Method of and device for creating thrust of liquid-propellant rocket engine

FIELD: rocketry.

SUBSTANCE: according to proposed method of creating thrust of liquid-propellant rocket engine with circulation of heat carrier based on taking of propellant components from tanks, increasing their pressure by pumps driven by turbine and introducing them into gas generator and combustion chamber, combustion of components in gas generator and chamber and creating thrust with ejection of combustion products through nozzle, when introducing component of propellant and products of their gasification into combustion chamber, tangential speed component is imparted to them and part of combustion products is replaced by heat carrier and in process of recirculation it is successively expanded at higher pressure of diverting part of nozzle, cooled, condensed in head exchanger-condenser, pressure is raised by pump and it is then delivered to near-critical part of nozzle to repeat the cycle. Liquid-propellant rocket engine with closed circuit of heat carrier contains chamber with mixing head and regenerative cooling duct, turbopump set with oxidizer and propellant pumps whose output main lines are connected with said mixing head of chamber and gas generator, and said closed circuit of heat carrier is formed with successively interconnected circulating pump, unit to introduce heat carrier to near-critical area of nozzle, heat exchanger-condenser, means to supply condensed component to input of circulating pump. According to invention, closed circuit is provided with section of diverting part of nozzle on which ring ribs made of heat-resistant material are secured over circumference.

EFFECT: reduced cost of launching of useful load into orbit.

5 cl, 3 dwg

 

The invention relates to the field of rocketry and can be used in liquid propellant rocket engines (LPRE), including the YARD.

In the space industry, there is the problem of reducing the cost of launch payloads, it is relevant, and work is underway to solve specific problems in providing solutions to this problem, including work on improving specific impulse.

There is a method of creating thrust rocket engine (Almasov VE other Theory of rocket engines. -M.: Mashinostroenie, 1980, s), based on the intake of the fuel components from the fuel tanks, the supply of these components under pressure sequentially in the gasifier and the combustion chamber, the combustion components with the creation of the high temperature products of combustion, emission under axial acceleration in the expanding portion of the nozzle of the engine. In implemented in accordance with the method LRE a closed pneumatic circuit is provided by the staged combustion cycle in the combustion chamber after the selection part of its energy to drive the turbine and accordingly pumps, which allows to some extent to increase a specific characteristic on the schema with the emission of low-temperature gas from the gas generator (open scheme with the loss of specific impulse with this release). This method is most widely used in commercial rocket-wear the spruce and determines the level of technology.

The disadvantage of this method is the inability to increase the specific impulse, i.e. the solution to the problem of reducing the cost of launch payloads at the level of the pressure in the chamber 20,0-25,0 MPa.

There is a method of creating thrust of the rocket engine and the device for its realization (AIAA paper # 2001-3964)in which to improve efficiency used by the air under the scheme of the engine Alchemist-combined engine.

Atmospheric air as an additional mass is injected from the output of a modified turbofan engine (after fans), and liquid hydrogen on Board the carrier. The cooling of high temperature to the ambient temperature is carried out in a heat exchanger air-to-air", where it transfers heat oxygen-depleted air is returned to the engine. The air is then dewatered and zakolerovat to intermediate cryogenic temperature turboexpander. The air zakolerovat to the temperature of liquefaction in consistently located heat exchanger and then fed to the separator for the separation of fractions with dual rotating column. In the separator carry out the separation of air, liquid oxygen (purity more than 90% with the inclusion of inert gases 10% and oxygen-depleted air (98% nitrogen). 90% of the liquid oxygen is about who will escaut through purification and accumulate on Board. Part of the waste oxygen-depleted air szhizhajut and used as working fluid separator and the rest as the refrigerant in the heat exchanger tweaking. At the exit of the heat exchanger fine tuning the main part of the oxygen-depleted air is compressed, cooled and expand on turboexpander for use as a cooler of the main cryogenic heat exchanger. After this oxygen depleted air is heated in the heat exchanger air-to-air re-injected into the air stream of the fan motor in the space surrounding the burner, as a diluent and working environment. At the entrance to the engine supplied liquid hydrogen is compressed to a pressure of about 92 atmospheres compressor of the engine, is gasified in the nitrogen liquefier (air, oxygen-depleted), expands in the turbine to further reduce the temperature, is passed through an additional coil cooling in the nitrogen liquefier and then last is used for cooling the oxygen-depleted air in the heat exchanger operating in reverse Brayton cycle. The cold side of the heat exchanger is part of the cryogenic heat exchanger for air. Hydrogen gas under high pressure is additionally heated waste heat from the fan motor and is used to drive the and turbine TNA engine. TNA is driven feed pump liquid hydrogen and liquid oxygen-depleted air before finally putting it in the burner fan motor.

To implement the method in the layout engine includes: a modified turbofan engine cryogenic heat exchangers, TNA and the separator media: liquid-air interface.

The disadvantage of this known process is an extremely complex process with multiple liquefaction educated partial mass cryogenic component of the working fluid, leading to costly development and low reliability of the process of creating traction in General, provided the implementation of the scheme in principle. This does not significantly decrease the cost of launch space payloads.

There is also known a method (Katorgin B. I., Chvanov VK, Arkhangelsk VI and other Liquid rocket engine (LRE) at cryogenic fuel closed loop drive turbopump Assembly (options), patent RU 2155273 C1, 18.08.1999, CL F 02 K 9/48, 9/64, 9/78, 64 G 1/40 MKI7- prototype). The method is based on a modified closed circuit (option schemes of senger). In particular, in accordance with the method is the extraction of the components of the fuel tanks missiles, increasing their pressure pumping means, the drive of which is turbine running on prod is the settlement gas producing core components the combustion components in the chamber with the exhaust of products of combustion sequentially through the combustion chamber and nozzle to create thrust. In this case, steam turbine used to drive TNA, at the outlet of the turbine is supplied in a closed loop working fluid drive turbine-heat carrier, which is directly vapor that condenses and is used to drive TNA.

This method is implemented in the device (see three-dimensional V.G. and other Analysis of the energy balance of the rocket engine with a closed steam circuit in the supply system. Proceedings KB Energomash 2002 - prototype).

The specified device at cryogenic fuel with a closed coolant circuit includes a camera with a mixing head and a tract of regenerative cooling, turbopump Assembly with pump oxidizer and fuel output line which is connected to the mixing head and the closed coolant circuit, which includes sequentially interconnected circulation pump, a tract of land regenerative cooling chamber, the inlet node of the coolant, providing a supply of condensed component to the input of the pump.

The disadvantage of this method create a thrust GDR and device for implementing the method is the inability to substantially reduce the cost of removing useful load, the arc of the orbit due to the limited increase the specific impulse of the engine thrust of this scheme.

The objective of the invention is to eliminate this drawback and to achieve a significant reduction in the value of payload delivery to orbit at a certain parallel decrease in energidepartementet basic units of the engine implementation increase the specific impulse.

This goal is achieved by a method of forming a thrust liquid rocket engine (LRE) with circulation of coolant based on the fence components, fuel tanks, increasing their pressure superficial means, the drive of which is carried by the turbine, the introduction into the gasifier and the combustion chamber, the combustion components in the gas generator and the camera, creating thrust with the emission of combustion products through the nozzle, according to the invention with the introduction of components and products of gasification in the combustion chamber give a velocity tangential component, replace part of the products of combustion cooled, and in the recycling process it consistently expand with the increased pressure on the extending part, nozzle, cooled, condensed in the heat exchanger-condenser, increase the pressure pump and submit it in the near critical part of the nozzle to repeat the cycle.

Upon cooling, the coolant expands with positive enthalpy on the blades of the working turbine pump drive th the medium of a closed loop.

As the coolant using the water.

Liquid propellant rocket engine with closed circuit fluid containing chamber of the mixing head and the path of regenerative cooling, turbopump Assembly with pump oxidizer and fuel output line which is connected to the specified mixing head chamber and the gas generator, and the specified closed coolant circuit is formed with sequentially connected to each other with a circulating pump, the inlet node of the coolant in the region near critical flow nozzle, heat exchanger-condenser, means for supplying condensed component on the inlet of the circulation pump according to the invention in a closed loop entered the area expanding portion of the nozzle where the lashing around the circumference of the annular ribs of the heat-resistant material.

The annular ribs on the expanding surface of the nozzle at a wall formed with gaps circle plots normally relative to the tangent to the critical section.

The proposed method is implemented in the device, the schema and the elements of which are presented in figures 1, 2, 3.

Figure 1 - graph of a possible increase of the pressure P (physical causes produce thrust and specific impulse at a given flow rate) with predominantly radial (radius R) the expansion of the working fluid with vzaimosvaz is authorized by lowering the braking pressure in the expanding portion of the nozzle and increase the static pressure at the swirling flow and at the races seals. The increase of thrust and specific impulse is achieved within the new sequence flow of the working fluid in a closed loop and the introduction of structural elements for braking of the flow ribs in native mode, the nozzle along the trajectory of the rocket flight mode nedonashivanie. Figure 1 in the coordinates of pressure (y-axis) and presents the following curves.

As - Processed and formalized the results of the experiments of the applicant. The upper curve is the envelope curve of the points of the maximum static pressure of the flowing fluid on the corresponding surface of the expanding portion of the nozzle of the rocket engine with irregular compression under braking conditions pererasseyaniya or nedonashivanie (provided in the latter case, the introduction of the ribs and end on the radius of the overlapping part of the bore). This corresponds to an increase traction and specific thrust (when achievable decreasing flow - ensure the achievement of the objectives of the invention).

B - Moderate increased static pressure of the flowing fluid in the expanding area of the nozzle of the rocket engine at the organization of the spin flow (known experimental data reflected in the periodical literature. See, for example, Sforzini Swirling flow in the nozzle JSR, vol.7#2 1970).

- Initial values of static is th pressure flowing coolant in same area expanding portion of the nozzle, moreover, this pressure creates thus the thrust of the rocket engine with conventional spreader pattern of creating traction within existing theoretical values and practices of rocket propulsion.

Figure 2 is a pneumatic circuit LRE, where:

1 - vortex combustor;

2 - mixing head;

3 - the fuel cavity of the head;

4 - cavity oxidation head;

5 - site supply of the working fluid of a closed loop create thrust;

6 - collector supply components for the refrigeration system;

7 - heat exchanger-condenser;

8 - transfer pump;

9 - turbine;

10 - closed loop coolant;

11 - oxidizer pump;

13 - fuel pump;

14 - nozzle;

15 - fin of a heat resisting material;

16 - line filling of a closed loop;

17 - return valve line fill;

18 - gasifier;

19 - turbine THA;

20 - tract regenerative cooling;

21 - output line of the pump oxidizer;

22 - output line of the fuel pump.

Figure 3 - view of the expanding portion of the nozzle 14 from the side of the gas flow with the representation of the location of edges of a heat resisting material 15 to increase the pressure in the parietal layer of the braking flowing gas and the pre-swirling flow in a clockwise direction from the nozzle. Uslon is not shown, the heat exchanger on the outer perimeter of the nozzle.

The proposed LRE with a closed coolant circuit includes a camera 1 with a mixing head 2 and tract regenerative cooling 20. It also includes turbopump Assembly 11 with the oxidizer pumps 12 and 13 fuel output whose output line 21 and 22 are connected with the cylinder, and the gas generator through the respective cavity oxidant 4 and 3 fuel injector head 2 by means of respective pipelines of the gas generator 18. The turbine 19 is used to convert a portion of thermal energy and drive these pumps 12, 13. The engine is formed of a closed loop 10 working body - fluid to create traction in the combustion chamber 1 and on the expanding surface of the nozzle 14. In the specified circuit 10 are connected in series between a circulating pump 8 driven turbine 9, the input node of the coolant in the region near critical flow nozzle 5, the surface area expanding portion of the nozzle 14, the heat exchanger-condenser 7, which feeds the condensed component on the inlet of the circulation pump 8. For cooling and condensing the coolant is supplied to the hotline of the heat exchanger-condenser 7 of the circuit 10. The inlet hot side connected to the output collector of the nozzle 14, and the inlet of the cold side of the heat exchanger-condenser 7 is connected to pressure the artery pump component, used for cooling the chamber, for supplying part of it on the heat exchanger and the combustion chamber 1. Within a closed loop 10 of the heat carrier the surface of the hot wall of the nozzle 14 (section extending part of the nozzle) is a surface to provide traction. This surface is one side of the conditional extending along the radius of the nozzle formal, non-annular channel - channel radial expansion of the gas coolant. Such conventional channel for radial expansion of the carrier is limited geometrically bottom border media: fluid and extending parallel to the Central part of the nozzle 14 of the combustion products. The expiration of the fluid enables the creation of traction and the heating medium supply parameters at the exit of the nozzle, providing a liquefaction of the specified coolant in the heat exchanger-condenser 7 installed on the output of the specified channel at the wall outlet area of the nozzle. At the specified hot wall of the expanding portion of the nozzle 14 lashing around the circumference of the ribs 15 of heat-resistant material. To achieve the maximum pressure rise ribs are pivoted educated breaks the circle plots, the surface of which is perpendicular to the local velocity vector of the gaseous coolant. To the maximum the scale the radius of the expanding portion of the nozzle 14 is entered in the input cylindrical hole hot side of the heat exchanger-condenser 7. To ensure the repeated passage of a coolant circuit (including a hot side heat exchanger 7 with the condensation of the specified fluid introduced transfer pump 8. The pump 8 is driven turbine 9 running on gas discharged from the chamber. When insufficient cooling and high temperature before condensation to a further reduction in the enthalpy can be achieved by triggering it on the rotor blades of the turbine 9. In the specified closed loop entered the highway filling path 16 with a check valve 17 to prevent leakage of the coolant from the circuit and providing recharge of the coolant when possible leak it out of closed loop at the end surface of the nozzle 14 during operation of the engine.

The operation of the rocket engine when creating a draft in accordance with this method, the proposed device is as follows. Components of fuel from the respective tanks (not shown) serves to increase the pressure at the outlet of the oxidizer pumps 12 and 13 fuel driven turbine 19, the output line of the pumps 21 and 22. In the combustion chamber 1 along with these components come under high pressure products of partial combustion from the gas generator 18 with the organization of the twist in the combustion chamber 1. Spraying and peremeshivaemogo fuel in the chamber 1 at a mixing head 2 is performed by feeding them from the cavities of the fuel and oxidizer 3 and 4, respectively, with additional giving input liquid or gaseous jets tangential component speed at the expense of putting them under a specific, non-perpendicular relative to the plane of fire face, the corner. In the cavity of the combustion chamber 1 carry out the oxidation of fuel components. At the same time to achieve a lower temperature, and guarantees of due to this condensation of the gaseous coolant with minimum dimensions units select other than the stoichiometric ratio of the components in the chamber with the increased total equivalent density components. In the area of critical section of the nozzle due to the geometric effects of the bounding walls on the flow of products of combustion with decreasing radius such wall intensify the vortex flow. When a tangential entry into the head combustion chamber components, gaseous products from the gas generator, and gasified in the cavity chamber of the heat carrier, which replaces part of the combustion products in the area of the critical section, optionally lock the throttle cross section in the critical region, which increases the static pressure at the periphery of the walls of the chamber and reduce the expense. As a result, when the swirling flow at the boundary of additional radial pressure gradient, centrifugal force and the presence of the tangential component of the introduced coolant, castingparemail with products of combustion, extend mainly along the radius with decreasing temperature and pressure on the wall of the nozzle 14 with the increase of the diameter specified is equivalent to the flow area of the nozzle. However due to the impact of the above edges 15 of the nozzle when braking on them fluid to subsonic speed at the races sealing is ensured by a significant increase pressure on adjacent areas adjacent the surface of the expanding portion of the nozzle (see figure 1, curve A). Accompanying the increase of pressure, temperature increase of the coolant directly for the critical section suppresses organization equivalent to the edge wall of the veil, that is entering the coolant in the liquid phase in place of the conditional substitution of the first edge of this veil with the formation of a shock wave and braking and the resulting cooling of the coolant. After expansion in the wall of the nozzle, the coolant is cooled in-line hot side of the heat exchanger-condenser 7 and condense. The heat transfer fluid is fed to the input of the pump 8 is driven in rotation by the turbine 9. The pressure of the coolant pump 8 increases to a value sufficient to introduce it in the near critical region of the camera 1. The circuit re-use of the carrier 10 closes, providing multiple thrust generation, thus providing the final reduction of the comp is required and, ultimately resulting significant reduction in costs.

As described above, the pressure increase of the fluid in the expansion of its expanding portion of the nozzle (shown in figure 1), is equivalent to increasing pressure on parietal races seals with traditional nozzle with axial acceleration mode pererasseyaniya. Through the organization of parietal spikes on the heat-resistant edges, organizing a shock wave on the equivalent barriers, improved near-wall pressure in the nozzle operating mode nedonashivanie specific to the nozzle significant portion of the flight trajectory of the missile (figure 1). The averaged pressure on the top of the experimental curve And significantly exceeds the equivalent pressure when accelerating expansion in the nozzle when the same is equivalent to (flow) the pressure in the chamber (lower curve) and provides increased traction characteristics within each of the repeated processes of expansion. theoretical principles behind the method and device described in the works of the applicant.

Thus, due to the reusable fluid with increasing pressure on the expanding portion of the nozzle and reduce consumption when creating thrust within one of the many cycles of achieving the goal of the invention is a substantial reduction in the cost of payload delivery to orbit.

1. The method which is the thrust of rocket engine with circulation, based on the fence components, fuel tanks, increasing the pressure of the pumping means, the drive of which is carried by the turbine, the introduction into the gasifier and the combustion chamber, the combustion components in the gas generator and the camera, creating thrust with the emission of combustion products through the nozzle, characterized in that the introduction of components and products of gasification in the combustion chamber give a velocity tangential component, replace part of the products of combustion cooled, and in the recycling process it consistently expand with the increased pressure on the expanding portion of the nozzle, cooled, condensed in the heat exchanger-condenser, increase the pressure pump and submit it to the near critical part of the nozzle to repeat the cycle.

2. The method according to claim 1, characterized in that the cooling fluid extends from the actuation of the enthalpy on the blades of the working turbine pump drive fluid of a closed loop.

3. The method according to claim 1, characterized in that the coolant water use.

4. Liquid propellant rocket engine with a closed cooling circuit that includes a camera with a mixing head and a tract of regenerative cooling, turbopump Assembly with pump oxidizer and fuel output line which is connected to the specified mixing head the second chamber and the gas generator, and the specified closed loop coolant is formed with a series-connected between a circulating pump, the inlet node of the coolant in the region near critical flow nozzle, heat exchanger-condenser, means for supplying condensed component on the inlet of the circulation pump, wherein in a closed loop entered the area expanding portion of the nozzle, where the lashing around the circumference of the annular ribs of the heat-resistant material.

5. Rocket engine according to claim 4, characterized in that the annular ribs on the expanding surface of the nozzle at a wall formed with gaps circle plots normally relative to the tangent to the critical section.



 

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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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