CN116696612B - Active heat pipe air cooling system for carrier rocket and design method - Google Patents
Active heat pipe air cooling system for carrier rocket and design method Download PDFInfo
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- 238000001816 cooling Methods 0.000 title claims abstract description 57
- 238000013461 design Methods 0.000 title claims abstract description 38
- 238000000034 method Methods 0.000 title claims abstract description 30
- 238000001704 evaporation Methods 0.000 claims abstract description 37
- 230000008020 evaporation Effects 0.000 claims abstract description 28
- 239000012530 fluid Substances 0.000 claims abstract description 23
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- 238000009833 condensation Methods 0.000 claims abstract description 19
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- 238000010438 heat treatment Methods 0.000 claims abstract description 5
- 238000012546 transfer Methods 0.000 claims description 24
- 239000007921 spray Substances 0.000 claims description 15
- 230000017525 heat dissipation Effects 0.000 claims description 12
- LFQSCWFLJHTTHZ-UHFFFAOYSA-N Ethanol Chemical compound CCO LFQSCWFLJHTTHZ-UHFFFAOYSA-N 0.000 claims description 6
- XLYOFNOQVPJJNP-UHFFFAOYSA-N water Substances O XLYOFNOQVPJJNP-UHFFFAOYSA-N 0.000 claims description 6
- 238000012360 testing method Methods 0.000 claims description 5
- DGAQECJNVWCQMB-PUAWFVPOSA-M Ilexoside XXIX Chemical group C[C@@H]1CC[C@@]2(CC[C@@]3(C(=CC[C@H]4[C@]3(CC[C@@H]5[C@@]4(CC[C@@H](C5(C)C)OS(=O)(=O)[O-])C)C)[C@@H]2[C@]1(C)O)C)C(=O)O[C@H]6[C@@H]([C@H]([C@@H]([C@H](O6)CO)O)O)O.[Na+] DGAQECJNVWCQMB-PUAWFVPOSA-M 0.000 claims description 3
- 229910000799 K alloy Inorganic materials 0.000 claims description 3
- WHXSMMKQMYFTQS-UHFFFAOYSA-N Lithium Chemical compound [Li] WHXSMMKQMYFTQS-UHFFFAOYSA-N 0.000 claims description 3
- ZLMJMSJWJFRBEC-UHFFFAOYSA-N Potassium Chemical compound [K] ZLMJMSJWJFRBEC-UHFFFAOYSA-N 0.000 claims description 3
- 229910052744 lithium Inorganic materials 0.000 claims description 3
- QSHDDOUJBYECFT-UHFFFAOYSA-N mercury Chemical group [Hg] QSHDDOUJBYECFT-UHFFFAOYSA-N 0.000 claims description 3
- 229910052753 mercury Inorganic materials 0.000 claims description 3
- BITYAPCSNKJESK-UHFFFAOYSA-N potassiosodium Chemical compound [Na].[K] BITYAPCSNKJESK-UHFFFAOYSA-N 0.000 claims description 3
- 229910052700 potassium Inorganic materials 0.000 claims description 3
- 239000011591 potassium Substances 0.000 claims description 3
- 239000003507 refrigerant Substances 0.000 claims description 3
- 150000003839 salts Chemical class 0.000 claims description 3
- 229910052708 sodium Inorganic materials 0.000 claims description 3
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Classifications
-
- F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
- F02—COMBUSTION ENGINES; HOT-GAS OR COMBUSTION-PRODUCT ENGINE PLANTS
- F02K—JET-PROPULSION PLANTS
- F02K9/00—Rocket-engine plants, i.e. plants carrying both fuel and oxidant therefor; Control thereof
- F02K9/96—Rocket-engine plants, i.e. plants carrying both fuel and oxidant therefor; Control thereof characterised by specially adapted arrangements for testing or measuring
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- Engineering & Computer Science (AREA)
- Chemical & Material Sciences (AREA)
- Combustion & Propulsion (AREA)
- Mechanical Engineering (AREA)
- General Engineering & Computer Science (AREA)
- Heat-Exchange Devices With Radiators And Conduit Assemblies (AREA)
Abstract
The invention provides an active heat pipe air cooling system for a carrier rocket and a design method, wherein the active heat pipe air cooling system comprises: the device comprises a deflector, a heat pipe and a power system; the evaporation section of the heat pipe is positioned in the fluid director, and the condensation section of the heat pipe is positioned in the power system; the fluid director is used for heating the evaporation section of the heat pipe, so that the liquid system working medium of the evaporation section is changed into a gaseous system working medium, and the gaseous system working medium flows to the condensation section of the heat pipe under the action of pressure difference; the power system is used for cooling the condensing section of the heat pipe, so that the gaseous system working medium of the condensing section is changed into a liquid system working medium, and the liquid system working medium flows back to the evaporating section of the heat pipe under the action of capillary force of the heat pipe. The embodiment of the invention adopts the active heat pipe cold air system, leads out the heat of the fluid director by adopting the heat pipe in a passive way, has the advantages of wide application range, reusability, low cost and quick response, and can meet the increasing emission requirement.
Description
Technical Field
The invention relates to the technical field of rocket engine testing, in particular to an active heat pipe air cooling system for a carrier rocket and a design method.
Background
With the gradual development of space exploration tasks, the launching task of the spacecraft has the characteristics of short period, high launching frequency and low cost requirement, and the high-temperature and high-pressure wake flow generated by engine jet flow has extremely high safety requirement on a launching system in the process from the ignition to the take-off of the carrier rocket.
The existing technical scheme is that a flow director is adopted to guide wake flow, so that wake flow aggregation at a transmitting device is avoided. In order to avoid damage to the launching equipment caused by high-temperature and high-pressure airflow flushing, the fluid director adopts a composite structure of high-temperature-resistant cement and other high-temperature-resistant materials. Although the passive heat protection scheme has the advantages of high reliability and simple structure, along with the increasing frequency of the emission task, the passive heat protection scheme often has the problems of long material manufacturing period and difficult replacement, and is difficult to adapt to the requirements of rapid and repeated emission.
Disclosure of Invention
In view of the above, an objective of the embodiments of the present invention is to provide an active heat pipe air cooling system and a design method for a carrier rocket, so as to solve the technical problems of long material manufacturing period and difficult replacement of the passive heat protection scheme adopted in the prior art.
To achieve the above object, in a first aspect, an embodiment of the present invention provides an active heat pipe air cooling system for a launch vehicle, the active air cooling system comprising: the device comprises a deflector, a heat pipe and a power system; wherein,
The evaporation section of the heat pipe is positioned in the fluid director, and the condensation section of the heat pipe is positioned in the power system;
The fluid director is used for heating the evaporation section of the heat pipe, so that a liquid system working medium of the evaporation section is changed into a gaseous system working medium, and the gaseous system working medium flows to the condensation section of the heat pipe under the action of pressure difference;
The power system is used for cooling the condensing section of the heat pipe, so that the gaseous system working medium of the condensing section is changed into a liquid system working medium, and the liquid system working medium flows back to the evaporating section of the heat pipe under the action of capillary force of the heat pipe.
In some possible embodiments, the flow director is a temperature equalizing plate, and the evaporation section of the heat pipe is welded to the flow director.
In some possible embodiments, the power system is an air-cooled radiator, condenser, or water-cooled tower.
In some possible embodiments, the air-cooled heat sink comprises two fans, a wind deflector, and fins, wherein,
The two fans are oppositely arranged, the fins are arranged between the two fans, and the wind shield is arranged between the two fans and wraps the outer sides of the fins.
In a second aspect, an embodiment of the present invention provides a design method for an active heat pipe air cooling system of a carrier rocket, where the design method includes:
determining wake parameters of the rocket engine, wherein the wake parameters comprise the temperature of an engine nozzle and the heat flow of the engine nozzle;
selecting a system working medium in the heat pipe according to the temperature of the engine spray pipe;
determining the number of the heat pipes according to the lengths of the flow directors, and determining the heat transfer capacity of a single heat pipe according to the number of the heat pipes and the heat flow of the engine spray pipe;
Determining parameters of the heat pipe according to the heat transfer capacity of the single heat pipe, the evaporating section temperature and the condensing section temperature of the heat pipe;
And calculating the mass flow of the power system according to the fact that the heat dissipation capacity of the power system is equal to the input heat of the flow guider, and selecting the model of the power system according to the mass flow by referring to a product manual.
In some possible embodiments, the selecting the system working medium in the heat pipe according to the temperature specifically includes:
When the temperature is higher than 450 ℃, the system working medium is sodium, potassium, lithium or sodium-potassium alloy;
when the temperature is more than 250 ℃ and less than 450 ℃, the system working medium is mercury or molten salt;
When the temperature is less than 250 ℃, the system working medium is water, ethanol or R134a refrigerant.
In some possible embodiments, the determining the heat transfer capability of a single heat pipe according to the number of heat pipes and the heat flow of the engine nozzle specifically includes:
calculating the input heat of the flow guider according to the heat flow and the area of the flow guider;
and determining the heat transfer capacity of a single heat pipe according to the number of the heat pipes and the input heat quantity of the flow guider.
In some possible embodiments, the mass flow rate is calculated by the formula:
Q=q 1＝C1*M1*(T2–T1), wherein Q is the input heat of the fluid director, Q 1 is the heat dissipation capacity of the power system, M1 is the mass flow, T 1 is the temperature of the system working medium at the inlet of the power system, T 2 is the temperature of the system working medium at the outlet of the power system, and C1 is the specific heat capacity of air.
In some possible embodiments, the design method further comprises:
checking whether the model of the power system meets the design requirement, and if so, performing flight test examination; if not, readjusting the model of the power system.
In some possible embodiments, the calculating the input heat of the flow director according to the heat flow and the area of the flow director specifically includes: and the heat flow of the engine spray pipe multiplied by the area of the flow director is the input heat of the flow director.
In a third aspect, an embodiment of the present invention further provides a carrier rocket, where the carrier rocket uses the active heat pipe air cooling system to perform heat protection.
The beneficial technical effects of the technical scheme are as follows:
The embodiment of the invention provides an active heat pipe air cooling system for a carrier rocket and a design method, wherein the active heat pipe air cooling system comprises the following components: the device comprises a deflector, a heat pipe and a power system; the evaporation section of the heat pipe is positioned in the fluid director, and the condensation section of the heat pipe is positioned in the power system; the fluid director is used for heating the evaporation section of the heat pipe, so that the liquid system working medium of the evaporation section is changed into a gaseous system working medium, and the gaseous system working medium flows to the condensation section of the heat pipe under the action of pressure difference; the power system is used for cooling the condensing section of the heat pipe, so that the gaseous system working medium of the condensing section is changed into a liquid system working medium, and the liquid system working medium flows back to the evaporating section of the heat pipe under the action of capillary force of the heat pipe. The embodiment of the invention adopts the active heat pipe cold air system, leads out the heat of the fluid director by adopting the heat pipe in a passive way, has wide application range, can be repeatedly used, has low cost and quick response, and can meet the increasing emission requirement.
Drawings
In order to more clearly illustrate the embodiments of the invention or the technical solutions in the prior art, the drawings that are required in the embodiments or the description of the prior art will be briefly described, it being obvious that the drawings in the following description are only some embodiments of the invention, and that other drawings may be obtained according to these drawings without inventive effort for a person skilled in the art.
FIG. 1 is a schematic diagram of an active heat pipe air cooling system for a launch vehicle according to an embodiment of the present invention;
FIG. 2 is a schematic view of an external structure of an air-cooled radiator according to the present invention;
FIG. 3 is a schematic view of an internal structure of an air-cooled radiator according to the present invention;
FIG. 4 is a flow chart of a design method of an active heat pipe air cooling system for a launch vehicle according to an embodiment of the present invention;
FIG. 5 is a schematic view of a launch vehicle deflector according to an embodiment of the present invention;
FIG. 6 is a schematic diagram of an active heat pipe air cooling system embodying the present invention;
FIG. 7 is a flow chart of a design of an active heat pipe air cooling system for a launch vehicle deflector according to an embodiment of the present invention.
Reference numerals illustrate:
1. a deflector; 2. a heat pipe; 3. a power system; 31. a blower; 32. a wind deflector; 33. a fin;
A. an engine nozzle.
Detailed Description
Features and exemplary embodiments of various aspects of the invention are described in detail below. In the following detailed description, numerous specific details are set forth in order to provide a thorough understanding of the invention. It will be apparent, however, to one skilled in the art that the present invention may be practiced without some of these specific details. The following description of the embodiments is merely intended to provide a better understanding of the invention by showing examples of the invention. In the drawings and the following description, at least some well-known structures and techniques have not been shown in detail in order not to unnecessarily obscure the present invention; also, the dimensions of some of the structures may be exaggerated for clarity. Furthermore, the described features, structures, or characteristics may be combined in any suitable manner in one or more embodiments.
Term interpretation:
thermal protection (thermalprotection): the basic purpose of thermal protection is to ensure the safety of the aircraft and to ensure that the internal payload or instrumentation is within the allowable temperature and pressure ranges.
Heat pipe (heatpipe): when the evaporation end of the heat pipe is heated, the liquid in the capillary core is heated and evaporated, the vapor flows to the condensation end under the action of pressure difference, the condensation end gives off heat to condense into liquid, and the liquid flows back to the evaporation end under the action of capillary force. The circulation is not completed, and the heat collection and transfer are realized.
And (3) a system working medium: in short, the "working substance" is a medium substance by which various power systems such as heat engines or thermal devices can perform the mutual conversion between heat energy and mechanical energy.
Blower (draughtfan): and the gas pressure is increased by means of the input mechanical energy, and the gas is discharged.
The embodiment of the invention provides an active heat pipe air cooling system of a carrier rocket fluid director and a design method thereof, wherein a heat pipe is an inactive high-efficiency heat exchange element, has the advantages of wide operating temperature range, compact structure, stable and reliable work, high safety and the like, and is applied to various fields of aerospace, energy sources, chemical industry and the like. In the field of aerospace, a heat pipe radiating system consisting of a plurality of heat pipes is used for managing the temperature of an aerospace vehicle; the heat pipe is combined into the heat pipe radiator, and the heat pipe radiator has high reliability which still does not influence the performance of equipment under the condition that part of the heat pipes fail;
The heat pipe consists of an evaporation section, a heat insulation section, a capillary core and a condensation section, and the working process of the heat pipe is that the evaporation section absorbs heat of a heat source to heat and evaporate an internal system working medium, and the system working medium is changed from a liquid state to a gaseous state; the working medium of the gaseous system exchanges heat with the cold source in the condensing section, and the gaseous system is changed into liquid state; the gaseous system working medium moves through density gradient generated by temperature difference of the evaporating section and the condensing section system working medium; the liquid system working medium of the condensing section moves to the evaporating section under the action of capillary force of the internal capillary core; and the circulation is carried out in this way, and the heat transport is completed.
Compared with the traditional active heat protection design, the embodiment of the invention adopts the integrated design method, adopts the deflector to set the heat pipe for passive heat conduction, and adopts an air cooling or water cooling mode for cooling; the active cooling system has the advantages of wide application range, reusability, low cost and quick response, and can meet the increasing emission requirements.
Example 1
FIG. 1 is a schematic structural diagram of an active heat pipe air cooling system for a launch vehicle according to an embodiment of the present invention, as shown in FIG. 1, the active air cooling system includes: the device comprises a deflector 1, a heat pipe 2 and a power system 3; the evaporation section of the heat pipe 2 is positioned in the fluid director 1, and the condensation section of the heat pipe 2 is positioned in the power system; the fluid director 1 is used for heating the evaporation section of the heat pipe 2, so that the liquid system working medium of the evaporation section is changed into a gaseous system working medium, and the gaseous system working medium flows to the condensation section of the heat pipe 2 under the action of pressure difference; the power system 3 is used for cooling the condensation section of the heat pipe 2, so that the gaseous system working medium of the condensation section is changed into a liquid system working medium, and the liquid system working medium flows back to the evaporation section of the heat pipe 2 under the action of capillary force of the heat pipe 2. By so cycling, heat transport is completed.
The embodiment of the invention adopts an active heat pipe cold air system, and leads out the heat of the fluid director 1 by adopting the heat pipe 2 in a passive way, so that the invention has the advantages of wide application range, reusability, low cost and quick response, and can meet the increasing emission demands.
In some embodiments, the flow director 1 may be a temperature equalizing plate, and the evaporation section of the heat pipe 2 is welded to the flow director. Specifically, after the deflector 1 is made into a temperature equalizing plate, a plurality of heat pipes 2 can be welded to the surface of the deflector 1 to transfer heat.
In some embodiments, power system 3 may be an air-cooled radiator or condenser or other device for cooling. Fig. 2 is a schematic external structure of an air-cooled radiator according to the present invention, and fig. 3 is a schematic internal structure of an air-cooled radiator according to the present invention, as shown in fig. 2 and 3, when the power system 3 is an air-cooled radiator, the air-cooled radiator includes two fans 31, a wind deflector 32 and fins 33, the fans 31 are used for providing cooling air, the wind deflector 32 performs air duct optimization to prevent air from leaking out, and unnecessary loss is caused; the fin 33 can increase the heat radiating area, improves the heat transfer effect, namely to the cooling effect of heat pipe condensation section, and two fans 31 set up relatively, and fin 33 sets up between two fans 31, and the deep bead 32 is connected with two fan inboard that set up relatively, and wraps up in the outside of fin 33, and the condensation section of heat pipe 2 is located fin 33 inside. According to the embodiment of the invention, the refrigerating capacity of the power system 3 can be flexibly adjusted by adjusting the thickness and the distance of the fins 33 so as to meet the system requirement, namely, the input heat of the flow guider 1 is equal to the heat dissipation capacity of the air-cooled radiator.
The working process of the active heat pipe air cooling system of the carrier rocket deflector provided by the embodiment of the invention is as follows:
During the take-off process of the carrier rocket, the heat flow generated by the engine spray pipe A flushes the deflector 1, and the evaporation section of the heat pipe 2 in the deflector 1 is heated; the system working medium of the evaporation section of the heat pipe 2 absorbs heat to become gas, moves to the condensation section to release heat under the action of pressure difference, and is condensed into liquid; the working medium of the liquid system flows back to the evaporation section under the action of capillary force in the heat pipe 2, and the circulation is performed in this way, so that the heat is transported; in the condensing section of the heat pipe 2, the low-temperature and high-speed air blown by the fan absorbs the heat of the heat pipe 2 and becomes high-temperature air to flow out, so that the cooling of the system is realized.
In the embodiment of the invention, the heat pipe 2 is arranged in the fluid director 1 to conduct heat passively, and the air cooling or water cooling mode is adopted to cool the heat pipe, so that the active heat pipe air cooling system has the advantages of wide application range, reusability, low cost and quick response, and can meet the increasingly-increased emission demands.
Example two
FIG. 4 is a flow chart of a design method of an active heat pipe air cooling system for a launch vehicle according to an embodiment of the present invention, as shown in FIG. 4, the design method comprises the following steps:
Step S11, determining wake parameters of the rocket engine, wherein the wake parameters comprise the temperature of an engine nozzle and the heat flow of the engine nozzle; the heat flow is heat sprayed out of the engine spray pipe in unit time, and the unit is kW.
Step S12, selecting a system working medium in the heat pipe 2 according to the temperature;
Specifically, because different engines correspond to different wake parameters, the embodiment of the invention can flexibly select the system working medium according to the wake parameters of the engines, and when the temperature is more than 450 ℃, the system working medium is sodium, potassium, lithium or sodium-potassium alloy; when the temperature is more than 250 ℃ and less than 450 ℃, the system working medium is mercury or molten salt; when the temperature is less than 250 ℃, the system working medium is water, ethanol or R134a refrigerant.
Step S13, determining the number of heat pipes according to the lengths of the flow directors, and according to the number of the heat pipes and the heat transfer capacity of a single heat pipe 2 of the heat flow of the engine spray pipe;
Step S14, determining heat pipe parameters according to the heat transfer capacity of the single heat pipe2, the evaporating section temperature and the condensing section temperature of the heat pipe 2;
Specifically, parameters of the heat pipe generally include capillary core, wall thickness, outer diameter, length, system working medium, heat transfer capacity parameters and the like of the heat pipe.
In step S15, the mass flow of the power system 3 is calculated according to the heat dissipation capacity of the power system 3 being equal to the heat absorption capacity of the fluid director 1, and the model of the power system 3 is selected according to the mass flow.
When the power system 3 is an air-cooled radiator, the mass flow of the power system 3 can also be called air quantity, a proper fan is selected according to the index of a fan product manual according to the air quantity and the air pressure, and the air pressure is the power provided by the fan, namely the pressure difference between the air outlet and the air inlet is met. Where the air volume is the flow rate of air (total volume) per unit of fixed time, typically including the air supply volume (output) and the fresh air volume (intake), can be used to indicate the capacity of the ventilation device being tested, and is typically calculated in cubic meters per second, cubic feet per second. The larger the air quantity is, the higher the heat dissipation capacity of the tested equipment is, and more heat can be taken away; wind pressure is the pressure or vortex of wind in a direction perpendicular to the direction of the air flow to which the device or object is subjected, in a fixed planar range.
In some embodiments, in step S13, the heat transfer capability of the single heat pipe 2 according to the number of heat pipes and the heat flow of the engine nozzle specifically includes the following steps:
Step S131, calculating the input heat of the deflector 1 according to the heat flow and the area of the deflector 1; specifically, the heat flow is input heat by multiplying the area of the flow director 1.
In step S132, the heat transfer capacity of the individual heat pipes 2 is determined according to the number of heat pipes 2 and the input heat.
Specifically, fig. 5 is a schematic structural diagram of a carrier rocket deflector according to an embodiment of the present invention, as shown in fig. 5, a is a partial perspective view of a junction between a heat pipe 2 and a deflector 1, and in fig. 5, B is a perspective view of the deflector, and a side surface of the deflector 1 is in a "herringbone" shape, so that heat of an engine nozzle can be split. The evaporation section of the heat pipe 2 extends into the deflector 1 from the side surface of the deflector 1, so in this embodiment, the number of heat pipes 2 needs to be designed according to the length of the side surface of the deflector 1, the length of the deflector 1 is 3000mm, in order to meet the heat dissipation requirement, the size requirement and the actual process condition, the diameter of the heat pipe 2 is initially selected to be 50mm, and the number N of the heat pipes 2 can be expected to be installed in the deflector 1 to be 50. At this time, if the input heat Q of the side surface of the deflector 1 is 100kW, the heat transfer capacity of the individual heat pipes 2 should be 2kW or more by dividing the input heat by the number of the heat pipes 2 (i.e., Q/N). In step S14, the design of the wick of the heat pipe 2 may be developed according to the requirement to meet the requirement of heat transfer capability.
Fig. 6 is a schematic diagram of an active heat pipe air cooling system according to the present invention, as shown in fig. 6, in some embodiments, the mass flow rate is calculated according to the formula: q=q 1＝C1*M1*(T2–T1), i.e. the input heat of the deflector 1 is equal to the heat dissipation capacity of the power system, wherein Q is the input heat of the deflector, Q 1 is the heat dissipation capacity of the second power system, M1 is the mass flow in kg/s, T 1 is the temperature of the system working medium at the inlet of the power system, T 2 is the temperature of the system working medium at the outlet of the power system, and C1 is the specific heat capacity of air in kj/kg/K.
In some embodiments, in order to ensure that the active heat pipe air cooling system designed by the design method can meet the heat dissipation requirement in practical application, the design method further includes: checking whether the model of the power system meets the design requirement, and if so, performing flight test examination; if not, readjusting the model of the power system.
For example, according to design indexes such as the input heat of the fluid director 1 and the temperature to be controlled, test and check are performed to check whether the fluid director can meet the design requirements, and if not, the signals of the fan, the thickness and the spacing of the fins and the like can be adjusted.
In order to enable those skilled in the art to better understand the technical solution provided by the embodiments of the present invention, the following describes in detail an active heat pipe air cooling system for a carrier rocket deflector provided by the embodiments of the present invention. FIG. 7 is a design flow diagram of an active heat pipe air cooling system for a launch vehicle deflector according to an embodiment of the present invention, as shown in FIG. 7, the design flow including an evaporation design stage, a condensation design stage, and a quality inspection stage;
Specifically, in the design stage of the evaporation section, simulation calculation or experimental research is required to be performed on rocket engine parameters, wake parameters of an engine are determined, the wake parameters comprise the temperature of an engine spray pipe and the heat flow of the engine spray pipe, the number of heat pipes is determined according to the length of the flow director, the diameter of the flow director and the length of the flow director are required to be smaller than or equal to each other, the length of the flow director can be measured or designed to be a certain length, namely, the number of heat pipes with the length of the flow director being capable of being provided with heat selection is ensured, the heat pipe parameters are selected according to the number of the heat pipes, the temperature of the engine spray pipe and the heat flow of the engine spray pipe, for example, the system working medium is determined according to the temperature of the engine spray pipe, the input heat of the flow director is calculated according to the heat flow of the engine spray pipe and the area of the flow director, the product of the heat flow of the engine spray pipe and the area of the flow director is recorded as the input heat of the flow director, the input heat is divided by the number of the heat pipes and is recorded as the heat transfer capacity of a single heat pipe, and the heat transfer parameters, for example, capillary core, wall thickness, outer diameter, length, heat transfer capacity and the heat are designed according to the heat transfer capacity parameters; for example, the heat transfer capability requirement of the individual heat pipes 2 is 5kW, and the inside of the heat pipes 2 can be carefully designed or the like according to this requirement.
In the design stage of the condensing section, the type of the fan is required to be selected according to the fact that the input heat of the fluid director 1 is equal to the heat dissipation capacity of the power system 3, and the thickness and the spacing of the fins are selected according to the type of the fan.
When the selection of the number of heat pipes 2, the wall thickness of the fins 33 and the spacing is completed, a quality inspection stage is entered to check whether the design meets the system requirements, i.e. whether the collection and transfer of heat on the deflector 1 can be achieved.
According to the embodiment of the invention, through an integral design method, the heat pipe 2 is arranged by the fluid director 1 to conduct heat passively; cooling by adopting an air cooling or water cooling mode; the active heat pipe cooling system designed by the design method has the advantages of wide application range, reusability, low cost and quick response, and can meet the increasing emission requirements.
In a third embodiment, the invention further provides a carrier rocket, and the carrier rocket adopts the active heat pipe air cooling system for heat protection. According to the carrier rocket provided by the embodiment of the invention, the active heat pipe air cooling system is adopted for conducting heat, namely, the active cooling is carried out in an air cooling mode. The active heat pipe air cooling system has the advantages of wide application range, reusability, low cost and rapid response speed, and can meet the increasing emission requirements.
In the description of the embodiments of the present invention, it should be noted that the orientation or positional relationship indicated by "upper, lower, inner and outer", etc. in terms are based on the orientation or positional relationship shown in the drawings, and are merely for convenience of describing the present invention and simplifying the description, rather than indicating or suggesting that the device or element in question must have a specific orientation, be configured and operated in a specific orientation, and thus should not be construed as limiting the present invention. Furthermore, the terms "first, second, or third" are used for descriptive purposes only and are not to be construed as indicating or implying relative importance.
The terms "mounted, connected, and coupled" in embodiments of the invention are to be construed broadly, unless otherwise specifically indicated and defined, for example: can be fixed connection, detachable connection or integral connection; it may also be a mechanical connection, an electrical connection, or a direct connection, or may be indirectly connected through an intermediate medium, or may be a communication between two elements. The specific meaning of the above terms in the present invention will be understood in specific cases by those of ordinary skill in the art.
While the invention has been described with reference to a preferred embodiment, various modifications may be made and equivalents may be substituted for elements thereof without departing from the scope of the invention. In particular, the technical features mentioned in the respective embodiments may be combined in any manner as long as there is no structural conflict. The present invention is not limited to the specific embodiments disclosed herein, but encompasses all technical solutions falling within the scope of the claims.
Claims (10)
1. An active heat pipe air cooling system for a launch vehicle, the active heat pipe air cooling system comprising: the device comprises a deflector, a heat pipe and a power system; wherein,
The evaporation section of the heat pipe is positioned in the fluid director, and the condensation section of the heat pipe is positioned in the power system;
The fluid director is used for heating the evaporation section of the heat pipe, so that a liquid system working medium of the evaporation section is changed into a gaseous system working medium, and the gaseous system working medium flows to the condensation section of the heat pipe under the action of pressure difference;
The power system is used for cooling the condensing section of the heat pipe, so that the gaseous system working medium of the condensing section is changed into a liquid system working medium, and the liquid system working medium flows back to the evaporating section of the heat pipe under the action of capillary force of the heat pipe.
2. The active heat pipe air cooling system of claim 1, wherein the flow director is a temperature equalization plate, and the evaporator section of the heat pipe is welded to the flow director.
3. The active heat pipe air cooling system of claim 1 wherein the power system is a condenser.
4. The active heat pipe air cooling system according to claim 1, wherein the power system is an air cooled radiator comprising two fans, a wind deflector and fins, wherein,
The two fans are oppositely arranged, the fins are arranged between the two fans, and the wind shield is arranged between the two fans and wraps the outer sides of the fins.
5. A design method for an active heat pipe air cooling system of a carrier rocket, the design method comprising:
determining wake parameters of the rocket engine, wherein the wake parameters comprise the temperature of an engine nozzle and the heat flow of the engine nozzle;
selecting a system working medium in the heat pipe according to the temperature of the engine spray pipe;
determining the number of the heat pipes according to the lengths of the flow directors, and determining the heat transfer capacity of a single heat pipe according to the number of the heat pipes and the heat flow of the engine spray pipe;
Determining parameters of the heat pipe according to the heat transfer capacity of the single heat pipe, the evaporating section temperature and the condensing section temperature of the heat pipe;
And calculating the mass flow of the power system according to the fact that the heat dissipation capacity of the power system is equal to the input heat of the flow guider, and selecting the model of the power system according to the mass flow.
6. The design method according to claim 5, wherein the selecting the system working medium in the heat pipe according to the temperature specifically comprises:
When the temperature is higher than 450 ℃, the system working medium is sodium, potassium, lithium or sodium-potassium alloy;
when the temperature is more than 250 ℃ and less than 450 ℃, the system working medium is mercury or molten salt;
When the temperature is less than 250 ℃, the system working medium is water, ethanol or R134a refrigerant.
7. The method of claim 5, wherein determining the heat transfer capability of a single heat pipe based on the number of heat pipes and the heat flow of the engine nozzle comprises:
calculating the input heat of the flow guider according to the heat flow and the area of the flow guider;
and determining the heat transfer capacity of a single heat pipe according to the number of the heat pipes and the input heat quantity of the flow guider.
8. The design method according to claim 5, wherein the mass flow rate is calculated by the formula:
Q=q 1＝C1*M1*(T2–T1), wherein Q is the input heat of the fluid director, Q 1 is the heat dissipation capacity of the power system, M1 is the mass flow, T 1 is the temperature of the system working medium at the inlet of the power system, T 2 is the temperature of the system working medium at the outlet of the power system, and C1 is the specific heat capacity of air.
9. The design method according to claim 5, further comprising:
checking whether the model of the power system meets the design requirement, and if so, performing flight test examination; if not, readjusting the model of the power system.
10. A launch vehicle characterized in that it is thermally protected by the active heat pipe air cooling system of any one of claims 1-4.
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| Publication number | Priority date | Publication date | Assignee | Title |
|---|---|---|---|---|
| CN111895700A (en) * | 2020-07-20 | 2020-11-06 | 中国人民解放军63921部队 | Water supply system for cooling and noise reduction system of rocket launching platform |
| CN215952354U (en) * | 2021-06-09 | 2022-03-04 | 浙江蓝箭航天空间科技有限公司 | Water circulation reusable water spraying noise reduction system for rocket launching |
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| Publication number | Priority date | Publication date | Assignee | Title |
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| US5803159A (en) * | 1997-05-23 | 1998-09-08 | The United States Of America As Represented By The Secretary Of The Army | Integrated fin-heat pipe |
| RU2270792C1 (en) * | 2004-08-05 | 2006-02-27 | Федеральное государственное унитарное предпряитие "Конструкторское бюро общего машиностроения им. В.П. Бармина" | Launch complex for preparation and launching of launch vehicles with spacecraft |
| EP2376763A2 (en) * | 2008-12-08 | 2011-10-19 | Firestar Engineering, LLC | Regeneratively cooled porous media jacket |
| CN114718759B (en) * | 2022-05-16 | 2023-04-07 | 北京航天试验技术研究所 | Cooling device and rocket engine test bench cooling system |
| CN218519862U (en) * | 2022-09-22 | 2023-02-24 | 浙江蓝箭航天空间科技有限公司 | Spacecraft supporting device |
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| Publication number | Priority date | Publication date | Assignee | Title |
|---|---|---|---|---|
| CN111895700A (en) * | 2020-07-20 | 2020-11-06 | 中国人民解放军63921部队 | Water supply system for cooling and noise reduction system of rocket launching platform |
| CN215952354U (en) * | 2021-06-09 | 2022-03-04 | 浙江蓝箭航天空间科技有限公司 | Water circulation reusable water spraying noise reduction system for rocket launching |
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