RU2178761C1 - Method of control of moment of momentum of spacecraft in the course of orbit correction and system for realization of this method - Google Patents

Abstract

FIELD: attitude control systems; control of orbit of center of mass of spacecraft. SUBSTANCE: proposed method includes determination required magnitude of rate of correction of spacecraft orbit and maintenance of preset attitude of spacecraft by means of powered gyroscopes in the course of correction. Preset area where vector of moment of momentum of spacecraft shall lie at the moment of completion of orbital correction is predicted at this. At absence of saturation in gyroscope system, accumulation of moment of momentum of spacecraft is predicted from the present moment of correction by steps in time taking into account disconnection of one of attitude-control motors or connection of pair of such motors. Magnitudes of vector of moment of momentum of spacecraft predicted at steps are summed-up with its preset magnitude. If total vector lies in preset area, control moment are applied to spacecraft from the moment of termination of prediction by disconnecting one of the above-mentioned motors. If these moments are not unloading moments, pair of motors not used in correction of orbit whose moment has maximum projection on direction opposite to vector of moment of momentum of spacecraft are started. No motor is disconnected at this. Provision is also made for control system including units required for realization of all operations. EFFECT: minimization of starts of attitude-control motors for unloading powered gyroscopes; reduced consumption of propulsive mass; reduced effect on spacecraft orbit. 3 cl, 8 dwg

Description

 The invention relates to the field of controlling the orientation and movement of the center of mass of spacecraft (SC) and is aimed at obtaining the maximum value of the correction speed of the SC orbit with minimal control errors while using jet orientation engines (DO) and power gyroscopes (SG) and ensuring a minimum number of system saturations SG during the execution of the spacecraft flight program after orbit correction, and, therefore, minimizing the flow of the working fluid and the impact on the orbit, due to the need to include of DO to unload the accumulated kinetic moment of SG.

 A known method of controlling a spacecraft equipped with jet engines with angled axes of the connected basis and displaced relative to the center of mass of the apparatus lines of action of rods (see [1]). In the proposed method, the control moments of the thrusts of jet engines (RD) installed by four equal-draft sections in two parallel base planes (north-east) and equidistant planes are applied to the spacecraft hull along three axes of the connected basis, providing a summation of the projections of these moments on axis of the required control torque and compensation for the remaining axes. In this case, the correction of the spacecraft’s orbit in two directions (connected axes, for example, “north-south” and “west-east”) is performed by the thrusts of the taxiway, the projections of which coincide with these directions (axes), and disturbing moments along these axes are mutually compensated. To create a moment opposite the perturbing along the third axis, RDs are used with mutual compensation of their moments and thrusts along the indicated axes. The block diagram of a system that implements an analogue method presented in [1] consists of a block of correction speed (BSC), a block of sensors of external information of the reference basis (BDVI), a block for determining control forces and moments (BOUSM), a block for generating control forces and moments (BFUSM), a block of determined pulse duration values DO (BOZPID), a block DO (BDO), a spacecraft body (CCA), a block of angular velocity sensors (BDUS), a block of linear acceleration sensors (BDLU), a block for setting the control parameters of a spacecraft (BZPUKA) ) In this case, the outputs of the BSK and BDVI are connected respectively to the first and second output of the BOSM, and the output of the BOSM is connected to the first input of the BOSM. The output of the BFUSM is connected to the first input of the BOZPID, and the output of the BOZPID is connected to the input of the BDO. In turn, the BDO is connected to the second input of the BOZPID. On the spacecraft body, BDO, BDUS and BDLU are installed. In this case, the BDLU output is connected to the third input of the BOSM, and the BDLU output is connected to the first input of the BSK. The first, second, third and fourth outputs of the BZPUK are connected respectively to the BDVI input, the second BSK input, the fourth BOUSM input and the second BFUSM input. This method and the control system developed on its basis make it possible to reduce the number of taxiways and fuel costs for the correction of the spacecraft’s orbit, but they do not solve the important issue of obtaining the maximum correction speed. In addition, in the proposed method, the control of the orientation of the spacecraft in the process of correcting the orbit using the RD is not very accurate, and this is essential, for example, for communication satellites, the orientation accuracy of which should be about 12 '.

в текущие моменты времени (t), проверяют выполнение условия принадлежности значений

области располагаемых значений вектора кинетического момента в системе силовых гироскопов (S) и в случае насыщения системы СГ в момент времени t_s определяют суммарное значение векторов управляющих моментов от двигателей ориентации при условии поочередного отключения каждого i-го двигателя, где i= 1, 2, . . . n - номера ДО, участвующих в коррекции орбиты, создают разгрузочный момент для системы силовых гироскопов двигателями ориентации, суммарный момент которых имеет наибольшую проекцию на направление, противоположное вектору

при соответствующем отключенном двигателе ориентации, а в случае, когда этот управляющий момент не является разгрузочными, для разгрузки силовых гироскопов включают ту пару двигателей ориентации, не участвующих в коррекции орбиты, момент которой имеет наибольшую проекцию на направление, противоположное вектору

при этом никакой из указанных i-ых двигателей не отключают, в процессе проведения коррекции орбиты и разгрузки силовых гироскопов прогнозируют изменения указанного суммарного вектора кинетического момента для случая коррекции орбиты с учетом всех работающих указанных i-ых двигателей на интервале от текущего момента времени разгрузки до расчетного момента времени окончания коррекции, суммируют указанные спрогнозированные изменения вектора с текущим значением суммарного вектора

определенным на момент начала прогноза

проверяют условие принадлежности полученных векторных сумм указанной области S и одновременно условие непринадлежности

области S.The closest analogue of the spacecraft control using reactive executive bodies (see [2]), used as a prototype, includes determining the required value of the spacecraft’s orbit correction speed, maintaining a given spacecraft orientation using power gyroscopes in the process of orbit correction by orientation engines, measurement in this case, the values of the kinetic moment vector in the SG system, according to the known values of the SC inertia moments, as well as the measured values of the angular velocity vector of the SC and the kinetic moment in the SG system limit the values of the total vector of the kinetic moment of the spacecraft

at current points in time (t), check that the condition for the values to belong

the range of available values of the kinetic momentum vector in the system of power gyroscopes (S) and in the case of saturation of the SG system at time t _s determine the total value of the vectors of control moments from the orientation engines provided that each i-th engine is turned off, where i = 1, 2, . . . n - DO numbers involved in orbit correction create an unloading moment for the system of power gyroscopes with orientation engines, the total moment of which has the greatest projection in the direction opposite to the vector

with the corresponding orientation motor switched off, and in the case when this control moment is not unloading, to unload the power gyroscopes, include that pair of orientation engines that are not involved in the orbit correction, the moment of which has the largest projection in the direction opposite to the vector

however, none of the indicated i-th engines is turned off, during the correction of the orbit and unloading of power gyroscopes, changes in the indicated total vector of the kinetic moment are predicted for the case of orbit correction taking into account all the working indicated i-th engines in the interval from the current moment of unloading to the calculated the time of the end of the correction, summarize the specified predicted changes in the vector with the current value of the total vector

defined at the time of the start of the forecast

check the membership condition of the obtained vector sums of the indicated domain S and simultaneously the non-membership condition

area S.

области S вплоть до завершения коррекции и в случае невыполнения этого условия повторяют разгрузку системы СГ при помощи двигателей ориентации, создающих указанный момент, или при помощи разгрузочной пары двигателей ориентации, не участвующих в процессе коррекции орбиты.If both of these conditions are not met at this point in time, they continue to carry out orbit correction with the simultaneous unloading of power gyroscopes, and if at least one of these conditions is fulfilled, the unloading of power gyroscopes is stopped by connecting the indicated disabled i-th engine to the orbit correction or shutdown process of the indicated discharge pair DO, after which they continue to verify that the conditions for the specified vector of total kinetic moment belong

region S until the completion of the correction and if this condition is not fulfilled, the unloading of the SG system is repeated using the orientation engines that create the specified moment, or using the unloading pair of orientation engines that are not involved in the orbit correction process.

 The block diagram of a system that implements the prototype method consists of the following blocks: a correction speed block (BSK), a block of sensors of external information of the reference basis (BDVI), a block for determining control forces and moments (BOSM), a block for generating control forces and moments (BFUSM ), a block of detectable pulse duration values DO (BOZPID), a block DO (BDO), a spacecraft body (CCA), a block of angular velocity sensors (BDUS), a block of linear acceleration sensors (BDLU), a block for setting control parameters of a spacecraft (BZPUKA), a block power gyroscopes (BSG), the unit for determining the value of the total kinetic moment vector (BOSVSKM), the unit for comparing the parameters of the domain of available values of the kinetic moment vector in the SG system and the total SG kinetic moment vector (BSPORZVKMSGSVKM), the unit for generating the current SG unloading time (BFVVSG), the block of predicted values of the total kinetic moment vector (BPZSVKM ), a unit for bringing the predicted values of the total vector of kinetic momentum to the initial conditions (BPPZSVKMNU), a unit for determining the time for the end of the correction mode (BOVORK ) BFUSM contains the following blocks: block selection DO to correct the orbit (BVDKO); a unit for forming an unloading moment for the SG system (BFRMSG); unit for determining discharge moments for the SG system (BORMSG); unit for unloading torque for the SG system (BVRMSG); a block for selecting a pair of DOs for unloading the SG system (BVPDRSG).

 In this case, the outputs of the BSK and BDVI are connected respectively to the first and second output of the BOSM, and the output of the BOSM is connected to the first input of the BOSM. The output of the BFUSM is connected to the first input of the BOZPID, and the output of the BOZPID is connected to the input of the BDO. In turn, the BDO is connected to the second input of the BOZPID. On the spacecraft body, BDO, BDUS and BDLU are installed. In this case, the BDLU output is connected to the third input of the BOSM, and the BDLU output is connected to the first input of the BSK. The first, second, third and fourth outputs of the BZPUK are connected respectively to the BDVI input, the second BSK input, the fourth BOUSM input and the second BFUSM input.

 The BSG input is connected to the second input of the BOSM, and the output of the BSG is connected to the first input of the BOSM. The first output of the BOZSVKM is connected to the first input of the BSPORZVKMSGSVKM, and the second output of the same block is connected to the first input of the BSPORZVKMSGSVKM. The second input BOZSVKM is connected to the output BDUS. The output of the BFTVRSG is connected to the third input of the BOZSVKM and the first inputs of the BPZSVKM and BOVORK. The first output of the BSPORZVKMSGSVKM is connected to the first input of the BFTVSG, and the second output to the third input of the BFUSM. The third output of the BSPORZVKMSGSVKM is connected to the fourth input of the BFUSM and the second input of the BFVVSG. The second input of the BSPORZVKMSGSVKM is connected to the fifth output of the BZPUK, and the fourth input of the same block is connected to the first output of the BPPZVKMNU.

 The output of the BPZSVKM is connected to the second input of the BPPZSVKMNU. The second input BPZSVKM is connected to the first output BOVORK, the third and fourth inputs of the same block are connected respectively with the third output BFUSM and the sixth output BZPUKA.

 The second input of the BOVORK is connected to the second output of the BSK, and the third input is connected to the second output of the BFUSM.

 The first input of the BVDKO is connected to the BOSM, and the second input to the fourth output of the BZPUK. The first output of the BVDCO is connected to the first input of the BFRMSG and the second input of the BORMSG. The second output of the same block is connected to the third output of the BFUSM. The first output of BFRMSG is simultaneously the first output of BFUSM, and the second output is the second output of BFUSM. The second input of BFRMSG is connected to the output of BVRMSG, and the third input of this block is the fourth input of BFUSM.

 The first input of the BORMSG is connected to the third input of the BFUSM, and the third input of the indicated block is connected to the fourth output of the BZPUK. The first output of the BORMSG is connected to the second input of the BVRMSG, and the second output of the same block is connected to the first input of the BVRMSG. The first input BVRMSG connected to the third input of BFUSM.

 The second input of the BVPDRSG is connected to the output of the BZPUK, and the third input is connected to the fourth input of the BFUSM. The first output of the BVDPRSG is connected to the first output of the BFUSM, and the second output of the same block is connected to the third output of the BFUSM.

При этом непосредственно значение

определяется в БОЗСВКМ по указанным измеренным значениям, поступающим на его второй вход с выхода БСГ. На первый вход БОЗСВКМ с БДУС поступает информация о значениях вектора абсолютной угловой скорости КА

В самом БОЗСВКМ имеются заранее заданные значения компонент тензора инерции КА

В результате в нем определяется значение вектора

Со второго выхода БОЗСВКМ значения

поступают на второй вход БСПОРЗВКМСГСВКМ. В БСПОРЗВКМСГСВКМ производится проверка выполнения условия принадлежности вектора кинетического момента КА допустимой области S. При этом параметры области S могут изменяться в зависимости от числа работающих в системе силовых гироскопов. Указанные изменения в БСПОРЗВКМСГСВКМ производятся при помощи БЗПУКА, шестой выход которого соединен с шестым входом БСПОРЗВКМСГСВКМ. По этому же каналу в БСПОРЗВКМСГСВКМ передается разрешение на поиск времени начала формирования кинетического момента КА. В случае невыполнения указанного условия с третьего выхода указанного блока на второй вход БФТВРСГ выдается значение времени t_S. Одновременно со второго выхода этого же блока на третий вход БФУСМ выдается значение вектора

Как только БФТВРСГ получает значение момента времени t_S, он тут же формирует на своем выходе значение t'₁ в соответствии с выражением

η = 1,2,3,..., где Δt - продолжительность минимального импульса разгрузочного момента

или

В свою очередь t'₁ передается на третий вход БОЗСВКМ, на третий вход и на четвертый вход БПЗСВКМ. По приходу в БОЗСВКМ значения t'₁ в нем происходит присвоение текущего значения вектора кинетического момента

значению

Значение вектора

поступает с первого выхода БОЗСВКМ на первый вход БППЗСВКМНУ и на третий вход БСПОРЗВКМСГСВКМ. В БППЗСВКМНУ оно служит в качестве начального условия в выражении

а в БСПОРЗВКМСГСВКМ оно необходимо для проверки условия

Значение t'₁, переданное в БПЗСВКМ, устанавливает нижнюю границу определенного интеграла, входящего в выражение

где

главный вектор управляющего момента от всех работающих ДО без учета отключений i-х ДО или пар ДО;

главный вектор управляющего момента всех внешних возмущающих сил. Оно же задает в БОВОРК момент времени t_k1 начала определения второй верхней границы указанного интеграла. Так, по приходу этого времени в БОВОРК из первого выхода БСК через первый вход БОВОРК считывается значение

Одновременно из первого выхода БФУСМ на второй вход БОВОРК приходит информация о номерах ДО, участвующих в коррекции орбиты. По указанным номерам в блоке определяется

Для этого используются известные силы тяги ДО и масса КА и производится расчет t_k1 по выражению

где

измеренное значение скорости коррекции КА на момент времени t'₁. Далее информация о значении t_k1 передается на пятый вход БПЗСВКМ.The control system operates as follows. At the beginning of the execution of the orbit correction mode from the second output of the BOSM to the second input of the BFUSM, requirements are issued for the formation of control forces F _i . Simultaneously, from the first output of the BOSM to the first input of the BSG, requirements are issued for the formation of control moments. BSG forms the law of control of the SG precession axes (see [3] p. 12 - 13) according to the required control moments. In this case, due to measurements carried out in the block of sine-cosine converters of the rotation angles of the precession axes, which is included in the BSG sucker, as well as from the known values of the kinetic moments of the gyro rotors, the value of the kinetic moment vector of the SG system is determined

In this case, the value

it is determined in BOZSVKM according to the specified measured values supplied to its second input from the output of the BSG. Information on the values of the vector of the absolute angular velocity of the spacecraft is received at the first input of the BOSSVKM from the BSS

In the BOSSCM itself there are predefined values of the components of the inertia tensor of the spacecraft

As a result, it determines the value of the vector

From the second output of BOZSVKM values

arrive at the second input of the BSPORZVKMSGSVKM. At BSPORZVKMSGSVKM, the conditions of membership of the SC kinetic moment vector of the admissible region S are checked. In this case, the parameters of region S can vary depending on the number of power gyros operating in the system. The indicated changes in the BSPORZVKMSGSVKM are made using the BZPUK, the sixth output of which is connected to the sixth input of the BSPORZVKMSGSVKM. On the same channel, permission is sent to the BSPORZVKMSGSVKM to search for the start time for the formation of the kinetic moment of the spacecraft. In case of failure to fulfill the specified condition from the third output of the specified block to the second input BFTVRSG issued the value of time t _S. At the same time, from the second output of the same block, the vector value is output to the third input of the BFUSM

As soon as BFTVRSG receives the value of the moment of time t _S , it immediately forms at its output the value t ' ₁ in accordance with the expression

η = 1,2,3, ..., where Δt is the duration of the minimum impulse of the unloading moment

or

In turn, t ' ₁ is transmitted to the third input of the BOZSVKM, to the third input and to the fourth entrance of the BPZSVKM. Upon the arrival of the value t ' ₁ in the MHSSMC, the current value of the kinetic moment vector is assigned in it

meaning

Vector value

It comes from the first output of the BOZSVKM to the first entrance of the BPPZSVKMNU and to the third entrance of the BSPORZVKMSGSVKM. In BPPZSVKMNU it serves as an initial condition in the expression

and in BSPORZVKMSGSVKM it is necessary to verify the conditions

The value of t ' ₁ , passed to the BZSVKM, sets the lower boundary of a certain integral in the expression

Where

the main vector of the control moment from all operating BSs without taking into account disconnections of i-BSs or pairs of BSs;

the main vector of the control moment of all external disturbing forces. In BOVORK, it also sets the time t _{k1 of the} start of determining the second upper boundary of the indicated integral. So, upon the arrival of this time in BOVORK, the value is read from the first output of BSC through the first entrance of BOVORK

At the same time, from the first exit of the BFUSM to the second entrance of the BOVORK, information is received about the numbers of subsidiaries involved in the correction of the orbit. By the indicated numbers in the block is determined

To do this, the known traction forces of DO and the mass of the spacecraft are used and t _{k1 is} calculated by the expression

Where

the measured value of the correction speed of the spacecraft at time t ' ₁ . Further, the information on the value of t _k1 is transmitted to the fifth input of the BPSVCM.

который он создает. Причем значение указанных управляющих моментов задается в БПЗСВКМ через первый его вход с седьмого выхода БЗПУКА. Для ДО, не участвующих в управлении,

По указанному каналу в БПЗСВКМ выдаются также значения, определяющие значения возмущающих моментов

(углы ориентации КА, положение солнечных батарей и т. д. ). С третьего выхода БФУСМ на третий вход БПЗСВКМ передаются также номера ДО для вектора

По значениям

в БПЗСВКМ определяется

Далее

суммируется с полученным там же

Затем производится интегрирование на интервале установленных границ. Полученные интегральные значения передаются со второго выхода БПЗСВКМ на второй вход БППЗСВКМНУ, где производится их суммирование с начальными значениями

Полученная сумма с выхода БППЗСВКМНУ передается на первый вход БСПОРЗВКМСГСВКМ, где производится проверка выполнения условия ее принадлежности области S. Если в БСПОРЗВКМСГСВКМ при проверке не выполняется это условие и выполняется условие принадлежности области S текущего значения вектора кинетического момента, то значение t_S подтверждается в БФТВРСГ и повторяется рассмотренный выше цикл разгрузки СГ на момент времени t'₂ и т. д.Information about DO numbers, originally involved in orbit correction, is also issued from the third output of the BFUSM to the third entrance of the BPZSVKM. In this case, each engine is assigned a control moment

which he creates. Moreover, the value of these control moments is set in BPZSVKM through its first input from the seventh output of the BZPUKA. For subsidiaries not involved in management,

On the specified channel in BPZSVKM also issued values that determine the values of disturbing moments

(spacecraft orientation angles, position of solar panels, etc.). From the third output of the BFUSM to the third input of the BPZSVKM, the DO numbers for the vector are also transmitted

By values

in BPZSVKM is determined

Further

sums up with received there

Then integration is performed over the interval of established boundaries. The obtained integral values are transferred from the second output of the BPZSVKMU to the second input of the BPPZSVKMNU, where they are summed with the initial values

The amount received from the output of the BSPZSVKMNU is transferred to the first input of the BSPORZVKMSGSVKM, where the condition of its belonging to the region S is checked. If the BSPORZVKMSGSVKM does not satisfy this condition and the condition that the region S belongs to the current value of the kinetic moment vector is verified, then the value of t _{S is} confirmed in the BFVVSR and the above-described SG discharge cycle is repeated at time t ′ ₂ , etc.

If, in the BSPORZVKMSGSVKM, when checking, the condition that the kinetic moment vector of the spacecraft S belongs to region S is fulfilled or the condition that the S region does not belong to the current value of the kinetic moment vector, then from the first output of the BSPORZVKMSGSVKM, the command to remove the SG unloading mode is issued to the fourth input of the BFUSM, and the first input of the BFTVSRG is to terminate the search mode of time t _z .

 The disadvantage of the control method and system described in the prototype is that at the end of the orbit correction process for the total kinetic moment of the spacecraft, the only condition is that it is in the allowable region, and a further change in the kinetic moment in the course of the spacecraft’s movement in the corrected orbit is not predicted. But the value obtained at the moment of the end of the orbit correction of the total value of the kinetic moment of the spacecraft can be such that during the execution of the next flight program this value will go beyond the allowable region S and, as a result, there will be a need to unload the SG system with the help of DO, which is not desirable , because it will lead to additional consumption of the working fluid and the deterioration of the orbit caused by the operation of DO.

 The problem solved by the proposed method and system is to control the kinetic moment of the spacecraft in the process of correcting the orbit of the spacecraft, which will ensure the formation at the time of completion of the correction of the required value of the vector of kinetic momentum of the spacecraft, which will allow the further execution of the flight program to minimize the number of inclusions of DO for unloading SG and, therefore, minimize both the flow rate of the working fluid and the effect on the orbit of the spacecraft.

в текущие моменты времени (t), проверяют выполнение условия принадлежности значений суммарного вектора кинетического момента космического аппарата

области располагаемых значений вектора кинетического момента в системе силовых гироскопов (S),
в отличие от известного способа прогнозируют заданную область (S_k), в которой должен находиться суммарный вектор кинетического момента космического аппарата

в момент времени завершения процесса коррекции орбиты, если в процессе коррекции орбиты насыщение системы силовых гироскопов не происходит, то с текущего момента времени коррекции (t_η), на которое имеются измерения вектора кинетического момента космического аппарата, включая начало коррекции, шагами по Δt до момента времени

прогнозируют накопление кинетического момента КА

с учетом отключения i-го двигателя ориентации или включения пары двигателей ориентации для формирования требуемого кинетического момента космического аппарата в момент времени

η = 1,2,3,..., где Δt- продолжительность минимального импульса момента

при условии включения пары двигателей ориентации или момента

при условии отключения i-го двигателя ориентации, участвующего в коррекции орбиты,

V_U - величина, определяющая заданную скорость коррекции орбиты,

текущее значение скорости коррекции орбиты на момент времени t_η, a_Σ- расчетное значение ускорения космического аппарата от двигателей ориентации, участвующих в коррекции орбиты, суммируют спрогнозированные значения вектора

с текущим значением суммарного вектора

определенным на момент начала прогноза t_η, проверяют выполнение условия принадлежности полученных значений суммарного вектора кинетического момента космического аппарата заданной области S_k, и если оно выполняется, то по достижению временем коррекции значения, равного

, определяют суммарное значение векторов управляющих моментов

от двигателей ориентации при условии отключения i-го двигателя, где i= l, 2, . . . n - номера двигателей ориентации, участвующих в коррекции орбиты, создают двигателями ориентации момент для системы силовых гироскопов, суммарный момент которых имеет наибольшую проекцию на направление, противоположное вектору

при соответствующем отключенном двигателе ориентации, а в случае, когда управляющие моменты

не являются разгрузочными, для формирования требуемого кинетического момента включают ту пару двигателей ориентации, не участвующих в коррекции орбиты, момент которой имеет наибольшую проекцию на направление, противоположное вектору

при этом никакой из указанных i-х двигателей не отключают.The problem is solved in that in the proposed method for controlling the kinetic moment of a spacecraft using reactive actuators, which includes determining the required value of the orbit correction speed of the spacecraft, maintaining a given orientation of the spacecraft using power gyroscopes in the process of correcting the orbit by orientation engines, measuring the vector of the kinetic moment in the system of power gyroscopes, according to the known values of the moments of inertia of the space app arata, as well as from the measured values of the angular velocity vector of the spacecraft and the kinetic moment in the system of power gyroscopes, determine the values of the total vector of the kinetic moment of the spacecraft

at current points in time (t), verify that the conditions for the values of the total vector of the kinetic moment of the spacecraft to belong

the range of available values of the kinetic moment vector in the system of power gyroscopes (S),
in contrast to the known method, a predetermined region (S _k ) is predicted in which the total vector of the kinetic moment of the spacecraft should be

at the time of completion of the orbit correction process, if the system of gyroscopes does not saturate during the correction of the orbit, then from the current moment of correction time (t _η ), for which there are measurements of the vector of the kinetic moment of the spacecraft, including the beginning of correction, in steps of Δt until time

predict the accumulation of the kinetic moment of the spacecraft

taking into account the shutdown of the i-th orientation engine or the inclusion of a pair of orientation engines to form the required kinetic moment of the spacecraft at a time

η = 1,2,3, ..., where Δt is the duration of the minimum momentum

subject to the inclusion of a pair of orientation engines or torque

provided that the i-th orientation engine participating in the orbit correction is turned off,

V _U - a value that determines a given speed correction of the orbit,

the current value of the orbit correction speed at the time t _η , and _Σ is the calculated value of the acceleration of the spacecraft from orientation engines participating in the correction of the orbit, the predicted values of the vector are summed

with the current value of the total vector

defined at the start of the forecast t _η , verify that the conditions of membership of the obtained values of the total vector of the kinetic moment of the spacecraft of the given region S _k are checked, and if it is satisfied, then when the correction time reaches a value equal to

determine the total value of the vectors of control moments

from orientation engines provided that the i-th engine is disconnected, where i = l, 2,. . . n - numbers of orientation engines participating in the orbit correction, create orientation moment by the orientation engines for the system of power gyroscopes, the total moment of which has the greatest projection in the direction opposite to the vector

with the corresponding orientation motor switched off, and in the case when the control moments

are not unloading, to form the required kinetic moment, they include that pair of orientation engines that are not involved in orbit correction, the moment of which has the largest projection in the direction opposite to the vector

however, none of the indicated i-th engines are turned off.

 The problem is solved in that in the control system of the kinetic moment of the spacecraft with the help of reactive executive bodies, containing a block of the orbit correction speed of the spacecraft, a block of sensors of external information of the reference basis, a block for determining control forces and moments, a block of orientation engines, a block of sensors of angular velocity of the space apparatus, a block of sensors of its linear accelerations, a block for setting parameters for controlling the spacecraft, a block for determining the end time of the correction mode ii, a unit for bringing the predicted values of the total vector of the kinetic moment to the initial condition, a block of power gyroscopes, a unit for determining the values of the total vector of the kinetic moment, a unit for generating control forces and moments, consisting of a unit for selecting orientation engines for correcting the orbit, a unit for generating the unloading moment for the power system gyroscopes, a unit for determining discharge moments for a system of power gyroscopes, a unit for selecting an unloading moment for a system of power gyroscopes, a unit in bores of a pair of orientation engines for unloading the system of power gyroscopes, a block of determined values of the duration of the thrust impulses of the orientation engines, the outputs of these correction speed blocks and external information sensors of the reference basis are connected respectively to the first and second inputs of the control force and moment determination unit, and the first output of the last specified the block is connected to the first input of the block forming the control forces and moments, the first output of the block forming the control forces and moments is connected to the first input of the block of determined values of the duration of the pulses of the thrust of the orientation engines, and the output of the last specified block is connected to the input of the block of the engines of the orientation, the output of the block of the engines of orientation is connected to the second input of the block of the determined values of the duration of the pulses of orientation motor, the output of the block of sensors of linear accelerations is connected to the first input of the block correction speed of the orbit of the spacecraft, and the output of the block of sensors of the angular velocity of the spacecraft is connected to the third input lok of determining the control forces and moments and the second input of the unit for determining the values of the total vector of the kinetic moment, and the second output of the unit for determining the control forces and moments is connected to the input of the unit of power gyroscopes, and the output of the latter is with the first input of the unit for determining the values of the total vector of kinetic moment, the second output the unit for determining the values of the total vector of the kinetic moment is connected to the first input of the unit for bringing the predicted values of the total vector of the kinetic moment to the initial Under the given conditions, the second output of the control forces and moments formation unit is connected to the third input of the correction mode end time determination unit, and the second input of the correction mode end time determination unit is connected to the second output of the spacecraft’s orbit correction speed unit, the first output of the spacecraft control parameters connected to the input of the block of external information sensors of the reference basis, the second output to the second input of the block of the orbit correction speed of the spacecraft, the third the output is with the fourth input of the control force and moment determination unit, the fourth output is with the second input of the control force and moment formation unit, the first input of the orientation engine selection unit for orbit correction being connected to the control force and moment determination unit, and the second input to the fourth by the output of the spacecraft control parameter setting block, the first output of the orientation engine selection block for orbit correction is connected to the first input of the unloading moment formation block for the power train system telescopes and the second input of the unit for determining the discharge moment for the system of power gyroscopes, the second output of the unit for generating the discharge moment for the system of power gyroscopes is simultaneously the second output of the unit for generating control forces and moments, and the first output is the first output of the block for generating control forces and moments, the second input of the block the formation of the discharge torque for the system of power gyroscopes is connected to the output of the block selection of the discharge torque for the system of power gyroscopes, the second input of the selection block and the unloading moment for the power gyroscope system is connected to the first output of the unloading moment determination unit for the power gyroscope system, the third input of the unloading moment determination unit for the power gyroscope system is connected to the fourth output of the spacecraft control parameter setting unit, the second output of the unloading moment determination unit for the power system gyroscopes connected to the first input of the block selection pair of orientation motors for unloading the system of power gyroscopes, and the second input after day - with the fourth output of the unit for setting the control parameters of the spacecraft, the first output of the unit for selecting a pair of orientation engines for unloading the system of power gyroscopes is connected to the first output of the unit for generating control forces and moments, in contrast to the known system, a unit for determining a given region of the kinetic momentum is additionally included, block comparing the current time of correction of the orbit of the spacecraft and the estimated time of the beginning of the formation of the total vector of the kinetic moment of the space ap arata, a unit for comparing the parameters of a given region of the kinetic moment of the spacecraft and the total vector of the kinetic moment of the spacecraft, a unit for setting the current time of formation of the kinetic moment of the spacecraft, a block of predicted values of the total vector of kinetic moment in the process of forming the total vector of kinetic moment, block for determining the time of the beginning of the formation of kinetic moment, and the first input of the unit for comparing the parameters of a given region of the kinetic m the spacecraft momentum and the total vector of the kinetic moment of the spacecraft is connected to the first output of the unit for determining the values of the total vector of the kinetic moment, the second input is the fifth output of the unit for setting the control parameters of the spacecraft, the third input is with the second output of the unit for determining the values of the total vector of kinetic moment, fourth input - with the first output of the unit for bringing the predicted values of the total vector of kinetic momentum to the initial conditions, the fifth input - with the output m of the unit for determining a given region of the kinetic moment, the sixth input — with the output of the unit for comparing the current time for correcting the orbit of the spacecraft and the estimated time for the formation of the total vector of the kinetic moment of the spacecraft, the first output of the same block is connected to the first input of the unit for setting the current time for the formation of the kinetic moment of the space apparatus, the second output of this block is connected to the third input of the block forming the control forces and moments, the third output is with the fourth input of the bl as for the formation of control forces and moments and the second input of the unit for setting the current time of formation of the kinetic moment of the spacecraft, the output of the unit for setting the current time of formation of the kinetic moment of the spacecraft is connected to the first input of the unit for determining the end time of the correction mode, the third input of the unit for determining the values of the total vector of kinetic moment and the first input of the block of predicted values of the total vector of kinetic moment in the process of formation of the total vector kinetic moment, the second input of the predicted value vector kinetic moment vector block during the formation of the total kinetic moment vector is connected to the first output of the correction mode end time determination unit, the third input - with the third output of the control forces and moments formation unit, the fourth input - with the sixth output unit for setting control parameters of the spacecraft, the first output of the same unit is connected to the second input of the unit for bringing the predicted values of the total vector kinetic moment to the initial conditions, the first input of the unit for determining the specified region of the kinetic moment of the spacecraft is connected to the seventh output of the control unit for controlling the spacecraft, the second input is the output of the angular velocity sensor unit, the third input is the output of the linear acceleration sensor unit, fourth input - with the output of the power gyroscopes unit, the first input of the unit for determining the start time of the formation of the kinetic moment is connected to the ninth output of the control parameter setting unit spacecraft, the second input - with the output of the block of angular velocity sensors, the third input - with the output of the linear acceleration sensor block, the fourth input - with the output of the power gyroscope block, the fifth input - with the second output of the block for determining the end time of the correction mode, the first input of the current comparison unit the time of correction of the orbit of the spacecraft and the estimated time of the beginning of the formation of the total vector of the kinetic moment of the spacecraft is connected to the eighth output of the control parameter set of the spacecraft apparatus, the second input - with the output of the unit for determining a given region of the kinetic moment, the second output of the unit for selecting orientation engines is connected to the third input of the predicted values of the total vector of kinetic moment in the process of forming the total vector of kinetic moment, the first input of the unit for determining the unloading moments for the system of power gyroscopes and the first input of the unit for selecting the discharge torque for the power gyro system is connected to the second output of the unit for comparing the parameters of a given region the kinetic moment of the spacecraft and the total vector of the kinetic moment of the spacecraft, the third input of the unit for selecting a pair of orientation engines for unloading the system of power gyroscopes and the third input of the unit for generating the unloading moment for the system of power gyroscopes are connected to the third output of the unit for comparing the parameters of a given region of the kinetic moment of the spacecraft and the total vector of the kinetic moment of the spacecraft, the second output of the block for selecting a pair of orientation engines for p zgruzki gyroscopes power system connected to a third input of the block of predicted values of the total angular momentum vector in the formation of the total angular momentum vector.

 The proposed method and system are devoid of the inherent disadvantages of the prototype, because in order to minimize the flow of the working fluid and the effect on the orbit, after the completion of the correction of the spacecraft’s orbit, due to the need to turn on the DO for unloading the SG system, when the spacecraft flight program is executed for a specified time interval the spacecraft kinetic moment at the end of the orbit correction process is brought into a given region, which is determined from the condition of minimizing the number of saturations of the SG system when the spacecraft flight program is executed, and, trace quently, and inclusions required for it to discharge.

(ССК). На фиг. 2 показан корпус КА, положение центра масс КА в плоскости XOZ ССК и направление действия векторов сил тяги ДО. На фиг. 3, 4, 5 приведены графики изменения кинетического момента КА при поддержании неизменной ориентации аппарата с помощью СГ с разными начальными условиями по кинетическому моменту КА. Фиг. 6 иллюстрирует пример реализации части системы управления. На фиг. 6 показана система управления, разработанная для реализации предлагаемого способа управления кинетическим моментом КА в процессе коррекции орбиты. На фиг. 7 изображены составляющие БФСУМ и их связь с остальными элементами предлагаемой системы управления. На фиг. 8 показан пример системы управления, разработанной для реализации предлагаемого способа.To clarify the essence of the proposed method are shown in FIG. 1, 2, 3, 4, 5, 6, 7, 8. In FIG. 1 shows the spacecraft body, orientation engines, and the coordinate system rigidly connected with the spacecraft

(SSK). In FIG. Figure 2 shows the spacecraft body, the position of the center of mass of the spacecraft in the XOZ SSK plane and the direction of action of the thrust force vectors DO. In FIG. Figures 3, 4, 5 show graphs of changes in the kinetic moment of the spacecraft while maintaining a constant orientation of the apparatus using SG with different initial conditions for the kinetic moment of the spacecraft. FIG. 6 illustrates an example implementation of a part of a control system. In FIG. 6 shows a control system designed to implement the proposed method for controlling the kinetic moment of the spacecraft in the process of orbit correction. In FIG. 7 shows the components of the BFSUM and their relationship with the remaining elements of the proposed control system. In FIG. 8 shows an example of a control system designed to implement the proposed method.

 To implement the proposed spacecraft control method, it is necessary first of all to determine a given region of the spacecraft kinetic moment, i.e., the area in which the spacecraft kinetic moment vector should be located after the completion of the orbit correction. The only condition that this region should satisfy is minimization of the number of saturations of the SG when controlling the orientation of the spacecraft in a given time interval. The problem of determining the region of the kinetic moment by modeling the motion of the spacecraft on this time interval is solved. Based on the simulation results, the region in which the vector of the kinetic moment of the spacecraft should be located is determined.

 Further, the proposed method is described in more detail.

Здесь

радиус-вектор КА,

вектор суммарного кинетического момента КА;

- кинетический момент системы СГ;

угловая скорость спутника;

векторы суммарной силы и суммарного момента внешних сил, действующих на КА соответственно;

тензор инерции КА; m - масса КА.To determine the specified region of the kinetic moment of the spacecraft, i.e., the region in which the vector of the kinetic moment of the spacecraft should be located after the completion of the orbit correction, mathematical modeling of the motion of the spacecraft from the moment of completion of the correction of the orbit at a given time interval, for example, until the next correction begins. The system of equations describing the mathematical model of the motion of the center of mass of the spacecraft and its motion relative to the center of mass can be written in the following form:

Here

spacecraft radius vector

vector of the total kinetic moment of the spacecraft;

- kinetic moment of the SG system;

satellite angular velocity;

vectors of total force and total moment of external forces acting on the spacecraft, respectively;

spacecraft inertia tensor; m is the mass of the spacecraft.

(где

вектор скорости КА), вектор его угловой скорости

варьируя исходные
значения вектора

и, следовательно,

, определим заданную область, в которой должен находиться кинетический момент КА на момент завершения процесса коррекции орбиты КА для минимизации расхода рабочего тела при управлении его ориентацией на задаваемом временном интервале. После определения заданной области кинетического момента, зная вектор состояния КА на момент завершения коррекции, т. е. , другими словами, имея результаты решения задачи, необходимо найти начальные значения этих величин, которые будут удовлетворять требуемому результату решения. Т. к. начальные значения вектора состояния КА на момент начала коррекции орбиты известны, то требуется определить начальные значения кинетического момента КА.Solving equations (1), knowing at the initial moment of time the initial values of the spacecraft state vector

(Where

spacecraft velocity vector), its angular velocity vector

varying the source
vector values

and therefore

, we determine the specified region in which the kinetic moment of the spacecraft should be at the time of completion of the process of correction of the orbit of the spacecraft to minimize the flow rate of the working fluid when controlling its orientation on a given time interval. After determining the given region of the kinetic moment, knowing the state vector of the spacecraft at the time of completion of the correction, i.e., in other words, having the results of solving the problem, it is necessary to find the initial values of these quantities that will satisfy the desired solution result. Since the initial values of the spacecraft state vector at the beginning of the orbit correction are known, it is necessary to determine the initial values of the kinetic moment of the spacecraft.

 To determine them, a simulation of the spacecraft orbit correction process is carried out. If during the orbit correction the saturation of the SG system occurs, then it is logical to try to combine the inevitable process of unloading the SG with the formation of such a kinetic moment of the spacecraft that will complete the correction of the orbit of the spacecraft without re-saturation of the SG system and, moreover, will make it possible to obtain a vector at the time of completion of the correction of the orbit kinetic moment of a spacecraft belonging to a given region. Modeling can be carried out, for example, using equations (1). During its implementation, the well-known characteristics of DOs, their influence on the change in the kinetic moment of the spacecraft, the degree of participation in the correction of the orbit are taken into account. The choice of DO should be such as to reduce the influence of the processes associated with the formation of the required kinetic moment of the spacecraft on the orbit correction.

(ССК), смещенное относительно начала ССК положение центра масс (⊕) и направление действия векторов сил тяги ДО D₉, D₁₀, D₁₁, D₁₂

Обозначим момент времени, в который произошло насыщение системы СГ, как t_S, а t_p - момент времени начала формирования кинетического момента КА, если в процессе коррекции орбиты насыщения системы СГ не было. Предположим, что на КА установлена система СГ с областью S, описанной сферой радиусом R_сф.Let us explain the essence of the proposed method by example. To increase the clarity of the solution, FIG. 1 and FIG. 2. In FIG. 1 and FIG. 2 shows the spacecraft body, orientation engines (D ₁ -D ₂₄ ), the coordinate system rigidly connected with the spacecraft

(SSC), the position of the center of mass (⊕) displaced relative to the beginning of the SSC and the direction of action of the traction force vectors DO ₉ , D ₁₀ , D ₁₁ , D ₁₂

Let us designate the point in time at which saturation of the SG system took place, as t _S , and t _p - the time moment of the onset of formation of the kinetic moment of the spacecraft, if during the correction of the orbit of saturation of the SG system there was no. Suppose that a spacecraft system with a region S described by a sphere of radius R _{sf is} installed on the spacecraft.

Suppose, for example, we need to carry out an orbit correction in the direction of the OX axis (see Fig. 1,2). To do this, turn on the engines D ₉ -D ₁₂ . Due to the displacement of the center of mass of the spacecraft by the engines 9, 11, a disturbing moment will be created. This moment violates the required orientation of the spacecraft. Compensation of the disturbing moment arising is carried out by the SG system, which makes it possible to use the thrust of the D ₉ -D ₁₂ engines to the maximum, and, therefore, increase the duration of the output of the correcting pulse and reduce the time of the orbit correction.

Здесь

полный кинетический момент КА;

кинетический момент системы СГ;

- угловая скорость КА;

- соответственно гравитационные моменты, вызванные влиянием на ССС гравитационных полей Земли, Луны и Солнца;

магнитный момент, обусловленный взаимодействием магнитного поля Земли и собственного магнитного момента КА;

орты векторов Земля-КА, Солнце-КА, Луна-спутник;

тензор инерции КА;
μ_E= 3.986032•10⁵км³/с²,
μ_S= 1.32715445•10¹¹км³/с²
μ_M= 4.90264•10⁵км³/с² - гравитационные параметры Земли, Луны и Солнца; R_E, R_S, R_M - радиус-векторы Земли, Луны и Солнца;

- момент от силы светового давления

которая возникает при попадании потока солнечного света на спутник и при его отражении. S - площадь поперечного сечения КА; Е₀ - мощность потока солнечного излучения; с - скорость света; r_* - средний радиус орбиты Земли; Δ- расстояние от КА до Солнца; k - коэффициент отражения света поверхностью КА;

собственный магнитный момент КА;

магнитное поле Земли.Suppose that after the correction of the orbit, the spacecraft should be in the mode of maintaining orientation until the next correction during the time interval Δt ′. Solving the system of equations (1), varying the initial values of the kinetic moment of the SG, we determine the region in which the kinetic moment of the SG should be in order to minimize the flow rate of the working fluid in the time interval Δt ′. As practice shows, in many cases, to predict the accumulation of the kinetic moment of the spacecraft, it is enough to solve the equations of the rotational motion of the spacecraft without a detailed simulation of the motion of the center of mass of the spacecraft. Then the system of equations (1), for example, for a geostationary satellite, can be written in the following form:

Here

full kinetic moment of the spacecraft;

kinetic moment of the SG system;

- angular velocity of the spacecraft;

- respectively, the gravitational moments caused by the influence on the SSS of the gravitational fields of the Earth, the Moon and the Sun;

magnetic moment due to the interaction of the Earth’s magnetic field and the spacecraft’s own magnetic moment;

the unit vectors of the Earth-KA, Sun-KA, Moon-satellite;

spacecraft inertia tensor;
μ _E = 3.986032 • 10 ⁵ km ³ / s ² ,
μ _S = 1.32715445 • 10 ¹¹ km ³ / s ²
μ _M = 4.90264 • 10 ⁵ km ³ / s ² - gravitational parameters of the Earth, the Moon and the Sun; R _E , R _S , R _M - radius vectors of the Earth, the Moon and the Sun;

- moment from the force of light pressure

which occurs when a stream of sunlight hits a satellite and when it is reflected. S is the cross-sectional area of the spacecraft; E ₀ - power flow of solar radiation; c is the speed of light; r _* is the average radius of the Earth’s orbit; Δ is the distance from the spacecraft to the sun; k is the light reflection coefficient by the surface of the spacecraft;

intrinsic magnetic moment of the spacecraft;

Earth's magnetic field.

As practice has shown, the forecast of the accumulation of the kinetic moment of the spacecraft obtained by solving equations (2) for the geostationary communications satellite "Yamal" gives good results. So, the solution will result in a certain region of the kinetic moment of the spacecraft (S _k ), we call it given, in which the kinetic moment of the spacecraft should fall upon completion of the correction of its orbit. After determining the given region of the kinetic moment, the spacecraft correction process is simulated using equations (1).

при условии включения пары двигателей ориентации или момента

при условии отключения i-го двигателя ориентации, участвующего в коррекции орбиты, для формирования заданного кинетического момента космического аппарата. Т. е. с текущего момента времени коррекции (t_η), на которое имеются измерения вектора кинетического момента КА, и до момента времени t_kη прогнозируем накопление кинетического момента с учетом отключения в момент времени

i-го ДО или включения пары ДО для формирования требуемого кинетического момента КА.If saturation of the SG system will not take place during the orbit correction process, then it is necessary to form the kinetic moment of the spacecraft by the time of completion of the correction of the orbit. For this, the time of the beginning of the formation of the kinetic moment is determined. The process of correction of the orbit of a spacecraft is simulated using expressions (1) and (2), starting from the initial moment of time of the correction. Having determined the time of completion of the correction process (t _k ), the value of the kinetic moment of the spacecraft at the time of completion of the correction of the orbit and knowing the dynamics of its change, find the increment of the vector of the kinetic moment of the spacecraft, which he must get in order to be in a given region S _k . Proceeding from this, one chooses DOs capable of reporting this increment in the shortest time, and having determined the duration of the formation mode of the kinetic moment of the spacecraft, they find the instant t _{p of} its beginning. The search t _{p is} performed in steps of Δt is the duration of the minimum momentum

subject to the inclusion of a pair of orientation engines or torque

provided that the i-th orientation engine participating in the orbit correction is turned off to form the given kinetic moment of the spacecraft. That is, from the current moment of correction time (t _η ), for which there are measurements of the vector of the kinetic moment of the spacecraft, and until the time t _{kη, we} predict the accumulation of the kinetic moment taking into account the shutdown at the time

of the ith DO or inclusion of a pair of DOs to form the required kinetic moment of the spacecraft.

Прогноз осуществляем с нулевыми начальными значениями по вектору

где

главный вектор управляющего момента от всех работающих ДО без учета отключений i-ых ДО или включения пар ДО;

главный вектор управляющего момента от всех работающих ДО с учетом отключений i-го ДО или включения пар ДО;

главный вектор управляющего момента всех внешних возмущающих сил. Значение t_kη определяется следующим образом:

где V_t - измеренное значение скорости коррекции КА на момент времени t_η. Спрогнозированное значение

суммируем с реальными начальными условиями по кинетическому моменту

полученными на момент времени t начала прогноза.

The forecast is carried out with zero initial values for the vector

Where

the main vector of the control moment from all operating BSs without taking into account the shutdown of the i-th BSs or the inclusion of pairs of BSs;

the main vector of the control moment from all operating DOs, taking into account the shutdown of the i-th DO or the inclusion of DO pairs;

the main vector of the control moment of all external disturbing forces. The value of t _kη is determined as follows:

where V _t is the measured value of the correction speed of the spacecraft at time t _η . Predicted value

summarize with the real initial conditions for the kinetic moment

obtained at time t of the beginning of the forecast.

Далее проверяем выполнение условия

и если оно выполняется, то t_p= t′_η и в этот момент времени с помощью выбранных ДО начинается управление кинетическим моментом КА, которое приведет его в момент завершения процесса коррекции орбиты в заданную область, иначе повторяем цикл расчетов на момент времени t₂ и т. д. до выполнения условия (7).

Next, check the condition

and if it is satisfied, then t _p = t ′ _η and at this point in time, using the selected DOs, control of the kinetic moment of the spacecraft begins, which will bring it at the moment of completion of the orbit correction process to a given area, otherwise we repeat the calculation cycle at time t ₂ and etc. until condition (7) is satisfied.

и вектор момента от ДО были противоположно направлены, причем чем больше тупой угол между указанными векторами, тем быстрее происходит формирование заданного значения кинетического момента. Вектора моментов при поочередном отключении ДО, участвующих в коррекции орбиты, определяются следующим образом:
- при отключении двигателя D₉:

- при отключении двигателя D₁₀:

- при отключении двигателя D₁₁:

- при отключении двигателя D₁₂:

где

- радиус-векторы точки приложения соответствующих тяг двигателей

где

(i= 9, 10, 11, 12) - радиус-векторы точки приложения соответствующих тяг двигателей

Должно выполняться следующее условие:

Выбрав те ДО, для которых условие (12) выполняется, обозначим их

определяем те ДО, при отключении которых суммарный момент

имеет наибольшую проекцию на направление, противоположное вектору

Математически это выражение можно записать в виде

где ←_→ - знак соответствия.DO for the formation of a given value of the kinetic moment are selected as follows. To achieve maximum efficiency, it is necessary that the vector

and the moment vector from the DO were oppositely directed, and the larger the obtuse angle between the indicated vectors, the faster the formation of a given value of the kinetic moment. The moment vectors during the successive shutdown of DOs involved in orbit correction are determined as follows:
- when turning off the engine D ₉ :

- when turning off the engine D ₁₀ :

- when engine D ₁₁ is turned off:

- when turning off the engine D ₁₂ :

Where

- radius vectors of the point of application of the respective engine rods

Where

(i = 9, 10, 11, 12) are the radius vectors of the point of application of the corresponding engine rods

The following condition must be met:

Having selected those DOs for which condition (12) is satisfied, we denote them

we determine those DOs, when disconnected which total moment

has the largest projection in the direction opposite to the vector

Mathematically, this expression can be written as

where ← _ → is the sign of compliance.

будет перпендикулярен вектору

или когда возмущающий момент от каждого из работающих для коррекции орбиты двигателей направлен по трем осям КА, а вектор

накапливается таким образом, что условие (12) выполнить невозможно. В таких случаях предлагается использовать пару ДО, не участвующих в коррекции орбиты. При этом двигатели, выполняющие коррекцию орбиты, не отключаются и сохраняется максимальная величина корректирующего импульса. Из возможных вариантов выбора пар ДО для разгрузки СГ включают ту из них, момент которой имеет наибольшую проекцию на противоположное направление вектора

т. е.In the general case, it is not always possible to generate a given value of the kinetic moment by selecting and disabling the i-th engine. This is due to the fact that a situation may arise in which each of the vectors

will be perpendicular to the vector

or when the disturbing moment from each of the engines working to correct the orbit is directed along the three axes of the spacecraft, and the vector

accumulates in such a way that condition (12) cannot be fulfilled. In such cases, it is proposed to use a pair of DOs not participating in orbit correction. At the same time, the engines performing orbit correction are not turned off and the maximum value of the correction pulse is preserved. Of the possible options for choosing DO pairs for unloading the SG include that one whose moment has the greatest projection in the opposite direction of the vector

i.e.

где р= 1, 2, . . . - номера пар ДО, удовлетворяющих условию формирования кинетического момента:

здесь

вектора моментов от р-ых пар ДО;

вектор разгрузочного момента, максимально удовлетворяющий условию формирования ДО.

where p = 1, 2,. . . - numbers of DO pairs satisfying the kinetic moment formation condition:

here

vector of moments from r-th pairs of DO;

vector of the unloading moment, which maximally satisfies the condition for the formation of BS.

In the case under consideration, DO pairs up to D ₁₅ and D ₆ , D ₁₆ and D ₅ or D ₁₃ and D ₈ , D ₇ and D ₁₄ can be used as DO pairs participating in the formation of the kinetic moment.

в момент завершения коррекции в заданную область кинетического момента КА Область поиска определяется временным интервалом (t, t_ks), где t_ks - момент завершения коррекции с учетом работы всех выбранных для коррекции ДО.Immediately at the beginning of the process of formation of the kinetic moment, the moment of its completion (t _Z ) is determined. This moment of time will be determined from the necessary sufficiency to complete the main process of correction of the orbit of the spacecraft and hit the vector

at the moment of completion of the correction to a given region of the kinetic moment of the spacecraft, the search area is determined by the time interval (t, t _ks ), where t _ks is the moment of completion of the correction, taking into account the work of all selected DOs for correction.

где V_u - заданная величина, определяющая скорость коррекции КА; V_ts - текущее значение кажущейся скорости коррекции на момент времени начала формирования кинетического момента t_p, определяется, например, как интегральная оценка измерений акселерометра с момента времени начала коррекции и до момента времени t_p; α_Σ- расчетное ускорение, получаемое КА от работающих ДО, участвующих в коррекции его орбиты.

where V _u is a given value that determines the speed of correction of the spacecraft; V _ts is the current value of the apparent correction speed at the time of the beginning of the formation of the kinetic moment t _p , is determined, for example, as an integral assessment of the accelerometer measurements from the time of the start of correction to the time t _p ; α _Σ is the calculated acceleration received by the spacecraft from operating DOs involved in the correction of its orbit.

 The search is performed with η steps, each of which differs from the previous one by Δt, i.e.

где Δt- продолжительность минимального импульса разгрузочного момента

или

Определив 1-ый момент времени как t'₁= t_p, осуществим прогноз изменений вектора

на интервале (t'₁, t_k1) по выражению:

где

главный вектор управляющего момента от всех работающих ДО без учета отключений i-ых ДО или пар ДО;

главный вектор управляющего момента от всех работающих ДО с учетом отключений i-го ДО или включения пар ДО;

главный вектор управляющего момента всех внешних возмущающих сил. Значение

известно, например, в результате тестовых включений ДО. Значение t_k1 определяется следующим образом:

где

измеренное значение скорости коррекции КА на момент времени t'₁.

where Δt is the duration of the minimum impulse of the unloading moment

or

Defining the first moment of time as t ' ₁ = t _p , we carry out a forecast of changes in the vector

on the interval (t ' ₁ , t _k1 ) by the expression:

Where

the main vector of the control moment from all operating BSs without taking into account disconnections of the i-th BS or pairs of BSs;

the main vector of the control moment from all operating DOs, taking into account the shutdown of the i-th DO or the inclusion of DO pairs;

the main vector of the control moment of all external disturbing forces. Value

It is known, for example, as a result of test inclusions of DO. The value of t _k1 is determined as follows:

Where

the measured value of the correction speed of the spacecraft at time t ' ₁ .

и далее суммируем спрогнозированное значение с реальными начальными условиями, полученными по выражению

на момент времени t'₁

Далее проверяем выполнение условия

и если оно выполняется, то t_z= t'₁, иначе выбираем t′₂= t_p+Δt и повторяем цикл расчетов на момент времени t'₂ и т. д. до выполнения условия (22).The forecast is carried out with zero initial values for the vector

and then we summarize the predicted value with the real initial conditions obtained by the expression

at time t ' ₁

Next, check the condition

and if it is satisfied, then t _z = t ' ₁ , otherwise we choose t ′ ₂ = t _p + Δt and repeat the calculation cycle at time t ′ ₂ , etc., until condition (22) is satisfied.

The formation of the kinetic moment of the spacecraft stops at the moment of completion of the correction of the orbit or when the correction time reaches the value of t _z .

где

радиус-вектор и вектор скорости КА соответственно.Consider the case of controlling the kinetic moment in the correction of the orbit of a geostationary communications satellite. Suppose that after completion of the orbit correction process, the spacecraft will have the following state vector in the Greenwich coordinate system:

Where

radius vector and spacecraft velocity vector, respectively.

The spacecraft must maintain its orientation in the orbital coordinate system during the correction of the orbit. This orientation is defined by the following Krylov angles: λ _Y = 0; λ _x = 0; λ _z = 270 ⁰ .
Suppose that after correction of the orbit, the spacecraft must maintain the same orientation unchanged for three days. Solving the system of equations (2), we obtain a given region of the kinetic moment, in which the kinetic moment of the spacecraft should appear after the completion of the orbit correction process. In FIG. Figures 3, 4, 5 show the resulting graph of the change in the vector of the kinetic moment of the spacecraft in the coordinate system associated with the spacecraft for three days maintaining the indicated orientation. It can be seen from the graph that the component Gz changes slowly over a given time interval, and therefore, the main constraints must be introduced on the other two components of the kinetic moment of the spacecraft. T. about. if the vector of the kinetic moment at the end of the correction process of the SC orbit will belong to the following region:
-8≤Gx≤5 Nms
-7≤Gy≤8 Nms
-5≤Gz≤5 Nms,
it will be possible to avoid saturation of the SG system during the execution of the further flight program.

 A block diagram of a system that implements the proposed method for controlling the kinetic moment of a spacecraft in the process of correcting its orbit is shown in FIG. 6, where the following designations are introduced: 1 - block correction speed (BSK); 2 - a block of sensors of external information of the reference basis (BDVI); 3 - block determining control forces and moments (BOUSM); 4 - block forming the control forces and moments (BFUSM); 5 - a block of determined values of the duration of the pulses DO (BOZPID); 6 - block DO (BDO); 7 - spacecraft body (KKA); 8 - block of angular velocity sensors (BDUS); 9 - block of linear acceleration sensors (BDLU); 10 - block setting the control parameters of the spacecraft (BZPUKA); 13 - block power gyroscopes (BSG); 14 is a block for determining the values of the total vector of kinetic moment (BOSSVKM); 16 is a block for generating the current SG unloading time (BFTVRSG); 11 is a block for bringing the predicted values of the total vector of the kinetic moment to the initial conditions (BPPZSVKMNU); 12 - block determining the end time of the correction mode (BOVORK). BFUSM 4 contains the following blocks, similar to those described in the prototype, shown in FIG. 7: 21 - block selection DO for orbit correction (BVDKO); 24 - block forming the discharge torque for the SG system (BFRMSG); 25 - block determining the unloading moments for the SG system (BORMSG); 22 - block selection of discharge torque for the SG system (BVRMSG); 23 - block selection pair DO for unloading the SG system (BVPDRSG). On the specified block diagram, in addition to the previously described blocks similar to those described in the prototype, the following are additionally introduced: 18 - block for determining a given region of kinetic moment (BOZOKM); 19 is a block comparing the current time of the correction of the orbit of the spacecraft and the estimated time of the beginning of the formation of the total vector of the kinetic moment of the spacecraft (BSTVKORVNFSVKM); 15 is a block comparing the parameters of a given region of the kinetic moment of the spacecraft and the total vector of the kinetic moment of the spacecraft (BSPIOKMSVKM); 16 - block setting the current time of formation of the kinetic moment of the spacecraft (BZTVFKMKA); 17 is a block of predicted values of the total vector of kinetic moment in the process of formation of the total vector of kinetic moment (BPZSVKMFSVKM), 20 is a block for determining the time of the start of formation of kinetic moment (BOVNFKM).

 At the same time, the outputs of BSK 1 and BDVI 2 are connected respectively to the first and second output of BOSM 3, and the output of BOSM 3 is connected to the first input of BFUSM 4. The first output of BFUSM 4 is connected to the first input of BOZPID 5, and the output of BOZPID 5 to the input of BDO 6. In turn, BDO 6 is connected to the second input of BOZPID 5. On the spacecraft 7 body, BDO 6, BDUS 8 and BDLU 9 are installed. In this case, the output of BDUS 8 is connected to the third input of BOUSM 3, and the output of BDLU 9 is connected to the first input of BSK 1. The first , the second, third and fourth outputs of the BZPUK 10 are connected respectively to the input of the BDVI 2, the second input of the BSK 1, the fourth input house BOUSM 3 and the second entrance of BFUSM 4.

 The input of the BSG 13 is connected to the second input of the BOSM 3, and the output of the BSG 13 is connected to the first input of the BOSVSKM 14. The first output of the BOSSVKM 14 is connected to the first input of the BSPIOKMSVKM 15, and the second output of the same block is connected to the first input of the BSPSVKMU 11 and the third input of the BSPIOSKMK 11 The second input of BOZSVKM 14 is connected to the output of BDUS 8. The output of BZTVKFKMA 16 is connected to the third input of BOZSVKM 14 and the first inputs of BPZSVKMFSVKM 17 and BOVORK 12. The first output of BSPIOKMSVKM 15 is connected to the first input of BZVVKMF 16 and the second input is third and 4. The third output of BSPIOKMSVKM 15 is connected to the fourth input BFUSM 4 and the second input BZTVFKMKA BSPIOKMSVKM 16. The second input 15 is connected to the fifth output BZPUKA 10, and the fourth input of the same block - a first output BPPZSVKMNU 11, the fifth input - yield BOZOKM 18, the sixth input - yield 19 BSTVKORVNFSVKM.

 The output of BPZSVKMFSVKM 17 is connected to the second input of BPPZSVKMNU 11. The second input of BPZSVKMFSVKM 17 is connected to the first output of BOVORK 12, the third and fourth inputs of the same unit are connected respectively to the third output of BFUSM 4 and the sixth output of BZPUK 10.

 The second input of BOVORK 12 is connected to the second output of BSK 1, and the third input is connected to the second output of BFUSM 4.

 The first input of BOZOKM 18 is connected to the seventh output of BZPUK 10, the second to the output of BDUs 8, the third to the output of BDLU 9, the fourth to the output of BSG 13. The first input of BSTVKORVNFSVKM 19 is connected to the eighth output of BZPUK 10, the second input to BSTVKORVNFSVKM 19 is connected BOZOKM 20.

 The first input of BOVNFKM 20 is connected to the ninth output of the BZPUK 10, the second to the output of the BDUS 8, the third to the output of the BDLU 9, the fourth to the output of the BSG 13. the fifth to the second output of the BOVORK 12.

 The first input of BVDKO 21 is connected to BOUSM 3, and the second input to the fourth output of BZPUK 10. The first output of BVDKO 21 is connected to the first input of BFRMSG 24 and the second input of BORMSG 25. The second output of the same unit is combined with the third output of BFUSM 4. First output of BFRMSG 24 is simultaneously the first output of BFUSM 4, and the second output is the second output of BFUSM 4. The second input of BFRMSG 24 is connected to the output of BVRMSG 22. The third input of BFRMSG 24 is connected to the fourth input of BFUSM 4.

 The first input of BORMSG 25 is connected to the third input of BFUSM 4, and the third input of the indicated block is connected to the fourth output of BZPUK 10. The first output of BORMSG 25 is connected to the second input of BVRMSG 22, and the second output of the same block is connected to the first input of BVPRSG 23. The first input of BVRMSG 22 is connected to the third input of BFUSM 4.

 The second input of the BVDPRSG 23 is connected to the output of the BZPUK 10, and the third input to the fourth input of the BFUSM 4. The first output of the BVDPRSG 23 is connected to the first output of the BFUSM 4, and the second output of the same block to the third output of the BFUSM 4.

 Consider examples of the implementation of these blocks.

 BSG 13 can be performed on the basis of two-stage SG. The scheme of this block and its description are presented in [3], p. 12-14.

 Blocks 11, 12, 14-25 and their functional connections can be implemented using microprocessor technology, for example, on the basis of one of the three-channel processors of the “Electronics” digital computer MS 1201.02-02 (see [4]) with additional input-output controllers .

 In FIG. 8 shows an example of such an implementation. The following notation was introduced: 27 - clock generator (TG), 26 - processor (P), 28 - address decoder (DSA), 29 - input-output device (UVV), SD - three-way bi-directional 16-bit data bus, ША - three-stable unidirectional 16-bit address bus, ШУ - control bus (10 lines of control signals).

 Blocks 11, 12, 14-25 are typed from standard read-only memory devices (ROM) with a capacity of 2 kB. The number of standard ROMs is determined by the volume of the algorithm to be solved in the block of the problem (see [5], pp. 115 - 117). Functional communications between blocks 11, 12, 14 - 25 are realized due to three-channel execution - via data, address and control buses.

 In addition to the connections indicated in the prototype, through the air-blast 29 a functional multi-channel communication 4, 5, 6, 7, 8, 9 of the outputs of the BZPUK 10 with the corresponding inputs of BFUSM 4, BSPIOKMSVKM 15, BPSZVKMPKO 17, BSTVKORVNFSVKM 19, BOZOKM 18, BOVNFKM 20.

 The implementation of this functional communication can be carried out using the equipment of the managing information-computer complex (UIVK) "Stack - 30".

 Through the UVV 29, the second output of BSK 1 is also connected with the second input of BOVORK 12, the output of BSG 13 with the first input of BOZSVKM 14, the first output of BOUSM 3 with the first input of BVDKO 21, the output of BDUS 8 with the second input of BOZSVKM 14, the connection of BOZOKM 18 with the fifth input BSPIOKMSVKM 15, communication of BSTVKORVNFSVKM 19 with the sixth input of BSPIOKMSVKM 15, communication of BOVNFKM 20 with the second input of BSTVKORVNFSVKM 19. In addition, from the second outputs of BFRMSG 24 and BVDPDRSG 23, information 5 is transmitted through its first input to VHF. The interface of these relationships is described in sufficient detail in [4], p. 33-35.

 The control system operates as follows. Before beginning the correction of the orbit of the BZPUK 10 from its seventh output to the first input, the BOZOKM 18 provides a control signal to determine the specified region of the kinetic moment of the spacecraft. Having received this signal, BOZOKM 18 receives at its second, third and fourth inputs from the output of the BDUS 8, the output of the BDLU 9 and the output of the BSG 13, respectively, the specified information for predicting the accumulation of the kinetic moment of the spacecraft (the vector of the apparent acceleration of the spacecraft, the angular velocity vector, the kinetic moment vector of the SG ) From the seventh output of BZPUK 10, information is also received that is necessary for calculating the disturbing moments and the region of permissible values of the kinetic moment of the SG (the position of the solar cells, the current orientation of the spacecraft, the state vector, the number of working SGs, the necessary information about those involved in the correction of the orbit of the DO, etc.). ) BOZOKM 18 predicts the process of correcting the orbit and determining a given region of the kinetic moment of the spacecraft by solving the system of equations (1) and transfers upon completion of this solution from its first output to the fifth input of the BSPIOKMSVKM 15 the parameters of the specified region of the kinetic moment of the spacecraft.

 Using BZPUK 10, the desired orientation mode is selected by turning on the required sensor, which is part of the BDWI 2 using a signal transmitted from the first output of the BZPUK 10 to the input of the BDWI 2 and confirming the choice of the orientation mode in BOUSM 3 from the fourth output of the BZPUK 10 to the first input of BOPSM 3 At the same time, in BSK 1 from the third output of BZPUK 10 to the first input of BSK 1, the spacecraft orbit correction parameters are set, including the magnitude and direction of the correction velocity vector. The task of constructing and maintaining orientation is solved by BOISM 3, which contains the kinematic contour of the motion control system (a more detailed description of the kinematic contour is presented, for example, when describing the application [3] for an invention). To do this, use the information received from the BDVI 2 and BDUS 8 at the third and fourth inputs of the BOUSM 3, respectively.

In addition, at the beginning of the execution of the orbit correction mode from the second output of the BOSM 3 to the second input of the BFUSM 4 (the BFUSM scheme is shown in Fig. 9, a detailed description of the operation is given in the prototype), requirements for the formation of control forces F _i are issued. BFUSM 4 in a certain way described in the prototype method, generates in its second output (in the case of the formation of a kinetic moment in the process of orbit correction, either from the second output of the BVPRSG 23 or from the first output of the BFRMSG 24 depending on how the formation of the required kinetic moment) to the first input of BOZPID 5 signals that are amplified there, stored and transmitted with a specified duration to the start-up valves BEFORE the output of BOZPID 5. In turn, from each DO goes to BOZPID 5 (on his second entry) receipt of the start of engine operation. As soon as the duration of the engine reaches a stored value, it is turned off (the supply of a control signal to the start valves is stopped).

При этом непосредственно значение

определяется в БОЗСВКМ 14 по указанным измеренным значениям, поступающим на его второй вход с выхода БСГ 13. На первый вход БОЗСВКМ 14 с БДУС 8 поступает информация о значениях вектора абсолютной угловой скорости КА

В самом БОЗСВКМ 14 имеются заранее заданные значения компонент тензора инерции КА

В результате в нем определяется значение вектора

в соответствии с последним выражением из (1).Simultaneously with the issuance of a requirement for the formation of control forces F _i from the first output of the BOSM 3 to the first input of the BSG 13, requirements for the formation of control moments are issued. BSG 13 forms the law of control of the SG precession axes (see [3] p. 12-13) based on the required control moments. Moreover, due to measurements carried out in the block of sine-cosine converters of the rotation angles of the precession axes, which is part of the BSG 13, as well as the known values of the kinetic moments of the rotors of the gyromotors, the value of the kinetic moment vector of the SG system is determined

In this case, the value

it is determined in BOZSVKM 14 according to the specified measured values received at its second input from the output of the BSG 13. At the first input of the BOZSVKM 14 with BDUS 8 receives information about the values of the vector of the absolute angular velocity of the spacecraft

In BOZSVKM 14 itself there are predefined values of the components of the inertia tensor of the spacecraft

As a result, it determines the value of the vector

in accordance with the last expression from (1).

 At the time of the start of the correction of the spacecraft’s orbit from the ninth exit of the BZPUK 10 to the first input of the BOVNFKM 20, the requirement is set for determining the time for the formation of the required value of the kinetic moment. Having received this signal, BOVNFKM 20 receives at its second, third and fourth inputs from the output of the BDUS 8, the output of the BDLU 9 and the output of the BSG 13, respectively, the initial information for predicting the accumulation of the kinetic moment of the spacecraft (the vector of the apparent acceleration of the spacecraft, the angular velocity vector, the kinetic moment vector of the SG ) From the seventh output of BZPUK 10, information is also received that is necessary for calculating the disturbing moments and the region of permissible values of the kinetic moment of the SG (the position of the solar cells, the current orientation of the spacecraft, the state vector, the number of working SGs, the necessary information about those involved in the correction of the orbit of the DO, etc.). )

согласно (4). Со второго выхода БСК 1 через второй вход БОВОРК 12 считывается значение

входящее в выражение (5). Одновременно из второго выхода БФУСМ 4 на третий вход БОВОРК 12 приходит информация о номерах ДО, участвующих в коррекции орбиты. По указанным номерам в блоке определяется a_Σ. Для этого используются известные силы тяги ДО и масса КА и производится расчет t_kη по выражению (5). Далее информация о значении t_kη передается со второго выхода БОВОРК 12 на пятый вход БОВНФКМ 20.BOVNFKM 20 predicts the process of correction of the orbit and accumulation of the kinetic moment of the spacecraft during this process over the time interval from the beginning of the correction of the orbit to the estimated time of its completion t _{kη The} forecast is carried out with zero initial values along the vector

according to (4). The value is read from the second output of BSK 1 through the second input of BOVORK 12

included in expression (5). At the same time, from the second output of BFUSM 4, the third input of BOVORK 12 receives information about the numbers of the subsidiaries involved in the orbit correction. Using the indicated numbers in the block, a _{Σ is} determined. For this, the known traction forces of DO and the mass of the spacecraft are used, and t _{kη is calculated} by expression (5). Further, information on the value of t _kη is transmitted from the second output of BOVORK 12 to the fifth input of BOVNFKM 20.

суммируем с реальными начальными условиями по кинетическому моменту

полученными на момент времени начала коррекции орбиты по (6). Если в результате моделирования в БОВНФКМ 20 выяснилось, что коррекция пройдет без насыщения системы СГ и при этом вектор кинетического момента КА в момент времени t_kη не будет принадлежать заданной области, то в БОВНФКМ 20 выполняется поиск времени t_p начала формирования заданного вектора кинетического момента КА.Predicted value

summarize with the real initial conditions for the kinetic moment

obtained at the time of the beginning of the correction of the orbit according to (6). If as a result of modeling in BOVNFKM 20 it turned out that the correction will take place without saturation of the SG system and the vector of the kinetic moment of the spacecraft at time t _kη will not belong to a given region, then in BOVNFKM 20 a search of time t _{p of the} beginning of the formation of the specified vector of the kinetic moment of the spacecraft .

при условии включения пары двигателей ориентации или момента

при условии отключения i-го двигателя ориентации, участвующего в коррекции орбиты). С текущего момента времени коррекции (t_η), на которое имеются измерения вектора кинетического момента КА, рассчитанные с использованием полученных от БСГ 13 с его выхода на четвертый вход БОВНФКМ 20 значений вектора кинетического момента силовых гироскопов и имеющемуся там значению тензора инерции КА по последнему выражению (1), и до момента времени t_kη прогнозируется накопление кинетического момента с учетом отключения в момент времени

i-го ДО или включения пары ДО для формирования требуемого кинетического момента КА по (3).Search t _p perform steps in Δt (duration of the minimum momentum

subject to the inclusion of a pair of orientation engines or torque

provided that the i-th orientation engine involved in the orbit correction is turned off). From the current moment of correction time (t _η ), for which there are measurements of the SC kinetic moment vector, calculated using 20 values of the kinetic moment vector of power gyroscopes and the SC inertia tensor value available there from the BSG 13 from its output to the fourth input of the BOVNFKM according to the last expression (1), and until the time t _{kη, the} accumulation of the kinetic moment is predicted taking into account the shutdown at the time

of the ith DO or inclusion of a pair of DOs to form the required kinetic moment of the spacecraft according to (3).

согласно (4). Значение t_kη определяется по (16). Спрогнозированное значение

суммируем с реальными начальными условиями по кинетическому моменту

полученными на момент времени t_η начала прогноза по (6).The prediction of the accumulation of kinetic moment by vector

according to (4). The value of t _kη is determined by (16). Predicted value

summarize with the real initial conditions for the kinetic moment

obtained at time t _{η the} beginning of the forecast according to (6).

и полученное значение t_p по выходу БОВНФКМ 20 передается в БСТВКОРВНФСВКМ 19 (на его второй вход). В противном случае повторяем цикл расчетов на момент времени t₂ и т. д. до выполнения условия (7).Next, the fulfillment of condition (7) is checked, and if it is satisfied, then

and the obtained value of t _p at the output of BOVNFKM 20 is transferred to BSTVKORVNFSVKM 19 (to its second input). Otherwise, we repeat the calculation cycle at time t ₂ , etc., until condition (7) is satisfied.

At the time of the start of correction, BZPUK 10 transmits from its eighth output to the first input of BSTVKORVNFSKKM 19 the value of the time of the beginning of correction, which is monitored during the correction process and after t _p arrives at BSTVKORVNFSKKM 19. At the point in time when the current orbit correction time becomes equal to the time t _{p the} start of the formation of the set value of the kinetic moment of the spacecraft BSTVKORVNFSVKM 19 from its output to the sixth input of the BSPIOKMSVKM 15 gives a command to start the formation of the required kinetic moment.

В самом БОЗСВКМ 14 имеются заранее заданные значения компонент тензора инерции КА

В результате в нем определяется значение вектора

в соответствии с последним выражением из (1). С первого выхода БОЗСВКМ 14 значения

поступают на первый вход БСПИОКМСВКМ 15. В БСПИОКМСВКМ 15 производится проверка выполнения условия (19). При этом параметры области допустимых значений кинетического момента СГ могут изменяться в зависимости от числа работающих в системе силовых гироскопов. Указанные изменения в БСПИОКМСВКМ 15 производятся при помощи БЗПУКА 10, пятый выход которого соединен со вторым входом БСПИОКМСВКМ 15.At the first input BOZSVKM 14 with BDUS 8 receives information about the values of the vector of the absolute angular velocity of the spacecraft

In BOZSVKM 14 itself there are predefined values of the components of the inertia tensor of the spacecraft

As a result, it determines the value of the vector

in accordance with the last expression from (1). From the first output of BOZSVKM 14 values

arrive at the first input of the BSPIOKMSVKM 15. In the BSPIOKMSVKM 15, the fulfillment of condition (19) is checked. In this case, the parameters of the region of permissible values of the kinetic moment of the SG can vary depending on the number of power gyros operating in the system. These changes in BSPIOKMSVKM 15 are made using BZPUK 10, the fifth output of which is connected to the second input of the BSPIOKMSVKM 15.

Как только БФТВРПОВФКМ 16 получает значение момента времени tp, он тут же формирует на своем выходе значение t'₁ в соответствии с выражением (17). В свою очередь t'₁ передается на третий вход БОЗСВКМ 14, на первый вход БОВОРК 12 и на первый вход БПЗСВКМПКО 17. По приходу в БОЗСВКМ 14 значения t'₁ в нем происходит присвоение

Значение вектора

поступает со второго выхода БОЗСВКМ 14 на первый вход БППЗСВКМНУ 11 и на третий вход БСПИОКМСВКМ 15. В БППЗСВКМНУ 11 оно служит в качестве начального условия в выражении (21), а в БСПИОКМСВКМ 15 оно необходимо для проверки условия принадлежности кинетического момента КА допустимой области, при нахождении его в которой обеспечивается управляемость КА. Значение t'₁, переданное в БПЗСВКМПКО 17, устанавливает нижнюю границу определенного интеграла, входящего в выражение (18). Оно же задает в БОВОРК 12 момент времени t_k1 начала определения второй верхней границы указанного интеграла. Так, по приходу этого времени в БОВОРК 12 из второго выхода БСК 1 через второй вход БОВОРК 12 считывается значение

входящее в выражение (19), одновременно информация о номерах ДО, участвующих в коррекции орбиты. По указанным номерам в блоке определяется a_Σ. Для этого используются известные силы тяги ДО и масса КА и производится расчет t_k1 по выражению (19). Далее информация о значении t_k1 передается по первому выходу БОВОРК 12 на второй вход БПЗСВКМПКО 17.In case of failure of condition (7) from the first output of the specified block to the first input BFTVRPOVFKM 16 the time value t _p is issued. At the same time, from the second output of the same block to the third input of BFUSM 4, the vector value

As soon as BFTVRPOVFKM 16 receives the value of the time instant tp, it immediately generates the value t ' ₁ at its output in accordance with expression (17). In turn, t ' ₁ is transmitted to the third input of BOZSVKM 14, to the first input of BOVORK 12 and to the first input of BPZSVKMPKO 17. Upon arrival at BOZSVKM 14, the value t' ₁ in it is assigned

Vector value

arrives from the second output of BOZSVKM 14 to the first input of BSPZSVKMNU 11 and to the third input of BSPIOKMSVKM 15. In BPPSVKMNU 11 it serves as an initial condition in expression (21), and in BSPIOKVSMKM 15 it is necessary to check the conditions of belonging of the kinetic moment of the spacecraft to the admissible region, when finding it in which the spacecraft is controlled. The value of t ' ₁ , transferred to BPZSVKMPKO 17, sets the lower boundary of a certain integral in expression (18). It also sets in BOVORK 12 the time instant t _{k1 of the} beginning of the determination of the second upper boundary of the indicated integral. So, upon the arrival of this time in BOVORK 12, the value is read from the second output of BSK 1 through the second input of BOVORK 12

entering into expression (19), at the same time information about the DO numbers involved in the orbit correction. Using the indicated numbers in the block, a _{Σ is} determined. For this, the known traction forces of DO and the mass of the spacecraft are used, and t _{k1 is} calculated by expression (19). Further, information on the value of t _k1 is transmitted along the first output of BOVORK 12 to the second input BPZSVKMPKO 17.

который он создает. Причем значение указанных управляющих моментов задается в БПЗСВКМПКО 17 через четвертый его вход с шестого выхода БЗПУКА 10. Для ДО, не участвующих в управлении,

По указанному каналу в БПЗСВКМПКО 17 выдаются также значения, определяющие значения возмущающих моментов

(углы ориентации КА, положение солнечных батарей и т. д. ). С третьего выхода БФУСМ 4 на третий вход БПЗСВКМПКО 17 передаются также номера ДО для вектора

По значениям

и

в БПЗСВКМПКО 17 определяется

Далее

суммируется с полученным там же

. Затем производится интегрирование на интервале установленных границ в соответствии с выражением (18). Полученные интегральные значения передаются с первого выхода БПЗСВКМПКО 17 на второй вход БППЗСВКМНУ 11, где производится их суммирование с начальными значениями

по (21). Полученная сумма с выхода БППЗСВКМНУ 11 передается на четвертый вход БСПИОКМСВКМ 15, где производится проверка выполнения условия (22). Если в БСПИОКМСВКМ 15 при проверке не выполняется условие (22) и выполняется условие принадлежности кинетического момента КА допустимой области, то значение t_p подтверждается в БФТВРПОВФКМ 16 и повторяется рассмотренный выше цикл разгрузки СГ на момент времени t'₂ и т. д.Information about the DO numbers, initially participating in the orbit correction, is also issued from the third exit of the BFUSM 4 to the third entrance of the BPZSVKMPKO 17 (either from the second exit of the BVDOK 21 or from the second exit of the BVDPRSG 23, depending on which DO participate in the correction of the orbit and the formation kinetic moment of the spacecraft). In this case, each engine is assigned a control moment

which he creates. Moreover, the value of these control moments is set in BPZSVKMPKO 17 through its fourth input from the sixth output of BZPUK 10. For subsidiaries not involved in the management,

On the specified channel in BPZSVKMPKO 17 are also given values that determine the values of disturbing moments

(spacecraft orientation angles, position of solar panels, etc.). From the third output of the BFUSM 4 to the third input of the BPZSVKMPKO 17, the DO numbers for the vector are also transmitted

By values

and

in BPZSVKMPKO 17 is determined

Further

sums up with received there

. Then, integration is performed over the interval of established boundaries in accordance with expression (18). The obtained integral values are transferred from the first output of the BPZSVKMPKO 17 to the second input of the BPPZSVKMNU 11, where they are summed with the initial values

by (21). The amount received from the output of the BPPZSVKMNU 11 is transferred to the fourth input of the BSPIOKMSVKM 15, where the condition (22) is checked. If during the verification, condition (22) is not fulfilled in the BSPIOKMSVKM 15 and the condition that the kinetic moment of the spacecraft belongs to an admissible region is satisfied, then the value of t _{p is} confirmed in BFTVRPOVFKM 16 and the above-mentioned SG unloading cycle is repeated at time t ' ₂ , etc.

If, in the case of BSPIOKMSVKM 15, condition (22) is fulfilled or the condition that the kinetic moment of the spacecraft belongs to an acceptable region is not satisfied, then from the third output of the BSPIOKMSVKM 15 to the fourth input of the BFUSM 4, a command is issued to remove the mode of formation of the kinetic moment of the spacecraft, to the second input of the BFTVRPOVFKM 16 - to terminate the search mode of time t _z .

 The proposed method and control system make it possible to control the kinetic moment of the spacecraft during the correction of the spacecraft’s orbit so that with a further flight program, the number of saturations of the SG system will be minimal, and therefore the number of power-ups for unloading the SG will also be minimal, and this in turn will lead to a decrease in the flow rate of the working fluid and the effect on the orbit of the spacecraft, caused by the operation of DO

Sources of information
1. Branets V. N. et al. "A control method for a spacecraft equipped with jet engines with rod action lines directed at an angle to the axes of the connected base and displaced relative to the center of mass of the apparatus, a system for implementing the method, a block of jet engines of the system." RU patent 2124461 C1.

 2. "A method of controlling a spacecraft using reactive executive bodies and a system for its implementation." Patent RU 2112716 C1.

 3. Kovtun B. C. et al. "A control system for the orientation of a spacecraft with power gyroscopes." Application 5032611/22 (012690), patent RU 2006430 C1.

 4. Technical description. ЩИ3.059.064-02 microcomputer "Electronics", 1990

 5. Petrosyan OA and others. Circuitry BIS ROM. - M.: Radio and communications. 1987 year

Claims (2)

1. A method of controlling the kinetic moment of a spacecraft using reactive actuators, including determining the required value of the orbit correction speed of the spacecraft, maintaining a given orientation of the spacecraft using power gyroscopes in the process of correcting the orbit by orientation engines, measuring the kinetic moment vector in the power system gyroscopes, the determination of the known values of the moments of inertia of the spacecraft, as well as the measured values in Ktorov spacecraft angular velocity and the angular momentum gyroscopes in the power system values total angular momentum vector of the spacecraft

, at current points in time (t), verification of the fulfillment of the condition for the values of the total vector of the kinetic moment of the spacecraft to belong

the range of available values of the kinetic momentum vector in the system of power gyroscopes, characterized in that they predict a given region (S _k ) in which the total vector of the kinetic moment of the spacecraft should be

at the moment of completion of the orbit correction process and provided that during the correction of the orbit of the saturation system of power gyroscopes does not occur, from the current moment of correction time (t _η ), for which there are measurements of the vector of the kinetic moment of the spacecraft, including the beginning of the correction, in η steps according to Δt until the time t _{kη, the} accumulation of the kinetic moment of the spacecraft is predicted

taking into account the shutdown of the i-th orientation engine or the inclusion of a pair of orientation engines to form the required kinetic moment of the spacecraft at a time

η = 1,2,3, ..., where Δt is the duration of the minimum pulse of the total control moment of orientation engines

subject to the inclusion of the specified pair of orientation engines or the total control moment of orientation engines

provided that the i-th orientation engine participating in the orbit correction is turned off,

V _U - a value that determines a given speed correction of the orbit,

the current value of the orbit correction speed at time t _η , and _Σ is the calculated value of the acceleration of the spacecraft from the orientation engines involved in the correction of the orbit, then the predicted values of the vector are summed

with the current value of the total vector

defined at the start of the forecast t _η , verify that the conditions for the obtained values of the total vector of the kinetic moment of the spacecraft of the given region S _k belong to and if it is satisfied, then go to the next ηth step, and if it is, then when the correction time reaches the value

determine the total value of the vectors of the specified control moments

from orientation engines provided that the i-th engine is disconnected, where i = 1, 2,. . . , n are the numbers of the orientation engines participating in the orbit correction, create by these orientation engines for the system of power gyroscopes the total moment that has the largest projection in the direction opposite to the vector

with the corresponding orientation motor switched off, and in the case when the specified control moments

are not unloading, to form the required kinetic moment, they include that pair of orientation engines that are not involved in orbit correction, the moment of which has the largest projection in the direction opposite to the vector

however, none of the indicated i-th engines are turned off.  2. A control system for the kinetic moment of the spacecraft using reactive executive bodies, comprising a block for correcting the speed of the orbit of the spacecraft, a block of sensors for external information of the reference basis, a block for determining control forces and moments, a block of orientation engines, a block of sensors for the angular velocity of the spacecraft, and a block of sensors for it linear accelerations, a block for setting the parameters for controlling the spacecraft, a block for determining the time for the end of the correction mode, a block for bringing predicted values values of the total kinetic moment vector to the initial conditions, the power gyroscope unit, the unit for determining the values of the total kinetic moment vector, the control force and moment formation unit, consisting of the orientation engine selection block for orbit correction, the unload moment formation unit for the power gyroscope system, and the unloading determination unit moments for a system of power gyroscopes, a unit for selecting an unloading moment for a system of power gyroscopes, a unit for selecting a pair of orientation engines for p the load of the system of power gyroscopes, the block of determined values of the duration of the pulses of the thrust of the orientation engines, the outputs of the indicated blocks of the correction speed and the sensors of the external information of the reference basis are connected respectively to the first and second inputs of the block determining the control forces and moments, and the first output of the last specified block is connected to the first input unit for forming control forces and moments, the first output of the unit for forming control forces and moments is connected to the first input of the block of defined values the duration of the thrust impulses of orientation engines, and the output of the last indicated unit is connected to the input of the orientation engine block, the output of the orientation engine block is connected to the second input of the block of determined values of the thrust impulse duration of orientation engines, the output of the linear acceleration sensor block is connected to the first input of the space orbit correction unit apparatus, and the output of the block of angular velocity sensors of the spacecraft is connected to the third input of the block determining the control forces and cops and the second input of the unit for determining the values of the total vector of the kinetic moment, and the second output of the unit for determining the control forces and moments is connected to the input of the unit of power gyroscopes, and the output of the latter is with the first input of the unit for determining the values of the total vector of kinetic moment, the second output of the unit for determining the values of the total vector kinetic moment is connected to the first input of the block bringing the predicted values of the total vector of kinetic moment to the initial conditions, the second output of the block control forces and moments are connected to the third input of the correction mode end time determination unit, and the second correction mode end time determination unit input is connected to the second output of the spacecraft’s orbit correction speed unit, the first output of the spacecraft control parameter setting unit is connected to the input of the external sensor unit information of the reference basis, the second output is with the second input of the block of the orbit correction speed of the spacecraft, the third output is with the fourth input of the block control forces and moments, the fourth output is with the second input of the control forces and moments forming unit, the first input of the orientation engine selection block for orbit correction being connected to the control forces and moments determining unit, and the second input with the fourth output of the space control parameter setting block apparatus, the first output of the block selection of orientation engines for correcting the orbit is connected to the first input of the block forming the unloading moment for the system of power gyroscopes and the second input of the block is determined the unloading moment for the system of power gyroscopes, the second output of the forming unit of the unloading moment for the system of power gyroscopes is simultaneously the second output of the forming unit of control forces and moments, and the first output is the first output of the forming unit of control forces and moments, the second input of the unit of forming of unloading moment for the system power gyroscopes connected to the output of the unit for selecting the discharge torque for the system of power gyroscopes, the second input of the block for selecting the discharge torque for the system with silt gyroscopes connected to the first output of the unloading moment determination unit for the power gyro system, the third input of the unloading moment determination unit for the power gyroscope system is connected to the fourth output of the spacecraft control parameter setting unit, the second output of the unloading moment determination unit for the power gyroscope system is connected to the first input block selection pair of orientation engines for unloading the system of power gyroscopes, and the second input of the latter with the fourth output of the rear unit ia control parameters of the spacecraft, the first output of the unit for selecting a pair of orientation engines for unloading the system of power gyroscopes is connected to the first output of the unit for generating control forces and moments, characterized in that it further includes a unit for determining a given region of kinetic momentum, a unit for comparing the current correction time the orbit of the spacecraft and the estimated time of the beginning of the formation of the total vector of the kinetic moment of the spacecraft, the unit for comparing the parameters given areas of the kinetic moment of the spacecraft and the total vector of the kinetic moment of the spacecraft, a unit for setting the current time of formation of the kinetic moment of the spacecraft, a block of predicted values of the total vector of kinetic moment in the process of forming the total vector of kinetic moment, a unit for determining the time for the start of formation of the kinetic moment, and the first input of the block comparing the parameters of a given region of the kinetic moment of the spacecraft and the total the vector of the kinetic moment of the spacecraft is connected to the first output of the unit for determining the values of the total vector of the kinetic moment, the second input is with the fifth output of the unit for setting the control parameters of the spacecraft, the third input is with the second output of the unit for determining the values of the total vector of kinetic moment, the fourth input is with the first output unit for bringing the predicted values of the total vector of kinetic momentum to the initial conditions, the fifth input - with the output of the unit for determining a given kin field moment, the sixth input - with the output of the unit for comparing the current time of the correction of the orbit of the spacecraft and the estimated time of the beginning of the formation of the total vector of the kinetic moment of the spacecraft, the first output of the same block is connected to the first input of the unit for setting the current time of the formation of the kinetic moment of the spacecraft, the second output of this the block is connected to the third input of the block forming the control forces and moments, the third output is to the fourth input of the block forming the control forces and moment s and the second input of the unit for setting the current time of formation of the kinetic moment of the spacecraft, the output of the unit for setting the current time of formation of the kinetic moment of the spacecraft is connected to the first input of the unit for determining the end time of the correction mode, the third input of the unit for determining the values of the total vector of the kinetic moment and the first input of the block of predicted values the total vector of the kinetic moment in the process of forming the total vector of the kinetic moment, the second input the predicted values of the total kinetic moment vector during the formation of the total kinetic moment vector is connected to the first output of the correction mode end time determination unit, the third input to the third output of the control force and moment formation unit, and the fourth input to the sixth output of the spacecraft control parameter setting unit , the first output of the same block is connected to the second input of the block bringing the predicted values of the total vector of kinetic moment to initial conditions ovium, the first input of the unit for determining the specified region of the kinetic moment of the spacecraft is connected to the seventh output of the unit for setting the control parameters of the spacecraft, the second input is with the output of the block of angular velocity sensors, the third input is with the output of the block of linear acceleration sensors, the fourth input is with the output of the power unit gyroscopes, the first input of the unit for determining the start time of the formation of the kinetic moment is connected to the ninth output of the unit for setting the control parameters of the spacecraft, the second input to the output ohm of the angular velocity sensor block, the third input - with the output of the linear acceleration sensor block, the fourth input - with the output of the power gyroscope block, the fifth input - with the second output of the correction mode end time determination unit, the first input of the comparison unit for the current time of the spacecraft’s orbit correction and calculated the time of the beginning of the formation of the total vector of the kinetic moment of the spacecraft is connected to the eighth output of the control unit settings of the spacecraft, the second input - with the output of the unit the determination of a given region of kinetic moment, the second output of the orientation engine selection block is connected to the third input of the predicted values of the total kinetic moment vector in the process of forming the total kinetic moment vector, the first input of the unloading moment determination unit for the power gyroscope system and the first input of the unloading moment selection unit for the system power gyroscopes are connected to the second output of the unit for comparing the parameters of a given region of the kinetic moment of space ap of the paratrope and the total vector of the kinetic moment of the spacecraft, the third input of the unit for selecting a pair of orientation engines for unloading the system of power gyroscopes and the third input of the unit for generating the unloading moment for the system of power gyroscopes are connected to the third output of the unit for comparing the parameters of the given region of the kinetic moment of the spacecraft and the total vector of kinetic moment spacecraft, the second output of the unit for selecting a pair of orientation engines for unloading the system of power gyroscopes inen with the third input of the block of predicted values of the total vector of kinetic moment in the process of forming the total vector of kinetic moment.

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