages. Among all the resistance increasing methods, the space deployable structure resistance increasing deorbiting method has a very small folding and unfolding ratio, meets the resistance increasing demand of dozens or even hundreds of square meters, is beneficial to space carrying, and has high application value. However, the space deployable structure resistance increasing deorbiting technology still has certain limitations, mainly including:

[0004] (1) The use of space deployable structure to complete the deorbiting task is limited to space debris with an orbital height of less than 600km. For space debris with an orbital height of more than 600km, the density of thin atmosphere molecules increases with the height, and the air resistance effect is very weak. The time cost required for air resistance deorbiting increases greatly. In order to realize the deorbiting requirement of debris within a given time, a larger space folding and unfolding mechanism is needed, and then a larger deorbiting device mass and volume cost is needed, so high-efficiency and low-cost deorbiting cannot be realized.

[0005] (2) The large-area space folding and unfolding mechanism is not mature enough in design and engineering application, especially the space structure which needs to be on-orbit for a long time and maintain shape, and the damage effect of space harsh environment also needs to be fully considered. There is certain technical risk in large-area resistance increasing deorbiting. SUMMARY

[0006] The present application aims to overcome the deficiency of the existing resistance increasing deorbiting technology in the adaptability of debris orbital height, and provides a method for deorbiting a composite space vehicle by active electric propulsion and passive air resistance deorbiting, solves the problems of long air resistance deorbiting time and large deorbiting load of high-orbit debris, and avoids the space environment adaptability problem caused by long-term on-orbit of the resistance increasing deorbiting device.

[0007] The above-mentioned object of the present application is mainly achieved by the following technical solutions:

[0008] A method for deorbiting a composite space vehicle by active electric propulsion and passive gas resistance, comprising the following steps:

[0009] (1) Obtain the deorbiting time T, the spacecraft mass M and the initial orbit height H; select the deorbiting handover height H1;

[0010] (2) Simulate and analyze the electric propulsion deorbiting process, set multiple electric propulsion thrusts, calculate the deorbiting task duration T1 and the electric propulsion deorbiting task mass M1 respectively, and draw the M1-T1 curve;

[0011] (3) Simulate and analyze the gas resistance deorbiting process, the initial height of the gas resistance deorbiting is the deorbiting handover height H1, the end height is 0, multiple surface-mass ratio parameters γ are set, the deorbiting duration T2 is calculated, wherein γ = S / M, S is the equivalent resistance area of deorbiting, and M is the mass of the task spacecraft; draw the γ-T2 curve;

[0012] (4) Draw the M2-T2 curve according to the γ-T2 curve obtained in step (3), M2 = M*γ*ρ, wherein M2 is the mass of the gas resistance deorbiting device, M is the mass of the task spacecraft, and ρ is the average equivalent surface density of the expansion device;

[0013] (5) The total duration of the task T = T1+T2, and the total mass of the deorbiting device M DEO =M1+M2; fix the total duration of the task T, calculate the minimum value of the total mass of the deorbiting device M DEO ;

[0014] (6) Traverse the deorbiting handover height H1, repeat steps (2) to (5), compare all the minimum values of the total mass of the deorbiting device M DEO , and obtain the deorbiting handover height H1 corresponding to the minimum value of M DEO , denoted as the optimal deorbiting handover height H1 and the total mass of the deorbiting device M DEO ;

[0015] (7) Deorbit according to the optimal deorbiting handover height H1 and the optimal total mass of the deorbiting device M DEO obtained in step (6).

[0016] After obtaining the optimal deorbiting handover height H1 and the total mass of the deorbiting device M DEO , simulate the orbit dynamics of the electric propulsion deorbiting process to determine the deorbiting time and working medium consumption parameters.

[0017] After obtaining the optimal deorbiting handover height H1 and the total mass of the deorbiting device M DEOAfter that, the configuration size and mass of the air resistance deorbit device are designed, and the design method is as follows:

[0018] (1) The resistance area of the air resistance deorbit device is determined;

[0019] (2) The configuration size and mass of the air resistance deorbit device are selected;

[0020] (3) According to the configuration size and mass of the air resistance deorbit device obtained in step (2), the precise dynamics simulation of the spacecraft deorbit process coupled with the attitude characteristics is performed, and the spacecraft deorbit time T2' is calculated;

[0021] (4) The error between the spacecraft deorbit time T2' obtained in step (3) and the deorbit time T2 is calculated, an error threshold is set, if the error is less than the error threshold, the configuration size and mass of the air resistance deorbit device are designed, if the error is greater than the error threshold, the configuration size and mass of the air resistance deorbit device are adjusted, and the operation in step (3) is repeated until the error meets the requirements.

[0022] In the step (3), the precise dynamics simulation needs to consider the earth flattening, atmospheric resistance and solar pressure.

[0023] In the step (5), the total mass M of the deorbit device DEO The calculation method is as follows: the M2-T2 curve is inverted and combined with the M1-T1 curve, the M2-T2 curve is translated according to the total deorbit time T, and then a vertical line is drawn relative to the points on the T1, T2 coordinates, when the distance between the intersection points of the vertical line and the M1-T1 curve and the M2-T2 curve is maximum, the total mass M of the deorbit device DEO is the minimum value.

[0024] A calculation system based on the above deorbit method, the deorbit time T, the spacecraft mass M and the initial orbit height H are input, and the optimal deorbit handover height H1 and the total mass M of the deorbit device DEO are calculated.

[0025] Compared with the prior art, the present application has at least the following beneficial effects:

[0026] (1) The aircraft deorbit method of the present application combines the advantages of space electric propulsion technology and passive air resistance deorbit technology, and divides the deorbit process of large space debris into active electric propulsion deorbit stage and air resistance passive deorbit stage. The advantages of high specific impulse, long life of active electric propulsion deorbit technology in high orbit and large area long-term air resistance deceleration of the unfolded structure in low orbit are fully utilized, and a more optimal deorbit scheme in terms of cost, deorbit capacity and mass cost is formed.

[0027] (2), the application improves the feasibility of the scheme of de-orbiting large space debris in high orbit, takes into account the high-efficiency and low-mass loss advantage of air resistance de-orbiting in low orbit height and electric propulsion in high orbit height, reduces the working time length of the space folding and unfolding mechanism in the harsh space environment, improves the overall space environment adaptability and reliability of the de-orbiting device, and improves the engineering level.

[0028] (3), the application introduces electric propulsion de-orbiting in high orbit, greatly reduces the design difficulty of the folding and unfolding device of air resistance de-orbiting, and provides a design idea for improving the overall reliability and engineering application level of the device.

[0029] (4), the embodiment of the application preferably considers the size and mass of the air resistance de-orbiting device, eliminates the error caused by the size difference, and the calculation result is more accurate. BRIEF DESCRIPTION OF DRAWINGS

[0030] Figure 1 It is a schematic diagram of the composite de-orbiting method of the application;

[0031] Figure 2 It is a de-orbiting method design process of the application;

[0032] Figure 3 It is a schematic diagram of determining the area-mass ratio parameters required for different de-orbiting times of the application;

[0033] Figure 4 It is a device total mass optimization principle diagram of the double de-orbiting stage of the application;

[0034] Figure 5 It is a schematic diagram of the air resistance de-orbiting device accounting iteration process of the application. DETAILED DESCRIPTION

[0035] The application will be described in further detail below in combination with the drawings and specific embodiments:

[0036] The design steps of the application are as shown in Figure 2 :

[0037] 1, clear de-orbiting time T, spacecraft mass M, initial orbit height H and other parameters: usually the design and implementation of the de-orbiting method needs to be carried out simultaneously with the design of the whole spacecraft, the mass of the de-orbiting spacecraft, the initial orbit height and the target de-orbiting time are usually obtained from the mission overall designer;

[0038] 2, determine the handover height H1 of electric propulsion and air resistance de-orbiting task: in order to fully optimize the composite space vehicle de-orbiting task, a series of de-orbiting handover heights H1 are designed, and the de-orbiting task load mass is optimized under each handover height H1;

[0039] 3, determine the electric propulsion de-orbiting task mass M1-de-orbiting time T1 curve:

[0040] (1) Orbit transfer, orbit maintenance and attitude adjustment tasks are performed by electric propulsion engine, and corresponding deorbit task is performed by electric propulsion capability and attached propellant. In order to perform the flight of deorbit task, additional electric propulsion propellant mass and container capacity structure mass are required, which is called electric propulsion deorbit task mass M1;

[0041] (2) Simulation analysis is performed on the electric propulsion deorbit process, and the initial height H and the handover height H1 of the spacecraft electric propulsion deorbit are fixed in the simulation scenario, as shown in Figure 1 , the deorbit task time T1 and the electric propulsion deorbit task mass M1 are obtained under different electric propulsion thrust sizes. Generally, increasing the thrust level will reduce the electric propulsion deorbit task time T1, and increase the propellant consumption, that is, increase the electric propulsion deorbit task mass M1;

[0042] (3) The task mass M1-deorbit task time T1 curve of electric propulsion deorbit is drawn;

[0043] 4. Synchronously determine the gas resistance deorbit device mass-deorbit time curve:

[0044] (1) The initial height of the gas resistance deorbit is the handover height H1, and the task spacecraft mass M;

[0045] (2) Preliminary simulation is performed on the gas resistance deorbit process, and the initial deorbit height H1 and the end height 0 are determined in the simulation scenario. The deorbit time length T2 under different area-mass ratio parameters γ (deorbit equivalent resistance area S / task spacecraft mass M) is calculated, that is, the area-mass ratio γ-deorbit time T2 curve, as shown in Figure 3 ;

[0046] (3) According to the task spacecraft mass M, the area-mass ratio parameter γ and the equivalent surface density ρ of the expansion device (the equivalent surface density ρ of the expansion device = the expansion device mass / the device resistance area, which is closely related to the configuration, mechanism design and material of the expansion device, and devices with the same configuration, size range and material have similar equivalent surface density), the gas resistance deorbit device mass M2-deorbit time T2 curve is determined. The gas resistance deorbit device mass M2 = task spacecraft mass M*area-mass ratio γ*average equivalent surface density ρ of the expansion device.

[0047] 5. For the determined deorbit task total time T = T1+T2, the total mass M DEO of the deorbit device is optimized, as shown in Figure 4 , the gas resistance deorbit curve is inverted, and the ordinate axis is shifted by a fixed distance relative to the ordinate axis of the electric propulsion deorbit curve, that is, the deorbit task total time T. All points on the horizontal coordinate are scanned, and the vertical line is drawn. The maximum distance between the intersection points of the two curves is the optimal state, and the mass cost is the lowest.

[0048] 6. Change different handover height H1, repeat steps 3-5, obtain optimal handover height and composite orbiting mass, and preliminarily determine orbiting scheme parameter system.

[0049] 7. After preliminarily determining the system overall parameters, it is necessary to perform pose-orbit coupling fine simulation on the working process of the orbiting device to verify the accuracy of the parameters. First, perform orbit dynamics simulation based on the propulsion control scheme of the electric propulsion engine to verify and determine the orbiting time T1 and the electric propulsion orbiting task mass parameter M1.

[0050] 8. Then, perform accounting of the air resistance orbiting device and the task parameters through the fine dynamics model of the pose-orbit coupling, as shown in the following formula: Figure 5

[0051] (1) Clearly define the resistance area requirement of the air resistance orbiting device;

[0052] (2) Clearly define the overall configuration and size parameters of the air resistance orbiting device;

[0053] (3) Bring the configuration parameters of the air resistance orbiting device into the fine dynamics model of the orbiting process, consider the effects of the earth flattening, atmospheric resistance and solar pressure and other perturbation forces, and develop accurate dynamics simulation of the spacecraft orbiting process coupled with the attitude characteristics, to calculate the orbiting time of the spacecraft under the given configuration size of the air resistance orbiting device;

[0054] (4) Judge the error of the calculated orbiting time and the target orbiting time, when the error is less than the threshold value, it is considered that the orbiting task requirements have been met, and the configuration size design of the air resistance orbiting device is completed, when the error does not meet the requirements, then further adjust the configuration size parameters of the air resistance orbiting device, and bring them into the fine dynamics model again for simulation, through multiple iteration optimization, obtain the configuration size and mechanism mass of the air resistance orbiting device that meet the task requirements.

[0055] The above is only the best specific embodiment of the present application, but the protection scope of the present application is not limited thereto, any change or replacement that can be easily thought of by those skilled in the art within the technical range disclosed by the present application should be covered within the protection scope of the present application.

[0056] The contents not described in detail in the specification of the present application are the known technology of the person skilled in the art.​

Claims

1. A method for deorbiting a spacecraft using a combination of active electric propulsion and passive air resistance, characterized in that: The following steps are involved: (1) Obtain the deorbit time T, spacecraft mass M and orbital initial altitude H; select the deorbit handover altitude H1; (2) Simulate and analyze the electric propulsion deorbit process, set multiple electric propulsion thrusts, calculate the orbit reduction mission duration T1 and the electric propulsion deorbit mission mass M1, and draw the M1-T1 curve; (3) Simulate and analyze the gas drag deorbit process. The initial height of the gas drag deorbit is the deorbit handover height H1, and the end height is 0. Set multiple surface-to-mass ratio parameters γ and calculate the deorbit time T2, where γ = S / M, S is the deorbit equivalent drag area, and M is the mass of the mission spacecraft; draw the γ-T2 curve; (4) Draw the M2-T2 curve based on the γ-T2 curve obtained in step (3), where M2 = M*γ*ρ, where M2 is the mass of the air-resistance deorbit device, M is the mass of the mission vehicle, and ρ is the average equivalent surface density of the deployment device; (5) Total mission duration T = T1 + T2, total mass of the deorbiting device M DEO =M1+M2; Fixed mission duration T, calculate the total mass of the deorbit device M DEO minimum value; (6) Traverse the deorbit handover height H1, repeat steps (2) to (5), and compare the total mass M of all deorbit devices obtained. DEO Minimum value, get the M with the smallest value DEO The corresponding deorbit handover height H1 is recorded as the optimal deorbit handover height H1 and the total mass of the deorbit device M DEO ; (7) The optimal deorbit handover height H1 and the optimal deorbit device total mass M obtained according to step (6) DEO Deorbit.

2. The deorbiting method for a spacecraft using a combination of active electric propulsion and passive air resistance according to claim 1, characterized in that: Get the optimal deorbit handover height H1 and the total mass M of the deorbit device DEO Finally, the orbital dynamics simulation of the electric propulsion deorbit process is carried out to determine the deorbit time and working fluid consumption parameters.

3. The deorbiting method for a spacecraft using a combination of active electric propulsion and passive air resistance according to claim 1, characterized in that: Get the optimal deorbit handover height H1 and the total mass M of the deorbit device DEO After that, the configuration, size and mass design of the air-resistance deorbit device are carried out. The design method is as follows: (1) Determine the resistance area of ​​the air resistance derailment device; (2) Select the configuration, size and mass of the air-resistance deorbit device; (3) Based on the configuration size and mass of the air-resistance deorbiting device obtained in step (2), perform an accurate dynamic simulation of the spacecraft deorbiting process coupled with the attitude characteristics, and calculate the spacecraft deorbiting time T2'; (4) Calculate the error between the spacecraft deorbit time T2' obtained in step (3) and the deorbit time T2, set an error threshold, and if the error is less than the error threshold, complete the configuration size and mass design of the air resistance deorbit device; if the error is greater than the error threshold, adjust the configuration size and mass of the air resistance deorbit device, and repeat the operation described in step (3) until the error meets the requirements.

4. The method for deorbiting a spacecraft using a combination of active electric propulsion and passive air resistance according to claim 3, wherein: In step (3), accurate dynamic simulation needs to take into account the earth's flattening, atmospheric drag and solar radiation pressure.

5. The deorbiting method for a spacecraft using a combination of active electric propulsion and passive air resistance according to claim 1, characterized in that: In step (5), the total mass M of the deorbiting device is DEO The calculation method is as follows: invert the M2-T2 curve and combine it with the M1-T1 curve, and translate the M2-T2 curve according to the total deorbit mission duration T; then draw a perpendicular line relative to the points on the T1 and T2 coordinates. When the distance between the perpendicular line and the intersection of the M1-T1 curve and the M2-T2 curve is the largest, the total mass of the deorbit device M DEO is the minimum value.

6. A computing system based on the deorbiting method according to any one of claims 1 to 5, characterized in that: Input the deorbit time T, spacecraft mass M and orbit initial height H, and calculate the optimal deorbit handover height H1 and the total mass of the deorbit device M DEO .

Images