Liquid-propellant rocket engine turbopump unit

FIELD: engines and pumps.

SUBSTANCE: this turbopump unit comprises turbine, oxidiser pump and fuel pump with impellers. In compliance with this invention, this turbine represents a birotatory design with two impellers. The latter have no nozzle diaphragms and run in opposite directions, each being communicated with oxidiser pump impeller and fuel pump impeller. Turbopump unit can incorporated an extra fuel pump, impellers of extra fuel pump and fuel pump being fitted on one shaft.

EFFECT: decreased centrifugal loads at turbine rotor.

2 cl, 2 dwg

 

The invention relates to rocket technology, particularly to a liquid-propellant rocket engines rocket engines, operating at cryogenic oxidizer and on hydrocarbon fuel.

Known liquid-propellant rocket engine for RF patent for the invention №2095607 designed for use in space boosters, speed boosters, and as the main engine of the spacecraft, includes a combustion chamber regenerative cooling channel, turbopump Assembly - TNA. TNA contains the pumps components - fuel and oxidant with the turbine on the same shaft, which introduced the capacitor. The output capacitor on the refrigerant line connected to the inlet into the combustion chamber and the entrance to the path of the regenerative cooling of the combustion chamber. The output of the capacitor to the line of fluid connected to the inlet to pump one of the components. The output of the pump of the same component in communication with the inlet of the condenser in the refrigerant line. The second input capacitor is in communication with the outlet of the turbine. The pump outlet of another component in communication with the entrance into the combustion chamber.

The disadvantage of TNA engine is the deterioration of the cavitation properties of the pump when the bypass condensate. Such a property of the pump inevitably leads to a reduction of the flow of one of the components of fuel through TNA, the fall of the thrust of the rocket several times and breakdown of the programme which we flight or to disaster.

The known method of operation of rocket engines and liquid-propellant rocket engine for RF patent for the invention №2187684. The method of operation of a liquid-propellant rocket engine is the component feeding fuel into the combustion chamber of the engine, the gasification of one of the components in the path of the cooling of the combustion chamber, supplying it on the turbopump turbine unit with subsequent discharge to the nozzle head of the combustion chamber. Part of the expense of one of the components of the fuel is directed into the combustion chamber, and the remaining part is gasified and is directed to the turbine of the turbopump assemblies. Exhaust turbines on the gaseous component is mixed with a liquid component flowing into the engine at a pressure higher than the vapour pressure of the resulting mixture. Liquid propellant rocket engine comprises a combustion chamber with a tract of regenerative cooling, pumps supply fuel components and turbine. Pumps and turbines arranged in two TNA: primary and booster. The engine contains a set of sequentially before the pump is one of the components of the main fuel turbopump Assembly booster pump turbopump Assembly and the mixer. The output of the main pump turbopump Assembly is connected with the nozzle head of the combustion chamber and the path of the regenerative cooling of the combustion chamber. The tract of regenerate the aqueous cooling in turn associated with turbines primary and booster turbopump units whose outputs are connected to the mixer.

The disadvantage of this scheme is that the heat removed in cooling the combustion chamber, may be insufficient to drive the turbopump Assembly of the engine of great power.

Known LRE by RF patent for the invention №2190114, IPC 7 F02K 9/48, publ. 27.09.2002, This rocket engine includes a combustion chamber with a tract of regenerative cooling, turbopump Assembly TNA pumps oxidant and fuel output line which is connected to the cylinder combustion chamber, the main turbine and the drive circuit of the main turbine. In the circuit drive the main turbine are connected in series between a fuel pump and a tract of regenerative cooling of the combustion chamber, which is connected to the input of the main turbine. The output from the turbine TNA is connected to the input of the second stage pump fuel.

This engine has a significant drawback. Bypass heated in the path of the regenerative cooling of the combustion chamber fuel input to the second stage pump fuel will lead to cavitation and to the effects mentioned above. Most LRE use such components in the fuel, the oxidant flow is almost always more fuel consumption. Therefore, for high-power W Is D, having great traction and a lot of pressure in the combustion chamber, this scheme is unacceptable, because the fuel consumption will not be enough to cool the combustion chamber and drive the main turbine.

Furthermore, it is not designed launch system rocket engine, ignition system components and fuel cut-off system rocket engine and clean the rest of the fuel in the path of the regenerative cooling of the combustion chamber.

Known liquid propellant rocket engine and how to run it on a Russian patent for invention №2232915, publ. 10.09.2003, (prototype), which contains the combustion chamber, turbopump Assembly, generator, system startup, means for igniting the fuel components and fuel line. The output of the oxidizer pump is connected to the input of the generator. The output of the first stage fuel pump connected to the channels of regenerative cooling chamber and from the mixing head. The output of the second stage of the fuel pump (fuel pump) connected to the flow regulator with electric drive. The other input of the regulator is connected with a starting tank with regular fuel. The output of the regulator is connected to the gas generator. The output of the gas generator connected to the turbine inlet turbopump unit, the output of which is connected to a mixing head. The flow regulator is supplied with hydraulic pre-stage, which is across the cavitating jet, hydrocele connected with starting a tank with regular fuel. Hydrocele connected with the second stage of the fuel pump. The choke installed on the output of the first stage fuel pump, done in conjunction with a controlled valve preliminary stages.

The disadvantage of this scheme is a fire or explosion TNA and missiles at launch or in flight due to low reliability of the seal between the turbine and the oxidizer pump, between the pump oxidizer and fuel, as well as between the fuel pump and additional fuel pump from a large differential pressure: 300...400 kgf/cm2for modern rocket engines. For example, when used as components of rocket fuel of hydrogen and oxygen the most insignificant leakage of these components lead to the formation of "explosive mixture" and almost always to the explosion of the rocket.

Objectives of the invention: reduction of size and weight TNA and reduction of the centrifugal loads on the turbine rotor.

The solution of the stated problem is achieved in the turbopump Assembly of liquid-propellant rocket engine, containing the turbine and pumps oxidant and fuel impellers, the fact that turbine made birotational and contains two working wheels made without nozzle apparatus for rotation in opposite directions, each of which are connected respectively with the impeller of the oxidizer pump and pump the fuel. Turbopump Assembly of liquid-propellant rocket engine may contain additional fuel pump, while impellers additional fuel pump and fuel pump installed on the same shaft.

The invention is illustrated in Fig.1 and 2, where:

- Fig.1 is a diagram of the first variant TNA,

- Fig.2 is a diagram of a second variant of TNA with an additional fuel pump.

Turbopump Assembly of liquid-propellant rocket engine TNA (Fig.1) contains the turbine 1 turbine impellers 2 and 3, without nozzle apparatus, the housing 4 and the inlet pipe 5, the oxidizer pump 6 impeller 7, the fuel pump 8 with the impeller 9, two shafts 10 and 11 installed on the supports 12, 13 and 14. On the shaft 11 mounted impellers 3 and 7, and the shaft 10 to the impellers 2 and 9 (details of the rotor).

Possible second option TNA (Fig.2), which further comprises an additional fuel pump 15 with the impeller 16 additional fuel pump 15, with the impeller 16 mounted on the shaft 10 together with the impeller 9 of the fuel pump 8.

The result is a real opportunity to design a TNA and first turbine smaller size and weight.

The application of the invention allowed:

1. To reduce the dimensions and weight of the TPU.

2. To ensure modularity, TNA.

3. To design the CE nodes TNA: the turbine and the pump, on the optimal parameters, including speed and coordinate rotation frequency due to the use of twin-shaft schema.

4. To increase the reliability of TNA by reducing the centrifugal force on the details of the turbine rotors.

1. Turbopump Assembly of liquid-propellant rocket engine, containing the turbine and pumps oxidant and fuel with the impellers, wherein the turbine is made birotational and contains two working wheels made without nozzle apparatus for rotation in opposite directions, each of which are connected respectively with the impeller pump oxidizer and fuel pump.

2. Turbopump Assembly of liquid-propellant rocket engine under item 1, characterized in that it contains an additional fuel pump, while impellers additional fuel pump and fuel pump installed on the same shaft.



 

Same patents:

FIELD: engines and pumps.

SUBSTANCE: proposed turbopump unit comprises oxidiser pump, turbine running on gas fuel, turbine bearing, system of seals separating said oxidiser pump from said turbine. Gas drainage between the system of seals and turbine, seal being fitted on turbine side Turbine bearing is arranged in chamber between said seal and turbine cavity.

EFFECT: decreased losses of separating gas in drainage, better dynamics of rotor.

4 cl, 4 dwg

FIELD: engines and pumps.

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1 dwg

FIELD: rocket engineering; production of the booster turbo-pump aggregates with an axial pumps used in them.

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EFFECT: the invention ensures the increased service life of the fixed bearing (5) and the service life of the whole aggregate.

3 cl, 1 dwg

FIELD: rocket engineering; production of the devices for the liquid propellant rocket engines.

SUBSTANCE: the invention is pertaining to the field of rocket engineering and may be used in the liquid propellant rocket engines (LPRE). The device for separation of the pump and the turbine of the booster turbo-pump aggregate of the LPRE consists of the pump (2), the turbine (3), the separating cavity (1) located between the pump (2) and the turbine (3) and the external intake tract (4). The separating cavity (1) is limited from the side of the pump (2) by the shaft gasket (5), which diameter is made smaller than the diameter of the shaft (10) in the area of the seat of the bearing of the turbine (11), and from the side of the turbine (3) - by the unloading disk (6) aligned with the turbine impeller (7). On the turbine impeller (7) there is the gasket of the unloading disk (8). The axial impeller of the pump (9) and the turbine impeller (7) are fixed on the shaft (10). From the direction of the turbine (3) the shaft (10) rests on the turbine bearing (11), which is brought out beyond the bounds of the separating cavity (1) and is installed from the direction of the pump (2). The cavity of the turbine bearing (12) which is adjoining the shaft gasket (5) is connected by the delivery channels (13) with the pump outlet (14). The offered device ensures the minimum losses of the power used for separation of the pump and the turbine, and also the effective refrigeration of the bearings by the liquid monophase hydrogen.

EFFECT: the invention ensures the minimum losses of the power used for separation of the pump and the turbine, the effective refrigeration of the bearings by the liquid monophase hydrogen.

2 cl, 2 dwg

FIELD: rocket engineering; production of the liquid propellant rocket engines.

SUBSTANCE: the invention is pertaining to the field of rocket engineering, in particular, to production of the liquid propellant rocket engines powered by the cryogenic oxidant and the hydrocarbon propellant. The liquid propellant rocket engine contains the combustion chamber with the tract of the regenerative cooling, the turbo-pumping aggregate with the turbine having the inlet and outlet trunks, and the pumps of the oxidant and the propellant, for which the outlet of the propellant p[ump is connected through the propellant valve to the combustion chamber, and the outlet of the oxidant pump through the oxidant valve is connected to the gas generator. At that the turbo-pump aggregate contains the additional propellant pump, which inlet is connected to the outlet of the propellant pump, and the outlet is connected to the gas generator through the high pressure pipeline, in which there is the high-pressure valve and the consumption regulator. In the trunk of the turbine there is the thrust regulator, to which the on-board trunk and the starting trunk with the return valve and the connector are connected. The method of the liquid propellant rocket engine starting provides for the spinning-up of the turbo-pump aggregate and opening of valves of the oxidant, the propellant, the propellant in the high-pressure trunk, run-up of the turbine conduct a compressed air from a land bulb, and the turbine spinning-up is exercised by the compressed air from the ground pressure vessel and the turbine drive at operation is exercised from the on-board vessel. The invention ensures simplification of the pneumatic-hydraulic circuit, the increased reliability, the increase of the power and the specific characteristics of the liquid propellant rocket engine, the decreased mass of the engines, the improved engine starting and cutoff and provision of the engine cleansing from the leavings of the propellant after the engine cutoff.

EFFECT: the invention ensures simplification of the liquid propellant rocket engine pneumatic-hydraulic circuit, the increased its reliability, power and specific characteristics, the decreased mass of the engine, the improved the engine starting, cutoff and cleansing from the leavings of the propellant after its cutoff.

4 cl, 1 dwg

FIELD: rocketry.

SUBSTANCE: invention relates to liquid-propellant rocket engines operating on cryogenic oxidizer and on hydrocarbon fuel. Proposed turbopump set of rocket engine contains the following parts of rotor of turbopump set mounted on shaft: oxidizer pump impeller, fuel pump impeller and turbine wheel arranged in housing of turbopump set and additional fuel pump with shaft and impeller of additional fuel pump. Design peculiarity of turbopump set is that magnetic clutch is installed between rotor of turbopump set and rotor of additional fuel pump. Driving disk of magnetic clutch is installed on shaft of turbopump set, and driven disk is mounted on shaft of additional fuel pump. Partition made on nonmagnetic material is found between driving and driven disks of magnetic clutch. Said partition is aligned with housing of additional fuel pump. Partition, driving and driven disks can be made spherical and/or provided with ribbing.

EFFECT: increased reliability of turbopump set.

5 cl, 3 dwg

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