Method for delivery of raw material to orbit, rocket power plant, rocket on its base, method for injection of space vehicles into geostationary orbit, transportation system for its realization and transportation-fueling system
FIELD: rocket-space equipment, mainly means and methods for water supply to low-orbital spacecraft.
SUBSTANCE: the offered method provides for use of the energy of formation of the raw material, in particular, of water from the fuel components for increasing the efficiency of the means of its injection into orbit. The offered rocket power plant has a chemical reactor, in which the given product is formed, as well as a heat-exchange unit, in which the heat of the chemical reaction is transferred to the fuel components. The latter results in the growth of the power plant specific impulse. The reaction product is cooled, and a condensate (water) is obtained which is accumulated in the storage tank. The offered rocket may use one of the cleared fuel tanks for accumulation of condensate. The offered transportation system includes the offered rocket, orbital station equipped with a system of water processing to fuel components, and means of delivery of the space vehicle to the station together with the non-filled boosting unit. The offered transportation-fueling station includes also an orbital fueling complex. Space vehicles injected into high-altitude orbits, in particular, into a geostationary orbit, as well as space vehicles returning on the Earth, may be refueled there. At injection of the space vehicle into a geostationary orbit the dependence of the efficiency of injection on the latitude of the cosmodrome is essentially reduced (by 2-3 times).
EFFECT: reduced cost of supply of the orbital stations and cost of injection of the space vehicle into a geostationary orbit, as well as into other trajectories, reduced dependence of the cost of injection of the space vehicle into a geostationary orbit on the latitude of the cosmodrome.
19 cl, 3 dwg
The invention relates to rocket and space engineering, in particular to liquid propellant rocket engines (LPRE), rocket propulsion plants (RDD) based on them, missiles, spacecraft (SC) into a geostationary orbit (GSO) and space transport and refueling systems.
A method of obtaining water and some other raw materials on Board the spacecraft, which consists in receiving them by interaction of initial chemical components in fuel cells with subsequent cooling of products of interaction ( - fuel cell, str). But fuel cells have a large specific dimensions and weight compared to the LRE. Energy intensity the most advanced electrochemical devices based on the fuel cell reaches 1.5-3.0 MJ/(kg installation) at time from several days, while the power density of the LRE usually lies in the range 1.5-2.5 MW/(kg LRE). Therefore, used on space vehicles fuel cells are thin and not used for the delivery of the raw product into orbit. Weight of raw product obtained during breeding along with power generation, can be no more than one per cent of the mass of the payload.
There is a method of delivery to the orbit of raw materials, on the example of water, consisting in the use of boosters, equipped with rocket engines for launch into orbit cargo ship with containers filled with water ( - "Progress", str).
In this way, the efficiency of delivery of water compared with General cargo may differ only by constructional features of the cargo compartment. In General, these features are related to devices for mounting and storage of cargo, and also to the use of fairings different sizes, due to the dimensions of the cargo, and not give significant differences in the efficiency of removal.
This method is closest to the proposed and selected as a prototype.
Famous rocket engines in which the working body receives energy from an external source, such as solar thermal rocket engine (RD), in which the working fluid passes through a radiant heat exchanger heated by the concentrated sunlight. The resulting high-temperature gas is fed into the engine compartment and out of the jet nozzle, creating thrust ( - solar rocket engine, str). Engines of this type have a significant drawback is the large size and great weight of the receiver-transducer of energy, referred to the unit of the generated thrust. For this reason, KMG external source e is ergie cannot be used as main RD.
Also known chemical rocket engine in which the working fluid is formed in the combustion chamber (CC) of the engine as a result of chemical reactions involving one, two or three components of consumption (CT), accompanied by release of energy - liquid propellant rocket engine, str). The most effective LRE closed circuit pump fuel components. For each component fuel rocket engine includes a high pressure pump driven by the turbine. The components of the fuel taken from the fuel tanks (TV) pumps through line low pressure, and served in the paths of high-pressure corresponding component, where the pressure of the pressure higher than the static pressure in the COP, through which fall within the COP. Part of the components used to create the working fluid of the turbine, which after passing through the last digets in the COP.
One or more such LRE together with supporting work systems, including fuel tanks and other pneumohydraulic related capacity, form RDU ( - rocket propulsion installation, str).
This RDU is closest to the proposed and selected as a prototype.
The disadvantage of the prototype are insufficiently high specific energy characteristics. If RDU to use modern rocket engines on the most effective keys is orodno-hydrogen fuel its characteristics will be able to provide the excretion PH simplest single-stage design scheme only at high perfection of the design, while the share of the mass of the payload is small. Technical and physical-chemical means to improve performance of LRE now practically exhausted. The most important characteristic of the engine is the flow rate of the working fluid from the nozzle, but LPRE it is limited by the heat of the chemical reaction CT.
Known boosters (PH), which is a missile racks with RDD-based LRE a closed circuit, including PH with three LRE, for example, multipurpose aerospace System (MAX) , which can be used for delivery to the space station as consumable components - fuel and substances necessary for life crew, and raw materials for the production of these components on the orbital station. PH contains the orbital stage with RDU, consisting of two or three engines, tank oxidizer (liquid oxygen)tank main fuel (liquid hydrogen) and the tank with the third component (kerosene). The third component increases the density of the fuel and allows you to increase the weight of the output of the load by reducing the weight of the structure. Flight of the PH takes place in two stages. On the first RDU R is operating in triple mode, in which spent fuel from all tanks and develops maximum thrust. When you run out of kerosene, begins the second phase of flight in which the RDU is running on two components, while the thrust is reduced and the specific impulse increases.
Presents the PH is closest to the proposed missile and selected as a prototype.
A prototype has two main disadvantages. The first is its efficiency, as PH General purpose, regardless of the type of payload. The second equipment of the missile more perfect LRE increases its effectiveness only to the extent caused by the improvement of engine parameters.
There is a method of SPACECRAFT at GEO using low-orbit fuel station to refuel the upper stage (RB) propellants ( - § 8,3, reusable RB option is found under the names of "orbit-to-orbit cargo ship" or "the interorbital tug"). At the fuel station can be delivered ready-made components of fuel or water, which is the raw material for the production of fuel components . The method consists of the delivery of the SPACECRAFT, together with the upper stage on an orbital fuel station, refueling the propellant components and further flight to GSO.
The presented method of excretion is the closest to the proposed method of excretion and selected as a prototype.
A known system of the SPACECRAFT to high-energy orbits, including GSO, including the orbital refueling system (PCD), orbital cargo device intended to supply the UGC, and the PH of one or more types, designed to launch geostationary satellites along with nezapravlennaya spreader blocks and orbital cargo unit on low orbit, ensuring their flight to UGC ( - § 6,2). The components of the fuel can be produced on the PCD of the raw materials of the product delivered orbital cargo unit.
Presents the exhaust system is most similar to that proposed and selected as a prototype.
Shows how to launch and SPACECRAFT at GEO has one significant drawback. Most space is at a non-zero latitude and, therefore, provide the orbit SPACECRAFT in low circular orbit with zero inclination. The flight to GSO with low circular orbits of various inclinations requires a different characteristic speed, as some of it is necessary to turn to the orbit plane. Comparative efficiency of flight to GSO orbits with different inclinations clearly depends on the difference between the required characteristic speeds. In this regard, the most effective is o f the s UGC in orbit with zero inclination. But because of the distance from the equator of the production base and the specific geographical conditions of the Equatorial zone, navigation support, supply and crew change UGC will require a significant investment in infrastructure and additional charges.
In addition, due to the fact that for the flight to the task required equipment KA additional systems, as well as additional spending fuel on manoeuvres when approaching from the UGC, the flight to GSO refueling on the task when using such PH has no advantages in efficiency before direct excretion (without using PCD).
The aim of the present invention is to reduce the cost of supply of orbital stations and the cost of the SPACECRAFT at GEO, and other trajectories, reducing the dependence of the value of the SPACECRAFT at GEO, the latitude placement of the spaceport.
The problem is solved in that the raw product get on Board the rocket during launch into orbit from the side of the fuel components by chemical interaction between them and subsequent cooling of the products of chemical interaction between components of the fuel channeled into the combustion chamber (NPL).
The problem is solved in that in the LRE RDDs are one or more heat exchange units, each of which is heated componenttemplate products of chemical interaction between the components parts fuel cooling and reduction in liquid state is heated component of the fuel, and the liquid products of combustion are collected in a separate container for further use. The components of the fuel are selected so that the liquefied products of their interaction was raw product, that could be used for the needs of the space station and were suitable for recycling components of the fuel using simple technologies.
The problem is solved by the creation of a specialized rocket to supply the orbital spacecraft of the raw product, while it is set RDD-based LRE with one or more heat exchange units, producing the final part of the flight liquid raw product, which merges in one of the emptied fuel tanks.
The problem is solved also by the fact that the AC together with the upper stage is delivered to the UGC, the upper stage dressed propellants produced by the UGC from raw product delivered specialized rocket tanker, and makes further flight to GSO, with raw product get on Board the rocket-tanker during launch into orbit from the side of the fuel components by chemical interaction between them and then cooled products of chemical interaction between components of the fuel on the monitor in the combustion chamber RDU, and bringing them into the condition, convenient for transportation.
The problem is solved also by the fact that the SPACECRAFT at GEO, we use the PH, the last stage which is delivered together with the AC on oke, disapprovals propellants, get on Board the task of the raw product delivered by rocket-tanker, and then makes a further flight to GSO.
The problem is solved also by the fact that the propulsion system orientation and orbit correction (hereinafter referred to corrective propulsion) missiles-tankers used to return to Earth compatible components of the fuel received by the UGC from raw product delivered these missiles tankers.
This invention is applicable to RD using liquid or gaseous CT and offers to get raw product in the remote control missile during flight of the products of interaction of chemical components stored on Board the rocket. The interaction products are cooled propellants and decompressions to the temperature and pressure corresponding to the storage conditions of raw product in the liquid state, and the heated components of the fuel channeled into the combustion chamber. The most simple and rational design of the missile is implemented using as chemical components components Topley is as remote from its fuel tanks, and the most effective and has acceptable properties of the components of fuel - liquid oxygen and liquid hydrogen (WW). This is due to the large gradoresurs WW, convenient physico-chemical properties of the product of the interaction of water and with a high calorific value of the reaction between the components:
The essence of the proposed technical solution is illustrated in the following graphics:
figure 1 shows the scheme of the tract fuel rocket propulsion;
figure 2 shows a diagram of the disposable rocket that uses the proposed RDU;
figure 3 shows a diagram of a rocket that uses the proposed RDU and consisting of a disposable external fuel tank and reusable block.
Diagram of the fuel supply system proposed RD RDD (figure 1) consists of a fuel pipeline of low pressure 1, the fuel pump 2, a high-pressure line 3 with the branch 4, the combustion chamber 5 with the cooling jacket 6, a fuel supply line 7 connecting the cooling jacket of the COP with the chemical reactor by the gas generator 8 with the outlet 9, which connects it with heat exchanger-condenser 10, consisting of a coolant circuit 11, the circuit of the heated body 12 and collector-drive condensate 13. The heat exchanger-condenser 10 is orientated in choru traction so, the collection drive condensate 13 is he at the bottom. At the exit of the collector-drive condensate 13 has a technological unit 14 consisting of a switch mains and separator gas and liquid phases. The input circuit of the heated body 12 is connected to the main high pressure 3. The high pressure gas line 15 connects the output circuit 12 to the input of the turbine 16, located on a common shaft with the pump 2 and forming with it turbopump Assembly. The gas pipe 17 connects the outlet of the turbine 16 with the combustion chamber 5. Mainline high pressure 3, the heating circuit 12 of the heat exchanger, the high pressure gas pipeline 15 and the pipeline 17 constitute a tract of high pressure fuel. There is a line of oxidizer high-pressure 18 connecting with the combustion chamber 5 and having a branch in the gas generator 8 - lead line oxidant 19. Process node 14 is connected through two pipelines. One of them is the return pipe 20 connects it with gas 17, the junction has a mixer 21, and in the pipeline - return valve 22. The second is the discharge pipe 23 extends from the engine and is controlled reducing valve 24. In the RDU provides the storage tank 25, which is connected to the discharge pipe 23.
The rocket (figure 2, 3) is composed of:
26 - tank of liquid hydrogen, 2 - oxidizer tank, 28 - tank dual use, 29 - fuel line liquid hydrogen, 30 - fuel line oxidant, 31 - fuel line to the third component, 32 - ternary LRE, 33 - off valve line of the third component 34 to the drain pipe of the third component, 35 - drain valve of the third component, 36 - unadvisedly part of the discharge line that has the starting valve 37.
A missile that has a reusable unit comprising a tank dual-use, and disposable tanks primary fuel and oxidant (figure 3) has in its composition: 38 - split node line of liquid hydrogen, 39 - split node line oxidant, 40 - protective case reusable apparatus 41 - explosive power connection.
To start RDU raw product on the rocket is missing. During the flight portion of fuel components is subjected to chemical interaction and the products of chemical interaction cooled components of the fuel channeled into the combustion chamber (NPL). Next, the cooled products of chemical interaction are throttling, and then form a complete liquid raw product, which is collected in one of the tanks RDU. By the end of the active site of removing the accumulation of the raw product is completely finished.
In RDU cancer of the s (1) fuel pipeline 1 is the pump 2, after which comes in line high pressure 3. Part of the fuel in one of the branches 4 line 3 is fed into the cooling jacket 5 of the nozzle and the combustion chamber 6, through which flows through fuel supply line 7 into the chemical reactor, the gas generator 8 communicates with the oxidant flowing through the supply line oxidant 19, forming products of combustion. The latter through pipe 9 serves in the coolant circuit 11 of the heat exchanger-condenser 10, where it is cooled and condensed. Cooling temperature corresponds to the liquid state of the products of the reactor when the storage pressure in the storage tank (less than 1 MPa). In the same circuit under the action of acceleration generated by the thrust of the engine, a separation of liquid and gas phase combustion products. The condensate coming down, falls into the collection drive condensate 13.
If the pressure of the combustion products above the critical (22 MPa for water), the coolant circuit 11 there is no boundary between liquid and gas phases, and this circuit does not work as a separator and as a hub, facilitating the lowering in the memory 13 a more chilled as a more dense part of the combustion products.
At the same time the main part of the fuel from the high-pressure line 3 is fed into the heating circuit 12, the heat transfer is thermal capacitor 10, forming the heated high-pressure gas. Last gas 15 flows into the turbine 16, which is used as working fluid. Leaving the turbine 16 of the exhaust gas through line 17 is discharged into the chamber 5, where it reacts with the oxidant, is also supplied to the camera via a high-pressure line 18, forming products of combustion, which creates a jet thrust.
At the initial stage of the flight the condensate from the collection drive 13 passes through a technological unit 14 in the return pipe 20, which falls in line 17 where it is mixed with fuel in the mixer 21. The pipe 20 has a check valve 22, preventing the flow of gas from the pipeline 17 in the coolant circuit 11 of the heat exchanger-condenser 10. At the time of completing the initial phase of the flight, the flow switch node 14 is triggered, after which the condensate begins to be drained from the engine in the storage tank 25 through the discharge pipe 23. The flow of condensate is controlled reducing valve 24. In the separator node 14 is the selection of the condensate residue gas phase, which is served in the return pipe 20. Before you stop the engine again actuates the flow switch node 14, tightly overlapping the discharge pipe 23 and directing the remainder of the combustion products through Britny pipe 20 into the pipeline 17.
If the pressure of the products of chemical reactor 8 above the critical value, the system remains operational, and the place of condensate pipelines are served chilled products chemical reactor 8, and the separator node 14 works as a hub, sharing more and less dense and, therefore, more or less cooled part products. In managed Panigale valve 24, the pressure of the cooled product is discharged in the storage tank 25 is fed liquid raw product.
Tract oxidant RDU can work on any of the known schemes, and also similar to the diagram above. Possible to work offline tract oxidant, and its interaction with the path of fuel through the exchange of components of the fuel and the products of the interaction of these components. Such schemes do not change the essence of the invention.
During engine operation, some parts of the paths of high-pressure components can contain a significant proportion of impurities, for example, products of the interaction of these components that are fed into the path of the coolant circuit of the heat exchanger. But in General, the nature of the environment (high-oxidizing and silnogazirovannykh paths for high pressure, respectively oxidizer and fuel) is stored, which justifies the use of a single name for a tract of each component.
Rocket, have th RDD-based ternary LRE, works in the following way. At the initial stage of the missile from a tank of liquid hydrogen 26, oxidizer tank 27 tank dual use 28 filled with the third component - fuel-based hydrocarbon fuel highways 29, 30 and 31 in three LRE 32 receives all three components of the fuel. After hydrocarbon fuel is used up, it triggers the shutoff valve 33, blocking line of the third component. After some time, actuates the drain valve 35, and the remnants of the third component is discharged into the environment through the drain pipe 34.
After closing the valve 35 at the time determined by the sequence diagram of the flight offers starting valve 37 and raw product begins to flow from the rocket engine 32 through the discharge pipe 36 into the tank dual use 28.
The way the SPACECRAFT at GEO is that the AC together with the upper stage delivered at UGC, where the upper stage disapprovals propellants produced by the UGC from the raw product, which is obtained on Board to deliver its missiles during launch into orbit from the side of the fuel components. The energy released during the formation of the raw product, is transmitted components of the fuel channeled into the combustion chamber RDU, contributing to its specific impulse. After doseproportional unit together with the AC takes a trip to GSO.
The system of the SPACECRAFT at GEO works as follows. Space tanker-based missiles, using a three-part LRE, delivers on UGC raw product in the required quantity, which is processed into components in the fuel. Obtained CT collected in petrol tanks UGC. The spacecraft is delivered to the UGC two-stage PH together with nezapravlennaya upper block, where the upper stage dressed necessary components. After refueling SPACECRAFT using a booster is displayed on the target orbit. The components of the fuel produced by the UGC, are also used for refueling returned to Earth KA, performs the task of ensuring PCD (delivery of cargo and crew).
The return portion of space tanker, forming together with the UGC transport-gas system at the start has fuel adjustment control is only required for the flight to the task. After arrival at the UGC fuel tanks corrective control gosupravlenie propellants produced by the UGC from raw product delivered space tanker, in the quantity necessary for return to Earth. After draining the raw product and refueling reusable space tanker does brake the descent in the atmosphere and landing in a specified area.
There may be some variations technical the decision, do not change the essence of the invention.
It is possible to install multiple heat exchange units on the highway one component (for example, for degreane gas which has passed through the turbine), or install one of the heat exchange unit on the fuel paths of the two fuel components.
Possible structural diagram of RD, which bypasses the heat exchanger laid bypass tract high pressure connecting with the principal before and after the heat exchanger, and in the joints installed flow switch fuel component. During the initial part of the flight component of the fuel passes through the bypass path bypassing the heat exchanger. At the beginning of the final mission phase, corresponding to the accumulation mode raw product, simultaneously with the switching of the flow of the fuel component in the primary path is switched heat exchanger Assembly. In this scheme there is no need for a return pipeline, and in case of failure of the heat transfer unit has the ability to switch it off during the continuous operation of the engine at which the rocket will be able to successfully complete flight cycle with the partial load.
The chemical interaction of the components is carried out in the chemical reactor, the simplest option which is the combustion chamber. But there may be other types of chemical reactors, for example, catalyti the definition, where the interaction takes place on the catalyst without the formation of a flame. Thus, the oxidation of hydrogen easily occurs on a Nickel or platinum catalyst.
In the proposed missile can be used, and two fuel component, then at least one of them must be posted not less than two fuel tanks.
All these variations have the same inventive entity with the option presented in the description.
The proposed method of delivery to orbit of raw product, rocket propulsion system, the missile is based on it, the way the SPACECRAFT at GEO, the transport system for its implementation and the transport-gas system combined to form a single inventive concept and meet the criteria of the invention.
The invention can be used for delivery to orbit the water and some other liquids, to supply the AC and the propellant components for SPACECRAFT at GEO, as well as other orbit flights, requiring reversal of the orbital plane. Water and its by-products can replace up to 50% or more of the output currently cargoes. In this case, propellants media are liquid oxygen and hydrogen, and payload - water. Calculations show that, without exceeding mastered currently teploprovodnyh loads on structural elements of the motor is El, it is possible to increase the specific impulse by 8-15% at the optimum ratio of components. The final output mass media increases by 10-12%, and the increase of payload can be in some cases up to 25% for partially reusable launch vehicle and more than 30% for a fully reusable launch vehicle.
The rocket using the proposed RDU will have additional benefits associated with the design feature. For example, the location of the payload (water) in the vicinity of the engine will reduce loading on the construction of fuel tanks, it does not require compliance with special rules electrophoresis, typical for normal goods. As payload is not loaded in PH before the start, and is made from components of fuel on Board the PH at the final stage of the flight, you can use it for his collection tank released the third component. The result is that the payload mass increases not only due to improvements RDU, but also by reducing the weight of the structure in PH there is no separate capacity to accommodate payload.
The proposed method of delivery of water has one distinctive feature: its effectiveness increases with the growth of final output mass of the rocket. Therefore, the use of additional what's useful devices, for example, reusable design elements, increases the comparative effectiveness of the proposed missiles before similar means of deducing General purpose. So, if reusable media General purpose reduces the unit cost of removing payload in 2 times in comparison with similar disposable media, and offer a more efficient rocket media General 1.1 times for single and 1.25 times for multiple modifications, the equipment of the missile similar with media General purpose reusable elements will increase its effectiveness in 2.27 times.
In General, the proposed missile can output 40-60% more payload than its prototype. This difference in cost will be more because of the options that simplify the design and preparation technology start-up based on the lesser of the required reliability. In addition, the cost of creating missiles can be significantly reduced because of her being noncritical dimensions for the task.
The method and system of the SPACECRAFT at GEO have one important attribute: when using converging possibilities and specific cost of removal for space located at different latitudes. For example, the optimal PH for use in the exhaust system follow the traveler: two-stage scheme, the first step is returned, is used many times; the second, non-refundable, is delivered to the UGC, disapprovals and used again when the flight to GSO. If this PH to start with the equator, then the second step must gain characteristic speed of not less than 3850 m/s, while for latitude start 51° it will be about 4800 m/s RN with the second stage, with a characteristic speed of 3850 m/s, ineffective due to strong perenaselenist first stage, and the use of the first step, weighted by the elements of salvation, will not always be able to carry a payload into orbit. In addition, geographical conditions at the equator unfavorable schema inference with landing the first stage on the route of flight. The most rational scheme of inference involves the return of the first steps for planting to the starting place. In this case, the second stage should have the ability to set the characteristic speed of not less than 5200 m/s Then the Equatorial PH in the best case will have the same proportions as the PH using other existing launch sites, and their capacity will differ slightly (due to the difference in the peripheral speed of rotation of the Earth at the point of start - no more than 250 m/s). So, first of all, the cost of putting this SC GSO will be expressed in different weight databrowsing fuel. Use the W proposed missile will result in what is the unit cost of fuel at the UGC will be less than the unit cost of removal of cargo using PH General purpose, which will lead to the convergence value of the SPACECRAFT at GEO for space located at different latitudes. In addition, for the "Equatorial" PH due to the strong perenaselenist block of the second stage can be limitations on the minimum mass refills.
If space tanker delivers water or other raw product on the task, and obtained after processing components fuel uses in his control when returning to Earth, the efficiency of the entire fuel system, including the space tanker and the task increases (for conventional means of removing such operation is economically meaningless). This is due to the fact that 1 kg neosupervital at the start of the component adjustment control will increase to 1.2 to 1.3 kg and more delivered raw product. Using generated at UGC fuel control is returned to the Land transport vehicles also reduces the cost of operating the PCD and, therefore, the cost to produce our fuel components.
Terminology in the description of the rocket, RDU and their components are borrowed from , the heat exchange unit and related devices - from .
Sources of information
1. "Space" - Encyclop the Diya. M., Soviet encyclopedia, 1985
2. "Aviation and space systems: a Collection of articles edited by Heating-Lozinsky and Aghbalyan. - M: IIA, 1997, ISBN 5-7035-2068-1
3. Cpptest "Space technology. Prospects of development. - M.: Moscow state technical University n.a. Bauman, 1997, ISBN 5-7038-1306-93.
4. Viewentry "space flight Mechanics in elementary form." - M.: Science, 3rd edition, 1980 - chap 7, §6, str.
5. "Great encyclopedic dictionary Polytechnic" or "technical dictionary". - M.: Great Russian encyclopedia, 1998; ISBN 5-85270-264-1
1. The method of delivery to orbit of raw product, such as water, using missiles, which consists in the fact that the raw product is obtained during a flight aboard a rocket from a part of the fuel components by chemical interaction between them and subsequent cooling of the products of chemical interaction, characterized in that the products of chemical interaction cooled components of the fuel channeled into the combustion chamber of the propulsion system of the missile.
2. Rocket propulsion system having at least one rocket engine with pump fuel supply system in which at least one component of the fuel in the path of high pressure, connects the output of the pump of this component with the entrance into the combustion chamber, built no less than the od of the n heat exchanger Assembly, including the heat exchanger, the contour of the heated body which is combined with a part of the tract of high pressure of this component, and a chemical reactor having a supply line for supplying the source components, the internal volume of which is connected to the coolant circuit of the heat exchanger, characterized in that it has at least one storage tank and the heat exchanging unit contains a device for removal of the cooled product of a chemical reactor of the coolant circuit of the heat exchanger in the storage tank.
3. Rocket propulsion system according to claim 2, characterized in that the heat exchange unit device for removal of the cooled product of a chemical reactor includes a lateral pipeline and a fixed adjustable lowering valve.
4. Rocket propulsion system according to claim 2, characterized in that at least one heat exchanger unit is laid Obvodny additional tract of high pressure, having a connection to the principal before and after this heat exchange unit, the ground connection is installed flow switch component fuel, and chemical reactors, heat exchange units is configured to launch during flight.
5. Rocket propulsion system according to claim 2, characterized in that it at least one is the supply line of a chemical reactor is connected with the fuel injection system of one of the propellant components.
6. Rocket propulsion system according to claim 3, characterized in that it has at least one inlet line of a chemical reactor is connected with the path of the high-pressure part of the fuel, as part of the discharge line, located between the heat exchanger and an adjustable reducing valve connected with the path of the high-pressure fuel return pipe, in this tract the high pressure in connection with the reverse pipeline is installed, the mixer, and at the junction of the outlet and return piping installed flow switch chilled product.
7. Rocket propulsion system according to claim 6, characterized in that on the return line check valve having a bandwidth in the direction of the path of the high-pressure fuel.
8. Rocket propulsion plant of any one of claim 2 to 5, characterized in that the heat exchanging unit includes a capacitor product formed in a chemical reactor, separator gas and liquid phases of the product, and device for removal of the cooled product of the reactor coolant circuit of the heat exchanger in the storage tank is configured to operate as a trap.
9. Rocket propulsion plant of any one of claim 2, 6, 7, characterized in that design, the I of the heat exchange unit includes a capacitor product, formed in the chemical reactor, the separator gas and liquid phases of the product, and device for removal of the cooled product of the reactor coolant circuit of the heat exchanger outside of the engine is arranged to operate as a trap.
10. Rocket propulsion system according to claim 9, characterized in that the discharge pipe has an additional separator gas and liquid phases, and the output of the gas phase piped to the path of high pressure fuel, which in this connection is selected mixer.
11. Rocket propulsion system according to claim 9, characterized in that the flow switch of the condensation product of a chemical reactor, located at the junction of the outlet and return pipes, provided with an additional separator gas and liquid phase with the gas phase additional separator is connected with a return pipe to the bypass switch condensate flow.
12. A rocket with rocket propulsion system with at least three fuel tanks, including at least one fuel tank, oboronyayuschihsya in flight before the two fuel tanks, abarognosis last, characterized in that the at least one fuel tank, abarognosis past two fuel tanks, abarognosis the settlement of the tournament, made tank dual purpose with the ability to accommodate both the fuel and the raw product, and rocket propulsion installation is made according to any one of claim 2 to 11, while it as storage tanks tanks used dual-purpose, and the fuel lines connecting the tanks, dual-purpose rocket engines, installed shut-off valves, in addition, the rocket is installed drain valves separating the volumes of these tanks, limited shut-off valves from the external environment, and device for removal of the cooled product of a chemical reactor connected to these tanks and provided with trigger valve.
13. Rocket on item 12, characterized in that the tank is equipped with a dual-purpose device for discharging raw product in the conditions of orbital flight, including sealed connecting node.
14. The rocket indicated in paragraph 12 or 13, characterized in that it is equipped with propulsion units system orientation and orbit correction, managed and unmanaged aerodynamic surfaces, the heat-shielding coating controls orbital flight, the atmospheric descent and soft landing.
15. The rocket indicated in paragraph 12 or 13, characterized in that the fuel tanks are not used for collecting raw product, made as separate from the rest of the rocket fuel BA is s, United with her explosive power connections and fuel lines and electrical cables, equipped with tear-off connectors.
16. Rocket through 15, characterized in that it specified the rest is made in the form of a solid block, equipped with propulsion units system orientation and orbit correction, managed and unmanaged aerodynamic surfaces, the heat-shielding coating controls orbital flight, the atmospheric descent and soft landing.
17. The way spacecraft in geostationary orbit, which consists in the delivery of LV spacecraft together with nezapravlennaya booster unit on the orbital complex, located in a low orbit, the filling of this upper stage fuel produced on-orbit refueling complex using raw product delivered by a missile, and further the flight in the geostationary orbit, wherein deliver on the orbital complex raw product is obtained during a flight aboard a rocket from a part of the fuel components by chemical interaction between them and then cooled products of chemical interaction between components of the fuel channeled into the combustion chamber of the propulsion system is aceti, and decompressional these products.
18. Transport system for spacecraft in geostationary orbit, including the orbital complex, equipped for reception of raw materials, processing them into components in the fuel, as well as for the reception, maintenance, and refueling of spacecraft and upper stages, at least one type of missile, intended to supply the orbital refueling complex commodity products, at least one type of carrier rockets, the last stage which is equipped with means for orbit-to-orbit flight to the orbital refueling complex and the means to provide refueling during orbital flight, characterized in that the missiles, designed to supply the orbital refueling complex raw products, performed on any of p-16.
19. Transport and refueling system for refueling on-orbit spacecraft and rocket stages, including the orbital complex, equipped for reception of raw materials, processing them into components in the fuel, as well as for the reception, maintenance, and refueling of spacecraft and rocket stages, at least one type of missile, intended to supply the orbital refueling complex commodity products, characterized in that the missiles,designed to supply the orbital refueling complex raw products, performed on any of PP, 16, and their propulsion systems orientation and orbit correction is compatible with at least one component of the fuel produced on-orbit refueling complex deliver missiles raw materials, and these tanks propulsion of missiles containing a compatible component of the fuel, provided with a device adapted for refueling these tanks in orbital flight conditions, including sealed connecting nodes.
FIELD: rocketry and space engineering; rocket pod engine plants.
SUBSTANCE: proposed engine plant includes propeller tanks (oxidizer tank and fuel tank), cruise engine, actuating members and high-pressure gas bottles. Oxidizer and fuel tanks are filled with low-boiling and high-boiling components, respectively. High-pressure gas bottles are installed in oxidizer tank. Rocket pod engine plant is provided with pipe lines mounted on fuel tank by means of brackets forming heat exchange unit. Pipe line inlets are communicated with outlets of high-pressure gas bottles and their outlets are communicated with actuating members of engine plant.
EFFECT: reduced mass and volume of high-pressure gas bottles and consequently reduced mass of rocket pod.
FIELD: rocket technology; heating gases using heat produced in nuclear fusion.
SUBSTANCE: proposed method is characterized in that gas is introduced in at least one chamber. The latter has wall coated with disintegrating material. This material is exposed to neutron flux to induce disintegration into fragments within chamber. Mentioned wall is cooled down on rear end relative to chamber and mentioned coating. In addition, device implementing this method is proposed. Gas heating device has at least one gas holding chamber. It has wall coated with disintegrating material and facility for exposing disintegrating material to neutron flux so as to induce and emit disintegration fragments within chamber. Device is designed to cool down mentioned wall on rear end of chamber and mentioned coating of disintegrating material. In addition, space engine using mentioned method for gas heating is proposed. This space engine has gas heating device and facility for exhausting hot gas into space to afford thrusting. Alternative way is proposed for gas heating by using nuclear fusion reaction suited to space engines for thrusting.
EFFECT: facilitated procedure of gas heating.
42 cl, 24 dwg
Known centrifugal propulsion (propeller, water-jet, vane) operate on the principle of the blades of the oars, racking up water
FIELD: rocketry and space engineering; rocket pod engine plants.
SUBSTANCE: proposed engine plant includes gas source, receiver, actuating members, pipe lines, truss, pneumatic board and other units and systems. Secured on hollow framework of pneumatic board are components of pneumo-hydraulic system of engine plant. According to first version, hollow hermetic rods of truss are used as receiver. Cavities of these rods are connected with gas source and with actuating members of engine plant by means of pipe lines. According to second version, hollow framework of pneumatic board is used as receiver. Cavity of framework of pneumatic board is connected with gas source and with actuating members of engine plant by means of pipe lines.
EFFECT: reduced mass of aircraft due to use of hollow structures as receiver.
FIELD: rocket and space engineering.
SUBSTANCE: invention relates to rocket pod propulsion systems. According to first design version, used as receiver of propulsion system of rocket pod is tower rod. Said rod is made tight and it contains blind spherical bottoms providing closed space in rod and adapter with union. According to second and third design versions, rocket pod tower is used as receiver of propulsion system. In this case tower is main tight and is provided with adapters with unions. According to second design version, channels are made in tower fittings providing connection of spaces in tower rods. According to third design version, tower is provided with pipelines connecting spaces of tower rods. According to fourth design version, rocket pod pneumatic panel skeleton is used as receiver of propulsion system. Said skeleton is welded of tubes and hollow fittings, is made tight and is provided with adapters with adapters with unions.
EFFECT: reduced mass of rocket pod propulsion system owing to use of hollow structures as receiver for operation of actuating members of pneumohydraulic systems of rocket pod.
5 cl, 4 dwg
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: space engineering; transportation of payloads in extravehicular space and in atmosphere.
SUBSTANCE: proposed flying vehicle has case, piston, shock absorber with control unit, jet engine and safety stops. Jet engine is rigidly connected with piston and safety stops are rigidly connected with case which is rigidly connected in its turn with shock absorber. Shock absorber is hydraulically connected with output of its control unit. Piston has projection located behind shock absorber and two flat projections having sections of lesser thickness in the middle. Case has two recesses receiving the said flat projections; safety stops are mounted between walls of sections of lesser thickness and walls of said recesses. At acceleration of vehicle, piston and jet engine rigidly connected with it perform reciprocating motion relative to case. Reciprocating motion is ensured by shock absorber control unit which ignites the gases contained in shock absorber; these gases repel piston from case. Return motion is performed due to operation of jet engine.
EFFECT: reduced overall dimensions; enhanced wear resistance of device.