RU2440281C1 - Method of controlling active flight of space object to be docked with passive spaceship - Google Patents

Abstract

FIELD: transport.

SUBSTANCE: invention relates to docking jobs, in particular, to docking of spaceship with international space station. Proposed method comprises using carrier rocket to place active spaceship (ASS) into target orbit with orbit ascending section altitude deviation but with preset latitude departure argument mismatch. Then, first maneuver of ASS is carried out in direction perpendicular to placing plane to change orbit ascending section altitude. Further, second maneuver is performed in opposite direction to eliminate mismatch in target orbit inclination originated after said first maneuver. Said maneuver is performed with ASS engine thrust throttling to create additional orbit ascending section altitude deviation and align orbital planes of objects to be docked. Formation of required initial angular mismatch between said objects, period approach of said objected decrease up to one pass.

EFFECT: decreased docking interval without additional orbit corrections.

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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 (PAC), 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 the convergence and subsequent docking of two space objects, the spacecraft is displayed in the plane of the orbit of the spacecraft. As a rule, the average FBO orbit height is higher than the average height of the AKO orbit, and therefore, after the AKO, it is transferred to the so-called phasing orbit, where the initial angular mismatch between the two objects is eliminated. 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 ballistic approach circuit), the AKO is transferred to the vicinity of the FFP using a two-pulse maneuver, where the approach is completed by automatic docking. This method of controlling an active spacecraft, used during the approach and docking of the spacecraft and cargo ships (GK) with the ISS, requires the active ship to 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, i.e. increase AKO autonomous flight before docking. This leads to an additional load on the crew of the PAC, which is forced to remain 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 convergence time factor becomes crucial.

A known method of controlling the movement of the AKO, mating with the FFP in a short time during one revolution, selected as a prototype. This method was developed to bring the American Gemini-11 RCC closer to the third stage of the Agen rocket. Using the advantageous location of the launch site of the cosmodrome at Cape Canaveral with the first and second stage launching sites located in the Atlantic Ocean, the FFP was first displayed on the ascending part of the orbit with an inclination of 28.84 °. Then, through the turn, but already on the descending part of the turn, the Gemini-11 spacecraft was withdrawn into the plane of the orbit of the FFP with a predetermined angular mismatch between objects of 12 °, allowing phasing to be completed and to be in the vicinity of the FFP in one revolution. This flight ended with a successful docking of the Gemini-11 rocket launcher with the Agena rocket stage 1 hour 34 minutes after 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 FSC already in orbit, providing the necessary initial phase angle will require the FFP to make numerous corrections for several months before the start of the AKO to form the docking orbit with the necessary accuracy. Given the busy schedule of manned spacecraft flights to the ISS (for example, only in 2010 4 flights of the Soyuz-TMA spacecraft and 5 space shuttle flights) [3], and the mass of the ISS (more than 340 tons), this will require an additional large amount of fuel to conduct 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, the preparation of conditions for a quick rapprochement will require too much time, incomparable with the flight time to the FFP [4].

The technical result of the invention is to reduce the duration of the approach to the FFP, in orbit for a long time, while maintaining an acceptable frequency of the dates of the launch of the AKO without additional corrections to the FFP. At the same time, depending on the launch vehicle’s energy, the frequency of launches into orbit for rendezvous and docking can be provided every day, which is significantly better than the existing two-day rendezvous scheme, which provides a frequency of once every two days.

The technical result is achieved due to the fact that in the method for controlling the movement of an active space object docked with a passive space object, including the launch of an active space object into a target orbit, in contrast to the known one, an active space object is brought into a target orbit with a deviation of an upward orbit that does not coincide in longitude with the orbit of a passive space object by the value Δλ of the _VU , but with a given mismatch with respect to the latitude argument ΔФ, after which, in the perpendicular direction to the lead-out plane, perform the first maneuver in the angular interval Δφ ₁ from the point of launch into the target orbit to change the longitude of the ascending node of the orbit by Δλ _VU1 by generating a pulse ΔV ₁ , and then in the opposite direction in the angular interval Δφ ₂ perform the second maneuver with throttling thruster active space object to eliminate the mismatch on the target orbit inclination Δi _1, effected after the first maneuver and create additional deviation Δλ longitude of ascending node _BV2 for MF pulse shaping ΔV t _2, resulting in the coincidence orbital planes abutting objects.

The technical result in the proposed control method is achieved by the fact that the time of launching the AKO into orbit is selected so that at the time of launching the difference in the arguments of the latitude of the AKO and the FFP ΔF is equal to the specified.

The invention is illustrated in figure 1 ÷ 7 and table 1 and 2.

Figure 1 shows the proximity of the spacecraft "Soyuz-TMA" with the ISS,

figure 2 presents the orbital elements and directions of maneuvers,

figure 3 represents the relative position of two orbits with the same inclination, differing in the position of the ascending node of the orbit,

figure 4 illustrates the dependence of the initial phase angle at the time of removal of the AKO from the position of the longitude of the ascending node of the orbit of the FFP at the time of its passage through the equator,

figure 5 represents a possible geometry of the relative position of the AKO and FFP at a given value of the initial phase angle and different positions of the plane of the orbit of the FFP,

Fig.6 shows the relative motion of the AKO in the orbital coordinate system of the FFP in the implementation of rapid convergence,

Fig.7 represents the lateral relative motion of the AKO in the orbital coordinate system of the FFP in the implementation of rapid convergence.

Table 1 presents the results of calculations on the use of the proposed method for the launch position of the launch vehicle with the coordinates:

45.92 ° N and 63.52 ° east (breeding in the ascending part of the coil),

table 2 presents the results of calculations on the use of the proposed method for the launch position of the launch vehicle with the coordinates:

51.88 ° N and 128.36 ° East (withdrawal in the descending part of the coil).

Figure 1 shows the well-known scheme of launching ACS, for example, the Soyuz-TMA spacecraft to the target orbit, which coincides with the orbital plane of the spacecraft on the example of the ISS with the subsequent transition to a phasing orbit, due to a two-pulse maneuver (V ₁ and V ₂ ) and culminating in transfer to the ISS vicinity also due to a two-pulse maneuver (V ₃ and V ₄ ). Figure 2 in the geocentric coordinate system presents orbital elements describing the orbital motion of the AKO and used in the description of the invention. The direction of acceleration f _{3 is} shown, perpendicular to the plane of the orbit of the AKO, providing a technical result. Figure 3 shows the relative position of two orbits with the same inclination. Moreover, i _R is the angle of intersection of two planes, Ψ and, respectively, 180 ° -Ψ are the arguments for the latitude of the intersection point of these planes, and Δλ _WU is the longitude mismatch of the ascending node of the orbit. Figure 4 demonstrates the determination of the initial phase angle depending on the longitude of the ascending node of the FAC orbit at the beginning of the starting daily turn of the ACF, when the latter (position 1) is brought into the plane of the FAC orbit. So, the value λ _VU0 = 12.7 ° (pos.2c ₁ ) corresponds to the initial phase angle ΔФ ₀ = 0 ° (pos.2c ₁ ), and λ _VU0 = 18.6 ° (pos.2c ₂ ) corresponds to the initial phase angle ΔФ ₀ = 90 ° (pos.2c ₂ ). Figure 5 presents three provisions of the FFP, differing in the value of the ascending node of the orbit, at the time of the withdrawal of the AKO. Longitude λ _VU0 corresponds to the coincidence of the planes of the orbits of the AKO (pos. 1) and the FFP with the nominal initial phase angle Φ (pos. 2 _nom ). Longitude λ _WU0 + Δλ _WU corresponds to the start time of the AKO, providing the initial phase angle Ф + ΔФ (pos.2 _{+ Δλ} ), where ΔФ = 180 ° -2Ψ is the correction to the nominal phase angle, and longitude λ _WU0 -Δλ _WU corresponds to the start time AKO, providing the initial phase angle Ф-ΔФ (pos. 2 _-Δλ ), where Ψ is the argument of the latitude of the point of intersection of two planes with the same inclination (Fig. 3). As an example, for the variant of the initial mismatch of the Δλ _VU0 -5 ° planes, the relative motion of the ACS in the FFP coordinate system (pos. 2) in the FFP orbit plane (Fig.6) and from the FFP orbit plane (Fig.7) is presented. Relative movement is shown from the moment of launch of the LV with AKO (pos. 3). Immediately after the withdrawal (pos. ₁ ), two maneuvers ΔV ₁ (pos. 4) and ΔV ₂ (pos. 5) are carried out to coordinate the planes of the orbits of the ACS and the ACS according to the proposed method, and then after the arrival of the ACS at the apogee of the orbit, the maneuver ΔV ₃ pos.6), translating AKO in the vicinity of FFP. In table 1 and table 2, for two different starting positions, the values of the required impulse ΔV and the initial mismatch in the inclination of the orbits Δi depending on the magnitude of the mismatch in longitude of the ascending node of the orbit Δλ of the _VU between the orbit of the ASD and the nominal, without changing the inclination, orbit of the AKO .

As a rule, for launching into orbit, a starting daily orbit is used, during which the plane of the FFP orbit passes through the launch site of the launch vehicle. Due to the restrictions on the areas of fall of the separating parts of the launch vehicle, the launch of the AKO into the plane of the orbit of the FFP is possible only 1 time per day. The next launch opportunity arises after the rotation of the Earth around its axis by ~ 354 °. The angular range of the daily turn in longitude of the ascending node is 22.5-24 ° depending on the orbit height of the FFP. This angular range corresponds to the phase spread of the initial angles between the AKO and the FFP from 0 to 360 °. Suppose that the phase angle ΔФ = 0 ° at the time of the removal of the ACS corresponds to the longitude of the ascending node of the orbit of the FFP at the time of passage of the equator of the starting daily turn λ _VU0 . If, for example, the longitude of the ascending node of the FSC orbit at the time of passage of the equator will differ by ~ 6 ° relative to λ _ВУ0 at the time of the removal of the ACS into the plane of the _FKO orbit (see Fig. 4), then the initial phase angle between the objects of the LF will be ~ 90 ° (FFP ahead of AKO).

When the AKO is put into orbit into the starting daily turn of any day with a given difference with the FFP with respect to the latitude argument ΔФ, it is almost impossible to provide the necessary value λ _{WU0 of the} longitude of the ascending node of the _FSC orbit for the planes of the objects to be joined. With the same inclination, the plane of the orbits of the objects will differ in the longitude of the ascending node of the orbit by Δλ of the _VU (Fig. 5). This circumstance will prevent the “soft” docking due to the high lateral speeds at the meeting point. Therefore, immediately after the withdrawal of the AKO, it is necessary to conduct a maneuver to eliminate the mismatch in Δλ of the _VU . In the event that, for example, the LV power is sufficient to counter the mismatch over λ of the _WF at 5 °, it will be possible to talk about an acceptable 170-degree phase range for starting the FFP. Currently, in order to implement the two-day scheme for the approach of the Soyuz-TMA PKK to the ISS, they are trying to provide a phase range of 140 ° and the frequency of launches is at least 1 time in two days. The 170-degree range is almost close to allowing launch once a day. An increase in the range along λ _VU0 to ± 6 ÷ 7 ° can unambiguously provide one attempt to start per day without additional corrections of the FFP orbit.

From spherical geometry, one can obtain the following relations [5]:

where i _R is the angle of mismatch of the planes of the orbits of the AKO and FFP,

i is the angle of inclination of both orbits,

Ψ is the argument of the latitude of the point of intersection of the planes of the orbits, and if the orbit of the AKO is to the left of the longitude of the ascending node from the orbit of the FFP, then the argument of the latitude of the point of intersection of the planes will be 180 ° -Ψ (see figure 3).

To correct the longitude of the ascending node and inclination, an impulse perpendicular to the plane of the orbit is used. In this case, if the impulse is conducted in the northern part of the revolution, then the direction along the vector of the kinetic moment of the orbit corresponds to a shift in longitude in the western direction and an increase in inclination. Otherwise: to shift the longitude eastward and reduce the inclination. The most favorable position in orbit for correcting the mismatch in the longitude of the ascending node of the orbits is the point of intersection of the planes AKO and FFP, which, as a rule, are located near the points of the apexes with Ф = 90 ° and Ф = 270 °, where Ф is the argument of the orbit latitude. Depending on the point of application of the pulse, its influence on the change in the orbital elements i and Ω, which determine the position of the orbit plane, is determined [6].

where f ₃ is the acceleration produced by the propulsion system (DU), in the direction perpendicular to the plane of the orbit (figure 3),

µ is the gravitational constant of a celestial body,

e is the eccentricity of the orbit,

ω is the argument of the latitude of the orbit,

a - semimajor axis of the orbit.

Given that we are considering a circumcircular motion with an eccentricity e ~ 0, as well as that the acceleration f ₃ is determined by the thrust of the PH R and the current mass M, which can be assumed constant, i.e.:

and correspondingly:

where V is the average speed of circular motion, formulas (3) and (4) can be simplified:

The argument of perigee ω is counted from the ascending node of the orbit to perigee. An eccentric anomaly is counted from the position of the perigee of the orbit (figure 2). Therefore, the argument of the latitude of the orbit will be associated with the angles E and ω by the following formula:

и производные по изменению орбитальных параметров i и Ω должны быть вычислены в точке

.Since the acceleration f ₃ has a limited value, the execution of any pulse will be carried out for some time Δt. During this time, the argument of the latitude of the orbit will change from Ф to Ф + Δφ. In order to simplify, we can assume that the pulse is applied in the middle of the angular interval Δφ, i.e. at point with latitude argument

and the derivatives with respect to the change in the orbital parameters i and Ω should be calculated at the point

.

Since λ _{WU is} similar to Ω, according to (7) we can write:

During the execution of the pulse ΔV _{1, the} inclination of the orbit i will receive a perturbation Δi ₁ , calculated from formula (6), as:

и производные по изменению орбитальных параметров i и λ_ВУ вычисляются в этой точке. Используя формулы (6) и (7), получим:The resulting incline mismatch must be countered by a second maneuver with a thrust vector directed in the opposite direction. If you carry out the second maneuver with the same thrust, then the effectiveness of this operation will be low, because the inclusion of the remote control will be close to the point of the first maneuver, and when correcting the inclination Δi we will get almost the same change in the plane of the orbit along Δλ of the _VU , but with the opposite sign. The way out of this situation is the maximum possible throttling of the thrust of the remote control and, accordingly, the shift of the interval of execution of the second maneuver to the region of more effective points of application of the impulse to change the inclination of the orbit. If we assume that thrust f ₃ decreases by a factor of k, then the angular execution interval of the second maneuver Δφ _{2 will} accordingly increase by approximately k times, adjusted for the efficiency coefficient associated with the shift of the point of application of the pulse along the orbit. In order to simplify, we accept that the pulse is applied in the middle of the interval, i.e. at point c

and the derivatives of the change in the orbital parameters i and λ of the _VU are calculated at this point. Using formulas (6) and (7), we obtain:

Since the _slave Δλ = Δλ _VU1 -Δλ _CU2 and Δi ₁ + Δi ₂ = 0, then using equations (9) ÷ (12), we can obtain the values of angular intervals - Δφ ₁ and Δφ ₃ solutions of the following system of transcendental equations (13) and (14):

Knowing the angular duration of the maneuvers Δφ ₁ and Δφ ₂ , we can obtain values for the velocity pulses ΔV ₁ and ΔV ₂

и

and

The proposed control method can be used when approaching and docking the PAC with the ISS orbital station. 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.

To calculate the start time of the ACS with a given initial angular mismatch ΔФ, it is necessary to have a state vector of the ACD at the time of the output of the ACS. The knowledge of this vector is provided by ground-based orbital measuring instruments of the Kama and Quantum type or by ASN-M on-board equipment using GPS and GLONASS measurements of satellite navigation systems.

Data on the time and the value of maneuvers for combining the planes of the orbits of the AKO and the airspace are calculated by the state vector of the spacecraft and entered into the flight mission of the spacecraft or spacecraft immediately before launching the spacecraft with spacecraft.

After the launch of the AKO into the target orbit, it is necessary to make the turns of the remote control of the last stage of the launch vehicle or AKO to issue a pulse in the direction perpendicular to the plane of the orbit. To control the orientation of the launch vehicle, either the cruising of the marching engines or their separate switching on (if there are several engines at the last stage) can be used to create a controlled moment along the yaw channel. For some LVs, special vernier engines can also be used for this purpose [7]. To minimize the effect of a turn on the parameters of the target orbit, it is advisable to conduct it with throttling thrust. If the reversal is carried out by AKO means, AKO orientation engines are used.

Performing additional maneuvers to align the planes of the orbits of the joined objects can be provided by LV means if the permissible launch mass of the LV exceeds the AKO mass, i.e. there is an excess of the withdrawn mass enclosed in the fuel of the launch vehicle or by the AKO proper if it has the required amount of fuel. In the case of performing a maneuver using AKO fuel or, if the launch vehicle has the ability to restart, it is possible to reduce the control method to only one maneuver by shifting the angular interval Δφ to perform the first maneuver in the vicinity of the intersection point of the planes of the orbits of the AKO and the ASO (see Fig. .four). In this case, the entire generated pulse will be used to eliminate the mismatch of the planes along the ascending node of the orbit without the occurrence of a mismatch in the inclination of the orbit Δi. In the case of AKO fuel sufficiency, the first maneuver performs LV immediately after launch, and the second maneuver of the AKO remote control in the vicinity of the passage of the equator in order to optimize fuel consumption.

To conduct a comparative analysis, calculations were performed on the use of the proposed method for two launch vehicles with coordinates:

No. 1: 45.92 ° N and 63.52 ° east (withdrawal to the ascending part of the coil)

No. 2: 51.88 ° N and 128.36 ° East (deduction in the descending part of a turn)

The AKO was launched into the target orbit with the parameters H _min / H _max = 135/440 km and the inclination i = 51.6 °. The initial spread in longitude of the ascending node was simulated for the following variants Δλ _WU = ± 5 °. The average acceleration f ₃ was ~ 1.5g. The throttle coefficient ДУ k for the second maneuver took the following values: k = 0.1, 0.05, 0.025. The case of the repeated inclusion of the remote control in the vicinity of the point of intersection of the planes of the AKO and FFP is also considered. The calculation results are summarized in tables 1 and 2, respectively, for the launch positions of launch vehicles No. 1 and No. 2. The calculations were carried out using the Satellite Tool Kit 8.0 software and confirm the receipt of the desired result. The best efficiency is achieved with maximum throttle control of the rocket launcher. The most optimal option from the point of view of fuel consumption appears if it is possible to re-enable the launch vehicle remote control in the vicinity of the intersection point of the AKO and FFP planes. A comparison of the two starting positions shows that for position No. 1, the use of the described control method is approximately 1.5 times more effective due to the launch of the AKO into a more optimal position in orbit (to the apex area).

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. The schedule of flights to the ISS in 2010 was approved. J. News of Cosmonautics, vol. 20, No. 1 (324), 2010, p. 25.

4. R. Murtazin, S. Budylov "Short Rendezvous Missions for Advanced Russian Human Spacecraft", 60 ^th International Astronautical Congress, 12 ^th -16 ^th October 2009, Daejeon, Republic of Korea.

5. James R. Wertz (2001) "Mission Geometry; Orbit and Constelation Design and Management", Space Technology Library.

6. J.E. Pollard (2000) "Simplified Analysis of Low-Thrust Orbital Maneuvers", Aerospace report No. TR-2000 (8565) -10.

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

A method for controlling the movement of an active space object docked with a passive space object

Claims (1)

A method for controlling the movement of an active space object docked with a passive space object, including the removal of the active space object into the target orbit, characterized in that the active space object is brought into the target orbit with a deviation in longitude of the ascending node of the orbit from the orbit of the passive space object by Δλ _VU , but with a given mismatch with respect to the latitude argument ΔФ, after which the first maneuver is performed in the angular inter Δφ ₁ from the point of launch to the target orbit to change the longitude of the ascending node of the orbit by Δλ _VU1 by generating a pulse ΔV ₁ , and then in the opposite direction in the angular interval Δφ ₂ perform the second maneuver with throttle thrust of the engine of the active space object to eliminate the mismatch in the inclination of the target orbit Δi _1, arising after the first maneuver, and to create additional deviation Δλ longitude of ascending node _BV2 by forming pulse ΔV _2, resulting in the coincidence orbital planes junction Mykh objects.

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