Spacecraft power plant
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.
The invention relates to rocket and space technology and can be used in cryogenic liquid rocket propulsion installation of the spacecraft.
When storing cryogenic fuels between runs of the engine is warming up design suction device of the tank and adjacent the cryogenic fluid. To avoid overheating of the cryogenic fluid and the formation of the vapor phase due to the inflow of heat leakage from the booster pump and through the insulation of the bottom of the bottom suction device thermostatically due to the evaporative heat exchanger type. As coolant in the heat exchanger is used cryogenic liquid coming from the tank through the metering device.
Typically, the heat exchanger is located on the outer surface of the bottom plate, as its location on the inner surface of the bottom plate worsening conditions of fluid flow when the fence it from the tank, increase hydraulic residues Nesebar fuel. To ensure heat transfer design to the coolant necessary for a good thermal contact channel heat exchanger with intake device. That are typically provided by welding or soldering.
Transfer of heat in design with its temperature control involves feeding cryogenic W is drasti of the cryogenic tank through the metering device in the heat exchanger, located on the outer surface of the suction device (see "Capillary system of sampling fluids from tanks spacecraft". Authors: C. C. Bagrov, A. C. Karpenkov, C. N. Poliaev, A. L. Sintsov, C. F. Shotover. Moscow, ESPC "Energomash", 1977, pp. 99-105) prototype.
On the surface of the drive intended for holding a cryogenic liquid such as liquid oxygen), is a cooling coil (heat exchanger). The coil (heat exchanger) is also on the lower bottom of the cryogenic tank, which prevents the flow of heat to drive the booster turbopump Assembly, as well as the response to an external heat gain to this part of the surface of the cryogenic tank.
Cooler in the heat exchanger is stored in the drive cryogenic component, which during the flight arrives from the drive through the throttle device in the channel of the heat exchanger, where the reduction of the saturation pressure and thus reduce the temperature appears in the temperature difference between the cooler and the design. Cryogenic component in the heat exchanger partially evaporates and is removed in the surrounding space. This device is a temperature of the evaporative type, in which the supply of heat to the heat exchanger located on the outer surface of the cryogenic tank and the cooler it is C is the result of thermal conductivity design suction device.
In the prototype heat exchanger is located on the outer surface of the structure suction device (bottom plate) of the cryogenic tank, without any irregularities. The positioning and fixing of the heat exchanger is made on a smooth surface.
The disadvantage of this solution is the following.
The presence on the outer surface of the bottom plate of any constructive protrusions or contains attachments (for example, elements of pneumatic systems, sensors and others) greatly complicates the provision of the necessary cooling efficiency of the cryogenic component due to the inability of the placement of the heat exchanger required dimensions and, consequently, insufficient thermal contact between the heat exchanger from the bottom bottom.
The objective of the proposed propulsion system of the spacecraft is providing the necessary cooling efficiency of the cryogenic component in the drive capillary type under grid separator when it is impossible to place the heat exchanger mentioned drive the required dimensions on the outer surface of the bottom of the bottom of the cryogenic tank.
The problem is solved due to the fact that in the engine installation of the spacecraft, including cryogenic tank with screen-vacuum thermal insulation, flow control valve,booster pump and an intake device of a cryogenic tank, moreover, the suction device is mounted on the bottom plate and contains the drive capillary-type heat exchanger under grid separator and a throttle device for supplying cryogenic fluid with a specified flow rate from the drive capillary-type heat exchanger, the inner cavity of the cryogenic tank body bottom plate made channel, the cross-sectional sizes which correspond to the maximum outside dimensions of the cross section of the heat exchanger. In the channel flush with the inner surface of the bottom plate is placed a heat exchanger. The output of the heat exchanger hermetically passes through the bottom plate and the outside of the cryogenic tank communicated with the piping pneumatic-hydraulic system propulsion system of the spacecraft.
As an example in Fig.1 presents a General view of the suction device of the cryogenic tank of the propulsion system of the spacecraft of Fig.2 shows the connection of the heat exchanger output with the lower bottom of the cryogenic tank of Fig.3 shows a view of the bottom plate from the internal cavity of the cryogenic tank, where:
1 - cryogenic tank;
2 - screen-vacuum thermal insulation;
3 - flow control valve;
4 - booster pump;
5 is a bottom plate;
6 - drive capillary type;
7 - heat exchanger;
8 - wire RA is the divisor;
9 - the throttle device;
10 - channel;
11 - fasteners;
12 - plate;
13 - wrap;
14 - the inner cavity of the cryogenic tank;
15 - the entrance of the heat exchanger;
16 - cavity drive capillary type;
17 - heat exchanger output;
18 - go nick;
19 - pipeline;
20 - blind threaded holes.
In the propulsion system of the spacecraft, including cryogenic tank 1 with screen-vacuum thermal insulation 2, the flow control valve 3, the booster pump 4 and an intake device of a cryogenic tank 1, and the suction device is mounted on the bottom plate 5 and contains the drive capillary type 6 with heat exchanger 7 under grid separator 8 and the throttle device 9 for supplying cryogenic fluid with a specified flow rate from the drive capillary type 6 into the heat exchanger 7, the inner cavity of the cryogenic tank 14 in the body of the bottom plate 5 is made of the channel 10, the dimensions of the cross section of which corresponds to the maximum outer dimensions of the cross section of the heat exchanger 7. Channel 10 flush with the inner surface of the bottom plate 5 is placed a heat exchanger 7. The output of the heat exchanger 17 is hermetically passes through the bottom plate 5 and the outer side of the cryogenic tank 1 communicates with the pipe 19 pneumatic-hydraulic system propulsion aerospace is th aircraft.
For example, for a cryogenic tank 1 made of aluminum alloy from the condition of structural strength in the presence of cryogenic tank 1 booster pump 4 and the isolating valve 3 of the bottom plate 5 has a profile with increasing thickness in the direction of the flange to cryogenic Baku 1 booster pump 4, the body of which is a spiral channel 10 to heat exchanger 7.
As one option, for example, to keep the heat exchanger 7 on the bottom plate 5 can be performed blind threaded holes 20, which by means of fixing elements 11 (for example, using screws with countersunk head) is fixed to the plate 12. Plate 12 repeats the profile of the bottom plate 5, adjacent to its surface, holding the heat exchanger 7 in the channel 10. In the plate 12 made the cut-outs 13, tells the cavity of the drive capillary type 16 channel 10 by reducing thermal resistance of heat transfer from the cryogenic liquid under grid separator 8 from the booster pump 4 and the external heat gain through screen-vacuum thermal insulation 2 to the cooler in the heat exchanger 7.
Between the heat exchanger 7 and the wall of the channel 10, there are technological gaps that while filling it with gas, have a high thermal resistance, which reduces the intensity of heat transfer designs from the flange booster pump 4 and outside the it heat gain through screen-vacuum thermal insulation 2 cryogenic tank 1 to the cooler in the heat exchanger 7. The presence of local gaps between the bottom plate 5 and the plate 12 also creates a thermal resistance that prevents the cooling liquid under grid separator 8.
These factors in the limited area of the bottom plate 5 under the grid separator 8 does not allow even by increasing the length of the heat exchanger 7 to solve the problem of providing the required temperature cryogenic liquid under grid separator 8 and temperature structure of the lower plate 5.
The presence of grooves 13 in the plate 12 allows you to fill cryogenic fluid gaps between the heat exchanger 7 and channel 10, as well as to fill gaps between the bottom plate 5 and the plate 12, while thermal resistance of the gap is reduced by more than an order of magnitude, and the total resistance to heat transfer to the cooler it is just a minor part. The length of the slots 13 in the plate 12 by ~10% greater than the distance between adjacent sections of the channel 10. The total area of the grooves 13 can be from 10 to 20% of the area of the plate. The grooves 13 should be evenly distributed on the surface of the plate 12 with the direction of their axes of symmetry to the axis of the cryogenic tank 1.
The input heat exchanger 15 passes through one of the cutouts 13 of the plate 12, and the output of the heat exchanger 17 passes through another cutout 13 of the plate 12 and is connected, for example, the adapter 18, which is tightly integrated into the lower is its bottom 5. The outer side of the cryogenic tank 1 to the adapter 18 is joined to the pipe 19 pneumatic-hydraulic system propulsion system of the spacecraft.
In addition, the plate 12 provides for cryogenic liquid when the fence from her cryogenic tank 1 with minimal resistance.
The heat exchanger 7, for example, can also be mounted using clamps.
The propulsion system of the spacecraft, including cryogenic tank 1 with screen-vacuum thermal insulation 2, the flow control valve 3, the booster pump 4 and an intake device of a cryogenic tank 1, and the suction device is mounted on the bottom plate 5 and contains the drive capillary type 6 with heat exchanger 7 under grid separator 8 and the throttle device 9 for supplying cryogenic fluid with a specified flow rate from the drive capillary type 6 into the heat exchanger 7, is as follows.
During filling of cryogenic tank 1, in which due to the boiling of the cryogenic liquid to fill the surfaces of the structural members, such as plate 12, the bottom plate 5 and the heat exchanger 7 is cooled to the saturation temperature of the liquid in the disposable pressure cryogenic tank 1 formed pairs POPs up. As the cooling design of the cryogenic fluid (e.g., through the slots 13 in the plate 12) fills the anal 10 in the bottom plate 5, where is the heat exchanger 7.
The heat exchanger 7 is hydraulically connected with the environment (space) through the pneumatic-hydraulic system propulsion system, and the pressure therein is below the saturation pressure of the cryogenic fluid under grid separator 8. After the throttle device 9 cryogenic liquid partially gasified, its temperature and the temperature of the channel 10 of the heat exchanger 7 is lower than the temperature of the cryogenic fluid under grid separator 8. Due to the difference of temperature is cooling and condensation of the gaseous phase of the cryogenic fluid in the gaps. Due to the fact that the saturation pressure of the cooled cryogenic liquid below the pressure of the cryogenic fluid located above the bottom plate 5, is filling the gaps of the cryogenic fluid. This significantly improves the heat transfer by thermal conductivity to the channel 10 of the heat exchanger 7 and the cooler in it. Therefore, during the stay of the propulsion system of the spacecraft in space conditions the heat from the booster pump 4, the external heat gain through screen-vacuum thermal insulation 2 to the bottom plate 5 and the heat from the cryogenic liquid under grid separator 8 is transferred to the coolant in the heat exchanger 7, vaporizing it. So is maintaining the required temperature range is the azone temperature cryogenic liquid under grid separator 8 in the intervals between the inclusions of the propulsion system.
The proposed propulsion system of the spacecraft provides the necessary cooling efficiency of the cryogenic component in the drive capillary type 6 under grid separator 8 by placing the heat exchanger 7 in the channel 10, is made in the body of the bottom plate 5 of the cryogenic tank 1 when it is impossible to place the heat exchanger 7 the required dimensions on the outer surface of the bottom plate 5 of the cryogenic tank 1.
The propulsion system of the spacecraft, including cryogenic tank with screen-vacuum thermal insulation, flow control valve, booster pump and an intake device of a cryogenic tank, and the suction device is mounted on the bottom plate and contains the drive capillary-type heat exchanger under grid separator and a throttle device for supplying cryogenic fluid with a specified flow rate from the drive capillary-type heat exchanger, characterized in that the inner cavity of the cryogenic tank body bottom plate made channel, the cross-sectional sizes which correspond to the maximum outside dimensions of the cross section of the heat exchanger, in the channel flush with the inner surface of the bottom plate is placed a heat exchanger, the output of the heat exchanger hermetically passes through the bottom plate and the outer side is by cryogenic tank communicated with the piping pneumatic-hydraulic system propulsion system of the spacecraft.
SUBSTANCE: invention relates to propellants for liquid, solid fuel and hybrid rocket engines, containing oxidiser and combustibles. Propellant oxidiser contains boron nitrate. Propellant contains saod oxidiser and combustibles, such as pure or bound boron, for instance boranes (diborane), beryllium borohydrate, boron carbide, borides of metals. In addition to large release of hydrogen born reacts with released nitrogen with formation of boron nitride and heat release 252.6 kJ/mol.
EFFECT: obtaining propellant oxidiser, which contains boron nitrate.
FIELD: machine building.
SUBSTANCE: hollow one-piece blank with bottom is pressed from a bar section, local bulges are provided on the bottom on the outer and/or inner side. The following operations are performed: machining of the blank internal and external surfaces, rotary drawing with thinning with permissible deformation degree, thermal treatment and finishing machining. The local bulges shall exceed the minimal thickness of the shell by at least 10 times.
EFFECT: improved strength characteristics.
FIELD: engines and pumps.
SUBSTANCE: proposed method comprises control over jet engine thrust vector, fuel flow running out from combustion chamber along Laval nozzle, lengthwise duplex control electromagnets mounted at nozzle outer expanded section, current MHD-generator installed at the nozzle throat, current stabiliser and rectifier and aircraft control system that controls said electromagnets. Control over jet engine thrust vector is ensured by deflection of plasma flow running out from combustion chambers with respect to nozzle mirror axis by electromagnetic field induced by control electromagnets. Plasma flow is known to consist of positively charge ions and electrons. Note here that bulk of ions is several orders of magnitudes higher than that of electrons. This defines the thrust vector by the direction of positive ion flow.
EFFECT: simplified design of aircraft, decreased weight, higher reliability of engine and aircraft.
4 cl, 2 dwg
FIELD: power engineering.
SUBSTANCE: propellant comprises fuel and oxidant. Versions of the fuel contain fuel and oxidant at the following ratios of components: beryllium boron hydride - 35.26%+-10%, ammonium dinitramide - 56.52%+-10%, beryllium - 8.22%+-5% or lithium boron hydride - 36.45%+-10%, ammonium dinitramide - 51.93%+-10%, lithium - 11.62%+-5%, or aluminium boron hydride - 24.1%+-10%, ammonium dinitramide - 58.84%+-10%, aluminium - 17.06%+-5%.
EFFECT: jet engine with such fuel from gases releases only pure hydrogen.
FIELD: engines and pumps.
SUBSTANCE: rocket propellant comprises fuel and oxidiser. Rocket propellant features the following composition at the following ratio of components in wt %: beryllium borane - 34.63±10, ammonia dinitramide - 55.50±10, beryllium hydride - 9.87±5, or beryllium borane - 23.78±10, ammonia dinitramide - 76.22±10, or lithium borane - 35.85±10, ammonia dinitramide - 51.06±10, lithium hydride - 13.09±5, or aluminium borane - 23.66±10, ammonia dinitramide - 57.76±10, aluminium hydride - 18.58±5, or decaborane - 39.64±10, ammonia dinitramide - 60.36±10. Other versions are produced using the reaction with ammonia (wt %): beryllium borane - 44.61±10, ammonia dinitramide - 35.75±10, ammonia - 19.63±5. All said reactions can be realised with the help of the other oxidiser, that is, nitrogen hexaoxide, N3O6.
EFFECT: emission of pure hydrogen solely.
SUBSTANCE: invention relates to rocket propellants for liquid, solid fuel and hybrid rocket engines. Rocket propellant contains fuel, which represent borazine, and oxidant. In presence of bound nitrogen in oxidant fuel additionally contains boron or its compounds, for instance, diborane, tetraborane, decaborane, beryllium borohydride, metal borides. As one of versions, rocket propellant contains as fuel 62.65±15.0% of borazine and oxidant oxygen 37.35±15.0%. As another version rocket propellant contains 34.56±13.0% of borazine, 51.52±13.0% of oxygen and 13.02±13.0% of boron. In all cases the sum of component ratios must constitute 100%.
EFFECT: invention makes it possible to obtain high energy fuel and reduce confuser of nozzle, that is, to lighten rocket engine.
FIELD: engines and pumps.
SUBSTANCE: proposed method comprises control over jet engine thrust vector, fuel flow running out from combustion chamber along Laval nozzle, lengthwise duplex control electromagnets mounted at nozzle outer expanded section, current MHD-generator installed at the nozzle throat, current stabiliser and rectifier and aircraft control system that controls said electromagnets. Control over jet engine thrust vector is ensured by deflection of plasma flow running out from combustion chambers with respect to nozzle mirror axis by electromagnetic field induced by control electromagnets. Plasma consists of positive-charge ions and negative-charge electrons. Note here that mass of ions is several orders of magnitude higher than that of electrons. This defines the thrust vector by the direction of positive ion flow.
EFFECT: simplified aircraft design, decreased weight and overall dimensions of jet engine.
2 cl, 1 dwg
FIELD: engines and pumps.
SUBSTANCE: liquid-propellant engine includes a combustion chamber, a gas generator and a turbopump unit with a turbine and pumps. The gas generator and the turbopump unit are installed above the combustion chamber in series one above another along its axis. The gas generator and the turbopump unit are made as one whole unit. As per the first version, the liquid-propellant engine includes the following in downward direction: a gas generator, a turbine, an oxidiser pump, a fuel pump, a multiplier and an additional fuel pump, and the turbine outlet is connected to the combustion chamber via a gas line. As per the second version, the liquid-propellant engine includes the following in downward direction: an additional fuel pump, a multiplier, a fuel pump, a gas generator and a turbine, and the turbine outlet is connected immediately to the combustion chamber.
EFFECT: increasing the rocket movement speed, improving its weight characteristics and increasing the flight range.
2 cl, 7 dwg
SUBSTANCE: invention relates to rocketry. Proposed vehicle comprises body, rocket engine with axially symmetric supersonic nozzle and truncated cone-shaped jacket made of soft thin-wall material, closed on body side and exposed on nozzle side to fold along lengthwise axis. Cone smaller exposed base extends beyond nozzle edge. Case is made form heatproof dense fabric. Fabric of smaller exposed base is connected with heatproof rigid ring. Jacket larger closed base has annular tubular channel communicated on one side with flexible tubular channels from heat-resistant dense fabric bound with jacket fabric and located along jacket one generatrix extending from larger base to smaller one and provided with gauged outlets in rigid ring area and, on the other side, with high-temperature gas source. Folded jacket is attracted to body by cords from thermally unstable material jointed with jacket smaller base rigid ring and located opposite flexible tubular channel outlets.
EFFECT: decreased weight and increased thrust.
FIELD: engines and pumps.
SUBSTANCE: proposed engine comprises onboard computer, electric power source, oxidiser turbo pump unit including main turbine, pumps and starting turbine, propellant gas generator, external compressed air bottle connected via high-pressure pipeline and external valve, fast-release connector, check valve and onboard pipeline to starting turbine, and igniters on combustion chamber and gas generator. In differs from known designs in that it incorporates two turbo pump units, that is, oxidiser turbo pump unit and propellant turbo pump unit. Gas generators are integrated with turbo pump units. Oxidiser turbo pump unit comprises oxidiser pump and extra oxidiser pump while propellant turbo pump unit comprises propellant pump. Turbo pump units and combustion chamber may be arranged in one plane in symmetry about combustion chamber lengthwise axis while their lengthwise axes are parallel with the latter. Shaft of said turbo pump units run in opposite directions. Said turbo pump units may feature equal weight. Combustion chamber is provided with bearing ring whereto one or two pairs of thrust vector control wires are connected. Oxidiser gas generator may be arranged between main turbine and oxidiser pump. Propellant gas generator may be arranged above second main turbine. Propellant gas generator sidewall allows regenerative cooling and comprises inner and outer spaced apart shells.
EFFECT: improved specific characteristics, higher reliability, multifold starting.
17 cl, 5 dwg
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: aircraft engineering.
SUBSTANCE: invention relates to in-flight control over spacecraft equipped with heat radiator and solar battery. Proposed process comprises spacecraft flight in orbit around the planet with solar battery turn to position corresponding to normal to solar battery working surface directed to the Sun. Spacecraft orbital orientation is constructed whereat solar battery spinning plate is parallel with spacecraft orbit plane while solar battery is located on the Sun side relative to orbit plane. Spacecraft orbit altitude and angle between direction to the Sun and spacecraft orbit plane are defined. Magnitude of said angle (β*) is defined whereat duration of turn shadow section equals the necessary time of radiator heat release in said turn. Orbit turns are defined wherein current magnitude of said angle is larger than β*. In said turns, solar battery is turned around crosswise and lengthwise rotation axes unless shadowing of solar battery radiator. Note here that minimum departure of orientation of solar battery working surface to the Sun. Spacecraft orbital flight is conducted in near-circle orbit at altitude not exceeding a definite design value.
EFFECT: higher efficiency of radiator with solar battery shadowed at whatever position of spacecraft on orbit turn.
SUBSTANCE: invention relates to control over spacecraft, particularly, to holding of geosynchronous spacecraft in preset are of stay and collocation with the other geostationary spacecraft. Proposed method comprises determination and correction of initial inclinations and longitude of injection orbit ascending node with allowance for epoch of spacecraft placing in orbit and term of its active existence. Note here that the time of beginning of operation in geostationary orbit when spacecraft orbit inclination reaches maximum permissible value iper. area. The latter corresponds to permissible reach in latitude at the boundary of nominal spacecraft stay area in altitude. Stable and minimum eccentricity magnitudes are defined. Eccentricity vector is corrected so that it equals the nominal value for spacecraft collocation and spacecraft orbit apse line is aligned with that of nodes. Spacecraft active collocation is executed at changing the inclination from 0 to iper without interaction with adjacent spacecraft control centres. At inclination larger than iper , eccentricity is increased to minimum with setting of Laplace vector in direction from the Sun. Note here that eccentricity vector is not corrected unless the end of spacecraft active existence term termination. At inclinations larger than iper, eccentricity vector equals modulo and is spaced apart relative to eccentricity vectors of the other spacecraft.
EFFECT: decreased power consumption for stay area and collocation of geostationary spacecraft.
SUBSTANCE: invention relates to control of movement of space vehicle (SV). According to the proposed method, thrusts of correction engines (CE) (control accelerations) are determined as per total changes in a period of SV revolution from correction to correction. The latter are performed with one and the same CE and a thrust level of those CE is estimated. For reliable understanding of thrusts of a pair of mutually opposite located CE, series control activations of the same pair are performed from time to time with equal pulses. Discrepancy as to total pulse of thrust is entered in equal parts with an opposite sign to the implemented pulses. As a result, reliable thrust levels of CE in operation are obtained.
EFFECT: reduction of costs and improvement of CE thrust determination accuracy as per the data of trajectory measurements, as well as improvement of SV orbit correction accuracy.
SUBSTANCE: invention relates to spacecraft (SC) motion control using solar radiation pressure forces distributed over SC working zones. The latter are formed as flat parallel optically transparent droplet flows. Distance between droplets of R radius in each flow in its lengthwise direction (Sx) and frontal-lateral direction (Sy) is divisible by
EFFECT: increase of efficiency of light pressure distributed external forces usage by means of decreasing their disturbing effect on relative SC motion.
3 dwg, 1 tbl
FIELD: aircraft engineering.
SUBSTANCE: flying launcher comprises cluster of tanks, fasteners, wing, engine and payload. Said cluster of tanks comprises two pairs of identical-volume cylindrical tanks with rocket liquid propellant of equal density and equal volume flow. Said four tanks are secured to each other by reinforcing belts to make parts of tanks with invariable center of gravity at efflux of propellant. Fasteners are secured to two tanks to allow attachment of the wing thereto. Linkage of tanks is arranged at top stage of square section and rounded angles.
EFFECT: decreased length of launcher.
4 cl, 5 dwg