A cycle precooling system for a rocket engine turbopump and a rocket

Abstract

The application provides a circulating precooling system for a rocket engine turbine pump, wherein high-pressure gas is introduced into the turbine through a starter to drive the turbine to rotate, so as to drive the first pump and the second pump to rotate, and pressure differences are generated in the interiors of the first pump, the fuel tank and the related conveying channels, and in the interiors of the second pump, the combustion-supporting agent tank and the related conveying channels. Under the action of the pressure differences, low-temperature fuel enters the first pump from the fuel tank and returns to the fuel tank after precooling is completed, and low-temperature combustion-supporting agent enters the second pump from the combustion-supporting agent tank to be precooled and returns to the combustion-supporting agent tank after precooling is completed. The application utilizes the original device inside the rocket to precool the turbine pump, and no additional precooling circulating pump or motor or other equipment is arranged, so that the turbine pump is fully precooled, the complexity and manufacturing cost of the system are reduced, the rocket is lightened, and the overall performance of the rocket is improved.

Description

TECHNICAL FIELD

[0001] The present application relates to the field of circulating refrigeration of aerospace equipment, in particular to a circulating pre-cooling system for a rocket engine turbine pump and a rocket. BACKGROUND

[0002] Pump pressure rocket engine is a jet engine which is not using outside air and is provided with propellant by the aircraft. The propellant is converted into kinetic energy of working medium in the rocket engine to form high-speed jet to discharge and generate thrust. The turbine pump is the heart of the rocket engine, which is composed of a turbine and a pump. The turbine and the pump are generally directly connected, and the turbine drives the pump to rotate to pressurize the liquid propellant.

[0003] The pump pressure rocket engine needs to be pre-cooled before starting, that is, the engine (mainly the turbine pump) and its delivery system are cooled by using propellant or other low-temperature medium to reach the required temperature for engine starting, so as to ensure the safe operation of the turbine pump. When not pre-cooled or insufficiently pre-cooled, the propellant is strongly vaporized in the delivery pipeline and the engine flow passage due to heating, which causes cavitation, flywheel rotation of the engine turbine pump, and causes fluctuations of pressure and flow, stall operation, oxygen-rich combustion, etc., resulting in the extension of the engine starting time and the thrust climbing time, and even complete failure of starting in severe cases. Therefore, it is very important to pre-heat the engine and its delivery system before ignition.

[0004] The pre-cooling methods of the turbine pump mainly include discharge pre-cooling and circulating pre-cooling, and the circulating pre-cooling can be divided into natural circulating pre-cooling and forced circulating pre-cooling. The forced circulating pre-cooling has the best effect, so currently a pre-cooling circulating pump is generally set to perform forced circulating pre-cooling. The pre-cooling circulating pump is a separate component specially set for pre-cooling, which is only used before engine starting, and needs to be additionally provided with a motor or a turbine as power, thereby increasing the complexity of the system. SUMMARY

[0005] Therefore, the technical problem to be solved by the present application is to overcome the defect that the rocket engine turbine pump needs to be provided with additional equipment for circulating pre-cooling, thereby increasing the complexity of the system, so as to provide a circulating pre-cooling system for a rocket engine turbine pump and a rocket.

[0006] To this end, the technical solution adopted by the present application is,

[0007] The present application provides a circulating pre-cooling system for a rocket engine turbine pump, which comprises a first pump, a second pump and a turbine arranged in the interior of the rocket engine, the turbine being used to drive the first pump and the second pump to operate, and further comprising,

[0008] A starter is provided with a gas outlet, the gas outlet is in communication with the gas inlet end of the turbine, so that the high-pressure gas in the starter drives the turbine to rotate.

[0009] a fuel tank, a first outlet and a first inlet are arranged on the fuel tank, the first outlet is communicated with the first inlet end of the first pump, the first outlet end of the first pump is communicated with the first inlet;

[0010] a combustion-supporting agent tank, a second outlet and a second inlet are arranged on the combustion-supporting agent tank, the second outlet is communicated with the second inlet end of the second pump, the second outlet end of the second pump is communicated with the second inlet.

[0011] Further, the first pump comprises a first shell, a first central rotating shaft arranged in the first shell and a plurality of first blades arranged around the first central rotating shaft, the first inlet end is arranged on the first shell in the axial direction of the first central rotating shaft, and the first outlet end is arranged on the first shell close to the edge of the first blades.

[0012] Further, the second pump comprises a second shell, a second central rotating shaft arranged in the second shell and a plurality of second blades arranged around the second central rotating shaft, the second inlet end is arranged on the second shell in the axial direction of the second central rotating shaft, and the second outlet end is arranged on the second shell close to the edge of the second blades.

[0013] Further, the turbine comprises a third shell, a third central rotating shaft arranged in the third shell and a plurality of third blades arranged around the third central rotating shaft, and the gas inlet end is arranged on the third shell close to the edge of the third blades.

[0014] Further, the turbine further comprises a gas outlet end, the gas outlet end is arranged on the third shell in the axial direction of the third central rotating shaft, so that the high-pressure gas is discharged from the turbine.

[0015] Further, the first central rotating shaft, the second central rotating shaft and the third central rotating shaft are the same central rotating shaft.

[0016] The first shell, the second shell and the third shell are not communicated with each other.

[0017] Further, a first temperature detection unit is arranged in the first shell to measure the temperature of the first pump, and a second temperature detection unit is arranged in the second shell to measure the temperature of the second pump.

[0018] Further, along the axial direction of the second central rotation axis, at least interval arrangement of first rotation unit and second rotation unit composed of several second blades are arranged on the second central rotation axis, the second liquid inlet end is arranged close to the second rotation unit, and the second liquid outlet end is arranged close to the blade edge of the first rotation unit.

[0019] Further, the first valve is arranged between the turbine and the starter to control the high-pressure gas from the starter into the turbine.

[0020] The starter is further provided with a gas charging port to charge the gas into the starter through the gas charging port.

[0021] The application provides a rocket adopting the circulating pre-cooling system.

[0022] The technical scheme of the application has the following advantages:

[0023] 1. The application provides a circulating pre-cooling system for a rocket engine turbine pump, the starter is provided with a gas outlet, the gas outlet is communicated with the gas inlet end of the turbine to drive the turbine to rotate by the high-pressure gas in the starter; the fuel tank is provided with a first outlet and a first inlet, the first outlet is communicated with the first liquid inlet end of the first pump, and the first liquid outlet end of the first pump is communicated with the first inlet; the combustion-supporting agent tank is provided with a second outlet and a second inlet, the second outlet is communicated with the second liquid inlet end of the second pump, and the second liquid outlet end of the second pump is communicated with the second inlet. The high-pressure gas in the starter is introduced into the turbine to drive the turbine to rotate, thereby driving the first pump and the second pump to rotate, so that the pressure difference is generated in the first pump, the fuel tank and the related conveying channels, and the pressure difference is generated in the second pump, the combustion-supporting agent tank and the related conveying channels. Under the action of the pressure difference, the low-temperature fuel flows out from the first outlet of the fuel tank, enters the first pump through the first liquid inlet end of the first pump to be pre-cooled, then flows out from the first liquid outlet end of the first pump and returns to the fuel tank through the first inlet of the fuel tank, so that the first pump is pre-cooled to the temperature required for starting the engine; the low-temperature combustion-supporting agent flows out from the second outlet of the combustion-supporting agent tank, enters the second pump through the second liquid inlet end of the second pump to be pre-cooled, then flows out from the second liquid outlet end of the second pump and returns to the combustion-supporting agent tank through the second inlet of the combustion-supporting agent tank, so that the second pump is pre-cooled to the temperature required for starting the engine. The starter, the fuel tank and the combustion-supporting agent tank are all original devices arranged in the rocket. The application pre-cools the turbine pump by using the original devices in the rocket, without additional pre-cooling circulating pump or motor and other equipment, so that the turbine pump is fully pre-cooled, the complexity and manufacturing cost of the system are reduced, the weight of the rocket is reduced, and the overall performance of the rocket is improved.

[0024] 2. The application provides a circulating precooling system for a rocket engine turbine pump, low-temperature fuel precooling a first pump and returning to a fuel tank, heat in the first pump being transferred to the fuel; similarly, low-temperature oxidizer precooling a second pump and returning to an oxidizer tank, heat in the second pump being transferred to the oxidizer. When the rocket engine is working, fuel and oxidizer flow out of the fuel tank and the oxidizer tank respectively into a thrust chamber and burn violently to become high-temperature and high-pressure gas, which is then accelerated into supersonic gas flow and sprayed out, producing thrust acting on the rocket engine to propel the rocket forward. The application converts heat in the first pump and the second pump into thrust of the engine, thereby realizing full utilization of heat. BRIEF DESCRIPTION OF DRAWINGS

[0025] In order to more clearly illustrate the technical solutions in the embodiments of the present application or the prior art, the following will briefly introduce the drawings needed to be used in the specific embodiments or prior art description. Obviously, the drawings described below are some embodiments of the present application, and those skilled in the art can obtain other drawings according to these drawings without any creative effort.

[0026] Figure 1 is a structure diagram of the circulating precooling system of the turbine pump in the embodiment 1 of the present application;

[0027] Figure 2 is a structure diagram of the turbine pump in the embodiment 1 of the present application;

[0028] Reference signs:

[0029] 1 - first pump; 1 - 1 - first liquid inlet; 1 - 2 - first liquid outlet; 2 - second pump; 2 - 1 - second liquid inlet; 2 - 2 - second liquid outlet; 3 - turbine; 3 - 1 - air inlet; 3 - 2 - air outlet; 4 - starter; 4 - 1 - air outlet; 4 - 2 - air inlet; 5 - fuel tank; 5 - 1 - first outlet; 5 - 2 - first inlet; 6 - oxidizer tank; 6 - 1 - second outlet; 6 - 2 - second inlet; 7 - first valve. DETAILED DESCRIPTION

[0030] The technical solutions of the present application will be described clearly and completely below in combination with the drawings. Obviously, the described embodiments are some of the embodiments of the present application, but not all the embodiments. Based on the embodiments in the present application, all other embodiments obtained by those skilled in the art without any creative effort belong to the protection scope of the present application.

[0031] In the description of the present application, it should be noted that the terms "center", "upper", "lower", "left", "right", "vertical", "horizontal", "inner", "outer" and the like indicate the orientation or positional relationship shown in the drawings, and are only for the convenience of describing the present application and simplifying the description, and do not indicate or imply that the device or element referred to must have a particular orientation, be constructed and operated in a particular orientation, and therefore cannot be understood as a limitation on the present application. In addition, the terms "first", "second", "third" are only for descriptive purposes and cannot be understood as indicating or implying relative importance.

[0032] In the description of the present application, it should be noted that unless otherwise explicitly specified and limited, the terms "mounting", "connecting", "connecting" should be understood broadly, for example, it can be fixedly connected, or it can be detachably connected, or integrally connected; it can be mechanically connected, or it can be electrically connected; it can be directly connected, or it can be indirectly connected through an intermediate medium; it can be the communication inside two elements. For those skilled in the art, the specific meaning of the above terms in the present application can be understood according to the specific circumstances.

[0033] In addition, the technical features involved in the different embodiments of the application described below can be combined with each other as long as there is no conflict.

[0034] Embodiment 1

[0035] The present embodiment provides a circulating pre-cooling system for a rocket engine turbine pump, as shown in Figure 1 Fig. 2, comprising a first pump 1, a second pump 2 and a turbine 3 arranged inside the rocket engine, the turbine 3 being used to drive the first pump 1 and the second pump 2 to operate, further comprising,

[0036] A starter 4 is provided with an air outlet 4-1, which is in communication with an air inlet end 3-1 of the turbine 3, so that the high-pressure gas in the starter 4 drives the turbine 3 to rotate; further, the starter 4 is a high-pressure gas cylinder starter 4, and the high-pressure gas is nitrogen;

[0037] A fuel tank 5 is provided with a first outlet 5-1 and a first inlet 5-2, the first outlet 5-1 being in communication with a first liquid inlet end 1-1 of the first pump 1, and a first liquid outlet end 1-2 of the first pump 1 being in communication with the first inlet 5-2; further, the fuel is liquid fuel, which can be one of liquid hydrogen, hydrazine, methyl hydrazine, unsymmetrical dimethyl hydrazine, kerosene, alcohol, liquefied natural gas; preferably liquefied natural gas.

[0038] A combustion-supporting agent storage tank 6 is provided with a second outlet 6-1 and a second inlet 6-2, the second outlet 6-1 is communicated with the second liquid inlet end 2-1 of the second pump 2, and the second liquid outlet end 2-2 of the second pump 2 is communicated with the second inlet 6-2; further, the combustion-supporting agent is liquid combustion-supporting agent, which can be one of liquid oxygen, dinitrogen tetroxide, hydrogen peroxide and nitric acid; preferably, the liquid oxygen.

[0039] The above-mentioned circulation pre-cooling system passes high-pressure gas into the turbine 3 through the starter 4 to drive the turbine 3 to rotate, thereby driving the first pump 1 and the second pump 2 to rotate, so that pressure difference is generated in the first pump 1, the fuel storage tank 5 and the related delivery channel, and pressure difference is generated in the second pump 2, the combustion-supporting agent storage tank 6 and the related delivery channel. Under the action of the pressure difference, the low-temperature fuel flows out from the first outlet 5-1 of the fuel storage tank 5 and enters the first pump 1 through the first liquid inlet end 1-1 of the first pump 1 to be pre-cooled, and then flows out from the first liquid outlet end 1-2 of the first pump 1 and returns to the fuel storage tank 5 through the first inlet 5-2 of the fuel storage tank 5, so that the first pump 1 is pre-cooled to the temperature required for starting the engine; the low-temperature combustion-supporting agent flows out from the second outlet 6-1 of the combustion-supporting agent storage tank 6 and enters the second pump 2 through the second liquid inlet end 2-1 of the second pump 2 to be pre-cooled, and then flows out from the second liquid outlet end 2-2 of the second pump 2 and returns to the combustion-supporting agent storage tank 6 through the second inlet 6-2 of the combustion-supporting agent storage tank 6, so that the second pump 2 is pre-cooled to the temperature required for starting the engine. The starter 4, the fuel storage tank 5 and the combustion-supporting agent storage tank 6 are all original devices provided inside the rocket. The present application pre-cools the turbine pump by using the original devices inside the rocket, without additionally setting a pre-cooling circulation pump or a motor and other equipment, so that the turbine pump is fully pre-cooled, the complexity and manufacturing cost of the system are reduced, and the rocket is lightened, thereby improving the overall performance of the rocket. When the rocket engine works, the fuel and the combustion-supporting agent flow out from the fuel storage tank 5 and the combustion-supporting agent storage tank 6 into the thrust chamber and are fiercely combusted to be converted into high-temperature and high-pressure gas, and then are accelerated into supersonic gas flow and sprayed out, thereby generating thrust acting on the rocket engine to push the rocket forward. The present application converts the heat in the first pump 1 and the second pump 2 into the thrust of the engine, thereby fully utilizing the heat.

[0040] Further, the starter 4 is further provided with a gas charging port 4-2, so that gas is charged into the starter 4 through the gas charging port 4-2.

[0041] As a specific embodiment, the first pump 1 comprises a first shell, a first central rotating shaft arranged in the first shell and a plurality of first blades arranged around the first central rotating shaft, the first liquid inlet end 1-1 is arranged on the first shell in the axial direction of the first central rotating shaft, and the first liquid outlet end 1-2 is arranged on the first shell close to the edge of the first blade; the second pump 2 comprises a second shell, a second central rotating shaft arranged in the second shell and a plurality of second blades arranged around the second central rotating shaft, the second liquid inlet end 2-1 is arranged on the second shell in the axial direction of the second central rotating shaft, and the second liquid outlet end 2-2 is arranged on the second shell close to the edge of the second blade; the turbine 3 comprises a third shell, a third central rotating shaft arranged in the third shell and a plurality of third blades arranged around the third central rotating shaft, and the gas inlet end 3-1 is arranged on the third shell close to the edge of the third blade; further, the turbine 3 further comprises a gas outlet end 3-2, and the gas outlet end 3-2 is arranged on the third shell in the axial direction of the third central rotating shaft, so that the high-pressure gas is discharged from the turbine 3.

[0042] Further, the first central rotating shaft, the second central rotating shaft and the third central rotating shaft are the same central rotating shaft, so that the turbine 3 drives the first pump 1 and the second pump 2 to operate;

[0043] The first shell, the second shell and the third shell are not communicated with each other; the first shell and the second shell are not communicated to prevent the fuel and the combustion-supporting agent from meeting to cause explosion, so as to ensure the safe and reliable operation of the turbine pump; specifically, the first shell and the second shell constitute a dynamic sealing system through a floating ring.

[0044] Further, the first shell is provided with a first temperature detection unit to measure the temperature of the first pump 1, and the second shell is provided with a second temperature detection unit to measure the temperature of the second pump 2.

[0045] As an optional embodiment, the structures of the first pump 1 and the second pump 2 can be the same or different, which is selected according to the actual situation. Specifically, as shown in Figure 2 the structure of the second pump 2 is that a first rotating unit composed of a plurality of second blades and a second rotating unit composed of a plurality of second blades are arranged on the second central rotating shaft at least in the axial direction of the second central rotating shaft, the second liquid inlet end 2-1 is arranged close to the second rotating unit, and the second liquid outlet end 2-2 is arranged close to the edge of the blade in the first rotating unit; the structure of the first pump 1 is that a third rotating unit composed of a plurality of first blades is arranged on the first central rotating shaft in the axial direction of the first central rotating shaft, the first liquid inlet end 1-1 is arranged close to the third rotating unit and in the axial direction of the first central rotating shaft, and the first liquid outlet end 1-2 is arranged close to the edge of the blade in the third rotating unit.

[0046] Further, a first valve 7 is arranged between the turbine 3 and the starter 4 to control the high-pressure gas from the starter 4 into the turbine 3; a second valve is arranged between the first liquid outlet 1-2 of the first pump 1 and the first inlet 5-2 of the fuel tank 5, and a third valve is arranged between the second liquid outlet 2-2 of the second pump 2 and the second inlet 6-2 of the oxidizer tank 6; before the first valve 7 is opened, the second valve and the third valve are opened to make the low-temperature fuel and the low-temperature oxidizer in the precooling process return to the fuel tank 5 and the oxidizer tank 6 smoothly; after the precooling is completed, the first valve 7, the second valve and the third valve are closed to make the fuel and the oxidizer enter the combustion chamber after the rocket engine is started, so as to generate the thrust acting on the engine to drive the rocket to move forward; further, the first valve 7, the second valve and the third valve are solenoid valves.

[0047] The embodiment also provides a rocket adopting the circulating precooling system.

[0048] Embodiment 2

[0049] The embodiment provides a rocket adopting the circulating precooling system in the above embodiment 1, in particular, the rocket adopts liquid propellant.

[0050] In addition, the working principle of the above circulating precooling system is as follows:

[0051] 1) Inflation: the first valve is opened, the high-pressure gas flows out from the gas outlet of the starter and enters the turbine from the gas inlet of the turbine;

[0052] 2) Rotation: the high-pressure gas blows the third blade of the turbine to rotate, and drives the first blade of the first pump and the second blade of the second pump to rotate, and then is discharged from the gas outlet of the turbine, so as to generate a pressure difference in the first pump, the fuel tank and the related conveying channels, and a pressure difference in the second pump, the oxidizer tank and the related conveying channels; in particular, the pressure of the first outlet of the fuel tank is higher than that of the first liquid inlet of the first pump, the pressure of the first liquid outlet of the first pump is higher than that of the first inlet of the fuel tank, the pressure of the second outlet of the oxidizer tank is higher than that of the second liquid inlet of the second pump, and the pressure of the second liquid outlet of the second pump is higher than that of the second inlet of the oxidizer tank.

[0053] 3) Precooling: under the action of the pressure difference, the low-temperature fuel flows out from the first outlet of the fuel tank and enters the first pump through the first liquid inlet of the first pump to be pre-cooled, and then flows out from the first liquid outlet of the first pump and returns to the fuel tank through the first inlet of the fuel tank; the low-temperature oxidizer flows out from the second outlet of the oxidizer tank and enters the second pump through the second liquid inlet of the second pump to be pre-cooled, and then flows out from the second liquid outlet of the second pump and returns to the oxidizer tank through the second inlet of the oxidizer tank.

[0054] 4) Closing: The first temperature detecting unit measures the temperature of the first pump, and the second temperature detecting unit measures the temperature of the second pump. When the first pump and the second pump are pre-cooled to the temperature required for starting the engine, the first valve is closed, and the pre-cooling is completed.

[0055] Obviously, the above embodiments are merely exemplary and not limiting. Based on the above description, other different forms of changes or variations can be made by those skilled in the art. All embodiments are not required to be exhaustive. The obvious changes or variations derived therefrom are still within the protection scope of the present application.

Claims

1. A circulating precooling system for a rocket engine turbine pump comprising a first pump, a second pump and a turbine disposed within the rocket engine, the turbine for driving the first and second pumps in operation, characterized by, Further comprising, a starter, which is provided with an air outlet, the air outlet being communicated with an air inlet end of the turbine, so that high-pressure gas in the starter drives the turbine to rotate, and a first valve is arranged between the turbine and the starter to control the high-pressure gas from the starter into the turbine; a fuel tank, which is provided with a first outlet and a first inlet, the first outlet being communicated with a first liquid inlet end of the first pump, and a first liquid outlet end of the first pump being communicated with the first inlet; an oxidant tank, which is provided with a second outlet and a second inlet, the second outlet being communicated with a second liquid inlet end of the second pump, and a second liquid outlet end of the second pump being communicated with the second inlet; wherein the starter supplies high-pressure gas to the turbine to drive the turbine to rotate, and the turbine drives the first pump and the second pump to rotate, so that the fuel in the fuel tank is returned to the fuel tank through the first pump, and the oxidant in the oxidant tank is returned to the oxidant tank through the second pump.

2. The cycle precooling system of claim 1, wherein The first pump comprises a first shell, a first central rotating shaft arranged in the first shell, and a plurality of first blades arranged on the first central rotating shaft, the first liquid inlet end is arranged on the first shell in the axial direction of the first central rotating shaft, and the first liquid outlet end is arranged on the first shell close to the edge of the first blades.

3. The cycle pre-cooling system of claim 2, wherein, The second pump comprises a second shell, a second central rotating shaft arranged in the second shell, and a plurality of second blades arranged on the second central rotating shaft, the second liquid inlet end is arranged on the second shell in the axial direction of the second central rotating shaft, and the second liquid outlet end is arranged on the second shell close to the edge of the second blades.

4. The cycle pre-cooling system of claim 3, wherein, The turbine comprises a third shell, a third central rotating shaft arranged in the third shell, and a plurality of third blades arranged on the third central rotating shaft, and the air inlet end is arranged on the third shell close to the edge of the third blades.

5. The recirculating pre-cooling system of claim 4, wherein, The turbine further comprises an air outlet end arranged on the third shell in the axial direction of the third central rotating shaft, so that high-pressure gas is discharged from the turbine.

6. The recirculating pre-cooling system of claim 4, wherein, The first central rotating shaft, the second central rotating shaft, and the third central rotating shaft are the same central rotating shaft. The first shell, the second shell, and the third shell are not communicated with each other.

7. The recirculating pre-cooling system of claim 2, wherein, The first shell is provided with a first temperature detection unit to measure the temperature of the first pump.

8. The recirculating pre-cooling system of claim 3, wherein, The second shell is provided with a second temperature detection unit to measure the temperature of the second pump.

9. The recirculating pre-cooling system of claim 3, wherein, In the axial direction of the second central rotating shaft, at least a first rotating unit composed of a plurality of second blades and a second rotating unit composed of a plurality of second blades are arranged on the second central rotating shaft at intervals, the second liquid inlet end is arranged close to the second rotating unit, and the second liquid outlet end is arranged close to the edge of the blades in the first rotating unit.

10. The cycle pre-cooling system of claim 1 or 2, wherein The starter is further provided with an air inlet to fill gas into the starter through the air inlet.

11. A rocket, characterized in that The circulating pre-cooling system of any one of claims 1-10 is included.

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