Rotation tether system maneuvering method for load transmission between different-plane tracks

Abstract

The invention discloses a rotating tether system maneuvering method for load transmission between different-plane orbits, and the method systematically designs a load throwing mechanism by constructing a double-body rotating tether system architecture: firstly, planning a load release mode based on the characteristics of a double-body rotating tether system; deducing speed increment and position increment parameters required for realizing the release mode; by calculating the total mass and length of the tether and combining the relative position vector and the velocity vector between the end body satellites required by load throwing, the key parameters of the throwing rotation surface are accurately calculated; and finally, constructing a rotating surface maneuvering control strategy of a full task period based on the parameters. Compared with a traditional coplanar orbit transfer scheme, the non-coplanar orbit transfer task of the effective load from the initial inclined low earth orbit to the target equator geostationary orbit is achieved in a breakthrough mode. The method has obvious technical reference value and engineering application prospect in orbit inclination angle adjustment maneuvering tasks in a satellite launching stage.

Description

Technical Field

[0001] The present application belongs to the field of aerospace science and technology, and specifically relates to a maneuvering method of a rotating tether system for transmitting loads between non-planar orbits. Background Art

[0002] In the aerospace sector, payload transportation is a critical task. Traditional methods rely on rocket pulse acceleration, but this approach has limitations in terms of fuel consumption, cost, and mission flexibility. Against this backdrop, the space rotating tether system has emerged. As a space-based combined vehicle connecting two spacecraft via a tether to jointly complete on-orbit missions, it offers unique advantages. Its excellent centrifugal stability and momentum exchange properties provide a reusable solution for on-orbit payload transportation. Consequently, the use of rotating tether systems for payload transfer has attracted widespread attention and research, showing promising development prospects.

[0003] The rotating tether system does not consume fuel during momentum exchange and can deliver payloads instantly. Compared to traditional rocket pulse acceleration, it has become a viable alternative to traditional rockets for out-of-plane orbital transport. Earlier, in 2000, Lorenzini published the paper "Mission analysis of spinning systems for transfers from low orbits to geostationary," which first proposed a complete scheme for payload transfer using a rotating tether system. The rotating tether system designed in this scheme can deliver a payload of 21,400 kg into GEO orbit within two years, consuming only 5,380 kg of propellant. Its payload to disposable consumable equipment mass ratio is much lower than that of traditional rocket transport payloads, laying the foundation for research on payload transfer using rotating tether systems.

[0004] Since then, researchers have continued to delve deeper, conducting more detailed studies on orbital transfer using rotating tethered systems. The 2016 paper, "Optimal Control of Payload Tossing Using a Space Tethered System," proposed a design for a rotating tethered system performing multi-stage payload transfer between coplanar elliptical orbits and an optimal control law for the drop phase, effectively mitigating the effects of the elliptical orbits on the rotating tethered system. Chinese patent application number CN202010089767.1 discloses a method for designing a tethered satellite transfer orbit between coplanar elliptical orbits and discusses collision and breakage prevention techniques during the drop phase.

[0005] However, existing technologies for payload transfer using rotating tether systems primarily focus on transporting payloads within the orbital plane. Previous research has focused on coplanar transport scenarios, where the rotational plane coincides with the orbital plane. The drop process provides no velocity increment perpendicular to the orbital plane, resulting in the payload arriving on an orbit coplanar with the rotating tether system's orbit. However, in practice, the task of using a rotating tether system to drop payloads onto an orbit that is not coplanar with the rotating tether system's center of mass orbital plane is a real and currently unresolved challenge, presenting a pressing challenge in this field. Summary of the Invention

[0006] Addressing the limitation of existing technologies that only allow for coplanar transport, this application provides a rotating tether system maneuvering method for payload transfer between out-of-plane orbits. Compared to traditional coplanar transfer schemes, this application can successfully transport a payload from an initially inclined low-Earth orbit to a final equatorial geostationary orbit, achieving the out-of-plane orbital transfer mission. This method has significant reference value for orbital inclination maneuvers during satellite launch.

[0007] In order to achieve the technical objectives of this application, this application mainly adopts the following technical solutions:

[0008] In one aspect of the present application, a method for maneuvering a rotating tether system for load transmission between skew tracks is provided, comprising the following steps:

[0009] S1: Construct a two-body rotating tether system, including a mother spacecraft and a daughter spacecraft connected by a tether, wherein the daughter spacecraft includes a payload and a capture / ejection mechanism;

[0010] S2: On the transfer orbit in which the rotating tether system operates, positioning the center of mass of the rotating tether system on the transfer orbit, positioning the sub-spacecraft at a target payload orbit position, and the transfer orbit being out of plane with the target payload orbit;

[0011] S3: Calculate the position increment Δr required for the payload to move from the transfer orbit to the out-of-plane target orbit based on the orbital roots of the rotating tether system and the target payload orbit. p and velocity increment Δv p ;

[0012] S4: Calculate the total mass of the tether based on the expected rate required for payload delivery, and calculate the tether length based on the total mass of the tether and the distance between the sub-spacecraft and the center of mass of the rotating tether system;

[0013] S5: Calculate the parameters of the jettisoning rotation surface according to the relative position vector and velocity vector between the end-body satellites required for payload jettisoning. The jettisoning rotation surface parameters include the rotation surface normal vector (n s ) C , angular rate of rotation and rotation angle Ψ s ;

[0014] S6: Design a rotational surface maneuvering strategy based on the casting rotational surface parameters, and control the rotational surface angle and the rotation angle to make the system maneuver from the initial rotational surface to the casting rotational surface, and satisfy the position increment Δr at the casting time. p and velocity increment Δv p Requirements, to achieve load transmission on non-planar tracks.

[0015] In one embodiment, the diameter of the tether gradually decreases from the center of mass to both ends.

[0016] In one embodiment, the tether limit breaking rate V tc satisfy:

[0017]

[0018] Where σ is the ultimate strength of the tether, f is the safety factor of the tether, and ρ is the tether density.

[0019] In one embodiment, in step S2, the mother spacecraft and the daughter spacecraft are respectively located on both sides of the center of mass of the rotating tether.

[0020] In one embodiment, in step S3, the position increment Δr p and velocity increment Δv p for:

[0021] (Δr p ) eq =(r tar ) eq -(r GTO ) eq ,(Δv p ) eq =(v tar ) eq -(v GTO ) eq ;

[0022] in,(*) eq Represents vector * in the Earth's equatorial inertial coordinate system OX eq Y eq Z eq The coordinate vector, v tar ,r tar represent the velocity and position vectors on the target payload track, v GTO ,r GTO are the velocity and position vectors on the orbit of the center of mass of the rotating tether system, respectively.

[0023] In one embodiment, the v tar,r tar ,v GTO ,r GTO They are:

[0024]

[0025] Where μ is the gravitational constant of the Earth, ||r GTO ||,||r tar || are the modulus of the position vectors of the rotating tether system center of mass orbit and the target load orbit, respectively, Ω GTO ,ω GTO ,i GTO ,θ GTO , e GTO are the right ascension of the ascending node, argument of perigee, orbit inclination, true anomaly and eccentricity of the transfer orbit, Ω tar ,ω tar ,i tar ,θ tar , e tar They are the right ascension of the ascending node, argument of perigee, orbit inclination, true anomaly and eccentricity of the target payload orbit.

[0026] In one embodiment, in step S4, the total mass m of the tether is t for:

[0027]

[0028] Among them, V t is the expected velocity of the load relative to the center of mass of the rotating tether system during casting, V tc is the ultimate breaking rate of the rotating tether, erf is the error function, m pay The quality of the catch / cast mechanism.

[0029] In one embodiment, in step S4, the tether length l is:

[0030]

[0031] Where m is the total mass of the rotating tether system, m1 is the mass of the mother spacecraft, and m t is the total mass of the rotating tether, l2 is the distance of the sub-spacecraft from the center of mass of the rotating tether system;

[0032] l2 is the relative position vector (ΔP) between the end-body satellites required by the payload C Modulus length: l2=||(ΔP) C ||;

[0033] The total mass m of the rotating tether system is:

[0034] m=m1+m2+m t ;

[0035] Among them, m1 is the mass of the mother spacecraft, m2 is the mass of the daughter spacecraft, and m t is the total mass of the rotating tether.

[0036] In one embodiment, in step S5, the relative position vector (ΔP) between the end-body satellites required by the payload during the jettison is C and velocity vector (ΔU) C for:

[0037]

[0038] in,(*) C Represents the coordinate vector of vector * in the transfer orbit motion coordinate system Cxyz, L O is the inertial coordinate system OX from the Earth's equator eq Y eq Z eq The rotation matrix to the transfer orbit inertial coordinate system OXYZ, L θ is the rotation matrix from the transfer orbit inertial coordinate system OXYZ to the transfer orbit motion coordinate system OXYZ, is the orbital angular velocity of the transfer orbit, Δr p , Δv p are the position increment and velocity increment required for the payload to enter the out-of-plane target orbit from the transfer orbit, respectively. (*) eq Represents vector * in the Earth's equatorial inertial coordinate system OX eq Y eq Z eq The coordinate vector of .

[0039] In one embodiment, in step S5, the rotation surface normal vector (n s ) C for:

[0040]

[0041] Where (ΔP) C , (ΔU) C are the relative position vector and velocity vector between the end-body satellites required by the payload during jettisoning;

[0042] Casting rotation angular rate for:

[0043]

[0044] Where l2 is the distance between the sub-spacecraft and the center of mass of the rotating tether system;

[0045] Casting rotation angle Ψ s for:

[0046]

[0047] in, is the rotation plane direction angle λ s ,η s The corresponding rotation matrix, λ s is the angle between the projection of the normal vector of the casting rotation surface on the GTO orbit and the vertical line, η s is the angle between the normal vector of the casting rotation plane and the GTO orbital plane; (ΔP) C is the relative position vector between the end-body satellites; x ΔP ,y ΔP ,z ΔP (ΔP) C The three component coordinates of .

[0048] In one embodiment, in step S6, the rotating surface maneuvering strategy includes:

[0049] Casting rotation plane direction angle λ s Strategy:

[0050]

[0051] Where, θ is the true anomaly of the GTO orbit, are the first and second derivatives of the true anomaly of the GTO orbit; λ0 is the angle λ in the direction of the casting rotation plane s Initial value.

[0052] Angle η of the casting rotation plane s Strategy:

[0053]

[0054] Casting rotation angle ψ s Strategy:

[0055]

[0056] Where η0 is the angle η of the casting rotation plane s Initial value; ψ0 is the casting rotation angle ψ s Initial value; a is acceleration, Angle ψ s It went through two stages. For the rotation acceleration stage, the initial rotation speed Maneuver to cast rotation speed This is the rotation holding phase, maintaining the casting rotation speed

[0057] The beneficial effects of this application are:

[0058] The present application discloses a method for maneuvering a rotating tether system for transmitting a load between eccentric orbits. The method can utilize a rotating tether system to drop a load onto a target load orbit that is eccentric to the rotating tether system's center of mass orbit, thereby replacing the traditional pulse orbit change method. Compared with the traditional eccentric orbit transmission method, the eccentric orbit transmission process of the present application is reusable, more energy-efficient, and does not require the load to have the ability to change orbits. The present application has certain reference significance and value for the eccentric orbit transportation mission of space payloads. BRIEF DESCRIPTION OF THE DRAWINGS

[0059] Figure 1 This is a flow chart of the method for maneuvering the rotating tether system in the embodiment of the present application.

[0060] Figure 2 This is a schematic diagram of a rotating tether in an embodiment of the present application;

[0061] Figure 3 This is a schematic diagram of the throwing method in the embodiment of the present application;

[0062] Figure 4 The rotating tether system and the track surface in the embodiment of the present application;

[0063] Figure 5 The trajectory of the payload relative to the earth in the embodiment of the present application;

[0064] Figure 6 It is the maneuvering trajectory of the rotating surface in the embodiment of this application. DETAILED DESCRIPTION

[0065] The technical solutions of the present application will be described clearly and completely below in conjunction with specific embodiments. However, those skilled in the art will understand that the embodiments described below are part of the embodiments of the present application, not all of them, and are only used to illustrate the present application and should not be considered to limit the scope of the present application. Based on the embodiments in the present application, all other embodiments obtained by those of ordinary skill in the art without creative work are within the scope of protection of the present application.

[0066] The present application provides a method for maneuvering a rotating tether system for transmitting payloads between non-planar orbits, including: designing a rotating tether system structure; designing a method for throwing the payload based on the rotating tether system structure; designing the velocity increment and position increment required for throwing based on the throwing method of the rotating tether system; further calculating the mass and length of the tether based on the velocity increment and position increment required for throwing; designing the throwing rotation surface parameters based on the velocity increment and position increment required for throwing; and designing the rotation surface maneuvering strategy in the entire mission process based on the rotation surface parameters during throwing. Compared with previous co-planar transfer schemes, this method can successfully transport the payload from the initially inclined low-Earth orbit to the final equatorial geostationary orbit, achieving the non-planar orbit transmission mission of the payload. This method has significant reference significance and reference value for the orbital inclination maneuvering mission during satellite launch.

[0067] In one embodiment, a method for maneuvering a rotating tether system for load transmission between different planes is provided, referring to Figure 1 As shown, the specific steps include:

[0068] Step 1: Design the structure of the rotating tether system

[0069] Specifically, a double-body rotating tether system is constructed, including a mother spacecraft and a daughter spacecraft connected by a tether, wherein the daughter spacecraft includes a payload and a capture / throwing mechanism.

[0070] This application uses a double-body rotating tether system structure, using a relatively large mother spacecraft and a daughter spacecraft as the end satellites of the tether connection, and using a tether to connect the mother spacecraft and the daughter spacecraft. Assume that the masses of the mother spacecraft, daughter spacecraft, payload, and capture / throw mechanism are m1, m2, and m cap 、m pay , the sub-spacecraft is the mass of the payload, capture / ejection mechanism and:

[0071] m2=m cap +m pay ;

[0072] Reference Figure 2 As shown in the figure, the tether's tension is greatest at the system's center of mass, gradually decreases along the tension distribution elsewhere, and is smallest at the capture / delivery end. Therefore, the tether's diameter gradually decreases from the center of mass to both ends. The tether's design is based on its ultimate strength, safety factor, and density to ensure that it will not break under the expected conditions of use. The ultimate rate at which the tether does not break is:

[0073]

[0074] Where σ is the ultimate strength of the rotating tether, f is the safety factor of the rotating tether, and ρ is the density of the rotating tether.

[0075] Step 2: Design the method of dropping the load based on the structure of the rotating tether system.

[0076] Determine the method of dropping the payload, position the center of mass of the rotating tether system on the transfer orbit (GTO) where the rotating tether system operates, and position the sub-spacecraft at the target payload orbit position, where the target payload orbit is not in the same plane as the GTO.

[0077] Specifically, refer to Figure 3 As shown, during the casting, the mother spacecraft and the daughter spacecraft are respectively located on both sides of the center of mass of the rotating tether system. The daughter spacecraft carries the payload and lands on a GEO orbit that is out of plane with the GTO, realizing out-of-plane orbit transmission.

[0078] Step 3: Design the velocity increment and position increment required for casting based on the casting method of the rotating tether system.

[0079] Specifically, based on the orbital roots of the rotating tether system and the target payload orbit, the position increment Δr required for the payload to enter the out-of-plane target orbit from the transfer orbit is calculated. p and velocity increment Δv p .

[0080] The minimum orbital inclination of the GTO orbit depends on the latitude of the launch site of the rocket carrying the rotating tether system. Therefore, it is usually different from the target delivery orbit, and the rotating tether system needs to deliver the payload in an out-of-plane manner. Assume that the orbital root element ascending node right ascension, argument of perigee, orbital inclination, true anomaly, semi-major axis and eccentricity of the GTO orbit are:

[0081] Ω GTO ,ω GTO ,i GTO ,θ GTO ,a GTO ,e GTO .

[0082] The orbital root elements of the target payload orbit are the right ascension of the ascending node, argument of perigee, orbital inclination, true anomaly, semi-major axis and eccentricity, respectively:

[0083] Ω tar ,ω tar ,i tar ,θ tar ,a tar ,e tar .

[0084] The steps for calculating the velocity increment and position increment required for payload delivery specifically include: calculating the velocity and position vectors on the GTO orbit and the target payload orbit based on the orbital roots of the two orbits; based on the difference in velocity and position vectors, determining the motion position and velocity requirements of the payload relative to the center of mass of the rotating tether system, and then obtaining the velocity increment and position increment required for delivery.

[0085] Furthermore, the position increment Δr p and velocity increment Δv p for:

[0086] (Δr p ) eq =(r tar ) eq -(r GTO ) eq ,(Δv p ) eq =(v tar ) eq -(v GTO ) eq ;

[0087] in,(*) eq Represents vector * in the Earth's equatorial inertial coordinate system OX eq Y eq Z eq The coordinate vector, v tar ,r tar represent the velocity and position vectors on the target payload track, v GTO ,r GTO are the velocity and position vectors on the orbit of the center of mass of the rotating tether system, respectively.

[0088] v tar ,r tar ,v GTO ,r GTO They are:

[0089]

[0090] Where μ is the gravitational constant of the Earth, r GTO ,r tar are the moduli of the position vectors of the rotating tether system center of mass orbit and the target payload orbit, respectively.

[0091] Step 4: Calculate the structural parameters of the tether, including the tether length and total mass.

[0092] Specifically, the total mass of the tether is calculated based on the desired rate and error range required for payload delivery.

[0093] Total mass of the tether m t for:

[0094]

[0095] Among them, V t is the expected velocity of the load relative to the center of mass of the rotating tether system during casting, V tc is the ultimate breaking rate of the rotating tether, and erf is the error function.

[0096] Based on the total mass of the tether, the distance from the daughter spacecraft to the center of mass, and the distance from the mother spacecraft to the center of mass of the rotating tether system, the length of the rotating tether system is determined to ensure the strength and stability of the tether during the casting process.

[0097] Specifically, the length of the rotating tether system is l, and the distance l1 from the mother spacecraft to the center of mass is:

[0098]

[0099] Where m is the total mass of the rotating tether system, m1 is the mass of the mother spacecraft, and m t is the total mass of the rotating tether, l2 is the distance between the sub-spacecraft and the center of mass of the rotating tether system;

[0100] Specifically, l2 is the relative position vector (ΔP) between the end-body satellites required by the payload. C Modulus length: l2=||(ΔP) C ||;

[0101] The total mass m of the rotating tether system is:

[0102] m=m1+m2+m t ;

[0103] Among them, m1 is the mass of the mother spacecraft, m2 is the mass of the daughter spacecraft, and m t is the total mass of the rotating tether.

[0104] It is understandable that during the out-of-plane transmission process, the expected speed of the sub-spacecraft should be less than the limit speed of the tether without breaking: V t <V tc .

[0105] Step 5: The relative position vector (ΔP) between the end-body satellites required for payload delivery C and velocity vector (ΔU) C , design the casting rotation surface parameters. Ensure that the casting rotation surface parameters can guide the rotation operation during casting, so that the payload enters the target orbit at the expected speed and direction.

[0106] Specifically, the casting rotation surface parameters include the casting rotation surface normal vector (n s )C , casting rotation angular rate and casting rotation angle Ψ s .

[0107] The relative position vector (ΔP) between the end-body satellites required by the payload during jettisoning C and velocity vector (ΔU) C for:

[0108]

[0109] in,(*) C Represents the coordinate vector of vector * in the transfer orbit motion coordinate system Cxyz, L O is the inertial coordinate system OX from the Earth's equator eq Y eq Z eq The rotation matrix to the transfer orbit inertial coordinate system OXYZ, L θ is the rotation matrix from the transfer orbit inertial coordinate system OXYZ to the transfer orbit motion coordinate system OXYZ, is the orbital angular velocity of the transfer orbit.

[0110] Normal vector of the casting rotation plane (n s ) C Position vector (ΔP) between the end-body satellites required by the payload C and velocity vector (ΔU) C calculate:

[0111]

[0112] Among them, n sx ,n sy ,n sz is the normal vector of the casting rotation surface (n s ) C Three coordinate components.

[0113] According to the normal vector of the throwing rotation surface (n s ) C Calculate the angle λ of the corresponding casting rotation plane from the GTO orbit to the target payload orbit s ,η s ,like Figure 4 As shown, the rotation plane direction angle λ s is the normal vector of the casting rotation surface (n s ) C The angle between the GTO orbit projection and the vertical line, the rotation plane direction angle η s is the normal vector of the casting rotation surface (n s ) C Angle with GTO track surface:

[0114]

[0115] Calculating the angular velocity of a cast for:

[0116]

[0117] Where l2 is the distance between the sub-spacecraft and the center of mass of the rotating tether system.

[0118] The position vector relative to the center of mass (ΔP) C Calculate the rotation angle ψ s :

[0119]

[0120] in, is the rotation plane direction angle λ s ,η s The corresponding rotation matrix, (ΔP) C is the relative position vector between the end-body satellites, x ΔP ,y ΔP ,z ΔP (ΔP) C The three component coordinates of .

[0121] Step 6: Based on the casting rotation surface parameters, design the rotation surface maneuvering strategy in the entire mission process.

[0122] Specifically, according to the design of the rotation plane maneuvering strategy based on the casting rotation plane parameters, the system is maneuvered from the initial rotation plane to the casting rotation plane by controlling the rotation plane angle and the rotation angle, and the position increment Δr is satisfied at the casting moment. p and velocity increment Δv p Requirements are met to achieve load transmission on different tracks.

[0123] Reference Figure 4 As shown, the maneuver process starts at the initial time t0 and ends at the final time t s End, and throw the payload out. λ0 is the angle λ of the throwing rotation plane at the beginning of the maneuver s The initial value of η0 is the angle η of the casting rotation plane at the beginning of the maneuver. s The initial value of ψ0 is the casting rotation angle ψ at the beginning of the maneuver s The initial value of are the derivatives of the above angles respectively; are respectively the casting rotation angle ψ of the casting rotation plane s and azimuth λ s ,η s The derivative of .

[0124] The maneuvering strategy from the initial rotation plane to the casting rotation plane is as follows:

[0125] Casting rotation plane direction angle λ s The strategy is:

[0126]

[0127] Where, θ is the true anomaly of the GTO orbit, are the first and second derivatives of the true anomaly of the GTO orbit, This ensures that there is no need to overcome the influence of orbital Coriolis force during maneuvers.

[0128] Angle η of the casting rotation plane s The strategy is:

[0129]

[0130] To ensure that the angle η during the maneuver s From the initial angle η0, Maneuver to the final angle η s ,

[0131] Casting rotation angle ψ s The strategy is:

[0132]

[0133] Where a is the acceleration, Angle ψ s It went through two stages. For the rotation acceleration stage, the initial rotation speed Maneuver to cast rotation speed This is the rotation holding phase, maintaining the casting rotation speed

[0134] Finally, we need to ensure that the rotation angle is the same as the rotation angle when throwing:

[0135]

[0136] Angle λ s , angle η s 、Rotation angle ψ s The strategy ensures that the rotating tether system can maneuver to the casting rotation plane and rotation position, and then provide accurate speed and position increments for the payload during casting, so that the payload can smoothly enter the target payload track.

[0137] Example

[0138] Step 1: Design the structure of the rotating tether system.

[0139] In this specific embodiment, a double-body rotating tether system structure is selected, and the masses of the mother spacecraft, payload, and capture / throwing mechanism are m1=100000 kg, m cap =50kg,m pay =300kg, then the total mass m2 of the sub-spacecraft and the rotating tether system is:

[0140] m2=m cap +m pay =350kg.

[0141] Dyneema material is used as the tether material. The rope length of the rotating tether system is l = 15000m, the safety factor is set to f = 1.75, the ultimate strength of the tether is σ = 1GPa, and the density is ρ = 0.99g / cm 3 , then the limiting velocity without the tether breaking is:

[0142]

[0143] Step 2: Design the method of dropping the load based on the structure of the rotating tether system.

[0144] When carrying out-of-plane payload transmission to GEO orbit, the center of mass of the rotating tether system with the target payload orbit being GEO orbit falls on the GTO orbit, the mother spacecraft and the daughter spacecraft are located on both sides of the center of mass respectively, and the daughter spacecraft falls on the GEO orbit.

[0145] Step 3: Design the velocity increment and position increment required for casting based on the casting method of the rotating tether system.

[0146] The selected GTO orbit corresponds to the Wenchang rocket launch site, and the orbital elements are:

[0147]

[0148] The payload target orbit is the GEO orbit, and the orbital root element is:

[0149]

[0150] According to the orbital roots of the rotating tether system orbit and the GEO orbit, the velocity v on the payload target orbit and the rotating tether system center of mass orbit tar ,v GTO and position vector r tar ,r GTO (unit: m and m / s):

[0151]

[0152] Velocity increment Δv required for the load p and position increment Δrp (unit: m and m / s):

[0153] (Δr p ) eq =[1500000] T ,(Δv p ) eq =[0-1248626] T .

[0154] The relative position vector (ΔP) between the end-body satellites required by the payload C and velocity vector (ΔU) C for:

[0155]

[0156] Among them, the rotation matrix L from the Earth equatorial inertial coordinate system to the GTO orbit inertial coordinate system is O , the rotation matrix L from the GTO orbit inertial coordinate system to the GTO orbit motion coordinate system θ and the orbital angular velocity of the GTO track for:

[0157]

[0158] Step 4: Calculate the total mass of the tether based on the expected rate required for payload delivery, and calculate the tether length based on the total mass of the tether and the distance between the sub-spacecraft and the center of mass of the rotating tether system.

[0159] The expected speed of the sub-spacecraft is:

[0160] V t =||(ΔU) C ||=1396m / s.

[0161] The total mass of the tether is:

[0162]

[0163] The total mass of the rotating tether system is the load:

[0164] m=m1+m2+m t =100820.0866kg.

[0165] The distance l2 from the sub-spacecraft to the center of mass is the relative position vector (ΔP) between the end-body satellites. C Modulus length:

[0166] l2=||(ΔP) C ||=15000m.

[0167] The length of the rotating tether system is l, and the distance l2 from the mother spacecraft to the center of mass is:

[0168]

[0169] During the out-of-plane transmission process, the expected speed of the sub-spacecraft must be less than the limit speed without the tether breaking:

[0170] V t =1396m / s<V tc =1861m / s.

[0171] Step 5: Design the parameters of the jettisoning rotation surface based on the relative position vector and velocity vector between the end-body satellites required for payload jettisoning.

[0172] Casting rotation plane vector (n s ) C for:

[0173]

[0174] According to the normal vector of the throwing rotation plane (n s ) C Calculate the casting rotation angle λ s ,η s :

[0175]

[0176] Calculate the angular rate of rotation of the casting rotation surface for:

[0177]

[0178] The position vector relative to the center of mass (ΔP) C Calculate the rotation angle ψ of the casting rotation plane s for:

[0179]

[0180] Step 6: Based on the casting rotation surface parameters, design the rotation surface maneuvering strategy in the entire mission process.

[0181] Starting from the initial time t0=0s, to the final time t s =6.6×10 5 s ends, and the payload is thrown out at the final moment.

[0182] The initial rotation surface parameters are:

[0183]

[0184] Casting rotation plane direction angle λ sThe strategy is:

[0185]

[0186] Angle η of the casting rotation plane s The strategy is:

[0187]

[0188] Casting rotation angle ψ s The strategy is:

[0189]

[0190] like Figure 5-6 As shown in the figure, the motion trajectory of the daughter spacecraft and the mother spacecraft relative to the center of mass of the rotating tether system gradually maneuvers from the initial rotation plane to the rotation plane at the time of jettison. The motion of the payload before jettison is composed of orbital motion on the GTO orbit and rotational motion around the center of mass of the rotating tether system. At the jettison position at apogee, the payload can smoothly enter the GEO orbit and orbit along the GEO orbit. The semi-major axis of the payload after jettison is a p =36000km, orbital inclination i p = 0rad and orbital angular momentum h p =[0.01.2996×10 11 ] T , which indicates that the orbital root element of the payload's orbit after being dropped is the same as that of the GEO orbit.

[0191] It can be seen from this specific embodiment that the maneuvering method of a rotating tether system for transmitting loads between non-planar tracks disclosed in this application can successfully achieve the non-planar transmission task of the payload and has strong application value.

[0192] Although the embodiments of the present application are described above in conjunction with the accompanying drawings, the present application is not limited to the above-mentioned specific embodiments and application fields. The above-mentioned specific embodiments are merely illustrative and instructive, and not restrictive. A person of ordinary skill in the art, guided by this specification and without departing from the scope of protection of the claims of this application, may also devise various forms, all of which fall within the scope of protection of this application.

Claims

1. A method for maneuvering a rotating tether system for load transmission between non-planar tracks, characterized in that: include: S1: Construct a two-body rotating tether system, including a mother spacecraft and a daughter spacecraft connected by a tether, wherein the daughter spacecraft includes a payload and a capture / ejection mechanism; S2: On the transfer orbit in which the rotating tether system operates, positioning the center of mass of the rotating tether system on the transfer orbit, positioning the sub-spacecraft at a target payload orbit position, and the transfer orbit being out of plane with the target payload orbit; S3: Calculate the position increment Δr required for the payload to move from the transfer orbit to the out-of-plane target orbit based on the orbital roots of the rotating tether system and the target payload orbit. p and velocity increment Δv p ; S4: Calculate the total mass of the tether based on the expected rate required for payload delivery, and calculate the tether length based on the total mass of the tether and the distance between the sub-spacecraft and the center of mass of the rotating tether system; S5: Calculate the parameters of the jettisoning rotation surface according to the relative position vector and velocity vector between the end-body satellites required for payload jettisoning, and the jettisoning rotation surface parameters include the jettisoning rotation surface normal vector (n s ) C , casting rotation angular rate and casting rotation angle Ψ s ; S6: Design a rotational surface maneuvering strategy based on the casting rotational surface parameters, and control the rotational surface angle and the rotation angle to make the system maneuver from the initial rotational surface to the casting rotational surface, and satisfy the position increment Δr at the casting time. p and velocity increment Δv p Requirements, to achieve load transmission on non-planar tracks.

2. The method for maneuvering a rotating tether system according to claim 1, wherein: The tether limit breaking rate V tc satisfy: Where σ is the ultimate strength of the tether, f is the safety factor of the tether, and ρ is the tether density.

3. The method for maneuvering a rotating tether system according to claim 1, wherein: In step S3, the position increment Δr p and velocity increment Δv p for: (Δr p ) eq =(r tar ) eq -(r GTO ) eq ,(Δv p ) eq =(v tar ) eq -(v GTO ) eq ; in,(*) eq Represents vector * in the Earth's equatorial inertial coordinate system OX eq Y eq Z eq The coordinate vector, v tar ,r tar Denote the velocity and position vectors of the payload target on the track, v GTO ,r GTO are the velocity and position vectors on the orbit of the center of mass of the rotating tether system, respectively.

4. The method for maneuvering a rotating tether system according to claim 3, wherein: The v tar ,r tar ,v GTO ,r GTO They are: Where μ is the gravitational constant of the Earth, ||r GTO ||,||r tar || are the modulus of the position vectors of the rotating tether system center of mass orbit and the payload target orbit, respectively, Ω GTO ,ω GTO ,i GTO ,θ GTO , e GTO are the right ascension of the ascending node, argument of perigee, orbit inclination, true anomaly and eccentricity of the transfer orbit, Ω tar ,ω tar ,i tar ,θ tar , e tar They are respectively the right ascension of the ascending node, argument of perigee, orbit inclination, true anomaly and eccentricity of the payload target orbit.

5. The method for maneuvering a rotating tether system according to claim 1, wherein: In step S4, the total mass m of the tether is t for: Among them, V t is the expected velocity of the load relative to the center of mass of the rotating tether system during casting, V tc is the ultimate breaking rate of the rotating tether, erf is the error function, m pay The quality of the catch / cast mechanism.

6. The method for maneuvering a rotating tether system according to claim 1, wherein: In step S4, the tether length l is: Where m is the total mass of the rotating tether system, m1 is the mass of the mother spacecraft, and m t is the total mass of the rotating tether, and l2 is the distance between the sub-spacecraft and the center of mass of the rotating tether system.

7. The method for maneuvering a rotating tether system according to claim 6, wherein: l2 is the relative position vector (ΔP) between the end-body satellites required by the payload C Modulus length: l2=||(ΔP) C ||; The total mass m of the rotating tether system is: m=m1+m2+m t 。 8. The method for maneuvering a rotating tether system according to claim 1, wherein: In step S5, the relative position vector (ΔP) between the end-body satellites required for the payload to be dropped is C and velocity vector (ΔU) C for: in,(*) C Represents the coordinate vector of vector * in the transfer orbit motion coordinate system Cxyz, L O is the inertial coordinate system OX from the Earth's equator eq Y eq Z eq The rotation matrix to the transfer orbit inertial coordinate system OXYZ, L θ is the rotation matrix from the transfer orbit inertial coordinate system OXYZ to the transfer orbit motion coordinate system OXYZ, is the orbital angular velocity of the transfer orbit, Δr p , Δv p are the position increment and velocity increment required for the payload to enter the out-of-plane target orbit from the transfer orbit, respectively. (*) eq Represents vector * in the Earth's equatorial inertial coordinate system OX eq Y eq Z eq The coordinate vector of .

9. The method for maneuvering a rotating tether system according to claim 1, wherein: In step S5, the normal vector of the casting rotation surface (n s ) C for: Where (ΔP) C , (ΔU) C are the relative position vector and velocity vector between the end-body satellites required by the payload during jettisoning; Casting rotation angular rate for: Where l2 is the distance between the sub-spacecraft and the center of mass of the rotating tether system; Casting rotation angle Ψ s for: in, is the rotation plane direction angle λ s ,η s The corresponding rotation matrix, λ s is the angle between the projection of the normal vector of the casting rotation surface on the GTO orbit and the vertical line, η s is the angle between the normal vector of the casting rotation plane and the GTO orbital plane; (ΔP) C is the relative position vector between the end-body satellites; x ΔP ,y ΔP ,z ΔP (ΔP) C The three component coordinates of .

10. The method for maneuvering a rotating tether system according to claim 9, wherein: In step S6, the rotating surface maneuvering strategy includes: Casting rotation plane direction angle λ s Strategy: Where, θ is the true anomaly of the GTO orbit, are the first and second derivatives of the true anomaly of the GTO orbit; λ0 is the angle λ in the direction of the casting rotation plane s Initial value; Angle η of the casting rotation plane s Strategy: Casting rotation angle ψ s Strategy: Where η0 is the angle η of the casting rotation plane s Initial value; ψ0 is the casting rotation angle ψ s Initial value; a is acceleration, Angle ψ s It went through two phases; For the rotation acceleration stage, the initial rotation speed Maneuver to cast rotation speed This is the rotation holding phase, maintaining the casting rotation speed

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