RU2759372C1 - Method for controlling the transport system when performing a flight to a high-energy orbit - Google Patents

Abstract

FIELD: space technology.SUBSTANCE: invention relates to the removal of space objects (SO) using upper stage rockets (USR) into high-energy orbits (for example, to the Moon) in several stages according to a two-launch scheme. The method involves putting the SO into a near-Earth orbit and docking with a near-Earth station (NES). The USR is placed separately from the SO into a near-Earth reference orbit, and then into a coplanar coelliptic orbit relative to the NES with a height difference of ΔН, at which a rapid approach of the spacecraft and the USR is performed, followed by their docking. With the help of the latter, an impulse is applied to the bundle of SO and USR to switch to a high-energy orbit. In case of failure of the launch of the USR, the SO is able to wait for the next launch window as part of the NES (for example, for departure to the Moon).EFFECT: increase in the reliability of the transport system to the level characteristic of a single-start scheme, due to the use of the NES.1 cl, 6 dwg

Description

The proposed control method can be used in space technology when approaching and docking a space object (SO), for example, a manned spacecraft (PC), with an upper stage (RB), launched into low-earth orbit, to perform maneuvers of a flight to high-energy orbits, for example, when flying to To the moon.

There is a known method for controlling the movement of an active spacecraft when approaching a passive spacecraft, for example, an orbital station (OS) [1. Murtazin R.F. A method for controlling the movement of a space object when approaching another space object. Patent RU No. 2657704 dated 06/14/2018], selected as an analogue, which provides convergence and docking in two turns. This method includes launching an active SC into an orbit, the plane of which intersects the orbital plane of a passive SC at an angle of 9, but with a given mismatch in the latitude argument ΔF. Then, in the vicinity of the point of intersection of the orbital planes of both spacecraft, a rendezvous impulse is performed, the characteristics of which are determined by the nominal parameters of the launching orbit, eliminating the misalignment of the planes. Then, using onboard measurements of the relative position of both VOs, a two-pulse maneuver is calculated for the transition of the active VO to the vicinity of the passive VO, where the approach ends with automatic docking. The urgent execution of the first rendezvous impulse, determined from the prelaunch calculations, makes it possible to reduce the rendezvous duration to two orbits.

FIG. 1 shows the extraction scheme used in the analogue, where: 1 - active KO, 2 - passive KO, 3 - the point of intersection of two planes at an angle of 9, where an impulse must be executed to eliminate the mismatch of the planes, 4 - the specified mismatch in the argument of latitude ΔF between objects at the moment of launching the active spacecraft into the spacecraft's orbital plane.

The disadvantage of this method is that after docking, the energy capabilities of the orbital station (OS) do not allow the transfer of the bundle to a high-energy orbit.

The known method of control of the American PC "Gemiya-11" during its approach and docking with the spacecraft (SC) -target "Agena", which also played the role of RB [2. World manned astronautics. History. Technique. People - ed. Yu.M. Baturina, M, RT Soft, 2005], selected as a prototype, including the launching of the spacecraft into a near-earth orbit, the rendezvous and docking with the RB, autonomously launched into the near-earth orbit, and the application of the transient impulse ΔV _PRL to the resulting bundle with the help of the RB for the transition to a high-energy orbit. The launch of the Gemini-11 spacecraft was carried out into the orbital plane of the spacecraft launched on the same day. The rendezvous was performed in one orbit using onboard measurements of the relative positions of the spacecraft and spacecraft to determine the characteristics of rendezvous pulses. After docking with the help of the engines of the Agena spacecraft, which was already performing the role of the RB, an impulse of 280 m / s was applied to the bundle, which ensured its transition to a high-energy orbit with an orbital apogee altitude of 1370 km.

FIG. 2 shows the flight diagram of the prototype of the spacecraft Gemini-11, which is the KO, with the spacecraft Agena, which is the RB, where: 5 - the reference orbit of the launch of the KO, 6 - the orbit of the RB, which is the target orbit for docking, 7 - the pulses of the transfer of the spacecraft to RB orbit, 8 - outbound impulse of the transition to a high-energy orbit ΔV _PRL , 9 - high-energy orbit. After launching KO 1 into reference orbit 5, pulses 7 were applied to it to transfer to target orbit 6, where the KO 5 was docked with RB 2. After docking due to the energy characteristics of the RB, a departure pulse ΔV _PRL 8 to enter a high-energy orbit 9.

The disadvantage of this method is the low reliability of the two-start circuit, because if the second spacecraft is not launched on time, an early termination of the flight of the first spacecraft will be required due to limitations on the autonomous flight resource.

The technical result of the invention is to improve the reliability of the transport system, which ensures the departure of the KO by a two-start scheme when using the OS.

The technical result is achieved due to the fact that in the method of controlling the transport system when performing a flight to a high-energy orbit, including placing the spacecraft into a near-earth orbit, approaching and docking with an RB, autonomously launched into a near-earth orbit, and applying to the resulting bundle of a transient impulse ΔV _PRL using RB to transfer to a high-energy orbit, in contrast to the known one, the SC is pre-docked with the OS, and the RB is put into a reference orbit that intersects the plane of the station's orbit at an angle θ, but with a given mismatch in the latitude argument ΔF, after which impulses ΔV _i are applied to the RB, where i-1,2, ..., the characteristics of which are determined by the nominal parameters of the reference orbit of the launching of the RB, for the transition to the target orbit, coplanar and coelliptical with respect to the station orbit with a difference in height by the value of ΔH, which is determined taking into account the actual value of ΔF, and then undock the space object from the station, to approach the RB on the target th orbit, impulses ΔV _j are applied to the SC, where j-1,2, ..., the characteristics of which are determined by the actual parameters of the target orbit, dock the SC with the RB and apply to the formed bundle a _{transfer pulse ΔV PRL} to go to a high-energy orbit with the help of RB.

Two-launch schemes are more economical due to the possibility of using carrier rockets with a lower carrying capacity. The possibility for the PC to standby as part of the OS for a successful launch of the RB allows you to save the resource for autonomous flight. The presented method provides for fast rendezvous and docking of the spacecraft with an RB, which has a limited operating life, with subsequent execution under optimal conditions for the launch of the spacecraft into a high-energy orbit.

The technical result in the proposed control method is achieved due to the fact that with the help of quasi-coplanar launching of the RB and rendezvous with the PC, previously docked with the OS, in an orbit coelliptical to the OS orbit with a difference in height ΔH, joint conditions are provided for the rapid convergence of the PC and RB and for optimal flight to the moon.

The essence of the invention is illustrated in FIG. 1 ÷ 6, where:

in fig. 1 shows a scheme of injecting an analogue into a reference orbit with a start time shift, which provides a mismatch of the orbital planes by an angle θ,

in fig. 2 shows the flight diagram of the prototype during the transition to a high-energy orbit,

in fig. 3 shows a diagram of the transition of the RB from the reference orbit of injection and SC from the orbit of the OS to the target orbit, the coelliptical orbit of the OS with the difference in height ΔH,

in fig. 4 shows a two-pulse scheme of the spacecraft transfer from the OS orbit to the target coelliptical orbit with a difference in height ΔH for docking with the RB,

in fig. 5 shows the graphs of the costs of the characteristic velocity of the RB for the rendezvous with the spacecraft, depending on the initial mismatch in the argument of latitude ΔФ between the objects for different values of the difference in height ΔН between the target coelliptical orbit and the orbit of the OS,

in fig. 6 shows the cyclogram of the proposed method for controlling the transport system when performing a flight to a high-energy orbit.

The following positions are marked on these figures:

1 - active SC, 2 - passive SC, 3 - the point of intersection of two planes at an angle θ, where an impulse must be executed to eliminate the mismatch of the planes, 4 - the specified mismatch in the argument of latitude ΔФ between the objects at the moment the active spacecraft is launched into the plane of the SC orbit, 5 - reference orbit for launching the SC, 6 - orbit of the RB, which is the target orbit for docking, 7 - pulses of the transfer of the SC to the orbit of the RB, 8 - departure pulse of the transition to a high-energy orbit ΔV _PRL , 9 - high-energy orbit, 10 - orbit of the OS, 11 - the difference in altitude AH between the target coelliptical orbit and the OS orbit, 12 are the pulses for the transfer of the SO from the OS orbit to the target orbit, 13 is the range of admissible values of the mismatch in the latitude argument ΔF, which allows the rendezvous in a coplanar setting, 14 is the range of admissible mismatch values in the argument of latitude ΔФ, which makes it possible to carry out the approach in a quasi-coplanar formulation.

Let us consider the proposed method using the example of a PC, which must be transferred to a departure trajectory to the Moon. The spacecraft is preliminarily put into orbit and docked with the OS, in which the crew expects a successful launch of the RB on the date of the appearance of ballistic conditions for departure to the Moon. After launch, the RB goes into orbit, the coelliptical orbit of the OS, and the PC undocks from the OS and also goes into the coelliptical orbit, where it joins the RB. _{After docking, a flight impulse ΔV PRL is} applied to the bundle with the help of the RB to fly off to the Moon.

The successful launch of the RB does not yet provide favorable conditions for the implementation of the rapid convergence of the RB and the PC that is at that time on the OS. So, for example, the discrepancy in the latitude argument ΔФ when approaching in two turns should be in a narrow optimal range of 15 ° [3.R. Murtazin "Two-turn diagram of the rendezvous of the Soyuz spacecraft with the International Space Station", w. Cosmonautics and Rocket Engineering TsNIIMash, 2017 (1) No. 94, p. 30-37], and the actual misalignment on the optimal date for the flight to the Moon can be in the range from 0 to 360 °.

To expand the conditions for the rendezvous, the quasi-coplanar launch described in [1] is used. In this case, the launch time of the RB is shifted so that the mismatch with the SC in terms of the latitude argument ΔF is in the optimal range, and the resulting angle of mismatch of the orbital planes of the SC and RB θ is partially eliminated by changing the launch azimuth of the launch vehicle (LV). After insertion into the reference orbit, pulses of the transfer to the target orbit are applied to the RB, the characteristics of which are determined from pre-launch calculations without measuring the parameters of the actual orbit. An additional expansion of the permissible phase conditions in ΔF can be obtained by varying the difference in height AH between the orbit of the OS and the coelliptical orbit. The transfer of the PC to a coelliptical orbit is performed in half a turn after undocking from the OS by applying two pulses, the characteristics of which are determined by the actual parameters of the RB orbit, and the PC and RB are docked. After docking, the departure pulse ΔV _PRL to the Moon is performed at the moment when the advance angle between the plane of the departure orbit and the Earth-Moon direction is 33-60 °, which provides a flight to the Moon in 2.5 - 4.5 days [4. "Fundamentals of spacecraft flight theory" ed. G.S. Narimanov, M, Mechanical Engineering, 1972].

FIG. 3 shows a scheme of rendezvous and docking in the target orbit. RB 2 is put into reference orbit 5 with a given mismatch in the argument of latitude ΔF 4 relative to the OS. Then pulses 7 are applied to the RB, the characteristics of which are determined by pre-launch calculations, to transfer to the target orbit 6, coelliptical to the orbit of OS 10 with a difference in height ΔH 11. Then KO 1 is undocked from the OS and pulses 12 are applied to it, the characteristics of which are determined by the actual parameters of the target orbit 6, for docking with the RB.

FIG. 4 in the orbital coordinate system (OSK) associated with the OS, a two-pulse scheme of the SC transfer from the OS orbit to the target orbit is presented in detail. After undocking the KO 1 from the OS, pulses 12 are applied to it, the characteristics of which are determined by the actual parameters of the target orbit 6, which nominally in the OCS should represent a horizontal straight line along the X-axis of the OCS with a difference in height ΔH 11.

On the graphs of the costs of the characteristic velocity of the RB when approaching the KO (Fig. 5), solid lines 13 indicate the costs corresponding to the launch of the RB into the plane of the OS orbit. In this case, there is no misalignment of the planes 3 shown in FIG. 1. The admissible range of the initial mismatch in the argument of latitude ΔF between RB and OS, corresponding to the horizontal "shelf", is rather narrow and does not exceed 15 °. Due to the quasi-coplanar launch, the extended range 14 is already 70 °. Varying the difference in the heights of the orbits of the OS and the target coelliptical orbit ΔН in the range of 5 ÷ 100 km, it is possible to expand it to 80 °, which with a high probability will provide joint conditions for fast approach and departure to a given high-energy orbit, for example, to the Moon.

After launching RB 2 into the reference orbit (Fig. 6), it is transferred to the target coelliptical orbit by three pulses 7, after which the KO 1 is undocked from the OS, with the help of two pulses 12 is transferred to the target orbit and docked with the RB. After docking, due to the power of the RB, a departure impulse 8 is performed to transfer the bundle to a high-energy orbit.

Let's look at an example. Let's say it is necessary to put the spacecraft into the lunar orbit. To fly to the Moon, an outbound impulse V _{OTL of} about 3150 m / s is required [4]. As a rule, the energy capabilities of the KO are not enough to execute such a large impulse, and RB is used for this. A simple solution to this problem is the joint launch of KO and RB. But a single-launch scheme will require the use of a super-heavy launch vehicle. The use of a two-start scheme with separate start of the spacecraft and the missile launcher makes it possible to significantly reduce the requirement for the thrust-to-weight ratio of the launch vehicle. For a flight to high-energy orbits, it is most effective to use RBs fueled with low-boiling components, for example, an oxygen-hydrogen pair. As disadvantages, it should be noted that the time interval between the launch of the RB and the execution of the departure pulse V _{OTL is} short and amounts to several hours. The moment of application of the outbound pulse is selected from the condition of the transfer to a given lunar orbit. This condition is most simply ensured when launching from the Earth by placing the SC and RB into an orbit, the plane of which satisfies the condition of departure to the Moon. Taking into account the functional limitations of RB in time, it is obvious that KO should be displayed first. The disadvantage of this is the low reliability of the circuit. if the RB is not launched on time, the KO will be forced to prematurely terminate the flight. The use of an operating system on which the KO can wait for the launch of the RB can improve reliability. In this case, the problem becomes the need to combine the conditions for the rapid convergence of SO and RB and for flight to the Moon. The last condition is repeated every 9-11 days [5. R. Murtazin "Space transport system to ensure the operation of a circumlunar orbital structure", Zh. Cosmonautics and Rocket Engineering TsNIIMash, 2017 (2) No. 95 p. 55-63], therefore, it is necessary to strive to ensure that, in the event of a non-launch in RB, the conditions for the rapid convergence of RB and SC awaiting the launch of RB on the OS are met in the next launch window to the Moon. For this, it is necessary to strive to increase the permissible phase range for fast convergence. For this, quasi-coplanar injection and variation in height ΔН of the target coelliptical orbit are used.

Let the initial phase angle be -15 ° on the planned date, which corresponds to the left border of the extended range 14 with a difference in height ΔH = 5 km. If the OS orbit is a multiple orbit, then after a multiple number of days the phase conditions for the rendezvous are repeated [4]. In an orbit of five-day multiplicity, the phase conditions shift by ~ 70 ° every day. Thus, after 10 plus 1 day, i.e. by the time of the appearance of the next window for departure to the Moon, the conditions for fast approach correspond already to the right boundary of the extended phase range 14 with a difference in altitude ΔH = 100 km.

The presented example of the implementation of the proposed method shows the possibility of implementing a two-launch scheme for flying off to the Moon while simultaneously providing conditions for the rapid convergence of the spacecraft and the RB.

The proposed method can be used to deliver a manned spacecraft to a polar circumlunar orbit.

Claims (1)

A method for controlling a transport system during a flight to a high-energy orbit, including injecting a space object into a low-energy orbit, approaching and docking with an upper stage autonomously launched into a low _{-energy orbit, and applying a transfer pulse ΔV PRL} to the resulting bundle with the help of an upper stage to transfer to a high-energy orbit , characterized in that the space object is pre-docked with the near-earth station, and the upper stage is put into a reference orbit crossing the station's orbit plane at an angle θ, and with a given mismatch in the latitude argument ΔФ, after which pulses ΔV _i are applied to the upper stage, where i = 1,2, ..., the characteristics of which are determined by the nominal parameters of the reference orbit of the injection of the upper stage, for the transition to the target orbit, coplanar and coelliptical with respect to the station orbit with a difference in height by the value ΔН, which is determined taking into account the actual value of ΔФ, and then undock a space object from the station, to approach the upper stage in the target orbit, impulses ΔV _j are applied to the space object, where j = 1,2, ..., the characteristics of which are determined by the actual parameters of the target orbit, the space object is connected to the upper stage and applied to the resulting bundle transient impulse ΔV _PRL for the transition to a high-energy orbit using the upper stage.

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