Method of control over program turn of accelerating unit
FIELD: aircraft engineering.
SUBSTANCE: invention can be used for control over program turn of accelerating unit with the help of fixed constant-thrust engines of orientation. Angular velocity is increased at acceleration and inertial flight and decreased to zero at deceleration and pulsed initiation of orientation engines. Level of fuel component in the tank that brings about the most tangible effect on turn dynamics is measured Angle mismatch and acceleration unit angular velocities are intermittently measured at turn as well as deflection of said fuel level from acceleration unit lengthwise axis Orientation engine are shut down at the ends of acceleration path and switched on at deceleration start path.
EFFECT: acceleration unit turn with damping of fuel components oscillations at ramming engine operation path section.
The invention relates to rocket and space technology, and in particular to methods of traffic control booster blocks (RB), providing dovodjenje SPACECRAFT) from the reference orbit to the target (usually geostationary) orbit, the implementation of the interorbital transitions and other operations with the SPACECRAFT.
In space technology known selected as a prototype method for managing software turn upper stage using a stationary engines orientation constant thrust, which consists in performing a set of angular velocity - acceleration, coasting, reducing the angular velocity to zero - braking and pulse enable engines orientation with decreasing angular velocity below a specified level (see ).
There is a method of managing software turn allows flat software reversal of the Republic of Belarus at a given angle at a given time with the lowest possible fuel consumption, engine orientation. Thus, the program spread, as a rule, is performed on the passive trajectory of the launch, when the components of the fuel in the tanks of the Republic of Belarus are in a state of weightlessness and not have a significant impact on the dynamics of the spread. However, the known method of control does not provide the desired quality control on the active sites when the engine is preload, creates thrust in the direction of the longitudinal axis of the Republic of Belarus. Sequence diagram of the operation of RB results in frequent inclusion of engines preload for the deposition of components of the fuel to the bottoms of the tanks for draining tanks and run engines. Often there is a necessity of implementation of a reversal of the Republic of Belarus on the flight with the engine running preload. In this case, the large mass of the propellant components are tucked to the lower ends of the fuel tanks and run-time software reversal perform transverse oscillations, providing on side walls of the tanks considerable force. Usually tanks are located in the rear part of the Republic of Belarus, the center of mass of RB are shifted to the bow, which is displayed on the orbit of the massive SPACECRAFT. Therefore, transverse vibrations of the fuel components create significant disturbing moments, the magnitude of which is comparable with the magnitude of the steering torque generated by the engine orientation. As a result, the process of turning when using the known method of control is accompanied by a significant "throw" on the corner (overshoot) and increased oscillatory, resulting in increased fuel consumption, engine orientation.
The task of the invention is to develop methods for the and management of the programme turn the upper stage, providing at the site of the work of engines preload execution of the reversal of the Republic of Belarus at a given angle without overshoot with simultaneous damping component fuels that have the greatest impact on the dynamics of software spread. For example, for the present and developed cryogenic upper stages such component of the fuel is an oxidizer (liquid oxygen), the mass of which is several times greater than the mass of fuel, in this case, because of the specifics of the design location of the tank oxidizer creates a large shoulder hydrodynamic forces about the center of mass of the Republic of Belarus.
The technical result of the invention is to optimize the timeline functioning of the Republic of Belarus due to the expansion capabilities of the control system in terms of combining various flight operations.
This technical result is achieved in that in the method of control software by reversal the upper stage using a stationary engines orientation constant thrust, which consists in performing a set of angular velocity - acceleration, coasting, reducing the angular velocity to zero - braking and pulse enable engines orientation with decreasing angular velocity below a specified level, in accordance with the invention, in the case of programmes is on spread,
combined with the engine preload, before reversing measure the level h has the most impact on the dynamics of the spread component of fuel in the tank, and in the process of turning periodically measure the misalignment angle Δϑ and angular velocity
The essence of the invention is illustrated in Fig.1-4.
Fig.1 - scheme of the upper stage during the implementation of the programme of turn.
Fig.2 - Program management reversal in function of dimensionless time τ.
Fig.3 - Line switching control in the phase plane of the dimensionless variable x and its derivative x' on the dimensionless time τ.
Fig.4 - Typical transient processes when implementing software reversal in accordance with the proposed control method.
Management software spread, combined with the engine preload, in a mathematical model of the control object, it is necessary to take into account fluctuations of the liquid fuel components. As an example, consider the mathematical model of the plane of the reversal of the Republic of Belarus on the pitch with the use of the pendulum is th model, describing the fluctuations of the oxidizer in the tank. The spatial equation of motion of RB as a rigid body with n mathematical pendulums are derived in . In this case, a flat turn with one pendulum these equations have the form (under the assumption of smallness of the angle of deviation of the pendulum from its equilibrium position)
ϑ is the angle of pitch of the Republic of Belarus;
sϑ- the angle of deviation of the pendulum from the longitudinal axis of the Republic of Belarus, equal to the deviation angle between the normal to the surface of the oxidizer from the longitudinal axis of the tank;
u - switch on the control motors of the Republic of Belarus, taking values of 1, 0, 1;
Here (see Fig.1)
m0- the mass of RB without taking into account the mass of the oscillating oxidant (weight solids);
IZ- moment of inertia of a rigid body;
m is the mass of the mother is through the point of pendulum equal to the mass of the oscillating oxidant;
P0- thrust preload directed along the longitudinal axis of the Republic of Belarus;
P - thrust per engine orientation, perpendicular to the longitudinal axis of the Republic of Belarus;
d - the distance from the suspension point of the pendulum to the center of mass of a rigid body;
l is the length of the thread of the pendulum;
xT- longitudinal coordinate of the center of mass of a rigid body in the base coordinate system;
xD- longitudinal coordinate of the points of application of the tractive force control motors in the base coordinate system.
The system of equations (1)-(2) has the following initial and final conditions
where ϑ0, ϑKrespectively defined initial and final values of the pitch angle;
tKthe time of completion of rotation.
Will pass to dimensionless variables
The system of equations (1)-(2) and boundary conditions (5)-(6) can be represented in the form
where the characters ' and “ mean respectively the first and second derivatives of the dimensionless time τ, and 2T is the dimensionless time reversal.
Keeping known in the prototype method, the sequence of operations during execution of the program spread (acceleration, coasting, braking), will choose a temporary control program as shown in Fig.2. The system (14) for the dimensionless time 2T will move from the initial conditions (11) in the final terms, determined by the formula
where σ=signx0. To ensure that the final conditions (16), it is enough to choose as the dimensionless time T one of the values
where i=1, 2, 3,..., while
The minimum dimensionless time reversal is provided by
The obtained values of the parameters T and ξ enable management software spread, shown is a of Fig.2,
as a function of time. However, for the technical implementation is desirable to provide the same control as a function of the state variables x and x', i.e. to carry out the synthesis of feedback control. The presence of feedback, as it is known, allows to compensate the influence of the disturbances associated with the influence of external factors, inaccurate knowledge of the parameters of the control object and the controller, the account in the mathematical model of secondary factors, etc. In the proposed method of control software by reversal the upper stage, just used the principle of feedback. For this purpose we have derived equations of lines switching in the phase plane (x, x'). When the engine is off orientation at the end of the segment acceleration and turning the opposite engine orientation at the beginning of the plot braking is carried out when the parameter x in the phase plane, respectively, first and second line switching (see Fig.3). The first equation of the line switch has the form (typical for this RB values of the dimensionless initial conditions
and the equation of the second line switch
To implement this control in advance before the flight RB calculate dependency ratios (13) linear functions (8) and (9) from the level h of the oxidizer in the tank. For this, we first calculated or experimentally determined dependence on h dimensionless parameters pendulum model the behavior of the liquid: the square of the dimensionless frequency of oscillation of the pendulum
During execution of the software spread periodically measure the angle ϑ and the angular velocity of
The results of mathematical modeling of process management software spread using the proposed method, shown in Fig.4 (a, b, C) show a good quality of management as at nominal values of the characteristics of the upper stage (a), and 5% of the energy dispersion in the values of the moments of inertia of the Republic of Belarus and t is GI engines preload (b, b).
Thus, thanks to the implementation of the proposed invention, the technical solutions, the task management software spread booster unit, providing at the site of the work of engines tucked up execution of the reversal of the Republic of Belarus at a given angle without overshoot with simultaneous oscillation components of the fuel in the tanks.
Sources of information
1. B. C. Rauschenbach, E. N. Turner. Attitude control of a spacecraft. M., "Nauka", 1974, pages 191-194.
2. A. W. Altshuler, B. A. Lobanov. The mathematical model of the spatial fluctuations of the liquid components of the fuel in the tanks of a space rocket on the active phases of flight. Aerospace equipment and technology. 2010, No. 2, pages 39-46.
3. K. S. Kolesnikov. The dynamics of the missile. M. "engineering", 2003
4. The radio wave transmitter "Microradar-21611" TU BY 190460725.003-2009. Manual REN.000-06. http://www.microradartest.com, , firstname.lastname@example.org.
The way to control software by reversal the upper stage using a stationary engines orientation constant thrust, which consists in performing a set of angular velocity - acceleration, coasting, reducing the angular velocity to zero - braking and pulse enable engines orientation with decreasing angular velocity is below the predetermined level, characterized in that in the case of Khujand is the software implementation of the spread, combined with the engine preload, before reversing measure the level h has the most impact on the dynamics of the spread component of fuel in the tank, and in the process of turning periodically measure the misalignment angle Δϑ and angular velocityturn the upper stage, the angle sϑand angular velocitythe deviation of a surface of the specified component in the fuel tank from the longitudinal axis of the upper stage, while the off orientation engines in the late part of the crackdown carried out when the x parameter values ƒ1(x'), starting the engines orientation at the beginning of the plot braking carried out when the x parameter values ƒ2(x'), where;switching functions, and the parameters x and x' is set as a linear function of the measured angles and angular velocities x=k(k1Δϑ+k2sϑ),with coefficients k, k1, k2determined according to a pre-calculated dependency of the level h of the component of fuel in the tank.
SUBSTANCE: invention relates to control of orientation of a space, in particular, a transport cargo ship (TCS) with stationary solar battery panels (SB). The method includes the rotation of TCS around a normal to the working surface of SB facing towards the Sun with an angular speed of at least 1.5 deg/sec. Meanwhile within the time interval of at least one round the components of the angular speed of TCS in a structural coordinate system are measured. Using the measured values, directions of the main central axes of inertia of TCS are determined. Among these axes an axis other than the axis of the minimum moment of inertia and making the minimum angle with the normal to the working surface of SB is found. The angle between the direction towards the Sun and the plane of the TCS orbit is determined. If this angle exceeds a certain value depending on the specified minimum angle and also - on the minimum and maximum SB currents, TCS is turned. Meanwhile the named found inertia axis is combined with the direction, perpendicular to the orbit plane and making a sharp angle with the direction towards the Sun. TCS is rotated around this axis towards the direction opposite to orbital rotation. During the rotation the current from SB is measured. At the achievement by the current of the minimum value TCS is again turned until the alignment of the named found axis of inertia of TCS with the named perpendicular direction and again the named rotation of TCS is performed.
EFFECT: ensuring of the necessary power return of SB in the mode of TCS rotation around one of its actual main central axes of inertia at maintaining of the axis of the minimum moment of inertia in the orbit plane.
FIELD: aircraft engineering.
SUBSTANCE: invention relates to aerospace engineering. Proposed method comprises supply spaceship spinning around perpendicular to solar battery working surface directed to the Sun at angular velocity of at least 1.5 degree/s. During said spinning at time interval of duration making at least one circuit supply angular velocity components are measured in structural system of coordinates. Measured magnitudes are used to define the directions of the main central axis of inertia of supply spaceship. Angle between direction to the Sun and spaceship orbit plane is defined. Spaceship orbit height is defined to determine the half-angle of the Earth disc visible from spaceship. In case said angle exceeds said half-angle the spaceship gravity orientation is constructed at aligning the axis of its minimum moment of inertia that makes the minimum angle with perpendicular to solar battery working surface, with direction to the Earth centre. Spaceship gravity orientation is maintained by spinning it around the axis of minimum moment of inertia at angular velocity defined from the condition of stability of the given gravity orientation of spaceship.
EFFECT: higher power yield of solar batteries owing to radiation reflected from the Earth at spaceship gravity orientation with spinning with allowance for actual main central axes of inertia.
FIELD: aircraft engineering.
SUBSTANCE: invention relates to aerospace engineering. Proposed method comprises supply spaceship spinning around perpendicular to solar battery working surface directed to the Sun at angular velocity of at least 1.5 degree/s. During said spinning at time interval of duration making at least one circuit supply angular velocity components are measured in structural system of coordinates. Measured magnitudes are used to define the directions of the main central axis of inertia of supply spaceship. Spaceship is turned to alignment of the central axis of inertia making the minimum angle with perpendicular to solar battery working surface with direction to the Sun. Spaceship is spun around spaceship around said axis to measure solar battery current. After current reaches minimum permissible magnitudes spaceship is, again, turned to align said axis of inertia with direction to the Sun. Again, spaceship is spun around said axis.
EFFECT: higher power yield of solar batteries owing to radiation reflected from the Earth at spaceship gravity orientation with spinning with allowance for actual main central axes of inertia.
SUBSTANCE: invention relates to aerospace engineering and can be used in spacecraft engines. Power plant comprises cryogenic tank with shield-vacuum heat insulation and channel with heat exchanger, flow control valve, booster pump, intake with capillary accumulator with heat exchanger and throttle and hydropneumatic system with pipeline. Channel cross-section sizes comply with maximum outer sizes of heat exchanger cross-section.
EFFECT: cooling of cryogenic component in capillary accumulator.
FIELD: aircraft engineering.
SUBSTANCE: invention relates to spacecraft control in atmosphere of planet by adjusting its aerodynamics. Spacecraft velocity in atmosphere at initial flight part increases (spacecraft flies toward conditional orbit pericentre). Atmosphere density is low yet to cause notable spacecraft deceleration. As spacecraft reaches atmosphere dense layers its velocity decreases to reach atmosphere enter velocity for angle of roll (γ) γ=π to be changed to γ=0. This manoeuvre allows changing the spacecraft to flight part with maximum aerodynamic performances. In flight with γ=0 continuous skip path is maintained whereat spacecraft velocity decreases monotonously. Maximum skip height reached, angle of attack o spacecraft increases, hence, spacecraft intensive deceleration occurs.
EFFECT: decreased final velocity at soft landing system operation, fuel savings.
FIELD: aircraft engineering.
SUBSTANCE: invention relates to spacecraft control in atmosphere of planet by adjusting its aerodynamics. Proposed method consists in selection of conditions for changing the angle of roll to zero at changing the spacecraft from isothermal descent section (IDS) to skip path. With spacecraft in IDS, angle of roll (γ) is, first, increased to decrease aerodynamic performances and to maintain constant temperature at critical area of spacecraft surface. As flight velocity decreases angle (γ) is decreased from its maximum. In IDS, increase in aerodynamics does not cause further temperature increase over its first peak. Therefore selection of the moment of changing to γ=0 allows efficient deceleration of spacecraft at the next step of flight. The best option is the descent of spacecraft of IDS when γ reaches its maximum. Here, angle of attack is set to correspond to maximum aerodynamic performances. This increases the duration of final flight stage and deceleration efficiency. Increase in angle of attach after descent from IDS and completion of climb results in increased in drag, hence, decrease in velocity at initiation of soft landing system.
EFFECT: minimised final velocity and maximum temperature at surface critical area, lower power consumption.
FIELD: physics; control.
SUBSTANCE: invention relates to controlling movement of a spacecraft fitted with a heat radiator and a solar panel. The method includes flying the spacecraft on an orbit around a planet and turning the solar panel in a position corresponding to the alignment of the normal to the working surface of the solar panel with the direction towards the Sun; performing orbital orientation of the spacecraft, where the plane of rotation of the solar panel is parallel to the plane of the orbit of the spacecraft and the solar panel is located relative to the plane of the orbit on the side of the Sun; determining the maximum value of the angle between the velocity vector of the spacecraft and the perpendicular to the transverse axis of rotation of the solar panel, passing through the surface of the radiator; determining the orbital altitude of the spacecraft and the angle between the direction towards the Sun and the plane of the orbit of the spacecraft; based on the orbital altitude and the angle, determining the orbit passes where the duration of the illuminated part of the pass exceeds the difference between the orbiting period of the spacecraft and the required duration of the heat release by the radiator on the pass; on the said orbit passes, when the spacecraft passes through the illuminated part of the pass, the solar panel is turned around the transverse axis of rotation until the intersection of the line passing through the region of the surface of the radiator facing the Sun and directed towards the Sun with the solar panel; turning the solar panel around the longitudinal axis of rotation until the angle between the normal to the working surface of the solar panel and the direction towards the Sun assumes a minimum value. The said solar panel rotations are performed within a calculated time interval.
EFFECT: high efficiency of the radiator by creating conditions for natural cooling thereof during eclipse of the solar panel for any altitude of an almost circular orbit of the spacecraft.
SUBSTANCE: invention relates to space cable systems (SCS) and can be used for the transfer of SCS to a spinning mode in the orbit plane without the application of jet engines. SCS development is executed from its initial compact state in the circular orbit by the repulsion of objects at a low relative speed. SCS end weights are connected by a cable, its length being varied by a cable feed-haul-in device arranged on one of the end objects. The objects are separated by a vector of local peripheral speed, for example, by a pusher. The objects are driven by a start pulse to separate the objects in practically free paths at the free feed of the cable. The cable development is terminated by the SCS transfer to a stable mode of associated pendulum motion at the stretched preset-length cable. At a definite range of angular phases of this mode the SCS objects are stretched by hauling in the cable at a definite constant speed. This results in changing the SCS into the spinning mode at a preset power integral and fixed final end of the cable.
EFFECT: relaxed weight-size constrictions of SCS, enhanced performances.
FIELD: physics, navigation.
SUBSTANCE: group of the inventions relates to control of angular motion of space vehicle (SV). The method includes additional generation of signals for assessment of orientation angle and angular velocity of rotation of space vehicle. Also the reference signals of the orientation angle, angular velocity and control assessment signal are generated. For the named orientation angle and angular velocity their differences with their assessed signals, and also the difference with their reference values are determined. The difference of control signal and its assessed value and, at last, the signal of correction of the signal of assignment of mathematical model and the signal of assessment of external noises using the respective formulas are determined. On this base the signals of assessment of orientation angle and angular velocity of space vehicle are determined, which are used for space vehicle control. The device in addition contains the reference model of the basic circuit of orientation of space vehicle and other necessary devices and connections.
EFFECT: improvement of orientation accuracy and operational reliability in case of failures of orientation angle sensor and sensor of angular velocity of space vehicle rotation.
2 cl, 2 dwg
SUBSTANCE: invention relates to controlling orientation of an artificial earth satellite with solar panels. The disclosed method includes performing necessary turning of the artificial earth satellite along with solar panels and, separately, the solar panels about a first and a second axis. The antenna of the artificial earth satellite is directed towards the earth and the normal to the solar panels is directed towards the sun. Independent programmed turns about the first and second axes of the artificial earth satellite are performed in intervals of uncertainty of orientation of the artificial earth satellite on shadow orbits. In different versions of said turns, after the first turn, the artificial earth satellite is held in an intermediate position and normal orientation of the artificial earth satellite is then restored. This improves the accuracy of predicting movement of the artificial earth satellite on shadow orbits and accuracy of measuring the range to the artificial earth satellite.
EFFECT: high accuracy of determining navigation-time data on navigation artificial earth satellites by consumers.
4 cl, 12 dwg
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.
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: 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 ΔФp=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.
FIELD: terminal control of motion trajectory of cryogenic stages injecting spacecraft into preset orbits by means of cruise engines.
SUBSTANCE: swivel combustion chamber of cruise engine is used for angular orientation and stabilization of cryogenic stage of spacecraft. Proposed method includes predicting parameters of motion of cryogenic stage at moment of cut-off of cruise engine; deviation of radius and radial velocity from preset magnitudes are determined; angle of pitch and rate of pitch are corrected and program of orientation of thrust vector for subsequent interval of terminal control is determined. By projections of measured phantom accelerations, angle of actual orientation of cruise engine thrust vector and misalignment between actual and programmed thrust orientation angles are determined. This misalignment is subjected to non-linear filtration, non-linear conversion and integration. Program of orientation of cryogenic stage is determined as difference between programmed thrust orientation angle and signal received after integration. Proposed method provides for compensation for action of deviation of cruise engine thrust vector relative to longitudinal axis of cryogenic stage on motion trajectory.
EFFECT: enhanced accuracy of forming preset orbit.
5 dwg, 1 tbl
FIELD: control of group of satellites in one and the same orbit or in crossing longitude and latitude ranges of geostationary orbit.
SUBSTANCE: proposed method consists in measurement of parameters of satellite orbits, determination of orbital elements, comparison of them with required ones and performing of correcting maneuvers with the aid of thrusters. Satellite inclination vectors are brought to circular areas of their permissible change which are spaced apart so that angle between line connecting the end of vector with center of its circular area and direction to Sun should exceed right ascension of Sun by 180°. According to first version, vectors of satellite eccentricity are shifted to similar circular areas so that similar line lags behind direction to the Sun by half angular displacement of vector over circumference of its natural drift within circular area. Then, distances between satellites are changed within required limits compensating for quasi-secular increment of inclination vector and correcting eccentricity vector so that at passing the center of interval between point of circumference entry of its natural drift to its circular area and point of exit from this area, line connecting the center of this circumference and center of circular area coincide with direction to the Sun. In case circular area of permissible change of each eccentricity vector is close to circumference of its natural drift (second version), said line for this area is matched with direction to the Sun and no correction is made in this case.
EFFECT: saving of propellant for correction; protracted flight of satellites at safe distance.
3 cl, 13 dwg
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.
FIELD: astro-navigation, control of attitude and orbital position of spacecraft.
SUBSTANCE: proposed system includes control computer, star sensor, Earth sensor, storage and timing device, processors for control of attitude, processing angular and orbital data, inertial flywheels and spacecraft orbit correction engine plant. Used as astro-orienters are reference and navigational stars from celestial pole zone. Direction of spacecraft to reference star and direction of central axis of Earth sensor to Earth center are matched with plane formed by central axes of sensors with the aid of onboard units. Shift of direction to reference star relative to central axis of Earth sensor is considered to be latitude change in orbital position of spacecraft. Turn of navigational star around reference star read off sensor base is considered to be inertial longitude change. Point of reading of longitude is point of spring equinox point whose hour angle is synchronized with the board time. This time is zeroed upon completion of Earth revolution. Stochastic measurements by means of static processing are smoothed-out and are converted into geographic latitude and longitude parameters. Smoothed inertial parameters are compared with parameters of preset turn of spacecraft orbit found in storage. Revealed deviations of orbit are eliminated by means of correction engine plant.
EFFECT: enhanced accuracy of determination of spacecraft attitude and orbital position; automatic elimination of deviation from orbit.
FIELD: spacecraft systems for supply of power with the aid of solar batteries.
SUBSTANCE: proposed method includes turning the solar battery panels to working position corresponding to matching of normal to illuminated surface of solar batteries with plane formed by axis of rotation of solar battery panels and direction to the Sun. Proposed method includes also determination of moments of the beginning of solar activity and arrival of high-energy particles onto the spacecraft surface. Then, density of fluxes of said particles is measured and the results are compared with threshold magnitudes. When threshold magnitudes are exceeded, solar battery panels are turned through angle between the said normal and direction to the Sun which corresponds to minimum area of action of particle fluxes on solar battery surfaces at simultaneous supply of spacecraft with electric power. When action of particles is discontinued, solar battery panels are returned to working position. Angle between direction to the Sun and axis of rotation of solar battery panels is measured additionally. In case threshold magnitudes are exceeded, solar battery panels are turned to magnitude of angle between normal to their illuminated surface and direction to the Sun which corresponds to minimum area of action of said particle fluxes on spacecraft surfaces (provided the spacecraft is supplied with electric power). System proposed for realization of this method includes units and their couplings for performing the above-mentioned operations. System is additionally provided with unit for measurement of angle between direction to the Sun and direction of axis of rotation of solar battery panels, as well as unit for determination of maximum current.
EFFECT: avoidance of lack of electric power on board the spacecraft at performing the "protective" turn from high-energy particle fluxes; possibility of using these measures for arbitrary orientation.
3 cl, 1 dwg
FIELD: spacecraft systems for supply of power with the aid of solar batteries.
SUBSTANCE: proposed method includes turning of solar batteries to the working position corresponding to matching of normal to their illuminated surface with plane formed by axis of rotation of solar battery panels and direction to the Sun. Proposed method includes also measurement of density of fluxes of solar electromagnetic radiation and high-energy particles determining the moments of beginning of solar activity and arrival of said particles to spacecraft surface. Additional measurement includes determination of appearance of signs of negative action of particle flux on spacecraft. During these moments, onboard solar batteries are charged to maximum level. When density of particle flux exceeds threshold magnitude, solar battery panels are turned through angle between said normal and direction to the Sun corresponding to minimum action of particle fluxes on solar battery surfaces. Discharge of storage batteries is hoped to close the energy gap on board the spacecraft. At minimum permissible level of storage battery charge, storage batteries are disconnected from load. When action of particles on spacecraft is discontinued, solar battery panels are returned to working position. System proposed for realization of this method includes units and their couplings for performing the above-mentioned operations. System is provided with unit for determination of current from solar batteries, unit for determination of moments of appearance of signs of negative action of high-energy particles on spacecraft and unit for setting the permissible level of charge of storage batteries.
EFFECT: reduction of negative action of high-energy particle flux on solar battery working surface due to maximum increase of angle of "protective" turn of solar batteries from direction of these fluxes to the Sun.
3 cl, 1 dwg
FIELD: electric power supply for spacecraft with the aid of solar batteries.
SUBSTANCE: proposed method includes turning the solar battery panels to working position corresponding to matching of normal to their illuminated surface formed by axis of rotation of solar battery panels and direction to the Sun. Proposed method includes also measurement of density of fluxes of solar electromagnetic radiation and high-energy particles followed by determination of moments of beginning of solar activity and arrival of high-energy particles to spacecraft surface. Method includes additionally measurement of spacecraft orbit altitude and angle between direction to the Sun and plane of spacecraft orbit. In case density of particle flux exceeds threshold magnitudes, solar battery panels are turned on illuminated surface of spacecraft orbit through angle (αs min) between said normal and direction to the Sun corresponding to minimum area of action of particle fluxes on spacecraft surfaces at supply of spacecraft with required amount of electric power. On shaded side of orbit, solar batteries are turned from direction of particle flux through maximum angle. When spacecraft escapes from shadow, reverse turn of solar battery panels is completed through said angle αs min. Upon completion of action of particle flux on spacecraft, solar battery panels are returned to working position. System proposed for realization of this method includes units and their couplings for performing the above-mentioned operations. System includes additionally unit for determination of intensity of spacecraft illumination, unit for measurement of spacecraft orbit altitude, unit for measurement of angle between direction to the Sun and spacecraft orbital plane, unit for control of turn of solar battery to position opposite to direction to the Sun, NO-gate and switch.
EFFECT: reduction of negative action of high-energy particle fluxes on solar battery working surface on shaded surface of orbit.
3 cl, 1 dwg