Holder of articles in weightlessness
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.
The invention relates to space technology, namely the means of ensuring the activities of the astronauts in weightlessness, including open space, in particular, to the means of securing items to the body of the spacecraft or between themselves.
Given the versatility, this lock can be effectively used in the diving operations, transport, and other industries production and business activity, as well as in domestic operations.
In the process of life and work in zero gravity, you need to fix, fix all the objects that surround astronaut or which it uses to avoid uncontrolled drift or separation from the space object and lost.
Feature activity in weightlessness is such that the astronaut has, as a rule, keep one hand on the handrail, performing at the same time, the target action with the other hand. On this principle - "one hand" - should be designed tools and fixtures for extravehicular activity. This principle applies to funds committed.
For fixing, there are various ways and means: mechanical, magnetic, adhesive; elastic fabric, elastic bands, ribbons, cords. The materials used for fixation, are Tr is a requirement to preserve their physical properties under the influence of space factors: vacuum temperature, radiation and other
One way of fixing is binding items fabric cords, ropes. The use of these funds, tying knots for astronaut in spacesuit under pressure, with limited mobility and reduced tactile sensitivity of the fingers in a glove, extremely time-consuming task with guaranteed positive outcome. Moreover, in practice, revealed the same effect as weakening, the blooming of nodes in zero gravity and attenuation throughout a leash.
A device for fixation of the structural elements on the body of the spacecraft (RF patent No. 1804423)providing a fixation band-shaped, flat design elements on two parallel rails of a space object using staples, U-shaped, one end of which is made rigid with the C-shaped profile, and the other end is made with a spring-loaded protrusion. This device ensures the fixation of the subject only to two parallel handrails and captures only a flat, ribbon-like objects and ensures the fixation of the subject in various configurations to any structural member of a spacecraft.
Known use in engineering wire to lock against unauthorized unscrewing the locking screw connections. The conclusion is s operation of twisting the wire in this case is carried out using tools, for example pliers, which is unacceptable due to the limited locomotor capabilities of the astronaut in a spacesuit (Handbook of metal. Volume 2, s. The motor cycle", Moscow, 1958) (prototype).
Object of the present invention is the provision of versatility, ergonomic quality, safety devices when securing items astronaut in a space suit in outer space due to the fixation of objects of various configurations to any structural elements - for example, the handrail of the space object, and performing the twisting lever with one hand while using the other hand to hold the handrail of the space object.
The solution is achieved in that the locking device includes a latch in the form of a wire in a non-metallic shell with rings on the ends, the diameter of which is commensurate with the size of the fingers narutoi gloves of the suit under excessive pressure, the wire material has the property of residual plastic deformation on the latch one end of which is rigidly fixed to the lever with the other end of the slotted hole to accommodate the latch, the width of the slot commensurate with the diameter of the lock.
Figure 1 shows a locking device.
Figure 2 shows the cross section of the latch.
Figure 3 shows the lever in the cut.
The piano is the site:
1 - latch;
2 - wire;
3 - non-metallic sheath;
4 - ring;
5 - lever;
6 - slot;
7 - fixed object;
8 - the handrail of the space object;
9 - twisting.
The latch 1, consisting, for example, annealed copper wire 2 has a non-metallic shell 3, for example, PTFE, and rings 4; lock 1 one end of which is rigidly fixed to the lever 5, which is on the other end of the slot 6.
The latch is used as follows. Astronaut in spacesuit inserts fingers in rings 4, Flex the latch 1 in the form of a fixed object 7 with a design element - handrail space object 8 to which the subject is attached, grasping the hand lever 5 and rotating it, ties up the loose ends of the clamp 1 in 9 twist, thus conducting the fixation of the object attached with the other hand on the handrail, which provides both the opportunity and the safety of weightlessness.
Due to the residual plastic deformation of the material of the retainer fixation of the subject is obtained a solid, stable, resistant to the influence of space factors, twisting can be easily disassembled by the astronaut, and the latch can be used again or repeatedly.
Steps to use the latch accessible and convenient for the execution of an astronaut in a space suit under overpressure conditions of weightlessness, that follows from the experience of implementing operations extravehicular activity.
The device for fixing objects in zero gravity, containing a latch in the form of a wire in a non-metallic sheath, characterized in that at the ends of the retainer ring provided along the diameter commensurate with the size of the fingers narutoi gloves of the suit under excessive pressure, the wire material has the property of residual plastic deformation on the latch one end of which is rigidly fixed to the lever, the other end having a slot, a width commensurate with the diameter of the retainer.
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
SUBSTANCE: invention relates to aerospace engineering and can be used at lunar rocket launching complexes (LLC). Nearby LCC, on lunar surface, arranged are thermal jacket with heat accumulators, pump station, solar batteries and storage battery, thermal jacket outer surface being coated with light-reflecting film while outer surface with heat-insulation panels. Heat accumulators are filled with liquid heat carrier to half of their volume to heat its by the heat of celestial bodies at opening the covers of said thermal jacket with the help of open/close system light sensors during natural lunar day. Temperature of LCC structure elements and rocket liquid propellant components is measured during natural lunar night LCC structure elements and rocket liquid propellant components are heated by pumping liquid heat carrier heated from celestial bodies via said components and elements from charged thermal accumulators into empty thermal accumulators during the entire natural lunar night with the help of pump station pumps supplied from solar batteries or storage battery.
EFFECT: higher reliability of heating system during long-term operation of LCC.
SUBSTANCE: claimed invention relates to thioethers, suitable for application in sealant composition, which contain the structure described by formula (I): -[-S-(RX)p-(R1X)q-R2-]n- (I), in which (a) each of R which can be identical or different stands for C2-10 n-alkylene group; C2-10 branched alkylene group; C6-8 cycloalkylene group; C6-14 alkylcycloalkylene; or C8-10 alkylarylene group ; (b) each of R1, which can be identical or different stands for C1-10 n-alkylene group; C2-10 branched alkylene group; C6-8 cycloalkylene group; C6-14 alkylcycloalkylene; or C8-10 alkylarylene group; (c) each of R2, which can be identical or different stands for C2-10 n-alkylene group; C2-10 branched alkylene group; C6-8 cycloalkylene group; C6-14 alkylcycloalkylene; or C8-10 alkylarylene group ; (d) X stands for O; (e) p has value in the range from 1 to 5; (f) q has value in the range from 0 to 5; (g) n has value in the range from 1 to 60; and (h) R and R1 are different from each other. Also described is method of obtaining such thioethers, sealant compositions which contain them, as well as aerospace flying apparatuses, which contain surface with coating containing thioethers, or a hole, sealed with a sealant composition, which contains thioethers.
EFFECT: obtaining novel thioethers, possessing excellent fuel resistance and resistance to higher temperature with reduced expenditures.
32 cl, 1 tbl, 2 ex
SUBSTANCE: invention relates to space engineering, particularly, to colonisation of space objects (SO). Spacecraft (SC) comprises landing module (long-operating base module LOBM)) (LM) and take-off module (TOM) LM comprises landing devices, sealed compartment with life support system, research hardware, devices for independent or towed motion over SO surface, sealed compartment with docking and TOM to-launch-changeover system, fuel tanks for TOM fuelling and means for docking with LOBM. TOM comprises rotary jets. TOM and LM are connected by fuel overflow conduits. Soft landing is carried out either manually or automatically in horizontal position with the help of onboard BM jet running on TOM fuel and TOM engines. TOM is replenished with fuel of TOM to be changed over into launch position and to fly away. LM is incorporated with LOBM.
EFFECT: expanded operating performances of LM.
4 cl, 3 dwg
SUBSTANCE: set of invention release to space engineering, particularly, to displacement in space using space resources and can be used for hitting the dangerous space objects (DSO) Proposed method comprises selecting the comet nucleus of one of mini-comet in orbits 6 approaching the Earth 3 as hitting space body (HSB). Rocket engine unit exploiting comet nucleus evaporant as a working medium is launched in space from the Earth and placed in orbit 8 to approach said HSB. Landing thereto is carried out at point 9. Said rocket engine unit is used to change said HSB from initial orbit 7 to trajectory 10 that ensures its collision with DSO 3 at point 11. This imparts to DSB the impulse that changes it from initial orbit 4 that implies the collision of the Earth at point 5 to safe orbit. Device to this end (not shown) comprises said rocket engine unit, snapper with comet matter evaporator, power plant (with solar concentrator and astronavigation device. With snapper penetrated into comet nucleus, evaporator sublimates nucleus volatile matters. Evaporated vapors are heated by solar concentrator effuse from rocket engine nozzle generates the thrust. Astronavigator sets required direction of thrust vector.
EFFECT: power savings in mission, lower operating period of engines, higher reliability.
4 cl, 2 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