Liquid propellant rocket engine

FIELD: engines and pumps.

SUBSTANCE: liquid propellant rocket engine comprising a combustion chamber, a turbopump set, a drainage cavity connected with a drainage pipeline, at the same time the drainage cavity is located between an oxidant pump and a turbine, and the drainage pipeline is equipped with a gas ejector, at the same time the gas ejector is connected by a pipeline with a cavity downstream the turbine. The pipeline comprises a valve and a throttle.

EFFECT: increased efficiency of a system of LPRE cavities draining and removal of fuel components from them.

2 cl, 2 dwg

 

The present invention relates to liquid propellant rocket engines (LPRE) and can be used in other areas of technology. In a rocket engine that uses two-component fuel, there is always the task of preventing the connection of the two fuel components in places where it is not provided by the circuit of the motor. This problem is traditionally solved in different ways:

for pipelines and dead-end cavities by blowing an inert gas;

- for units of automation and TNA - installing the seals, the introduction of barrier cavities with drainage leaks fuel components into the environment through a special drain piping. So designed the units most famous LRE, for example, on the engine ALO-137 - prototype (see "Foreign aircraft and rocket engines", CIAM, 1971, page 467). In most cases, the use of these measures is sufficient to ensure the safe operation of the engine. However, in some cases there is a necessity to prevent leakage of oxidant into the cavity of the turbine and the discrepancy between the magnitude of the leakage and capacity of the drainage pipe. Thus arises the need for drainage of the leakage and its intensification. Use purge for this purpose is not always possible, and sometimes backfires, as hasproduced requires certain migratory areas and can "push" drained component from the drainage channel. This is a disadvantage of the known technical solutions.

Known liquid-propellant rocket engine for RF patent for the invention №2484284, IPC P02K 9/42, publ. 10.06.2013,

The drawback is the large weight of a venting system due to the presence of a massive cylinder of compressed gas.

The task of creating the present invention is to reduce the weight of the engine and increase the effectiveness of the system of drainage cavities rocket engine and remove the fuel components of them, are accumulated due to unauthorized leaks.

Achieved technical result - the reduction of the weight of the engine.

This goal is achieved in liquid rocket engine containing a combustion chamber, turbopump Assembly, drain cavity, connected to a drainage pipe, with a drainage cavity is located between the oxidizer pump and turbine, and the drain pipe is provided with a gas ejector connected with a cylinder of compressed gas so that the gas ejector is connected to the cavity for the turbine.

The invention is illustrated in the drawings, figures 1 and 2, where:

figure 1 shows the scheme of the engine,

figure 2 - option power scheme ejector.

The diagram shown in figure 1, the engine consists of a chamber 1, fed oxidizing gas from the gas generator 2, which, in turn, is powered by a pump forming part of TNA 3 (the ASAS 4 fuel and oxidizer pump 5). Turbine 6, fed by a gas generator, is located between the gas generator 2 and the camera 1. The fuel pump 4 is also connected to the camera 1 of the engine. Drainage cavity 7 is located between the oxidizer pump 5 and the turbine 6, it docked drain pipe 8, which has the ejector 9, functioning from the high-pressure gas, which is collected by the pipeline 10 from the cavity 11 and the turbine 6.

The pipeline 10 may include a valve 12 and the throttle 13. (2) the Input line of the oxidizer 14 and 15 fuel docked to the inputs of the pumps 4 and 5.

The engine works as follows. The fuel input line 14 enters the pump 4 and then into the chamber 1. The oxidant on the input line 15 enters the oxidizer pump 5 and from there into the gas generator 2. There, in the gas generator 2 receives a portion of fuel from the pump 4. In the gas generator 2 is the combustion process, combustion gases enter the turbine 6, bringing it into rotation. Turbine 6, in turn, causes the rotation of the pumps 4 and 5. The gas after the turbine 6 into the chamber 1, where it digiguide and out through the nozzle, creating thrust. Pressure components fuel pumps 4 and 5 is increased and, accordingly, increases the pressure in the gas generator 2 and the camera 1. The engine goes on the current mode.

To prevent leakage of oxidant from the pump 5 in polot the turbine serves as a drainage cavity 7 with the drain pipe 8. For more effective removal of leaks installed the ejector 9, which after the opening of the valve 12 due to the effect ejection sucks oxidant from the pump 5, removing it from the leak.

Thus, the execution of the drainage cavity between the oxidizer pump 5 and the turbine 6 and the supply and drainage pipeline gas ejector contributes to a more effective removal of leakage of oxidant from the pump, bypassing the turbine cavity 6. The use of gases selected from the turbine 6, allows to refuse from a container of compressed gas and to reduce the weight of the engine.

1. Liquid propellant rocket engine containing a combustion chamber, turbopump Assembly, drain cavity, connected to a drainage pipe, with a drainage cavity is located between the oxidizer pump and turbine, and the drain pipe is provided with a gas ejector, characterized in that the gas ejector is connected by a pipeline with a cavity for the turbine.

2. Liquid propellant rocket engine under item 1, characterized in that the pipe contains a valve and a choke.



 

Same patents:

FIELD: chemistry.

SUBSTANCE: described is a method of increasing energy characteristics of liquid rocket engine, working on fuel components being liquid oxygen and hydrocarbon combustible, with kerosene with a liquid additive, representing a solution of highly molecular polyisobutylene (PIB) with a medium-viscosity molecular weight from 3.1·106 to 4.9·106 in kerosene in an amount, providing concentration of polyisobutylene in kerosene from 0.015% to 0.095% of kerosene weight, being applied as the hydrocarbon combustible; cutting of an impeller of a combustible pump of the engine turbopump unit is performed, with an external diameter of the impeller D2 being determined by formula D2=D1(1BC1+A)0,5, D1 is an external diameter of a working wheel of a standard combustible pump; A is a relative increase of the combustible pump pressure in operation with the PIB; B is a relative decrease of hydroresistance of a tract of a chamber regenerative cooling caused by the PIB impact; C=ΔPcoolΔPp is a ratio of hydroresistance of the tract of the chamber regenerative cooling to pressure of the pump for the supply of the component without the PIB, in order for the value of a weight ratio of components (Km) when the engine operates in nominal and forced modes with an application of kerosene with the liquid additive PIB to remain equal to the value Km when the engine works on pure kerosene.

EFFECT: increase of energy characteristics of the LRE.

2 dwg, 3 tbl

FIELD: engines and pumps.

SUBSTANCE: proposed engine comprises liquid or solid propellant fuel wherein oxidiser and/or combustible includes fixed nitrogen as well as fine or fixed boron. Note here that boron atoms-to-nitrogen atoms ratio makes 1:1 with departure of ±20%. Rocket fuel has fuel excess relative to oxidiser.

EFFECT: higher fuel heat generation.

9 cl

FIELD: engines and pumps.

SUBSTANCE: invention relates to liquid-propellant rockets, accelerating units and can be used at starting the engines when liquid-propellant store residues do not exceed 3% of initial value. Proposed method consists in gasification of liquid residues of unusable fuel reserve in oxidiser and combustible tanks, generation of braking pulse by their combustion in combustion chamber of gas rocker engine and high-rate blowdown of combustion products into space. In compliance with this invention, solid-propellant gas generating compounds (SPGGC) are used for gasification of unusable rocket propellant reserve. SPGGC is fed to oxidiser tank with excess oxygen while SPGGC with limited content of oxygen is fed to fuel tank. Note here that chemical composition and mount of SPGGC at minimum possible residues of unusable rocket propellant components are defined proceeding from preset characteristic speed: ΔVΣpreset=ΔVresSPGGC+ΔVSPGGCSPGGC, where ΔVΣpreset is characteristic speed, ΔVresSPGGC is pulse developed owing to minimum unusable residues of rocket propellant in both tanks required for their oxidation, ΔVSPGGCSPGGC is the pulse developed only by combustion of SPGGC gases in gas rocket engine. Proposed device comprises engine with oxidiser and fuel tanks, tank supercharge system, has rocket engine with power supply system and rocket propellant component residues gasification system. Note here that engine plant is equipped with solid-propellant gas generators with their outlets connected with gas feed devices. Said gas generators are equipped with pyro membranes fitted in fuel tanks with residues of liquid rocket fuel.

EFFECT: higher efficiency.

3 cl, 1 dwg, 1 tbl

FIELD: engines and pumps.

SUBSTANCE: proposed method proceeds from gasification of rocket propellant components (RPC) to be fed into combustion chamber. Note here that after outage of liquid-propellant mid-flight engine RPC gasification system is actuated. Supercharge gas is fed into balloons with extra RPC. Redox gas generators are used to feed heat carriers into tanks with RPC residues depending upon specific fuel kept in tanks.

EFFECT: higher power efficiency and environmental safety, better operating performances.

1 dwg, 1 tbl

FIELD: engines and pumps.

SUBSTANCE: proposed method consists in feeding fuel components in combustion chamber via coaxial aligned-jet nozzle communicating oxidiser chamber with combustion zone and including sleeve covering with clearance the nozzle to communicate fuel zone with combustion zone arranged in mixing head in concentric circles to make central and peripheral zones. Note here that at first stage mode oxygen is fed into combustion chamber via hollow nozzle with developed outlet surface of aligned-jet nozzle while hydrogen is fed via shaped gap between said nozzle and said sleeve and kerosene is fed via channels made in said sleeve. Note also that outlet section of said channels opens into combustion chamber while their outlet communicates with kerosene unit chamber. In operation at second and further stages oxygen is fed into combustion chamber via nozzle with developed outlet surface of aligned-jet nozzle while hydrogen is fed via shaped gap between said nozzle and said sleeve.

EFFECT: better fuel mix mixing.

2 dwg

FIELD: engines and pumps.

SUBSTANCE: proposed engine comprises gas generator, turbo pump unit, feed and adjustment assemblies, mixing head including housing, unit to feed oxidiser, mainly, oxygen, primary fuel feed tank, extra fuel feed tank and fire bottom unit. Coaxial aligned-jet nozzles making central and peripheral zones are arranged in said units in concentric circles. Said coaxial aligned-jet nozzles include hollow tip communicating oxidiser zone with fire zone, sleeve covering said tip with clearance to communicate the primary fuel unit with fire zone. Note here that tips of at least central zone nozzles have radial grooves at their outlets composed of alternating ledges and recesses. Note also that the sleeve between tip ledges has channels with outlets open to fire zone and inlets communicated with extra fuel unit chamber. Note that outer profile of said channels is equidistant with tip profile.

EFFECT: better fuel mix mixing.

5 dwg

FIELD: engines and pumps.

SUBSTANCE: coaxial spray atomiser comprises case with hollow nozzle with outlet provided with radial grooves composed of alternating ledges and recesses to communicate oxidiser chamber with combustion zone and sleeve covering said nozzle to communicate hydrogen chamber with combustion zone. Note here that said radial grooves are arranged to make perimeter of jet central part limited by beam generatrices not exceeding 3s while beam length equals 2.3-2.5 s where s is beam thickness. Note also that beam number amounts to three and that the sleeve between nozzle ledges has channels with outlets open to fire zone and inlets communicated with second fuel unit chamber.

EFFECT: better fuel mix mixing.

2 dwg

FIELD: engines and pumps.

SUBSTANCE: proposed method consists in feeding fuel components in combustion chamber via coaxial aligned-jet nozzle communicating oxidiser chamber with combustion zone and including sleeve covering with clearance the nozzle to communicate fuel zone with combustion zone arranged in mixing head in concentric circles to make central and peripheral zones. At first-stage mode, oxygen is fed in combustion chamber via nozzle with developed outlet surface of coaxial aligned-jet nozzle. Hydrogen is fed via shape clearance between aforesaid nozzle and sleeve while kerosene is fed via channels made in said sleeve. Note here that outlet profile of said channels is equidistant to nozzle profile. Note also that outlet part of said channels opens into combustion chamber while inlet part communicates with kerosene unit chamber. At second-stage mode and that of other stage, oxygen is fed in combustion chamber via hollow nozzle with developed outlet surface of coaxial aligned-jet nozzle. Hydrogen is fed via shaped clearance between nozzle and said sleeve.

EFFECT: better fuel mix mixing.

2 dwg

FIELD: engines and pumps.

SUBSTANCE: proposed engine comprises gas generator, turbo pump unit, feed and adjustment assemblies, mixing head including housing, unit to feed oxidiser, mainly, oxygen, primary fuel feed tank, extra fuel feed tank and fire bottom unit. Coaxial aligned-jet nozzles making central and peripheral zones are arranged in said units in concentric circles. Said coaxial aligned-jet nozzles include hollow body communicating oxidiser zone with fire zone, sleeve covering said body with clearance to communicate the primary fuel unit with fire zone. Note here that bodies of at least central zone nozzles have radial grooves at their outlets composed of alternating ledges and recesses. Note also that the sleeve between body ledges has channels with outlets open to fire zone and inlets communicated with extra fuel unit chamber.

EFFECT: three-component fuel rocket engine, better mix formation.

5 dwg

FIELD: engines and pumps.

SUBSTANCE: proposed method of feeding three-component liquid-propellant engine, mainly, oxygen-kerosene-hydrogen, consists in feeding said components into combustion chamber via aligned jet atomisers. Said atomizers comprise hollow nozzle connecting oxidiser chamber with combustion zone and sleeve covering said nozzle with clearance to communicate fuel chamber with combustion zone, both being arranged in mixing head in concentric circles to make central and peripheral zones. In first-stage mode, oxygen is fed into combustion chamber via said hollow nozzle with developed exit surface of aligned jet atomiser. Hydrogen is fed via shaped clearance between nozzle and sleeve of aforesaid atomiser. Kerosene is fed via sleeve channels. Note here that outlets of said channels open into combustion chamber. Their inlets are communicated with kerosene chamber. In first-stage mode, oxygen is fed into combustion chamber via said hollow nozzle with developed exit surface of aligned jet atomiser. Hydrogen is fed via shaped clearance between nozzle and sleeve of aforesaid atomiser.

EFFECT: higher efficiency of mixing.

2 dwg

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

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.

1 dwg

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

FIELD: aircraft industry; rocketry.

SUBSTANCE: invention relates to design of liquid-propellant rocket engines. Proposed liquid-propellant rocket engine without afterburning of generator gas contains regenerative cooling chamber 1, turbopump set 2 with gas generator 3 to drive turbine 4, two flow rate controls and two nozzles 9, 10 installed in pressure main lines 11, 12 of pumps of turbopump set 2. Sensing elements of spools 5, 6 of controls communicate through pipelines with inputs of nozzles 9, 10 and their minimum sections. According to invention servo-actuate restrictor 14 of control, playing the part of thrust control, installed in feed main line 12 of one of propellant components into gas generator 3. Restricting element of servo-actuated restrictor 14 communicates through pipeline 21 with pressure main line 12 of pump of said component after nozzle 10, and pipeline 22 delivering second component into gas generator 3 is connected with pressure main line 11 of pump of said component after servo-actuated restrictor 13 of control playing the part of propellant components flow rate ratio control.

EFFECT: improved energy-mass ratios of engine, provision of constant propellant components flow rate through engine and thrust irrespective of ratio of components passing through engine.

1 dwg

FIELD: engines and pumps.

SUBSTANCE: invention relates to rocket-propelled vehicles, particularly, to the gas duct of liquid-propellant rocket engines with after-burning. The aforesaid gas duct comprises the outlet manifold of the main turbo-pump unit, a bent pipeline and a swinging assembly. The aforesaid bent pipeline is coupled with the outlet manifold and the said swinging assembly is linked with the engine chamber. Note here that the aforesaid swinging assembly is furnished with a bi-degree universal joint and the joint of the swinging assembly with the engine chamber and bent pipeline represents a flange coupling incorporating a metal T-shape gasket furnished with a load-bearing ring with two flexible springs provided with mountain-like ledges. Note also that the aforesaid one-piece bent pipeline is made from a heat-resistant nickel-alloy, while the bent pipeline flange represents a load-bearing belt with a developed end face surface for the engine frame support to be attached thereto. The aforesaid T-shape taper gasket springs feature the thickness varying over their length, while their length L-to-mean thickness δ ratio makes L/δ ˜8 to 10 and the angle α of the spring taper surface inclination to the flange coupling axis makes 1.5 to 2.5 degrees. The flexible spring OD including the aforesaid mountain-like ledges exceeds the ID of the flange coupling sealing surfaces by 0.1 to 0.2 mm. All parts of the gas duct are made from the EK-61 heat-resistant nickel alloy. The propose invention allows a higher tightness of the fixed joints and pipelines carrying high-temperature high-pressure oxidising medium.

EFFECT: improved performances due to ease of uncoupling gas duct from engine chamber and bent pipeline.

7 cl, 4 dwg

FIELD: engines and pumps.

SUBSTANCE: in method for compensation of differences in physical properties of fuel components based on matching of operation modes of universal liquid-propellant rocket engine supply units, according to invention for generator-free engine with separate turbine pump (TP) during its transfer from hydrogen to liquefied natural gas (LNG) (methane), at first fuel (LNG, methane) flow is increased to required value for provision of reliable cooling of chamber, after cooling prior to fuel supply to turbine of TP its total flow is divided into two parts, one of which is supplied to TP turbine, and the other one is discharged, at that after TP passing, fuel fission process is repeated, at that its one part is sent for combustion in combustion chamber, and the other is discharged or sent for further use. Discharged parts of fuel flow may be used as working fluid, for instance, for steering nozzles, for turbine of engine swinging system, for supercharging of tanks, repeatedly as working fluid of chamber fuel and/or propellant pump. Invention provides for operation of engine both on fuel components "oxygen+hydrogen" and also on fuel "oxygen+liquefied natural gas" (methane).

EFFECT: reduced cost of engine and expanded field of its application.

7 cl, 4 dwg

Rocket engine unit // 2381378

FIELD: engines and pumps.

SUBSTANCE: invention relates to rocketry and can be used in designing rocker carrier first stages with multi-tank propellant compartments with wrap-around arrangement. Engine unit comprises multi-tank propellant compartment and fluid propellant rocket engines, every engine being communicated, via feed lines, with adjoining tanks. One of the engines communicates, via feed lines and booster pump units, with all tanks.

EFFECT: synchronised utilisation of propellant components from like tanks without introducing disturbing torques to rocket.

1 cl, 1 dwg

FIELD: engines and pumps.

SUBSTANCE: invention relates to rocketry, particularly to liquid propellant rocket engines operated on three fuel components, i.e. cryogenic oxidiser, hydrocarbon fuel and liquid hydrogen. Proposed engine comprises at least one combustion chamber with jet nozzle, regenerative cooling system, gas generator, and turbopump unit comprising turbine, oxidiser pump and fuel pumps. It differs from known designs in that said turbopump unit comprises two fuels pumps and two extra fuel pumps designed to operate on first fuel and second fuel. Note here that second fuel pump and additional second fuel pump are arranged below oxidiser pump. Downstream of fuel pumps, first and second fuel valves are arranged connected, via electric line, with synchronisation device. Proposed engine incorporates also control unit connected with aforesaid synchronisation device. Method of operation of above described engine comprises feeding fuel and oxidiser into gas generator and combustion chamber, igniting them and exhausting combustion products via jet nozzle. In compliance with this invention, first fuel utilised, second fuel is fed into gas generator and combustion chamber. Prior to feeding second fuel, fuel pipelines and nozzle regenerative cooling systems are blown down to remove first fuel residues.

EFFECT: improved operating performances of liquid propellant engine in wide range of flight conditions at various altitudes.

5 cl, 3 dwg

FIELD: engines and pumps.

SUBSTANCE: invention relates to rocketry, particularly to liquid propellant rocket engines operated on three fuel components, i.e. cryogenic oxidiser, hydrocarbon fuel and liquid hydrogen. Proposed rocket comprises first- and second-stage rocket units connected in parallel, oxidiser and fuel tanks coupled by power assemblies and equipped with at least one first-stage engine and one second-stage engine. In compliance with this invention, second-stage unit comprises second fuel tank, every second-stage engine incorporates combustion chamber and fuel feed turbopump unit. Proposed engine comprises at least one combustion chamber with jet nozzle, regenerative cooling system, gas generator, and two turbopump units comprising turbine, oxidiser pump and fuel pumps. In compliance with this invention, outlets of all pumps communicate, via gas duct, with gas generator outlet communicates with every combustion chamber. Method of operation of above described engine comprises feeding fuel and oxidiser into gas generator and combustion chamber, igniting them and exhausting combustion products via jet nozzle. In compliance with this invention, first fuel utilised, second fuel is fed into gas generator and combustion chamber. Prior to feeding second fuel, fuel pipelines and nozzle regenerative cooling systems are blown down to remove first fuel residues.

EFFECT: higher thrust-to-weight ratio, improved operating performances.

12 cl, 8 dwg

FIELD: engines and pumps.

SUBSTANCE: invention can be used during development of liquid-propellant engines (LPI) for carrier rockets (CR). Method consists in the fact that acceleration pulse is created owing to combustion of fuel components in ignition device (ID) and supply of its combustion products to chamber nozzle. Ignition device is tripped after the required pulse is obtained by the mixture in CR tank. At that, combustion products are supplied to the combustion chamber nozzle together with their ballasting, e.g. with fuel which is first passed through the cooling path of the chamber. The proposed method is implemented in LPI containing combustion chamber with ID, nozzle, turbo-pump unit, automation and control units, which, according to the invention, is equipped with an additional line with the valve for ballasting of ID combustion products, which connects the outlet of the cooling path of combustion chamber to its mixing head.

EFFECT: simplifying the design and reducing power consumption.

3 cl, 1 dwg

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