RU2490181C1 - Method of control over active space object to be docked to passive space object - Google Patents

Abstract

FIELD: transport.

SUBSTANCE: invention relates to space engineering and may be used for docking of two space objects, one active and another passive. Active space objected is placed in reference orbit to define characteristics of approach pulses via nominal parameters of said reference orbit to be applied to active space object in initial orbit. Then, characteristics of approach pulses are defined from actual parameters of active space object and applied to next orbits.

EFFECT: faster approach to passive space object.

3 dwg, 1 tbl

Description

The proposed control method can be used in space technology during the approach and subsequent docking of two space objects located in a near-circular orbit of a celestial object, for example, a manned spacecraft launched by a carrier rocket (LV) as an active space object (AO) and the international space station (ISS) , as a passive space object (FFP).

A known control method, selected as an analogue, in which to ensure convergence and subsequent docking of two space objects, the AKO is displayed in the plane of the orbit of the FFP. As a rule, the average FBO orbit height is higher than the average height of the AKO orbit, and therefore, after the derivation and determination of the parameters of the orbit, the first two approach pulses are calculated, with which the AKO is transferred to the so-called phasing orbit, at which the initial angular mismatch between the two objects. This is due to the fact that the angular velocity of rotation around the celestial body in AKO is higher than in FFP. After several turns (the number is determined by the selected approach ballistic scheme) and determination of the current orbit parameters, a two-pulse maneuver is calculated, with which the AKO is transferred to the vicinity of the FFP, where the approach is completed by automatic docking. This method of controlling an active spacecraft, used when approaching and docking manned and cargo spacecraft with the ISS, requires that the active spacecraft be brought into the orbit of the ISS [1]. In this case, the initial phase angle between space objects is arbitrary, which requires a certain time for its elimination. In addition, a certain amount of time with this method is spent on performing standard technological operations, such as determining the orbit from the data of the radio monitoring of the orbit from ground-based measuring points (NPCs) and the subsequent determination of the characteristics of the convergence pulses in the Flight Control Center, as well as transmitting the received data to the on-board computer complex ( BVK) AKO. Due to restrictions on the visibility zones of NPCs, as only a part of the flight turns are observable, it is necessary to divide the implementation of such technological operations into several days, tying them to visibility zones. This, in turn, requires an increase in autonomous AKO flight before docking (currently it is about 34 turns or almost 50 hours). This circumstance leads to an additional load on the crew of a manned missile, forced to stay in the cramped conditions of a limited spacecraft for a long time and is almost completely unacceptable when implementing a rescue mission on the ISS, when the approach time factor becomes decisive.

A known method of controlling the movement of AKO, mating with FFP, selected as a prototype. This method was developed for the rapprochement of the American Gemini-11 AKO with the third stage of the Agen launch vehicle in a short time during one revolution. Using the advantageous location of the launch site of the cosmodrome at Cape Canaveral with the first and second stage launch sites located in the Atlantic Ocean, the FFP was first displayed in the ascending part of the orbit with an inclination of 28.84 °. Then, through the orbit, but already in the descending part of the orbit, to the reference orbit coinciding with the orbital plane of the FFP, the Gemini-11 AO was displayed with a given angular mismatch between objects of about 4 °. Such an initial phase angle made it possible to complete the phasing of the AKO and to be in the vicinity of the ASW in one turn. 50 minutes after the launch, the Gemini-11 spacecraft was at the peak of its orbit and about 50 km behind the ASC in the zone of operation of the onboard radar of proximity, which allows one to obtain the actual parameters of the relative state vector of the AEC in the orbital system associated with the AEC. After the application of approach pulses, the characteristics of which were obtained from the actual state vector, the AKO was brought into the vicinity of the FFP. This flight ended with a successful docking of the Gemini-11 spacecraft with the Agena rocket stage 1 hour 34 minutes after the launch, which is by far the best result in manned space exploration [2].

The main disadvantage of this control method is that its implementation is possible only by sequentially launching the FFP and AKO into orbit with a fixed inclination for a certain time to provide the necessary initial phase angle between objects with high accuracy of ± 0.5 °. In the case of approaching and docking with the ISS or another FFP already in orbit, providing the necessary initial phase angle will require the FFP to make numerous consecutive corrections before the start of the AKO to form the docking orbit with the necessary accuracy. Given the intensive schedule of flights to the ISS for spacecraft providing transport operations, such a prelaunch formation of the ISS orbit will require an additional large amount of fuel for maneuvers. The backup start date for AKO during the formation of the phase range of ± 0.5 ° is practically unrealizable. In the case of the mission of the rescue ship, preparing the conditions for a quick rapprochement will require too much time, incomparable with the flight time to the FFP [3].

The technical result of the invention is to reduce the duration of the approach to three turns with a FFP in orbit for a long time with the expansion of the phase range to 20 °, which allows to ensure the launch date of the AKO without additional corrections to the FFP.

The technical result is achieved due to the fact that in the method of controlling the motion of an active space object that is docked with a passive space object, including the launch of the active space object into the reference orbit, the determination of the characteristics of the approach pulses and the subsequent application of the approach pulses to the active space object, in contrast to the known, previously determine the characteristics l of the approach pulses from the nominal parameters of the reference orbit of the derivation, where l = 1, 2, ..., which are dyval the first coil after the removal, and then determine the characteristics of the convergence m of pulses of the actual parameters of the orbit of the active object space, where m = 1, 2, ..., which is applied in subsequent convergence coils.

The technical result in the proposed control method is achieved due to the fact that the transition to the phasing orbit begins almost immediately after launching into the reference orbit using l approach pulses, the characteristics of which are predefined in accordance with the nominal parameters of the reference orbit of the AKO launch. In this case, no time is required to wait for the determination of the actual parameters of the reference orbit of the launch of the AKO.

The invention is illustrated in figures 1-3 and table 1, where figure 1 shows the proximity of AKO "Soyuz-TMA" with the ISS,

figure 2 shows the sequence diagram of the basic technological operations during the approximation of the Soyuz-TMA with the ISS with the existing two-day approach scheme,

figure 3 shows the sequence diagram of the basic technological operations during the implementation of the proposed method in a three-turn proximity circuit,

Table 1 presents the results of calculations based on accuracy determined by deviations from the aiming state vector at the aiming point of convergence and fuel consumption, expressed in characteristic speed (in m / s) when using the proposed method in a three-turn approaching scheme in comparison with the existing two-day approaching scheme.

Figure 1 shows a well-known launching scheme, for example, Soyuz-TMA, to the target orbit, coinciding with the orbital plane of the FFP using the example of the ISS. After the withdrawal, the initial phase angle Φ between the AKO and the FAC is eliminated due to the transition of the AKO to the phasing orbit (item 1) using a two-pulse maneuver (V ₁ and V ₂ ) and the subsequent flight in this orbit until the start of the autonomous section, providing a transition to the vicinity The ISS also due to the two-pulse maneuver (V ₃ and V ₄ ).

Figure 2 shows, as an example, a sequence diagram of the main technological operations during the implementation of the two-day approach circuit of the Soyuz-TMA ACS with the ISS. As can be seen from this figure, immediately after the removal (pos. 1), the orbit is determined on two turns, after which, in the visibility zone of the third turn (pos. 2), data are entered on the first two-pulse approach maneuver (V ₁ and V ₂ ) (pos. 3), ensuring the transition of AKO to the phasing orbit. Then, in a day, this technological chain, consisting of determining the orbit of the AKO, storing data for the maneuver and performing the maneuver, is repeated with the issuance of a correcting impulse V _cor (item 4). The purpose of the impulse V _cor is to eliminate errors in the forecast of the state vector of AKO accumulated during the first day of the flight. A day later, this technological chain is repeated again with the second two-pulse maneuver (V ₃ and V ₄ ) (pos. 5), after which the Soyuz-TMA is in the vicinity of the ISS and an autonomous flight section begins (pos. 6).

Figure 3 presents the sequence diagram of the proposed method in a three-turn proximity circuit. In contrast to the two-day scheme, the transition to the phasing orbit occurs immediately after the withdrawal (pos. 1) and the subsequent end of the NPC zone at the 1st turn (pos. 2). In this case, two pulses (l = 1, 2) are used to transfer to the phasing orbit (V ₁ and V ₂ ) (item 3). Data on the timing and characteristics of the maneuver pulses for transition to the phasing orbit are calculated by the FMS state vector and the nominal AKO launch vector and are entered into the AKO flight mission immediately before the launch of the LV with the AKO or transferred to the IAC AKO immediately after the launch to the reference orbit. The location of the first two pulses is located between the zones of the NPCs of the first and second daily turns. This allows one, on the one hand, to determine the true parameters of launching the AKO into the reference orbit without disturbing pulses, and, on the other hand, to ensure the transmission of the current state vector and characteristics of two correcting pulses m = 1, 2 (V ₃ and V ₄ ) to the BVK AKO (pos. 4 ) The need for introducing into the approach circuit and subsequent execution of two correcting pulses m = 1, 2 (V ₃ and V ₄ ) is formed due to the difference between the real reference orbit of the AKO and the nominal reference orbit of the lead, which leads to the appearance of a six-dimensional slip vector at the aiming point of convergence. Since each pulse of the corrective maneuver has three components, then two pulses make it possible to completely eliminate all six parameters of the miss vector. The zone of two correcting impulses is located between the end of the visibility zone of NIPs at the second daily turn and the beginning of the autonomous approach section and should be at least half a turn. The length of the zone of at least half a turn is justified by the fact that it is necessary to eliminate the lateral mismatch between the real orbit of the AKO and the orbit of the FFP. Since the optimal point for this operation is one of the two intersection points of both orbits (AKO and FFP) [4], the duration of the zone for executing pulses in the half-turn completely guarantees the location of this point in it. The correcting pulses m = 1, 2 (V ₃ and V ₄ ) are calculated by the actual AKO removal vector obtained from the measurements of the orbit in the visibility range of NIPs at the first daily turn, taking into account the nominal performance of the first two phasing pulses l = 1, 2 (V ₁ and V ₂ ). At the end of the execution of the correcting pulses m = 1, 2 (V ₃ and V ₄ ), the AKO is in the vicinity of the ISS and an autonomous section (pos. 5) similar to that shown in figure 2 begins, with the remaining pulses m = 3,4,5 ...

Pulse calculations are carried out using the correction equation [5], in relative motion centered in the FFP

$$\left[\begin{array}{c}{r}_k\\ V_k\end{array}\right]=M_{k0}$$

$$\left[\begin{array}{c}{r}_0\\ V_0\end{array}\right]+{\displaystyle \sum _i{A}_{ki}\Delta V_i}$$

- конечный вектор, после решения задачи сближения,Where $$\left[\begin{array}{c}{r}_k\\ V_k\end{array}\right]$$

- the final vector, after solving the approximation problem,

- вектор состояния в начале сближения, $$\left[\begin{array}{c}{r}_0\\ V_0\end{array}\right]$$

is the state vector at the beginning of the approach,

M _k0 is the forecast matrix in relative motion from the moment of approaching t ₀ to the moment of approaching end t _k ,

A _ki is the correction matrix in relative motion for the ith pulse ΔV _i .

,The state vector at the end point should correspond to the aiming state vector: $$\left[\begin{array}{c}{r}_k\\ V_k\end{array}\right]=\left[\begin{array}{c}{r}_{PR}\\ V_{PR}\end{array}\right]$$

,

.In the problem of convergence of two space objects, as a rule, the meeting problem is used when $$\left[\begin{array}{c}{r}_{PR}\\ V_{PR}\end{array}\right]=\left[\begin{array}{c}0\\ V_{PR}\end{array}\right]$$

.

.If the final vector does not match the selected aiming vector, a miss vector is formed $$\left[\begin{array}{c}\delta r\\ \delta V\end{array}\right]=\left[\begin{array}{c}{r}_k\\ V_k\end{array}\right]-\left[\begin{array}{c}{r}_{PR}\\ V_{PR}\end{array}\right]$$

.

The maneuvering interval is used to determine the approach pulses. Each combination of points of application of approach pulses corresponds to a certain combination of pulses. ΔV ₁ , ..., ΔV _n . The actual value of the approach pulses and their application points is determined from the variety of solutions of the correction equation by calculating and minimizing the functional F

$$F=\Delta V_{\mathrm{one}}+\cdots +\Delta V_n={\displaystyle \sum _{i=\mathrm{one}}^n\Delta V_i=\min }$$

Data on the corrective maneuver are laid down in the IAC AKO in the visibility range of NPCs in the second round. In principle, the parameters of the corrective maneuver can be calculated autonomously by the BVK AKO using the actual state vector of the AKO.

After completing the corrective two-pulse maneuver, the stage of the autonomous section begins, which is located approximately a revolution before the sighting point of the meeting. An autonomous section is carried out using a system of mutual measurements of AKO and FFP, operating with a limited range and providing a relative state vector. At the end of the autonomous section, AKO and FFP are docked in the visibility zone of the fourth round. From the cyclogram presented in figure 3, it can be seen that the duration of the flight from hatching to docking in the proposed method will be three turns or about 4.5 hours.

The proposed control method can be used when approaching and docking the ACS with the ISS orbital station or any other ACS. It is especially useful to use this method during a rescue mission, when the time factor for the delivery of the crew on board the ISS becomes crucial.

Table 1 presents the results obtained by statistical modeling 300 of the convergence process: the mathematical expectation (m) and standard deviation (σ) of the impact state vector at the impact approach point in the relative orbital coordinate system rnb [5], as well as fuel consumption, expressed in characteristic speed (in m / s). Additionally, in a separate line, the maximum (m + 3σ) fuel consumption is presented. The proposed method was considered when performing a three-turn approach and, for comparison, the implementation of a standard two-day approach scheme.

Upon receipt of the data in Table 1, it was assumed that the AKO is launched into the target orbit with the parameters Н _min / Н _max = _200/240 km and inclination i = 51.6 °. The initial spread along the excretion vector was simulated, corresponding to a spread over the period of ± 22 ^S , and on the side ± 3.5 angles. min., which corresponds to the scatter of the parameters of the reference orbit during the launch by a specific launch vehicle used in the launch of the AKO [6]. The ISS orbit corresponds to the planned altitude for 2012 - 408 km. The initial phase angles for the proposed three-turn circuit (17) and for the standard two-day circuit (215 °) corresponded to the optimal values for each circuit. The calculations were carried out using the software for calculating the approximation problem for AKO with the ISS orbital station.

As can be seen from Table 1, the σ of the lateral components of the state vector at the target point (b, V _b ): 0.22 km and 0.27 km in b and 0.39 m / s and 0.31 m / s in V _b , are very close in both proximity schemes. Also, the marginal fuel consumption for a rapprochement of 114.9 m / s and 113.2 m / s are almost identical. But, in the proposed method, compared with the two-day scheme, σ of the state vector at the aiming point in the orbit plane (r, n, V _r , V _n ) is much smaller. So, if the component r in both schemes is approximately equal to σ (0.24 km and 0.25 km), then the component n σ is one and a half times (0.83 km and 1.24 km), and the component V _n σ is half as much (0.27 m / sec and 0.55 m / s), and for the component V _r σ is less than three times less (0.18 m / s and 0.63 m / s). This effect is realized due to a significantly smaller portion of the flight before issuing the state vector parameters in the BVK AKO before the start of the autonomous portion and, therefore, to a much lesser influence of disturbing factors (inaccuracies in pulse processing, dynamic operations, atmospheric influence, etc.). This, in turn, provides more accurate kinematic conditions before the start of the work of the algorithm of the autonomous approach section and contributes to its reliable operation. All of the above confirms the possibility of using this method in the task of convergence of two space objects.

It should also be noted that the range of initial phase angles in a three-turn circuit with the proposed control method, which was determined when modeling the approximation problem, is about 20 °. If the actual AKO parameters were expected after determining the orbit from the radio control data of the orbit from NIPs, as implemented in the standard two-day approach scheme, the transition to the phasing orbit would be completed a turn later, which would accordingly narrow the possible range of initial phase angles to 5 °

Table 1 A method for controlling the movement of an active space object docked with a passive space object The proposed method (three-turn approach circuit) r, km n, km b, km V _r , m / s V _n , m / s V _b , m / s Fuel consumption, m / s m -0.007 -0.007 -0.008 0.022 -12.53 0.026 104.42 σ 0.24 0.83 0.22 0.18 0.27 0.39 3.49 marginal fuel consumption (m + 3σ) 114.9 m / s Two-day rapprochement scheme r, km n, km b, km V _r , m / s V _n , m / s V _b , m / s Fuel consumption, m / s m 0.001 -0.205 0.020 0.118 -12.66 -0.014 101.61 σ 0.25 1.24 0.27 0.63 0.55 0.31 3.87 marginal fuel consumption (m + 3σ) 113.2 m / s

Bibliography

1. Wigbert Fehse (2003) "Automated Rendezvous and Docking of Spacecraft", Cambrige University press

2. NASA Press kit (1966) "Project Gemini-11" - prototype, http://www.scribd.com/doc/11483557/Gemini-11-Press-Kit

3. R. Murtazin, S. Budylov "Short Rendezvous Missions for Advanced Russian Human Spacecraft", Acta Astronautica 67 (2010) 900-909

4. Elyasberg P.E. "Introduction to the theory of flight of artificial Earth satellites", Moscow, Nauka, 1965

5. R.F. Appazov, O.G. Sytin, "Design Methods for the Trajectories of Earth Carriers and Satellites", Moscow, Nauka, 1987

6. Steven J. Isakowitz (2004) "International Reference Guide to Space Launch Systems" Forth edition

Claims (1)

A method for controlling the motion of an active space object docked with a passive space object, including the launch of the active space object into the reference orbit, the determination of the characteristics of the approach pulses and the subsequent application of the approach pulses to the active space object, characterized in that they first determine the characteristics l of the approach pulses from the nominal parameters the reference orbit of the launch, where l = 1, 2, ..., which are applied on the first turn after the launch, then determine akteristiki m pulses convergence of the actual parameters of the orbit of the active object space, where m = 1, 2, ..., which is applied in subsequent convergence coils.

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