Method of control of travel of space liquid-fuelled rocket after command for disabling of mid-flight engine of burnt-out stage
FIELD: aircraft engineering.
SUBSTANCE: invention can be used for control of travel of space liquid-fuelled rocket (SLFR). After the command for disabling of the mid-flight engine (MFE) of the burnt-out stage MD is switched to the reduced-thrust stage and MD is completely disabled, the rolling travel of the rocket is controlled using two pairs of gas nozzles, the moment of MD final disabling is forecasted, one of pairs of gas nozzles is enabled before the forecasted moment of MD final disabling for creation of the control rolling moment, the pair of gas nozzles are disabled at forecasted moment, and the value of the period of work of pair of gas nozzles is determined before the flight depending on the moment of inertia of the rotating part of the turbo-pump unit with the allowance for attached mass of fuel components with reference to the axis of rotation, an absolute value of the roll moment, created by each pair of the gas nozzles at their enabling, an absolute value of angular speed of rotation of the turbo-pump unit rotor at the reduced-thrust stage, angle between the axis of rotation of the turbo-pump unit rotor and the longitudinal axis of the rocket.
EFFECT: invention allows to improve safety of SLFR flight.
The invention relates to rocket and space technology, namely to methods of motion control liquid space rocket (ILV) on the flight phase after the command on off switch mounted in the gimbal of the main engine (MD) spent the ILV stage and before the separation of that degree.
In rocket technology known selected as a prototype method of controlling liquid movement ILV after a shutdown command and the MD exhaust stage, consisting in the translation of the CBMs at reduced thrust, and the final shutdown of the CBMs by stopping delivery of the fuel components in the combustion chamber .
The disadvantage of this method is that its implementation operates on the ILV disturbing moment, which can lead to LV twist around its longitudinal axis. This point, due to the braking of the rotating parts of the turbopump Assembly (TNA) and carries their weight fuel components.
To parry disturbing moments on a roll at LV tandem layout mid-flight with one engine usually used for more steering motors or remote gas nozzle roll, which produces gas for the generator MD. However, the effectiveness of the gas nozzle is directly proportional to thrust sustainer engine and by the time that stops fridaysaturday TNA, and LV becomes the maximum angular velocity about the roll, the gas nozzle are already inefficient and not able to fend off the perturbations caused by the rotation of the LV around its longitudinal axis. Therefore, given the systematic nature of the disturbance, opposing him from the moment of the gas nozzle must be created in advance.
The object of the present invention to provide a method for motion control of LV after a shutdown command and the MD exhaust stage, providing counteracting the disturbing torque due to the braking of the rotating parts of TNA and thereby prevents undesirable twisting of the LV on the roll, which in combination with other adverse conditions may result in "folding" part of the gyrostabilized platform and emergency flight termination ILV.
The technical result of the invention is to increase the safety of the flight to LV.
Said technical result is achieved in that in the method of controlling the movement of liquid ILV after a shutdown command and the MD of spent stages, including the translation of CBMs for reduced thrust and the final shutdown of the CBMs by stopping delivery of the fuel components in the combustion chamber, in accordance with the invention, in the case where the motion control ILV roll OS�carried out using two pairs of gas nozzles, gas to the gasifier which produces MD, carried out at the beginning of the forecast time (t0final cutoff MD, in advance, a time interval (Δt) to the predicted point in time (t0final shutdown of the CBMs include one pair of gas nozzles, creating a control torque roll, the sign of which is opposite to the sign of the angular velocity of rotation of the rotor TNA, then turn off the specified pair of gas nozzles in the predicted time (t0), while the value of time interval (Δt) of the work of this pair of gas nozzles is determined before the beginning of the missile by the formulawhere IP- the moment of inertia of the rotating parts of the TNA with regard to added masses of propellant relative to the axis of rotation,MX- the absolute value of the roll moment created by each pair of gas nozzles at their inclusion, Ω is the absolute value of the angular velocity of rotation of the rotor turbopump on the reduced thrust, φ is the angle between the axis of rotation of the turbopump rotor and the longitudinal axis of the LV.
The essence of the invention is illustrated in Fig.1.
Fig.1 - typical dependence of the parameters of angular motion of the LV from time after a shutdown command and the MD.
As an example, consider a possible implementation of the proposed method of control of SR light class t�PA Korean missiles "KSLV-I". The combustion of the sustainer stage I engine this LV is installed in the gimbal. Motion control of LV in pitch and yaw is accomplished by deflection of the camera by means of two electro-hydraulic servos. To control the movement of the roll used 4 gas nozzle, which are included in pairs. At the end of stage I and the control system sends the command to shut down its MD, after which the engine is transferred to the reduced thrust (finite thrust - KCT), constituting 38% of nominal. In this mode, you can still control the movement of LV in pitch and yaw by turning the combustor MD. In the channel of the roll mode, the FTC used the gas nozzle, albeit with reduced efficiency. The turbopump rotor, feeding fuel components into the combustion chamber, on the mode of KST rotates with angular velocity Ω≈12600 rpm, the angle between the axis of rotation of the rotor and the longitudinal axis of the LV is φ=6°. Rotating turbopump possess angular momentum (kinetic), the projection of which on the longitudinal axis of SR is equal to K1=IPΩcosφ, where IP- the moment of inertia of the rotating parts of the turbopump relative to the axis of rotation with regard to added masses of propellant (mass components between the shoulder blades pump the fuel and oxidizer), approximately equal to the 0.375 kg·m2.
When finished�relatively off the MD stops supply of the fuel components in the combustion chamber. Angular velocity of rotation of the rotor turbopump approximately 0.9 to fall to zero. The law of conservation of angular momentum of a closed system of LV-turbopump rocket will acquire angular momentum relative to the longitudinal axis, is equal to K2=IXωX0where IX- moment of inertia of LV relative to the longitudinal axis of approximately $ 16500 kg·m2at the end of the sustainer stage I engine. While (K2=K1, ie, LV, begin to rotate around the longitudinal axis with an angular velocity
For the numeric values of the parameters of the angular velocity of rotation will be about 1.7 g/S. For the repayment of this angular velocity must include one of the two pairs of gas nozzles, namely, that which generates control torque roll, the sign of which is opposite to the sign of the angular rotation speed of the turbopump. This pair of gas nozzles should be fixed on time
where MX- the absolute value of the roll moment created by each pair of gas nozzles at their inclusion (MX≈114 kg·m when running MD on the reduced thrust). For these parameter values, the inclusion of a pair of nozzles will be Δt≈0,44 S.
Therefore, in accordance with the method of control to prevent unwanted rotation of LV on the roll, it is proposed p�ed the beginning of the flight by the formula (2) to determine the duration of the time interval Δt, during which should work a pair of gas nozzles. In flight, at the end of stage I is the forecast of time t0the final engine is turned off. This prediction can be done either on the basis of information coming into the control system from the sensors of the levels of the oxidant and fuel, installed in the tanks of the rocket and allows to determine the mass of the remaining fuel components, either on the basis of a comparison of the current values of the trajectory parameters (speed, altitude, etc.) with their program values (method of implementation of the prognosis depends on the accepted principles of the system of missile guidance). After the time t0final off the MD predicted, in advance, for the time interval Δt to the time t0in accordance with the method of administration includes one pair of gas nozzles, creating a control torque required sign (opposite sign to the angular rotation speed of the turbopump). Off this pair of gas nozzles at time t0.
Fig.1 presents the results of mathematical simulation of angular motion of the LV in the channel of the roll, since the inclusion of a pair of nozzles roll. Shown:
- the time dependence of the angular velocity of rotation of the rotor TNA Ω·0.001, rad/s;
- the time dependence of the angular velocity in�of Amenia ILV roll (longitudinal axis) ω Xg/C;
- the time dependence of the roll angle γ LV, g;
- the time dependence of the team on the inclusion of one of the pairs of gas nozzles u (u takes on the values -1, 0, +1. u=-1 means includes a pair of gas nozzles, creating a negative point on a roll, u=0 means that the gas nozzle off). From Fig.1 shows that the proposed control method is already using 1.4 with after turning the gas nozzle reduces the angular velocity ωXalmost to 0, the absolute value of the roll angle γ LV does not exceed 1,1°.
Thus, thanks to the implementation proposed in the invention of the technical solution solves the problem of parry disturbing torque due to the braking of the rotating parts of TNA, and achieved technical result of the invention - improving the safety of flight ILV.
Sources of information
1. V. I. Fedoseev. Basic techniques of rocket flight. M.: "Nauka", 1981, p. 139.
A method of controlling liquid movement of a space rocket after a shutdown command and the sustainer engine exhaust level which contains the translation of the main engine at reduced thrust, and the final shutdown of the main engine by stopping delivery of the fuel components in the combustion chamber, characterized in that in the case where the motion control of the missile roll OSU�applied using two pairs of gas nozzles, gas to the gasifier which produces sustainer engine, carried out at the beginning of the forecast time (t0final shutdown of the main engine, in advance, a time interval (Δt) to the predicted point in time (t0final shutdown of the main engine, include one pair of gas nozzles, creating a control torque roll, the sign of which is opposite to the sign of the angular velocity of rotation of the rotor of the turbopump unit, then turn off the specified pair of gas nozzles in the predicted time (t0), while the value of time interval (Δt) of the work of this pair of gas nozzles is determined before the beginning of the missile by the formulawhere IP- the moment of inertia of the rotating parts of the turbopump unit with regard to added masses of propellant relative to the axis of rotation, MX- the absolute value of the roll moment created by each pair of gas nozzles at their inclusion, Ω is the absolute value of the angular velocity of rotation of the rotor of the turbopump unit at reduced thrust, φ is the angle between the axis of rotation of the rotor of the turbopump unit and the longitudinal axis of the rocket.
FIELD: physics, navigation.
SUBSTANCE: group of inventions relates to interorbital, including interplanetary, flights of rocket propelled spacecraft. A method of constructing an optimal spacecraft trajectory is based on solving a two-point boundary value problem of the Pontryagin maximum principle and taking into account characteristics of the macro- and microstructure of the cost function. The latter can be the time of flight or fuel consumption during flight. Analytical bases for efficient search of initial domains of values of Lagrange multipliers at each iteration are established. This facilitates the construction of a series of sub-optimal solutions which converge to an optimal solution. A corresponding algorithm yields the optimal solution last or, in case of unattainability thereof (due available resources of the spacecraft) a solution close to optimal. An electronic processor for implementing the method and a spacecraft with said processor are also disclosed.
EFFECT: faster operation, improved convergence, low qualification requirements and wider field of use of the disclosed algorithm and accompanying equipment.
16 cl, 7 dwg
FIELD: physics, atomic power.
SUBSTANCE: invention relates to atomic power engineering and space-rocket engineering. The spacecraft nuclear propulsion system comprises a heater - gas-cooled nuclear reactor, a cooler, a recuperative heat exchanger, a pipe system with a gaseous working medium, coaxial turbine-compressor-electric power generator, electric jet engines, an automatic control system with measurement and control means. The number of loops of the turbine-compressor-electric power generator with equal electric power is a multiple of two with opposite direction of rotation of rotors of the turbine-compressor-electric power generator in each pair, wherein the pipe system connects the output of the heater - gas-cooled nuclear reactor with the input of each turbine, and the output of the turbine with the input of the channel of the heated gaseous working medium of its recuperative heat exchanger, the output of the channel of the heated gaseous working medium of the recuperative heat exchanger with the input of its cooler, the output of the cooler with the input of its compressor, the output of the compressor with the input of the channel of the cold gaseous working medium of its recuperative heat exchanger, the output of the channel of the cold gaseous working medium of each recuperative heat exchanger with the input of the heater - gas-cooled nuclear reactor.
EFFECT: high efficiency and reliability of the spacecraft nuclear propulsion system.
19 cl, 2 dwg
SUBSTANCE: invention relates to space engineering and can be used in carrier rockets.. Proposed rocket comprises head unit with payload, parallel separable rocket stages with multichamber engines with fuel tanks shaped to torus, tapered tail, short central body at first stage, single trough-like nozzle at second stage, bottom part composed of outer and inner cones composed by outer and inner surfaces of short central body shell and inner surface of single trough-like nozzle shell. Fuel tanks and single trough-like nozzle are arranged inside short central body between first-stage tanks.
EFFECT: decreased bottom resistance, higher specific pulse.
5 cl, 9 dwg
SUBSTANCE: invention relates to space engineering, particularly, to astronaut operation in weightlessness. Proposed holder comprises retainer composed by wire (made of afterflow material) in non-metallic sheath, ring at retainer end in diameter comparable with sized of fingers of inflated space-suit glove, lever with opening in diameter comparable with retainer diameter.
EFFECT: higher safety of articles retention in open space.
SUBSTANCE: invention relates to space engineering, particularly, to astronaut operation in weightlessness. Proposed holder comprises retainer composed by wire (made of afterflow material) in non-metallic sheath, rings at retainer end in diameter comparable with sized of fingers of inflated space-suit glove.
EFFECT: higher safety of articles retention in open space.
FIELD: engines and pumps.
SUBSTANCE: pulse is obtained by ejection of gasified liquid residues of unused components of rocket propellants (RP). Pulse is generated by combustion of unused components of rocket propellants (RP) on rocket gas engine combustion chamber. Volume of unused propellant residues is limited to divide a second heat carrier mass flow rate into parts, one being fed in tank section confined by the screen while another portion being fed into tank second part. Amount of fed heat carrier is defined proceeding from evaporation of residual propellant component drops. Device for withdrawal of separable carrier rocket section comprises oxidiser and propellant tanks, tank supercharging system, rocket gas engine with feed and gasification systems. It incorporates feed lines with acoustic radiators (calculated proceeding from minimum mass loses for gasification by preset amounts of propellant and pressure). Said separation screen is calculated proceeding from surface tension force.
EFFECT: reduced power consumption for gasification.
3 cl, 4 dwg
SUBSTANCE: invention relates to space engineering and can be used for attachment and separation of cluster-configuration of carrier rocket. Proposed device comprises air operated pusher, attachment assembles and lock. Air operated pusher comprises cylinder with rod equipped with turn keys, spherical joint with ball lock and retainer piston, structural rod secured at bearing structure nearby wall second stage. Cylinder comprises extra cavity for rod pull-in.
EFFECT: higher reliability, decreased weight.
3 cl, 9 dwg
FIELD: aircraft engineering.
SUBSTANCE: invention relates to design and thermal control of spacecraft in weight of up to 100 kg launched as parallel payloads. Spacecraft unpressurised parallelepiped-like container has cellular panels (3, 4, 5) with instruments (2) installed threat. Heat from instruments (2) is uniformly distributed over said cellular panels by means of manifold heat pipes (6). Note here that instruments are stabilised thermally. Notable decrease in instrument heat release switches on the electric heaters at upper cellular panel (3). This allows a tolerable temperature of instruments to be ensured by cellular panel and heat pipes (6). Lower cellular panel (4) is directed towards the Earth and represents a radiator design. Upper and lower panels are interconnected by adjustable diagonal struts (8). Shield-vacuum heat insulation (9) is arranged at lateral faces of instrument container without cellular panel. Said insulation is arranged at screen structure secured at cellular panel, on inner side of solar battery panes (1).
EFFECT: decreased weight, enhanced performances on mini- and micro-spacecraft.
SUBSTANCE: invention relates to cosmonautics and can be used for safeguarding Earth against collision with dangerous cosmic body. Moon launch missile system comprises launching table located directly on the Moon surface, thermal casing placed on launching table and having opening cover at the top, mirrored outer surface and inner surface covered with heat insulating material (teflon, polytetrafluoroethylene, polychlorotrifluoroethylene, crystalline copolymer of ethane with tetrafluoroethylene), temperature-control system with heat accumulators and heater, power source, jet-propulsion solid-fuel missile with payload of 5-9 tons and takeoff mass of 20-30 tons. The launching table in the central part has translating cover to exhaust gases during missile takeoff.
EFFECT: invention permits to improve the Earth safety against collision with dangerous cosmic body.
SUBSTANCE: invention relates to space engineering and can be used for increasing the radiation safety of manned spaceship crew. Spaceship comprises shuttle unit, working compartment, power plat with fuel store and adaptor stage. The latter is provided with hatches with tight covers and is arranged inside fuel tank to communicated working compartment with shuttle unit. At increased radiation level the crew moves into adapter stage to be isolated by covers.
EFFECT: higher radiation safety.
2 cl, 1 dwg
FIELD: rocketry and space engineering; cryogenic stages of space rockets.
SUBSTANCE: according for first version, oxidizer supply unit is shifted in transversal direction and is secured in lower point of convex part of lower head plate of oxidizer tank, thus forming additional space in inter-tank compartment in axial direction; this space is used for displacement of cruise engine together with fuel tank towards oxidizer tank. According to second version, oxidizer supply unit is secured on concave part of lower head plate of oxidizer tank. Full suction of oxidizer from tank is performed by means of passages of intake unit introduced into concave part of lower head plate of oxidizer tank and used for coupling the lower zone of oxidizer tank with oxidizer supply unit inlet.
EFFECT: improved mass characteristics due to reduction of overall dimensions in length.
FIELD: rocketry and space engineering; designing artificial satellites.
SUBSTANCE: proposed spacecraft has modules where service equipment is arranged and modules where target equipment and command and measuring devices are located. Optical devices of target equipment of infra-red range with cooled elements are mounted in central module. Radio equipment of on-board repeater is arranged in side modules whose position is changeable relative to position of central module. Optical and command and measuring devices are mounted on one frame at reduced coefficient of linear thermal expansion; they are combined with central module through three articulated supports. Cooled elements of optical devices are connected with radiators located beyond zone of thermal effect; service equipment module is provided with solar batteries having low dynamic effect on accuracy of spacecraft stabilization. Besides that, this module is provided with plasma engine whose working medium excludes contamination of said optical devices.
EFFECT: enhanced accuracy of spacecraft stabilization; electromagnetic compatibility of systems.
FIELD: rocketry and space engineering; adapters for group launch of spacecraft.
SUBSTANCE: proposed adapter has body consisting of two parts: one part is made in form of load-bearing body with platform for placing the spacecraft on one end and with attachment frame on other end; other part is made in form of load-bearing ring secured on payload frame and provided with attachment frame. Attachment frames of load-bearing body and load-bearing ring are interconnected by means of bolted joints fitted with two rubber washer shock absorbers each; one of them is mounted between surfaces of attachment frames to be coupled and other is mounted between opposite surface of attachment frame of load-bearing body and metal washer laid under bolt head. Diameter of metal washer exceeds diameter of rubber washer shock absorber; spacecraft attachment units are secured on platform of load-bearing body by means of bolted joints with rubber washer shock absorbers mounted between platform surfaces to be coupled and spacecraft attachment units.
EFFECT: reduction of dynamic vibration and impact loads due to extended range of varying dampening properties of adapter.
6 dwg, 1 tbl, 1 ex
FIELD: future space engineering; interstellar flights.
SUBSTANCE: proposed method is based on use of reactive thrust of spacecraft rocket engines in their maneuvering in gravity field of black hole. Kerr (rotating) black hole, i.e. its ergosphere may be selected for the purpose. Several separate spacecraft are directed in succession to gravity field of black hole ensuring stable exchange of information among them (for example, by radio or light channel). Provision is made for acceleration of spacecraft to relativistic speeds and obtaining information on effect of such speeds and accelerations on physical processes, equipment and living beings (at safe flying out of sphere of influence of black hole), as well as verification of theories of black holes.
EFFECT: enhanced efficiency.
FIELD: rocketry and space engineering; upper stages of launch vehicles injecting payloads from reference orbit into working orbits.
SUBSTANCE: proposed cryogenic stage includes cruise engine, oxidizer tank, toroidal fuel tank, inter-tank compartment, truss for connection with payload and truss for connection with launch vehicle. Toroidal fuel tank is made in form of lens in cross section with bottoms changing to frames. Tank is coupled with said trusses and inter-tank compartment through outer frame forming load-bearing system for taking-up external inertial loads.
EFFECT: reduction of total longitudinal clearance and mass of cryogenic stage; increased zone of payload under launch vehicle fairing.
FIELD: rocketry and space engineering; scientific and commercial fields.
SUBSTANCE: proposed method includes placing payloads on injection facility, launching the launch vehicle, separation of injection facility from launch vehicle and injection of injection facility into geocentric orbit where said payloads are separated from injection facility. Main payload is placed on injection facility directly of body of accompanying payload; this body combines its functions with functions of main load-bearing member of adapter system for placing the main payload. After separation of injection facility from launch vehicle, additional acceleration of injection facility is performed and injection facility is injected into reference orbit and then it is shifted to geocentric orbit where main and accompanying payloads are separated. Accompanying payload is separated from injection facility after main payload is at safe distance without waiting for complete turn of main payload. Spacecraft in facility injecting the artificial satellites into geocentric orbit are placed in succession on injection facility beginning with lower one. Main payload in form of one or several spacecraft is placed on body of lower spacecraft through separation device. Body of lower spacecraft combines its functions with functions of adapter load-bearing member for placing the main payload.
EFFECT: increased mass ratio of launch vehicle and injection facility; extended functional capabilities.
3 cl, 2 dwg
FIELD: space engineering; spacecraft for descent in atmosphere of planet.
SUBSTANCE: proposed spacecraft has case with foldable wings and/or stabilizers provided with deployment mechanisms. In folded state at deceleration of spacecraft in atmosphere, said wings and/or stabilizers are covered with separable frontal heat shield which is oval in shape in projection on plane perpendicular to longitudinal axis of spacecraft. Side surfaces of tail section of spacecraft case with wings and/or stabilizers (and some other members) may be covered with separable aerodynamic flaps which form conical surface. After deceleration at initial stage of descent, shield is separated and wings (stabilizers) deploy to working position. Proposed spacecraft has high aerodynamic properties and is provided with reliable protection against aerodynamic and thermal loads at deceleration at high supersonic flight speeds.
EFFECT: low cost of servicing.
4 cl, 13 dwg
FIELD: construction of large-sized structures in space; space engineering.
SUBSTANCE: proposed settlement includes production, living and auxiliary rooms built from lightened modules which are combined in single complex of cylindrical shape with tunnel located along its main longitudinal axis. Said tunnel is embraced by three bodies: main body, body of communication chambers and body of transfer chamber. Pressurized passages are provided between these chambers. Gravitational drive mounted on tunnel is used for rotating the complex in order to form artificial gravity in all rooms. Facing secured on outer surface of guards is used for protection against adverse effect of space. Guards of main body are made from torous members assembled from enlarged space building modules. Main body may be provided with hollow longitudinal and radial stiffening members fastened together and secured to said torous members and to tunnel.
EFFECT: reduced labor consumption and time required for assembly of space structure.
2 cl, 8 dwg
FIELD: space engineering.
SUBSTANCE: proposed method includes joint assembly of payload and launch vehicle for forming space launch vehicle which is equipped with apogee stage with solid-propellant engine plant. Carrier-aircraft is coupled with space launch vehicle and launch vehicle is raised by this aircraft to preset altitude, then launch vehicle is separated and solid-propellant engine plants of three boost stages are started in succession; launch vehicle is injected into preset near-earth orbit and payload is separated from launch vehicle at preset point of trajectory in preset direction. In the course of flight of launch vehicle upon discontinuation of operation of engine plants of boost stages and completion of first boost leg, ballistic pause is performed at motion of space launch vehicle over ballistic trajectory at climbing the required altitude of orbit. Upon completion of ballistic pause at second boost leg engine of apogee stage is started and space launch vehicle is injected into preset near-earth orbit at respective velocity increment and compensation of error during operation of boost stages. Aircraft rocket space complex includes 1st class aerodrome, carrier-aircraft and space launch vehicle. Masses of boost and apogee stages are selected at definite ratio. Provision is made for transportation container for delivery of space launch vehicle to aerodrome. Telemetric information measuring and tracking points are located on aeroplanes; they are made in form of mobile radio unit for reception of external information.
EFFECT: reduction of distance from launch site of space launch vehicle to point of separation of payload.
18 cl, 11 dwg
FIELD: space engineering.
SUBSTANCE: proposed method includes transportation of space launch vehicle to launching position, preparation for launch, raising the space launch vehicle to preset altitude by carrier-aircraft, separation from carrier-aircraft, stabilization of space launch vehicle and starting the engine plant of first boost stage. Space launch vehicle is transported to launching position in transportation-and-operation container. Then, container is transferred by means of crane to erection trolley, detachable compartments are dismantled and space launch vehicle is transported to carrier-aircraft. Space launch vehicle is secured to carrier-aircraft by means of locks of carrier-aircraft. Space launch vehicle is equipped with boost stages with solid-propellant engine plants, stabilization unit and units for attachment of launch vehicle to carrier-aircraft. It is also equipped with separable tail fairing and lattice stabilizers made in form of cylindrical panels which are secured on it. After bringing the space launch vehicle to preset altitude, locks of carrier-aircraft are opened by command and lattice stabilizers of tail fairing are opened simultaneously. After preset pause, before separation of space launch vehicle, tail fairing with lattice stabilizers is separated from space launch vehicle. Proposed method makes it possible to reduce launch mass and ensure stabilization on flight leg of safe distance from carrier-aircraft till moment of start of 1st stage engine plant.
EFFECT: extended field of application.
7 cl, 5 dwg