Space and missile system
SUBSTANCE: invention relates to space and missile equipment and can be used for final carrier rocket stages. A space and missile system (SMS) includes a carrier rocket with a final stage with an outer housing compartment with an intermediate power support frame with outer and inner frames connected to each other by means of fastening elements; a space vehicle with the main fairing with an end frame. The outer diameter of the intermediate support power frame corresponds to the diameter of the end frame of the main fairing.
EFFECT: invention allows attachment of different standard sizes of main fairings with carrier rockets without increasing the time for assembly and preparation for SMS launching.
The invention relates to rocket and space technology, namely the means of removing devices for space applications at specified orbits.
In a recent application in rocket and space systems, spacecraft, large volume demanded the development of a nose cone with a larger diameter, resulting in a transition of a head fairing to the last stage of launch vehicles, having a smaller docking port diameter, is carried out using a reverse cone head fairing.
Known missile and space system according to the patent of the Russian Federation 2351510 consisting of a launch vehicle and spacecraft with nose fairing prototype.
To dock with the nose fairing on the last stage of the launch vehicle is mounted a removable compartment and an external diameter which mates with the nose fairing diameter greater than the diameter of the last stage of the carrier rocket and goes beyond the current restrictions, rail and air transport, which leads to the need of the removable compartment to transport to the launch site separately from the last stage of the carrier rocket, and then to carry out their Assembly. This increases Assembly time and preparation for launch rocket-space systems - the disadvantage of the prototype.
The object of the proposed invention is the creation of space-rocket system�eat which provides docking of various sizes nose cone of the launch vehicle, having at the last stage of smaller diameter of the connecting compared to the nose fairing, without increasing Assembly time and preparation for launch rocket-space systems.
The object is achieved in that in the missile and space system containing a booster, which includes the last stage with the outer hull compartment, spacecraft and fairing, the composition of the shell compartment of the last stage of the carrier rocket power introduced intermediate support frame is rigidly connected at its end with end frame payload fairings, forming a one-piece in-flight connection, and the outer diameter of the intermediate power of the reference frame corresponds to the diameter of the end frame payload fairings, with power intermediate support frame consists of outer and inner frames, coupled together with fasteners through the outer shell of the body compartment of the last stage of the carrier rocket
Fig.1 shows a space-rocket system, Fig.2 shows the connection of a head fairing with the last stage of the carrier rocket, where:
3. fairing;4. external Cabinet compartment of the last stage of the carrier rocket;
5. the last stage of the carrier rocket;
6. power intermediate support frame;
7. end frame payload fairings;
8. outer frame;
9. the inner frame;
12. the separation device.
Proposed rocket-space system consisting of booster 1, which has produced the last step 5 with the outer housing compartment of the last stage of the carrier rocket 4, spacecraft 2 with nose fairing 3, the composition of the shell compartment of the last stage of the carrier rocket 4 power introduced intermediate support frame 6 is rigidly connected at its end with end frame head fairing 7, forming a one-piece in-flight connection, and the outer diameter of the power intermediate support frame 6 corresponds to the diameter of the end frame payload fairings, 7, at that power intermediate support frame 6 consists of 8 external and 9 internal frames, coupled together by means of fastening elements 11 through the outer shell 10 shell compartment of the last stage of the carrier rocket 4.
The introduction of power intermediate support frame 6 is provided dock operated launch vehicles 1 (with a diameter of last stupas�no booster 5 less than the diameter of the lower frame payload fairings 8) operated with the head fairing 3 different sizes, with outer diameter power intermediate support frame 6 corresponds to the diameter of the lower frame payload fairings 7 and is located within the existing dimensional constraints, railway and air transport.
In addition, unshared in flight, the power connection intermediate support frame 6 with a sufficiently powerful end frame payload fairings 7 provides reliable transfer of load from the head fairing 3 on the outer Cabinet compartment of the last stage of the carrier rocket 4.
Department head fairing 3 from booster 1 is manufactured on shared flight to the junction of the head fairing 3 with the end frame payload fairings 7.
The proposed space-rocket system operates as follows.
In the operation of rocket and space systems fairing 3 prior to his separation from the launch vehicle 1 transmits the load on the junction of the head fairing 3 with a force intermediate the supporting frame 6 of the last stage of the carrier rocket 5.
After passing the space rocket system dense layers of the atmosphere fairing 3 is divided in the longitudinal direction at two folds, production�tsya compartment folds of the head fairing 3 from end of frame payload fairings 7 using the separation device 12. End frame payload fairings 7 remains at the last stage of the booster 5, and after separation of the booster 1 from the last stage booster 5 is separated from her Cabinet compartment last stage of the rocket carrier 4 together with the end frame payload fairings 7.
The implementation of this proposal in missile and space system allows for the docking of various sizes fairings with 3 launchers 1 having at the last stage of the carrier rocket 5 with a smaller diameter nose fairing 3, without increasing Assembly time and preparation for launch rocket-space systems.
Rocket-space system containing a booster, which includes the last stage with the outer hull compartment of the last stage of the launch vehicle, spacecraft with nose fairing, characterized in that the composition of the shell compartment of the last stage of the carrier rocket power introduced intermediate support frame is rigidly connected at its end with end frame payload fairings, forming a one-piece in-flight connection, and the outer diameter of the intermediate power of the reference frame corresponds to the diameter of the end frame payload fairings, with power intermediate support frame consists of the outer and inner�th frames, coupled together with fasteners through the outer shell of the body compartment of the last stage of the carrier rocket.
SUBSTANCE: invention relates to systems for payload delivery to upper air and higher. Proposed system (1) comprises tubular rocket launching cart (2) with friction drives of cable-rope path (26) displacing below two-pin hinge (63) secured to ground and lifted to coaxial portable tube (124, 143) extending to three main fastened cables/ropes (27), their weight being equalized by balloons (164). Then, said cart displaces to connector station (166) retained above the ground in stratosphere by two secondary cables/rope (184) suspended from fastening frame (162) to tensioning of said balloons. Said cart is retained by end grip (196) of said cart directed in two secondary and two tertiary cables/rope (186) to be lifted by bottom lifter (198) guides by secondary cables. Said bottom lifter is retained by top lifter (168) suspended from fastening frame of tensioning balloons. Said cart hooked by lifting ring (183) guided by secondary cables/ropes upward revolves in necessary direction with release of rocket and, fact, recoilless ejection during free fall of said cart downward with engine ignition at safe distance.
EFFECT: safe, non-polluting reusable system.
50 cl, 67 dwg
FIELD: aircraft engineering.
SUBSTANCE: rocket cryogenic upper stage (RCUS) designed according to the tandem layout comprises a fuel tank with an instrument compartment and transitional system for fastening of a spacecraft, an oxidizer tank (OT), intertank spacer, RCUS mid-flight engine (MFE), an intermediate compartment, fire and explosion prevention system, thermal mode maintaining system with the unit of demountable connections of communication with the land equipment and separable inlet pipelines, manifolds for purging of stagnant zones and device for maintaining of the thermal mode of the zone and RCUS equipment, a sealing diaphragm, the detachable head fairing (HF) with windows for detachment of the fire and explosion prevention system and devices for maintaining of the thermal mode of gases for purging of RCUS zone, additional thermal insulation of RCUS zone, a part of separable inlet pipes of manifolds with demountable joints and the unit of demountable connections for communication with the land equipment, an intertank spacer, conjugated with the intertank frame for fastening OT with MFE and conjugated to the top spacer of the separated intermediate compartment with the units of connection and separation with US and HF.
EFFECT: invention allows to improve fire and explosion safety of upper 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.
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: 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