Rotating tether system maneuvering method for out-of-plane inter-orbit payload transfer

Abstract

The application discloses a rotating tether system maneuvering method for out-of-plane inter-orbit load transmission, which systematically designs a load throwing mechanism by constructing a double-body rotating tether system framework. Firstly, a load release mode is planned based on the characteristics of the double-body rotating tether system, and then the required velocity increment and position increment parameters for realizing the release mode are derived. By calculating the total mass and length of the tether, and combining the relative position vector and velocity vector between the end-body satellites required for load throwing, the key parameters of the rotating plane for throwing are accurately calculated. Finally, the rotating plane maneuvering control strategy for the whole task cycle is constructed based on the above parameters. Compared with the traditional coplanar orbit transfer scheme, the application breaks through the out-of-plane orbit transmission task of effectively loading from the initial inclined low earth orbit to the target equatorial geostationary orbit. In the orbit inclination adjustment maneuvering task in the satellite launching stage, it has significant technical reference value and engineering application prospect.

Description

Technical Field

[0001] This application belongs to the field of aerospace technology, specifically relating to a method for maneuvering a rotating tether system for load transfer between skewed orbits. Background Technology

[0002] In the aerospace field, payload transportation is a critical mission. Traditional methods often 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 spacecraft that connects two spacecraft via a tether to jointly complete on-orbit missions, it possesses unique advantages. Its excellent centrifugal stability and momentum exchange characteristics provide a reusable solution for on-orbit payload transportation. Therefore, the method of payload transfer using a rotating tether system has attracted widespread attention and research, demonstrating broad development prospects.

[0003] Spinning tether systems require no fuel during momentum exchange and can deliver payloads instantaneously. Compared to traditional rocket pulse acceleration, they represent a feasible alternative for transporting payloads to different orbits. Lorenzini first proposed a complete scheme for payload transfer using spinning tether systems in his 2000 paper, "Mission analysis of spinning systems for transfers from low orbits to geostationary." This design enabled a spinning tether system to deliver a 21,400 kg payload into GEO orbit within two years, consuming only 5,380 kg of propellant. The payload-to-disposable equipment mass ratio was significantly lower than that of traditional rocket payloads, laying the foundation for research on spinning tether system payload transfer.

[0004] Subsequently, researchers continued to delve deeper, conducting more detailed studies on orbital transfer using rotating tethered systems. The 2016 publication, "Optimal Control of payload tossing using space tethered system," presented a scheme design for multi-stage payload transfer using a rotating tethered system between coplanar elliptical orbits and the optimal control law for the drop phase, effectively avoiding the impact of the elliptical orbit 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 anti-collision and anti-breakage technologies during the drop process.

[0005] However, current technologies for load transfer using rotating tether systems are mostly focused on load transfer tasks within the track plane. That is, previous research has largely centered on coplanar transport schemes, where the rotation plane coincides with the track plane, and the drop process does not provide a velocity increment perpendicular to the track plane, resulting in the payload reaching a track coplanar with the rotating tether system's track. However, in practical applications, the task of transferring loads using rotating tether systems to tracks that are out of plane from the rotating tether system's center of mass is a real challenge that remains unresolved and is a pressing problem to be solved in this field. Summary of the Invention

[0006] To address the limitations of existing technologies that only allow for coplanar transport, this application provides a rotating tether system maneuvering method for payload transfer between disparate orbits. Compared to traditional coplanar transfer schemes, this application can successfully transport the payload from an initially inclined low Earth orbit to a final equatorial geostationary orbit, achieving the payload transfer mission between disparate orbits. This method has significant reference value and implications for orbital inclination maneuvering missions during satellite launches.

[0007] To achieve the technical objectives of this application, the following technical solutions are mainly adopted:

[0008] In one aspect of this application, a method for maneuvering a rotating tethered system for load transfer between skewed tracks is provided, comprising the following steps:

[0009] S1: Construct a twin-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 / drop mechanism;

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

[0011] S3: Based on the orbital elements of the rotating tether system and the target load track, calculate the position increment Δr required for the load to move from the transfer track to the non-planar target track. p and velocity increment Δv p ;

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

[0013] S5: Calculate the launch rotation surface parameters based on the relative position vectors and velocity vectors between the end-body satellites required for payload launch. The launch 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 plane maneuvering strategy based on the throwing rotational plane parameters. By controlling the rotational plane angle and rotation angle, the system maneuvers from the initial rotational plane to the throwing rotational plane, and satisfies the position increment Δr at the throwing moment. p and velocity increment Δv p The requirement is to achieve load transfer 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's ultimate 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 density of the tether.

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

[0020] In one implementation, 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 This represents the vector * in the Earth's equatorial inertial coordinate system OX. eq Y eq Z eq The coordinate vector, v tar ,r tar Let v represent the velocity and position vectors on the target payload's trajectory, respectively. GTO ,r GTO These represent the velocity and position vectors on the orbit of the center of mass of the rotating tethered system, respectively.

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

[0024]

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

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

[0027]

[0028] Among them, V t V is the desired velocity of the relative rotation of the tether system's center of mass required for the load during throwing. tc Let m be the ultimate breaking rate of the rotating tether, erf be the error function, and m be the ultimate breaking rate of the rotating tether. pay The quality of the capture / throwing mechanism.

[0029] In one embodiment, in step S4, the length l of the tether 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 Let l1 be the total mass of the rotating tether, and l2 be the distance between the subspacecraft and 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] Where m1 is the mass of the mother spacecraft, m2 is the mass of the daughter spacecraft, and m t This is the total mass of the rotating rope.

[0036] In one implementation, in step S5, the relative position vector (ΔP) between the end-body satellites required for the payload during launch is determined. C And velocity vector (ΔU) C for:

[0037]

[0038] in,(*) C L represents the coordinate vector of vector * in the coordinate system Cxyz of the transfer orbit. O From the Earth's equatorial inertial coordinate system OX eq Y eq Z eq The rotation matrix L to the transfer orbit inertial coordinate system OXYZ θ Let be the rotation matrix from the transfer orbit inertial coordinate system OXYZ to the transfer orbit motion coordinate system OXYZ. For the orbital angular velocity of the transfer orbit, Δr p Δv p These represent the position increment and velocity increment required for the load to move from the transfer orbit to the target orbit in a different plane, respectively. (*) eq This represents the vector * in the Earth's equatorial inertial coordinate system OX. eq Y eq Z eq The coordinate vector.

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

[0040]

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

[0042] Throwing rotation angular rate for:

[0043]

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

[0045] Throwing rotation angle Ψ s for:

[0046]

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

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

[0049] Projectile rotation direction angle λ s Strategy:

[0050]

[0051] Where θ is the true anomaly angle of the GTO orbit. λ0 represents the first and second derivatives of the true anomaly angle of the GTO orbit; λ0 represents the angle λ0 of the launch rotation plane. s Initial value.

[0052] Projectile rotation direction angle η s Strategy:

[0053]

[0054] Projection rotation angle ψ s Strategy:

[0055]

[0056] Where η0 is the angle η of the launching rotation plane. s Initial value; ψ0 is the launching rotation angle ψ s Initial value; 'a' is acceleration. Angle ψ s It went through two phases. For the rotation acceleration phase, the initial rotation speed is... maneuver to throwing rotation speed During the spin-holding phase, maintain the casting spin speed.

[0057] The beneficial effects of this application are as follows:

[0058] This application presents a rotating tether system maneuvering method for load transfer between non-planar orbits. This method utilizes a rotating tether system to project a load onto a target load orbit that is non-planar to the orbit of the rotating tether system's center of mass, thus replacing the traditional pulse orbit-changing method. Compared to traditional non-planar orbit transfer methods, this application's non-planar orbit transfer process is reusable, more energy-efficient, and does not require the load to have orbit-changing capabilities. This application has certain reference value and significance for the non-planar orbit transportation of space loads. Attached Figure Description

[0059] Figure 1 This is a flowchart illustrating the motorization method of the rotating tether system in the embodiments of this application.

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

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

[0062] Figure 4 The rotating tether system and the track surface are shown in the embodiments of this application;

[0063] Figure 5 This is the trajectory of the payload relative to the Earth in the embodiments of this application;

[0064] Figure 6 This is the motion trajectory of the rotating surface in the embodiments of this application. Detailed Implementation

[0065] The technical solution of this application will be clearly and completely described below with reference to specific embodiments. However, those skilled in the art will understand that the embodiments described below are only some embodiments of this application, not all embodiments, and are only used to illustrate this application, and should not be regarded as limiting the scope of this application. Based on the embodiments in this application, all other embodiments obtained by those skilled in the art without creative effort are within the scope of protection of this application.

[0066] This application provides a maneuvering method for a rotating tether system for payload transfer between non-planar orbits, comprising: designing the rotating tether system structure; designing the payload delivery method based on the rotating tether system structure; designing the required velocity and position increments for delivery based on the delivery method; further calculating the mass and length of the tether based on the required velocity and position increments; designing the delivery rotation plane parameters based on the required velocity and position increments; and designing the rotation plane maneuvering strategy throughout the mission based on the rotation plane parameters during delivery. Compared to previous coplanar transfer schemes, this method can successfully transport the payload from an initially inclined low Earth orbit to a final equatorial geostationary orbit, achieving the non-planar orbit payload transfer mission. This method has significant reference value and implications for orbital inclination maneuvering missions during satellite launches.

[0067] In one embodiment, a method for maneuvering a rotating tethered system for load transfer between skewed tracks is provided, referring to... Figure 1 As shown, the specific steps include:

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

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

[0070] This application selects a two-body rotating tethered system structure, using a relatively large mother spacecraft and daughter spacecraft as the terminal satellites of the tethered connection, and employing a tether to connect the mother spacecraft and daughter spacecraft. Let the masses of the mother spacecraft, daughter spacecraft, payload, and capture / drop mechanism be m1, m2, and m3, respectively. cap m pay The sub-spacecraft consists of the mass of its payload, capture / drop mechanism, and:

[0071] m2=m cap +m pay ;

[0072] Reference Figure 2 As shown, the tether's tension is greatest at the system's center of mass and gradually decreases along the tension distribution elsewhere, with its diameter being 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 it does not break under the intended use conditions. The limiting 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: Based on the structure of the rotating rope system, design the method for throwing the load.

[0076] The method of payload release is determined. On the transfer orbit (GTO) of the rotating tether system, the center of mass of the rotating tether system is positioned on the transfer orbit, and the sub-spacecraft is positioned at the target payload orbit position, which is out of plane from the GTO.

[0077] Specifically, refer to Figure 3 As shown, during the launch, the mother spacecraft and the daughter spacecraft are located on opposite 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 not in the same plane as the GTO, thus achieving cross-plane orbital transfer.

[0078] Step 3: Based on the throwing method of the rotating tether system, design the required speed increment and position increment for throwing.

[0079] Specifically, based on the orbital elements of the rotating tether system and the target load orbit, the position increment Δr required for the load to move from the transfer orbit to the non-planar target orbit is calculated. p and velocity increment Δv p .

[0080] The minimum orbital inclination of a GTO orbit depends on the launch site latitude of the rocket carrying the rotating tether system. Therefore, unlike the target payload transfer orbit, the rotating tether system typically requires out-of-plane payload transfer. Assume the right ascension of the ascending node, argument of perigee, orbital inclination, true anomaly, semi-major axis, and eccentricity of the GTO orbit are as follows:

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

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

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

[0084] The steps for calculating the velocity and position increments required for load release specifically include: calculating the velocity and position vectors on the GTO track and the target load track based on the track elements; determining the motion position and velocity requirements of the load relative to the center of mass of the rotating tether system based on the difference between the velocity and position vectors, and thus obtaining the velocity and position increments required for release.

[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 This represents the vector * in the Earth's equatorial inertial coordinate system OX. eq Y eq Z eq The coordinate vector, v tar ,r tar Let v represent the velocity and position vectors on the target payload's trajectory, respectively. GTO ,r GTO These represent the velocity and position vectors on the orbit of the center of mass of the rotating tethered system, respectively.

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

[0089]

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

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

[0092] Specifically, the total mass of the tether is calculated based on the expected rate and error range required for load throwing.

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

[0094]

[0095] Among them, V t V is the desired velocity of the relative rotation of the tether system's center of mass required for the load during throwing. tc Let be the ultimate breaking rate of the rotating tether, and erf be the error function.

[0096] Based on the total mass of the tether, the distance from the sub-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 deployment 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 Let l1 be the total mass of the rotating tether, and l2 be the distance between the subspacecraft 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] Where m1 is the mass of the mother spacecraft, m2 is the mass of the daughter spacecraft, and m t This is the total mass of the rotating rope.

[0104] Understandably, during heterogeneous transport, the expected velocity of the sub-spacecraft should be less than the limiting velocity at which the tether will not break: V t <V tc .

[0105] Step 5: Determine the relative position vector (ΔP) between the end-body satellites required for payload delivery. C And velocity vector (ΔU) C Design the launch rotation parameters. Ensure that the launch rotation parameters can guide the rotation operation during launch, so that the load enters the target trajectory at the expected speed and direction.

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

[0107] The relative position vector (ΔP) between the end-body satellites required for payload delivery. C And velocity vector (ΔU) C for:

[0108]

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

[0110] Throwing the normal vector of the rotating surface (n) s ) C The position vector (ΔP) between the end-body satellites required by the payload. C And velocity vector (ΔU) C calculate:

[0111]

[0112] Where, n sx ,n sy ,n sz For the normal vector of the launching rotation surface (n) s ) C Three coordinate components.

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

[0114]

[0115] Calculate the throwing rotation rate for:

[0116]

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

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

[0119]

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

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

[0122] Specifically, based on the parameters of the launching rotation plane, a rotation plane maneuvering strategy is designed. By controlling the rotation plane angle and rotation angle, the system maneuvers from the initial rotation plane to the launching rotation plane, and satisfies the position increment Δr at the launching moment. p and velocity increment Δv p The requirement is to achieve load transfer on non-planar tracks.

[0123] Reference Figure 4 As shown, the maneuver process begins at the initial time t0 and ends at the final time t0. s The process ends, and the load is ejected. λ0 is the angle λ of the ejection rotation plane at the start of the maneuver. s The initial value, η0, is the angle η of the launching rotation plane at the start of the maneuver. s The initial value, ψ0, is the launching rotation angle ψ at the start of the maneuver. s initial value, These are the derivatives of the angles mentioned above; The launching rotation angle ψ of the launching rotation plane is respectively. s and azimuth λ s ,η s The derivative of .

[0124] The maneuvering strategy from the initial rotating plane maneuver to the jetting rotating plane is as follows:

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

[0126]

[0127] Where θ is the true anomaly angle of the GTO orbit. Let the first and second derivatives be the true anomaly angles of the GTO orbit. This ensures that the effects of the Coriolis force on the track do not need to be overcome during the maneuver.

[0128] Projectile rotation direction angle η s The strategy is:

[0129]

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

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

[0132]

[0133] Where a is acceleration, Angle ψ s It went through two phases. For the rotation acceleration phase, the initial rotation speed is... maneuver to throwing rotation speed During the spin-holding phase, maintain the casting spin speed.

[0134] Ultimately, the rotation angle must be the same as the rotation angle at the time of 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 throwing rotation plane and position, thereby providing the load with accurate speed and position increments during throwing, so that the load can smoothly enter the target load track.

[0137] Example

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

[0139] This specific embodiment selects a twin-body rotating tether system structure, with the mass of the mother spacecraft, payload, and capture / drop mechanism being m1 = 100,000 kg, m cap =50kg, m pay =300kg, then the subspacecraft mass m2 and the total mass m of the rotating tether system are:

[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 = 1 GPa, and the density is ρ = 0.99 g / cm³. 3 Then the limiting speed at which the rope will not break is:

[0142]

[0143] Step 2: Based on the structure of the rotating rope system, design the method for throwing the load.

[0144] During the non-plane transfer of payloads to GEO orbit, the center of mass of the rotating tether system with the target payload orbiting GEO orbit falls on the GTO orbit, with the mother spacecraft and the daughter spacecraft located on opposite sides of the center of mass, and the daughter spacecraft falling on the GEO orbit.

[0145] Step 3: Based on the throwing method of the rotating tether system, design the required speed increment and position increment for throwing.

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

[0147]

[0148] The target orbit for the payload is a GEO orbit, and the orbital elements are:

[0149]

[0150] Based on the orbital elements of the rotating tethered system track and the GEO track, the velocities v on the target load track and the center of mass track of the rotating tethered system are... tar ,v GTO and position vector r tar ,r GTO (Units: m and m / s):

[0151]

[0152] The required velocity increment Δv for the load p and position increment Δrp (Units: 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 orbital inertial coordinate system O The rotation matrix L from the GTO orbital inertial coordinate system to the GTO orbital motion coordinate system. θ Orbital angular velocity of GTO orbit for:

[0157]

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

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

[0160] V t =||(ΔU) C ||=1396m / s.

[0161] The total mass of the rope 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 heterogeneous transport, the desired velocity of the sub-spacecraft must be less than the tether-free velocity limit.

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

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

[0172] Throwing rotating surface vector (n) s ) C for:

[0173]

[0174] Based on the normal vector (n) of the launching rotation plane s ) C Calculate the direction angle λ of the launching rotation plane. s ,η s :

[0175]

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

[0177]

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

[0179]

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

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

[0182] The initial parameters of the rotating surface are:

[0183]

[0184] Projectile rotation direction angle λ sThe strategy is:

[0185]

[0186] Projectile rotation direction angle η s The strategy is:

[0187]

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

[0189]

[0190] like Figure 5-6 As shown, the trajectories of the sub-spacecraft and the parent spacecraft relative to the center of mass of the rotating tether system gradually maneuver from the initial plane of rotation to the plane of rotation at the time of release. The pre-release payload motion consists of orbital motion in the GTO orbit and rotational motion around the center of mass of the rotating tether system. At the release position at the apogee, the payload smoothly enters the GEO orbit and performs orbital motion along the GEO orbit. The semi-major axis of the payload after release is a. p =36000km, orbital inclination i p =0 rad and orbital angular momentum h p = [001.2996×10 11 ] T This indicates that the orbital elements of the orbital trajectory after the payload is released are the same as those of the GEO orbital trajectory.

[0191] As can be seen from this specific embodiment, the rotating tether system maneuvering method for load transfer between different-plane tracks disclosed in this application can successfully achieve the task of transferring effective loads between different planes, and has strong application value.

[0192] Although the embodiments of this application have been described above in conjunction with the accompanying drawings, this application is not limited to the specific embodiments and application fields described above. The specific embodiments described above are merely illustrative and instructive, not restrictive. Those skilled in the art can make many other forms based on the guidance of this specification and without departing from the scope of protection of the claims of this application, and these are all within the scope of protection of this application.

Claims

1. A method for maneuvering a rotating tethered system for load transfer between skewed tracks, characterized in that, include: S1: Construct a twin-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 / drop mechanism; S2: On the transfer orbit in which the rotating tether system operates, the center of mass of the rotating tether system is positioned on the transfer orbit, the sub-spacecraft is positioned at the target payload orbit position, and the transfer orbit is out of plane with the target payload orbit; S3: Based on the orbital elements of the rotating tether system and the target load track, calculate the position increment Δr required for the load to move from the transfer track to the non-planar target track. p and velocity increment Δv p ; S4: Calculate the total mass of the tether based on the expected rate required for payload ejection, and calculate the tether length based on the total mass of the tether and the distance between the center of mass of the subspacecraft and the rotating tether system; S5: Calculate the launch rotation surface parameters based on the relative position vectors and velocity vectors between the end-body satellites required for payload launch. These parameters include the launch rotation surface normal vector (n). s ) C Throwing rotation angular rate and the throwing rotation angle Ψ s ; S6: Design a rotational plane maneuvering strategy based on the throwing rotational plane parameters. By controlling the rotational plane angle and rotation angle, the system maneuvers from the initial rotational plane to the throwing rotational plane, and satisfies the position increment Δr at the throwing moment. p and velocity increment Δv p The requirement is to achieve load transfer on non-planar tracks.

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

3. The method for maneuvering a rotating tethered system according to claim 1, characterized in that, 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 This represents the vector * in the Earth's equatorial inertial coordinate system OX. eq Y eq Z eq The coordinate vector, v tar ,r tar Let v represent the velocity and position vectors on the target orbit, respectively. GTO ,r GTO These represent the velocity and position vectors on the orbit of the center of mass of the rotating tethered system, respectively.

4. The method for maneuvering a rotating tethered system according to claim 3, characterized in that, The v tar ,r tar ,v GTO ,r GTO They are respectively: Where μ is the Earth's gravitational constant, ||r GTO ||,||r tar || represents the magnitudes of the position vectors of the center-of-mass orbit of the rotating tethered system and the target load orbit, respectively, Ω GTO ω GTO i GTO θ GTO e GTO These are the right ascension of the ascending node, argument of perigee, inclination, true anomaly, and eccentricity of the transfer orbit, respectively. tar ω tar i tar θ tar e tar These are the right ascension of the ascending node, argument of perigee, orbital inclination, true anomaly, and eccentricity of the target orbit, respectively.

5. The method for maneuvering a rotating tethered system according to claim 1, characterized in that, In step S4, the total mass m of the tethering rope t for: Among them, V t V is the desired velocity of the relative rotation of the tether system's center of mass required for the load during throwing. tc Let m be the ultimate breaking rate of the rotating tether, erf be the error function, and m be the ultimate breaking rate of the rotating tether. pay The quality of the capture / throwing mechanism.

6. The method for maneuvering a rotating tethered system according to claim 1, characterized in that, In step S4, the length l of the tether is: Where m is the total mass of the rotating tether system, m1 is the mass of the mother spacecraft, and m t Let l1 be the total mass of the rotating tether, and l2 be the distance between the subspacecraft and the center of mass of the rotating tether system.

7. The method for maneuvering a rotating tethered system according to claim 6, characterized in that, 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 tethered system according to claim 1, characterized in that, In step S5, the relative position vector (ΔP) between the end-body satellites required for payload delivery is determined. C And velocity vector (ΔU) C for: in,(*) C L represents the coordinate vector of vector * in the coordinate system Cxyz of the transfer orbit. O From the Earth's equatorial inertial coordinate system OX eq Y eq Z eq The rotation matrix L to the transfer orbit inertial coordinate system OXYZ θ Let be the rotation matrix from the transfer orbit inertial coordinate system OXYZ to the transfer orbit motion coordinate system OXYZ. For the orbital angular velocity of the transfer orbit, Δr p Δv p These represent the position increment and velocity increment required for the load to move from the transfer orbit to the target orbit in a different plane, respectively. (*) eq This represents the vector * in the Earth's equatorial inertial coordinate system OX. eq Y eq Z eq The coordinate vector.

9. The method for maneuvering a rotating tethered system according to claim 1, characterized in that, In step S5, the normal vector (n) of the throwing rotation surface s ) C for: Where, (ΔP) C ,(ΔU) C These are the relative position vector and velocity vector between the end-body satellites required for the payload during deployment; Throwing rotation angular rate for: Where l2 is the distance between the subspacecraft and the center of mass of the rotating tether system; Throwing rotation angle Ψ s for: in, Let λ be the direction angle of the plane of rotation. s ,η s The corresponding rotation matrix, λ s Let η be the angle between the projection of the normal vector of the launch rotation plane onto the GTO orbit and the vertical. s The angle between the normal vector of the launch rotation plane and the GTO orbital plane; (ΔP) C x is the relative position vector between the end-body satellites; ΔP ,y ΔP ,z ΔP For (ΔP) C The three component coordinates.

10. The method for maneuvering a rotating tethered system according to claim 9, characterized in that, In step S6, the rotating surface maneuvering strategy includes: Projectile rotation direction angle λ s Strategy: Where θ is the true anomaly angle of the GTO orbit. λ0 represents the first and second derivatives of the true anomaly angle of the GTO orbit; λ0 represents the angle λ0 of the launch rotation plane. s Initial value; Projectile rotation direction angle η s Strategy: Projection rotation angle ψ s Strategy: Where η0 is the angle η of the launching rotation plane. s Initial value; ψ0 is the launching rotation angle ψ s Initial value; 'a' is acceleration. Angle ψ s It went through two phases; For the rotation acceleration phase, the initial rotation speed is... maneuver to throwing rotation speed During the spin-holding phase, maintain the casting spin speed.

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