Method for missile take-off from aircraft for orbit injection of payload

FIELD: rocketry, applicable at an air start, mainly of ballistic missiles with liquid-propellant rocket engines.

SUBSTANCE: the method consists in separation of the missile with a payload from the carrier aeroplane and its transition to the state with initial angular parameters of motion in the vertical plane. After separation the missile is turned with the aid of its cruise engine, preliminarily using the parachute system for missile stabilization. The parachute system makes it possible to reduce the duration of the launching leg and the losses in the motion parameters (and the energy) in this leg. To reduce the missile angular bank declination, the strand of the parachute system fastened in the area of the missile nose cone is rehooked. To reduce the time of missile turning towards the vertical before the launcher, the cruise engine controls are preliminarily deflected to the preset angles and rigidly fixed. By the beginning of missile control in the trajectory of injection this fixation is removed. In the other modification the missile turning is accomplished by an additional jet engine installation. It is started depending on the current angular parameters of missile motion so that by the beginning of controlled motion in the trajectory of injection the missile would have the preset initial angular parameters of motion.

EFFECT: enhanced mass of payload injected to the orbit.

4 cl

 

The invention relates to rocket technology and can be used to launch missiles, mainly ballistic liquid powered by engines made by, with the aim of placing a payload into orbit.

The missile before launch is placed in the cargo compartment of the aircraft in a horizontal position, as a rule, the head part against the direction of its flight. From aircraft missile catapulted using a special energy resources.

Start sector begins with the moment of separation of the rocket with a payload from the aircraft and ends with the moment of withdrawal of missiles in a vertical plane at a specified angular motion parameters. The parameters of the rocket motion, obtained at the end of the start sector are the initial conditions for the implementation of a software rocket motion to output payload into orbit.

The most important requirement for the start sector is ensuring the safety of the aircraft in the process of starting the main engine (MD) missiles. This is achieved by removing the missile from the aircraft at the time of launch of the MD at the desired distance, for example, that can be provided by the magnitude of the time delay command to switch on the engine.

In the process of implementing a spread of missiles at the start sector losses occur by the values of the vertical, horizontal costal is, considering the initial velocity of flight, range and altitude of the center of mass of the rocket (especially vertical components). This increases the cost of fuel when operating the main engine to compensate for these losses, which eventually leads to reduction of weight delivered to orbit payloads. Therefore, the maximum reduction in the duration of the maneuver spread of missiles and, as a consequence, the minimization of these losses are one of the main tasks solved by the motion of a rocket on the site launch.

The number of known ways to launch missiles from aircraft. In the method described in the patent of Russian Federation โ„–2068169 (priority from 24.08.1992 g), is considered the start using a special platform, which laid the rocket. The missile platform is removed from the cargo compartment through the exhaust chute. Then force parachute platform with the rocket deploys a channel pitch to the position at which the air flow begins to run on the rocket on the side opposite the platform, separate the platform from the rocket and lead her through the exhaust chute, launch the rocket engines and using controls stabilize at the specified path.

In the method of removing a payload into space (patent RF โ„– 2159727 priority from 07.12.1999,) basically dealt with the provision of the necessary parameters of movement of the aircraft to the moment cancer of the s from the media.

According to the invention the preparation and takeoff multimode aircraft responsice produce from the airfield, the least remote from the zone of the launching rocket, the safest route of flight. Flying in this area is carried out in the most range. At the approach to the area of the plane rasandik gains altitude and supersonic speed. In a given geographical point the aircraft performs "the hill" and separates the booster upon reaching the desired pitch attitude. Thus preferably provide a zero angle of attack booster at the time of launch. Next plane rasandik transfer mode command-measuring point for support of the payload to its output into orbit. After that, the plane returned to the landing airfield.

Technical essence closest to the proposed invention is a Method of controlling aerospace system for removing payload" (patent RF โ„–2160214 priority from 29.07.1999 g), which is selected as a prototype. In this way after arrival at maximum cruise mode in the area startup booster carry out dive the aircraft carrier and at the time set them to the maximum flight speed is transferred to pitch up to the maximum allowable angle of attack. Then transferred to the angle of attack is, giving close to zero normal overload. Options cabrerana such that the plane at the moment of separation from him rocket with a payload has the speed, the height and inclination of the flight path, giving maximum payload, and normal overload is close to zero.

The Department reported the rocket speed lag from the aircraft to a safe distance at the time of its engines. A rocket with a payload deploy using engines after them on or before their inclusion with additional jet-set in a position other than vertical angle of 10-30 degrees in the vertical plane in the direction of removal.

Shown in prototype operations on realization of a spread of missiles after its release from the aircraft have considerable disadvantages, which are as follows.

The firing of the main engine (engines) missiles is carried out at a safe distance from the aircraft, the magnitude of which will depend on the time of uncontrolled movement of the missile on the site launch. A safe distance is usually taken equal to 250...300 m, which, for example, for any of the examined below missiles will correspond to the time interval from the release of the missile from the aircraft to enable the main engine (the appearance of thrust) - not less than 6 seconds. With regard to the of osobennosti output of main engine at full thrust, the time of readiness for operation of the steering actuator when the steering machine can provide turn controls the motor with the set (working) speed relaying in accordance with control commands) unmanaged area of movement may increase to ˜8, i.e. the beginning of the spread of missiles at a given angle using only the main engine after turning it on can be carried out not earlier than this time.

Missiles launched from aircraft, as a rule, are characterized by a static instability. After exiting the plane operates disturbing aerodynamic moment, focused on the dive. Rocket by the end of unmanaged land movement deviates in the vertical plane at substantial angles to the side of the dive (more than 90 degrees from the vertical), which results in further protracted process of the spread of missiles at a given angle, to the higher loss values of the motion parameters of the rocket at the end of the start sector and, as a consequence, lower energy opportunities rockets for launching a given weight of the payload.

Proposed in the prototype version of the reversal of the rocket using a jet-set also has its drawback. Spread with it is to turn on the main engine. With edutella, from the end of the additional jet-set, coinciding with the moment of switching on the main engine, and prior to the possible angular spread due to the thrust of the main engine, there is an unmanaged site, the duration of which, as mentioned above, may take up to 2 seconds. During this time, under the action of disturbing the aerodynamic torque, depending on the implemented modes of motion taking into account wind effects may be directed towards the predetermined spread of missiles and against him, the rocket to the launch control using cruise engine will not occupy a given angular position in space.

For withdrawal of missiles in a given situation will require additional time main engine, which will lead to an increase in the duration of the start sector and the deterioration of conditions for output payload into orbit. The duration of the movement of missiles at the start sector will be minimal when the output of a rocket at a desired angular position in space will coincide with the beginning of its control by the main engine.

Improving conditions output payload into orbit by reducing the duration of the start sector and reduce losses on the parameters of motion of a rocket is the main task, p is why the invention.

The problem is solved in that in the known method of launch from plane to output the payload into orbit, including the process of overcoming missile with a payload of the aircraft, the firing of the main engine of the first stage of the rocket to a safe plane distance, the spread of missiles using the main engine after its inclusion in a specified angular position in a vertical plane prior to the implementation of its software movement or a similar reversal of the rocket with an additional jet-set, introduces the following additional operations.

In a variant of the reversal of the rocket using the main engine in the process of entering the missile from the aircraft uses a system of parachutes, stabilize the rocket due to the pull of the parachute system, to the commencement of the controlled motion of the rocket detach the parachute system controls the main engine deploy a missile in a vertical plane to given values of angular motion parameters (e.g., angle, angular velocity in the channel pitch), and then stabilize the missile in relation to the software path.

In a variant of the spread of missiles with additional jet-set control angular motion parameters of the rocket after exiting the aircraft and depending on their led is in command onboard control system include the installation of additional reactive with a time delay relative to the time of release of the missile from the aircraft, which provides the output of a rocket to the specified values of the angular parameters of movement in the vertical plane to the commencement of the controlled motion using the main engine, to the same point in time stop for more jet-set, then use the main engine stabilize the missile in relation to the software path.

To reduce the angular declinations missiles in the channel roll introduces the operation, which consists in the fact that by the time the missile from the aircraft are perceptu strand parachute system, enshrined in the area of the head of the rocket, with the original mounting point below the longitudinal axis of the rocket in its vertical plane of symmetry, the attachment point above the axis in the same plane.

To reduce the time reversal of a rocket in a vertical position before the start of the rocket rejects swing Central nozzle of the propulsion engine of a first stage or control the camera in the schema of the engine with a fixed Central nozzle at a given angle (angles) towards a point in the channel pitch to pitch up rockets, they are fixed rigidly in the deflected position, with the inclusion of cruise engine deploy a missile in the vertical direction, the fixation with these authorities to remove the PTO is NTU early controlled motion of the rocket.

The introduction of the stabilizing parachute system provides the necessary dynamic position of the missile before its reversal using the main engine, which leads to reduction in the duration of the start sector. Strand parachute system is attached to the head of the rocket, and its thrust is directed along the velocity vector flow. On the site of uncontrolled movement of the parachute keeps the rocket from the "stall" in the dive, and when the upper supporting its strands rocket (above the longitudinal axis) also helps to reduce its angle of declination in the channel Bank. This is due to the fact that after separation of the missile from the aircraft are being implemented, as a rule, negative angles of attack (-171...-179 degrees, the longitudinal axis of the missiles aimed at its fore part, and the normal - axis down, the output of the missile from the aircraft - head part against the direction of flight), in which occurs the vertical component of thrust parachute system, pointing up. Such power in the event of a roll angle of the missile creates a moment about its longitudinal axis directed in the direction of decreasing this angle. When the lower fastening strand parachute system to the rocket this effect will be reversed.

Sometimes when the motion of a rocket in flight strand parachute system, in order to ensure the security of the process of landing, yordany be mounted with the bottom point of the rocket. This is because near the ceiling of the rear fuselage of the aircraft, where the rockburst-hazardous area, strand parachute can make circular oscillations with amplitude, which is not precluded by its contact with structural elements of the aircraft, which is unacceptable. The mounting for the lower end of the rocket increases the required gap between the oscillating strand and rockburst-hazardous area design the tail of the plane. The gap increases as the missile from the aircraft. Therefore, to ensure the safety of the process output missiles from aircraft and preservation of the above-mentioned effect of reducing the variance of its channel roll provided by the operation percepti strand parachute system from the "bottom" point of attachment "top".

The angle of the preliminary deflection controls the main engine is determined at the design stage of the rocket, and the operation of their prior rejection and fixing are carried out either at the factory or at the airport before loading missiles on the plane.

In the proposed method, spread with additional jet-set output missiles in the final angular position with the required values of angular motion parameters (end of section start) is to a known point in time, namely the beginning of the missile guidance using archivage engine. Thus, in comparison with the prototype, reduced the duration of the start sector and improve the conditions of the conclusion of the payload into orbit. When you implement the specified operation of the spread of missiles, there is no need to use additional reactive installation of a movable nozzle (nozzles) and the steering actuator, which would complicate the design and increase the weight of the rocket. It is sufficient to use the engine turn with a fixed nozzle, the axis of which is perpendicular to the longitudinal axis of the missile and is located in the vertical plane of its symmetry.

The motor mount pivot rocket is made in such a way as to create the necessary constant torque in the direction of turn of the missile to pitch up. To turn the rocket in the channel pitch was strictly in the vertical plane is used for another block, roll, consisting of two pairs of stationary nozzles on and off which is functional, depending on the parameters of the angular movement of the missile about its longitudinal axis.

The additional reactive system (engine reversal and block Bank) occurs or until complete combustion of fuel, or it is turned off forcibly by command from the control system. Energy characteristics of the jet-set are selected at the design stage R is chum.

Comparison of the known method start with the proposed is considered on the example of calculations of the start of one of the space rockets from an aircraft carrier an-124 aircraft.

Rocket weighing ˜100 tons of liquid fuel payload is located in the cargo compartment of the aircraft in a horizontal position, the head part against the direction of its flight. When you start moving missiles in the cargo compartment of the motion parameters of the aircraft was following values:

- speed of flight - 650 km/h (relative to air);

the flight altitude is 10000 m;

- the angle of trajectory - 24 degrees (from horizon);

- the pitch - 22.7 degrees (from horizon).

The time of movement in the plane ˜2.5 seconds.

Control the missile in the pitch and yaw at the expense of the deviation of the Central nozzle of the upper stage main engine, and in the channel roll - through block roll.

On site start on a rocket can operate wind disturbance.

The firing of the main engine (the beginning of the emergence of thrust) was carried out at a safe distance of 250 m, which corresponded to the time of flight of the missile ˜6 from the moment of exit from the aircraft. The start time for the implementation of controlled motion through the main engine was ˜8 C.

On the site launch, the missile must be transferred from the initial angular position in simple is ansto, obtained at the time of its separation from the aircraft, at the specified destination, necessary for implementation of the program of motion of the missile-to-target output payload into orbit. When you exit the airplane, the rocket takes a position close to the horizontal, and the head part is directed in the opposite direction from the direction of flight of the aircraft carrier. The direction of removing payload coincides with the direction of flight of the aircraft.

In the calculations it was assumed that the angles of deflection of the rocket from the vertical at which the warhead missiles aimed in the opposite direction from the direction of removal, are positive and in the direction of excretion is negative.

Given a finite angular position of the missile in the vertical plane is defined at the design stage and, as a rule, the angle of the rocket from the vertical is in the range 0...-30 deg. If necessary, may be restrictions on the magnitude and the sign of the angular velocity of the rocket. In this example it was assumed that at the end of the site launch, the missile must be launched at angle 0...-10 degrees.

In the proposed method, a spread of missiles using the main engine were used prior to the deviation of the Central nozzle of the engine 4 deg (at an operating angle 7 deg) and a stabilizing parachute system with summary the th area of the domes 28 m 2and resistance factorn=0,9. Below are the comparative evaluation of the main parameters of movement in a vertical plane on the start sector in the process of its output at a given angle with the main engine on offer and the known methods of spread. Values in parentheses correspond to the way of the spread of the prototype. Inclination angles measured from the vertical, and the time from the release of the missile from the aircraft.

The motion parameters of the rocket on the site start:

- deviation of the missile in the channel pitch at the time of exit from the plane, deg109 (110)
- the angle of the rocket in the channel pitch when turning on the main engine (6), deg93 (152)
- maximum angle of deviation of the missile in the channel pitch, deg115 (157)
- the angle of the rocket in the channel pitch to the commencement of the controlled motion using the main engine (8), deg71 (151)
- time out missiles at a given angle (end of section start)10,4 (12,5)
is the mass of fuel consumed the main engines on the start sector, kg1500 (2400)
- the value of the vertical velocity of a rocket compliance the state at the time of exit from the aircraft and at the end of section start m/s51,9 (51,9) - -19,8 (-58,5)
the altitude of the rocket at the time of exit from the aircraft and at the end of the start sector, m10125 (10125) - 10150 (9970)

From the above results it is seen that in the way of the spread of missiles only using the main engine are observed increased loss values of the motion parameters of the rocket on the start sector, which it is estimated will not allow to comply with the requirement to output a given payload into orbit. For example, the output payload on a circular orbit will be less than desired on ˜140 kg. Introduction stabilizing parachute and pre-deflection nozzle of the engine reduces this loss values of the parameters of motion of a rocket to the values that are requirements for the conclusion of a payload into orbit.

Introduction operation percepti strand parachute system from the bottom with respect to the longitudinal axis of the rocket attachment point on the top reduces the magnitude of the angular deviation of the missile in the channel Bank. For example, for the rockets top fastening strand parachute system reduces the maximum deflection angle it in the channel roll when driving in the area starting from 36 to 14 degrees.

This operation is useful for missiles launched from aircraft, which can be increased angular deviation is about to roll up values close to the bounding values from the viewpoint of stabilizing the motion of the rocket. Limiting value is determined by the angles of pumping gyrostabilized platform, used in the control system of the missile.

Preliminary deviation of the governing bodies of the first stage engine (at a given angle 4 deg) allows to reduce the duration of the maneuver spread of missiles at ˜1 C.

As an additional jet-set to turn the missile was accepted installation consisting of a motor reversal and block roll, composed of two pairs of stationary nozzles. Engine turn after turn creates a constant point in the channel pitch equal to 81400 kgf m, and pull each nozzle block roll adopted 100 kgf. The time until complete burnout is ˜5s.

After the release of the missile from the plane of the nozzle block roll on and off functionality, depending on the parameters of the angular movement of the missile about its longitudinal axis.

In the known method the motor reversal is open until the start of the cruise engine. Further spread of missiles, as noted above, can be performed using the main engine is not earlier than 2 from the moment of its activation.

In the proposed method, the motor reversal is activated with a time delay is otnositelno of the release of the missile from the aircraft, which depends on the current angular parameters of movement, the value of the specified target turn angle and time to achieve this angle. The time to reach the specified angle (the duration of the start sector) will be equal to the start time control of the rocket using the main engine. To this point in time complete the work the engine turn and block roll. In this case not specified above unmanaged section of the rocket.

The turn-on delay time motor reversal is calculated according to the following functionality inherent in on-Board control system:

where

M - constant torque developed by the motors turn;

Jz1- moment of inertia of the missile relative to its transverse axis passing through the center of mass of the rocket;

ϑ, ωz1- controlled current values of the angle and angular velocity of the rocket in the channel pitch after exiting the plane (ϑ - measured from vertical);

ϑk- end the set value of the pitch angle of the rocket relative to the vertical at the end of the start sector (in our example ϑk=0...-10 deg);

T - the time of the end section of the start (the beginning of the controlled motion of the rocket with the help of main engine), T = 8;

t - the current time of movement (from the moment of exit from the aircraft).

In EMA t=t Hwhere the condition f(tH)=0 is the time delay for motor reversal.

When the above trigger conditions of the comparative assessment of the implementation of the spread of missiles with additional reactive install (engines turn and block roll)proposed in the known and proposed ways to start.

In the known method the motor reversal is enabled at a given point in time (at a given time delay) relative to the moment of exit from the aircraft, which is determined at the design stage of the complex. Taking into account the characteristics of missiles, conditions start acting on the rocket perturbations and decide on the site of the start of the task this time was ˜2 for all possible modes of motion of the rocket.

In the proposed method, the motor reversal is activated with a time delay, the value of which depends on the current values of the angular parameters of movement, i.e. will vary depending on the specific implementation mode of motion. For this example implementation start time delay, in accordance with the above functionality, turned out to be equal to 3.45 C. In the prototype engine turn terminates at the time of main engine, and in the proposed method, it continues to spread missiles until the achala its controlled movement using the main engine.

To 6 seconds into the flight (after inclusion of main engine) angle and the angular velocity of the rocket in the channel pitch has reached the following values;

for known way ϑ=22,9 hail, ωz1=-42,5 deg/s;

for the proposed method ϑ=77,8 hail, ωz1=-36,0 deg/S.

8 seconds into the flight (beginning-of-control rocket with

main engine) these options amounted to;

for known way ϑ=-9,3 hail, ωz1=15.1 deg/s;

for the proposed method ϑ=-11,0 hail, ωz1=-37,5 deg/S.

As can be seen, in the proposed method, a rocket to the beginning of the control by the main engine reaches the desired value of the angle and angular velocity in the channel pitch, aimed in the direction of payload delivery. The movement of the missile on the start sector is 8 seconds.

In the known method start because of the presence of unmanaged land a rocket to the specified point in time becomes a significant positive angular velocity in the channel pitch against launch a payload into orbit. For compensation before the beginning of realization of the program of movement will require additional time main engine, which will lead to an increase in the duration of the start sector to 10.2 with and, as consequence, to decrease energy opportunities RA is Yety to output payload into orbit. For example, the mass of fuel consumed on the start sector, compared with the proposed method, more on 1240 kg

On the estimates in the proposed method, start, compared with the prototype, the weight of the output payload is increased by ˜13%.

Thus, the proposed method launch missiles from aircraft allows, in comparison with the known, to reduce the loss by the values of the motion parameters of the rocket on the start sector and increase the mass delivered to orbit payloads.

1. The way you launch from an aircraft for launching a payload into orbit, including the output of a rocket with a payload of the aircraft, the firing of the main engine of the first stage at a safe from the aircraft the distance, the spread of missiles using the main engine after its inclusion in a specified angular position in a vertical plane prior to the implementation of its software movement, characterized in that in the process of entering the missile from the aircraft uses a parachute system, stabilize the rocket due to the thrust of this system to the commencement of the controlled motion of the rocket detach the parachute system controls the main engine deploy a missile in a vertical plane to the specified the values of angular motion parameters, and then stabilize the missile relative to programme the th path.

2. The method according to claim 1, characterized in that by the time the missile from the aircraft are perceptu strand parachute system, enshrined in the area of the head of the rocket, with the original mounting point below the longitudinal axis of the rocket in its vertical plane of symmetry, the attachment point above the axis in the same plane of symmetry.

3. The method according to claim 1, characterized in that before the start of the rocket rejects swing Central nozzle of the propulsion engine of a first stage at a given angle or control of the camera, the diagram of the engine with a fixed Central nozzle at specified angles in towards a point in the channel pitch to pitch up rockets, they are fixed rigidly in the deflected position, and since the thrust of the main engine at the turn of the missile in the direction of the vertical release from these controls to the commencement of the controlled motion of the rocket.

4. The way you launch from an aircraft for launching a payload into orbit, including the output of a rocket with a payload of the aircraft, the firing of the main engine of the first stage at a safe from the aircraft the distance, the spread of missiles with additional jet-set at a desired angular position in a vertical plane prior to the implementation of its software movement, trichosis fact, after exiting the aircraft control angular motion parameters missiles and depending on their values on the command Board system management include additional reactive installation with a time delay relative to the time of release of the missile from the aircraft, providing the output of a rocket to the specified values of the angular parameters of movement in the vertical plane to the beginning of its controlled movement using the main engine, stopping in addition, the additional reactive installation, and then use the main engine stabilize the missile in relation to the software path.



 

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5 dwg, 1 tbl

FIELD: cosmonautics, applicable in space activity - space exploration, exploration of the solar system, observation of the Earth from the space, at which it is necessary to determine the space co-ordinates of the space vehicles and the components of their flight velocity vectors.

SUBSTANCE: the method consists in the fact that in the intermediate orbit simultaneously with determination of the co-ordinates of the space vehicle (SV) at initial time moment t0 by signals of the Global Satellite Navigation Systems the determination and detection of radiations at least of three pulsars is carried out, and then in the process of further motion of the space vehicle determination of the increment of full phase Δะคp=Δϕp+2·π·Np of periodic radiation of each pulsar is effected, the measurement of the signal phase of pulsar Δϕp is determined relative to the phase of the high-stability frequency standard of the space vehicle, and the resolution of phase ambiguity Np is effected by count of sudden changes by 2·π of the measured phase during flight of the space vehicle - Δt=t-t0; according to the performed measurements determined are the distances covered by the space vehicle during time Δt in the direction to each pulsar and the position of the space vehicle in the Cartesian coordinate system for the case when the number of pulsars equals three is determined from expression where Dp - the distance that is covered by the space vehicle in the direction to the p-th pulsar; Δt - the value of the difference of the phases between the signal of the p-th pulsar and the frequency standard of the space vehicle, measured at moment Tp - quantity of full periods of variation of the signal phase of the p-th pulsar during time Δϕp; Np - column vector of the position of the space vehicle at moment Δt; - column vector of the space vehicle position at initial moment t0; -column vector of estimates of space vehicle motions in the direction cosines determining the angular position of three pulsars.

EFFECT: provided high-accuracy determination of the space vehicle position practically at any distance from the Earth.

2 dwg

FIELD: space engineering; on-board terminal control facilities of cryogenic stages with non-controllable cruise engines.

SUBSTANCE: parameters of motion of cryogenic stage at moment of cruise engine cutoff are predicted and radius of deviation of radius and radial velocity of cryogenic stage from their preset magnitudes are determined. Signals for correction of pitch angle and rate of pitch are shaped for compensation of said deviation. Pitch angle correction signal is limited at preset level and its excess above this level is determined. When signal is shaped for limitation, addition to correction pitch rate signal is formed. This correction is equal to product of said excess by ratio of functions of sensitivity of radial velocity of cryogenic stage to pitch angle and rate of pitch. Resultant pitch rate correction signal is formed as sum of this signal determined without taking into account pitch angle correction signal limitation and addition. Thus, priority follow-up of velocity error is ensured at limited pitch angle correction.

EFFECT: enhanced accuracy of forming preset orbit due to reduction of disturbance level on angular stabilization loop.

9 dwg, 1 tbl

FIELD: space engineering; designing spacecraft motion control systems.

SUBSTANCE: proposed method is performed by information of orientation unit to Sun by introducing the orbit parameters into on-board computer followed by calculating the Sun position in observation field of orientation unit for each point of orbit for orientation of axes in orbital coordinate system; search angular velocity is set for spacecraft to ensure capture of Sun by observation field of orientation unit, after which angular velocity is decreased to zero ensuring position of Sun in observation field of orientation unit. Then spacecraft is turned in such way that Sun should move to required initial point; turning the spacecraft to preset points is continued for each orbital point.

EFFECT: reduced mass; simplified construction of spacecraft due to reduced number of instruments and units; extended field of application.

3 dwg

The invention relates to rocket and space technology and can be used to create launch vehicles (LV), including conversion, for a spacecraft in low earth orbit

The invention relates to space technology, and more particularly to management of orbital maneuvers boosters with lively marching rocket engines

The invention relates to automatic control systems nonstationary, mainly space objects

The invention relates to automatic control systems nonstationary, mainly cosmic objects

The invention relates to automatic control systems nonstationary, mainly cosmic objects

FIELD: space engineering; designing spacecraft motion control systems.

SUBSTANCE: proposed method is performed by information of orientation unit to Sun by introducing the orbit parameters into on-board computer followed by calculating the Sun position in observation field of orientation unit for each point of orbit for orientation of axes in orbital coordinate system; search angular velocity is set for spacecraft to ensure capture of Sun by observation field of orientation unit, after which angular velocity is decreased to zero ensuring position of Sun in observation field of orientation unit. Then spacecraft is turned in such way that Sun should move to required initial point; turning the spacecraft to preset points is continued for each orbital point.

EFFECT: reduced mass; simplified construction of spacecraft due to reduced number of instruments and units; extended field of application.

3 dwg

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