CN112983683A - Device and method for rocket engine high-altitude simulation test - Google Patents
Device and method for rocket engine high-altitude simulation test Download PDFInfo
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- CN112983683A CN112983683A CN202110392640.1A CN202110392640A CN112983683A CN 112983683 A CN112983683 A CN 112983683A CN 202110392640 A CN202110392640 A CN 202110392640A CN 112983683 A CN112983683 A CN 112983683A
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- water spray
- rocket engine
- cooling
- cooler
- spray condenser
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- 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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- G—PHYSICS
- G01—MEASURING; TESTING
- G01M—TESTING STATIC OR DYNAMIC BALANCE OF MACHINES OR STRUCTURES; TESTING OF STRUCTURES OR APPARATUS, NOT OTHERWISE PROVIDED FOR
- G01M15/00—Testing of engines
- G01M15/14—Testing gas-turbine engines or jet-propulsion engines
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- F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
- F05—INDEXING SCHEMES RELATING TO ENGINES OR PUMPS IN VARIOUS SUBCLASSES OF CLASSES F01-F04
- F05D—INDEXING SCHEME FOR ASPECTS RELATING TO NON-POSITIVE-DISPLACEMENT MACHINES OR ENGINES, GAS-TURBINES OR JET-PROPULSION PLANTS
- F05D2220/00—Application
- F05D2220/80—Application in supersonic vehicles excluding hypersonic vehicles or ram, scram or rocket propulsion
Abstract
The invention provides a device and a method for a rocket engine high-altitude simulation test, relates to the field of rocket engine ground simulation high-altitude tests, and can improve flame cooling efficiency in the high-altitude simulation test. The device comprises a rapid cooler, a water spray condenser, a ground test bed and a vacuum pumping system, wherein the rapid cooler is communicated with a spray pipe of the rocket engine and comprises an ice cooling module, and the ice cooling module is made of ice; the ice cooling module is provided with a cooling channel, the water spray condenser is communicated with the rapid cooler, airflow at 100 +/-20 ℃ enters the water spray condenser to be cooled, the ground test bed is provided with a rocket engine and is communicated with the rapid cooler and the water spray condenser; the vacuumizing system is communicated with the ground test bed, the rapid cooler and the water spray condenser. The method of the present disclosure is based on the above-described apparatus. The method uses the alkaline ice medium of the rapid cooler to rapidly cool the flame to about 100 ℃, and shortens the flame cooling time in the high-altitude simulation test.
Description
Technical Field
The disclosure relates to the field of ground simulation high-altitude tests of rocket engines, in particular to a device and a method for high-altitude simulation tests of rocket engines.
Background
The high-altitude simulation test of the rocket engine is a ground test which is necessary to be carried out by the rocket engine, and the high-altitude simulation test is to create an environment similar to a high-altitude condition in ground test equipment, so that the rocket engine works in the environment and carries out various tests such as performance, reliability, working life and the like.
The thrust of the rocket engine is continuously increased along with the rise of the altitude until the maximum thrust is reached in a vacuum environment, the environmental pressure of the rocket engine in the vacuum environment is very low, and the high altitude characteristic of the rocket engine can be really simulated only by establishing a low-pressure vacuum environment at the corresponding altitude during a ground high altitude simulation test.
In the prior art, when a rocket engine high-altitude simulation test is performed, in order to protect relevant test equipment, a water injection cooling mode is often adopted to cool the flame of the rocket engine. However, the efficiency of flame water injection cooling is extremely low.
In order to shorten the flame cooling time in the high-temperature simulation high-altitude simulation test, a device and a method for the high-altitude simulation test of the rocket engine are urgently needed to realize the rapid cooling of the flame.
Disclosure of Invention
The embodiment of the invention provides a device and a method for a rocket engine high-altitude simulation test, which are used for rapidly cooling the flame, which is discharged from a rocket engine spray pipe, of about 3000 ℃ to about 100 ℃ by using an alkaline ice medium of a rapid cooler, so that the flame cooling time in the high-altitude simulation test is shortened.
In order to achieve the above purpose, the embodiment of the invention adopts the following technical scheme:
in one aspect, an apparatus for high altitude simulation test of rocket engine is provided, which includes:
a rapid cooler in communication with a nozzle of the rocket engine, the rapid cooler comprising an ice cooling module of a material of ice; the ice cooling module is provided with a cooling channel, flame generated by the rocket engine is sprayed out from the spray pipe and flows through the cooling channel to contact ice around the cooling channel in real time, and the flame flows through the cooling channel to be cooled into airflow at 100 +/-20 ℃;
the water spray condenser is communicated with the rapid cooler, 100 +/-20 ℃ of airflow enters the water spray condenser for cooling, and the airflow cooled by the water spray condenser is discharged to the outer side of the device;
the ground test bed is provided with a rocket engine and is communicated with the rapid cooler and the water spray condenser;
the vacuumizing system is communicated with the ground test bed, the rapid cooler and the water spray condenser; the ground test bed, the rapid cooler and the water spray condenser are all in corresponding high vacuum states.
In some embodiments, further comprising:
and the pressurizing module is installed between the rapid condenser and the water spray condenser, and the air flow flows through the pressurizing module to increase the pressure and then enters the water spray condenser for cooling.
In some embodiments, the evacuation system comprises: at least one group of vacuumizing modules, wherein each group of vacuumizing modules comprises first-stage to Nth-stage vacuum pumps, and N is more than or equal to 2;
the first-stage to Nth-stage vacuum pumps are connected in series in sequence;
different groups of vacuum-pumping modules are arranged in parallel;
all evacuation modules all with quick cooler ground test bench with the water spray condenser intercommunication, every evacuation module all can be right quick cooler ground test bench with the water spray condenser evacuation.
In some embodiments, further comprising:
the drying module is used for drying the gas flowing out of the water spray condenser and comprises a drying agent, and the drying module is installed between the water spray condenser and the vacuumizing system; the vacuum-pumping system works so that the drying module is also in a corresponding high-altitude vacuum state.
In some embodiments, the number of drying modules is the same as the number of evacuation modules of the evacuation system;
and one drying module is arranged between the water spray condenser and each vacuumizing module.
In some embodiments, further comprising: the M sub-cooling systems are uniformly arranged around the rapid cooler and are circularly distributed around the central line of the rapid cooler;
each sub-cooling system is communicated with the rapid cooler, and a pressure limiting valve is arranged at the communication position of each sub-cooling system and the rapid cooler;
and when the flame pressure in the rapid cooler exceeds the threshold pressure of the pressure limiting valve, the flame exceeding the threshold pressure enters the sub-cooling system for rapid cooling.
In some embodiments, the sub-cooling system comprises: a sub-cooler and an evacuation subsystem;
the rapid cooler, the sub-cooler and the vacuumizing subsystem are communicated with each other, and the vacuumizing subsystem is used for vacuumizing the sub-cooler, so that the sub-cooler is in a corresponding high-altitude vacuum state.
In another aspect, a method for high altitude simulation test of rocket engine is provided, including:
starting a vacuumizing system, vacuumizing the ground test bed, the rapid cooler and the water spray condenser to a corresponding high-altitude vacuum state, continuously pumping flames of the ground test bed, the rapid cooler and the water spray condenser by the vacuumizing system, and continuously maintaining the vacuum state by the vacuumizing system;
starting the rocket engine, and spraying flame at about 3000 ℃ from a spray pipe of the rocket engine;
the rapid cooler cools the flame into airflow of 100 +/-20 ℃, and the rapid cooler performs first harmless treatment on the flame while cooling the flame;
the water spray condenser continuously cools the air flow, and the water spray condenser cools the air flow and simultaneously carries out secondary harmless treatment on the air flow;
and discharging the airflow subjected to the second harmless treatment from the vacuum-pumping system.
In some embodiments, the starting the evacuation system to evacuate the ground test stand, the rapid cooler, and the water spray condenser to a vacuum state of a corresponding altitude comprises:
starting a plurality of groups of vacuumizing modules of a vacuumizing system, and vacuumizing the ground test bed, the quick cooler and the water spray condenser by the plurality of groups of vacuumizing modules simultaneously;
the multistage vacuum pumps of each group of vacuumizing modules are connected in series, and simultaneously vacuumize the ground test bed, the rapid cooler and the water spray condenser;
the multistage vacuum pump simultaneously extracts airflow out of the rocket engine high-altitude simulation test device.
In some embodiments, the starting rocket engine, the rocket engine nozzle emitting a flame at about 3000 ℃, comprises:
starting the rocket engine, and enabling the rocket engine to swing on the ground test bed;
the jet pipe of the rocket engine sprays flame at about 3000 ℃, and the flame swings along with the swing of the rocket engine;
and in the region exceeding the threshold pressure in the rapid cooler, opening a pressure limiting valve corresponding to the region, and allowing the flame exceeding the threshold pressure to enter a sub-cooling system corresponding to the region for cooling and discharging.
In the present disclosure, at least the following technical effects or advantages are provided:
1. according to the embodiment of the invention, the alkaline ice medium of the rapid cooler is used for rapidly cooling the flame at about 3000 ℃ discharged by the rocket engine nozzle to about 100 ℃, so that the problem of low water injection cooling efficiency of the flame is effectively solved, and the beneficial effect of shortening the flame cooling time in the high-altitude simulation test is realized.
2. The rapid cooler can reduce the flame temperature and greatly reduce the flame flow speed, and because the specific impulse of a rocket engine is 2900 m/s-3000 m/s and the specific impulse of the flame (airflow) flowing out of the rapid cooler in the horizontal direction is 30 m/s-40 m/s, the flame speed is reduced by more than 50 times after the rapid cooler is cooled, the vacuumizing capacity of a vacuumizing system is greatly reduced, and the vacuum pump power of the vacuumizing system is reduced.
Drawings
In order to more clearly illustrate the embodiments of the present invention or the technical solutions in the prior art, the drawings needed to be used in the description of the embodiments of the present invention or the prior art will be briefly introduced below, and it is obvious that the drawings in the following description are some embodiments of the present invention, and for those skilled in the art, other drawings can be obtained according to these drawings without creative efforts.
FIG. 1 is a first schematic diagram of an apparatus for high altitude simulation testing of a rocket engine;
FIG. 2 is a second schematic diagram of an apparatus for high altitude simulation testing of a rocket motor;
FIG. 3 is a schematic view of the flow direction of the rapid cooler of the apparatus;
FIG. 4 is a schematic diagram of an example of an application of the apparatus for rocket motor high altitude simulation test;
FIG. 5 is a schematic diagram of a partial flame admission sub-cooling system during oscillation of the rocket engine in the apparatus;
FIG. 6 is a flow chart of a method for high altitude simulation testing of a rocket motor;
reference numerals: 1-a ground test bed; 2-a rapid cooler; 21-a first cooling stage; 211 — a first cooling channel; 22-a second cooling section; 221-a second cooling channel; 3-vacuum pumping system; 31-a vacuum-pumping module; 311-primary vacuum pump; 312-a secondary vacuum pump; 313-a tertiary vacuum pump; 4-a rocket engine; 5-a water spray condenser; 6-a pressurization module; 7-vacuum tank; 8-a drying module; 9-a sub-cooling system; 91-sub-cooler; 92-an evacuation subsystem; 93-a pressure limiting valve; a-open state; b-closed state.
Detailed Description
The present disclosure is described in detail with reference to the embodiments shown in the drawings, but it should be understood that these embodiments are not intended to limit the present disclosure, and those skilled in the art should understand that the functional, methodological, or structural equivalents of these embodiments or substitutions may be included in the scope of the present disclosure.
In the description of the embodiments of the present disclosure, it is to be understood that the terms "central," "longitudinal," "lateral," "upper," "lower," "front," "rear," "left," "right," "vertical," "horizontal," "top," "bottom," "inner," "outer," and the like are used in the orientations and positional relationships indicated in the drawings, which are merely for convenience in describing the invention and to simplify the description, and are not intended to indicate or imply that the referenced devices or elements must have a particular orientation, be constructed and operated in a particular orientation, and are therefore not to be construed as limiting the invention.
Furthermore, the terms "first," "second," "third," and the like are used for descriptive purposes only and are not to be construed as indicating or implying relative importance or implicit to a number of indicated technical features. Thus, a feature defined as "first," "second," etc. may explicitly or implicitly include one or more of that feature. In the description of the invention, the meaning of "a plurality" is two or more unless otherwise specified.
The terms "mounted," "connected," and "coupled" are to be construed broadly and may, for example, be fixedly coupled, detachably coupled, or integrally coupled; can be mechanically or electrically connected; they may be connected directly or indirectly through intervening media, or they may be interconnected between two elements. The specific meaning of the above terms in the creation of the present invention can be understood by those of ordinary skill in the art through specific situations.
The embodiment of the present disclosure provides a device for rocket engine high altitude simulation test, please refer to fig. 1, fig. 1 is a schematic diagram of the device for rocket engine high altitude simulation test, the device includes a rapid cooler 2, a water spray condenser 5, a ground test bed 1 and an evacuation system 3, wherein:
the rapid cooler 2 is communicated with a spray pipe of the rocket engine 4, the rapid cooler 2 comprises an ice cooling module, and the material of the ice cooling module is ice; the ice cooling module is provided with a cooling channel, flame generated by the rocket engine 4 is sprayed out from the spray pipe and flows through the cooling channel to contact ice around the cooling channel in real time, and the flame flows through the cooling channel to be cooled into airflow at 100 +/-20 ℃;
the water spray condenser 5 is communicated with the rapid cooler 2, 100 +/-20 ℃ of airflow enters the water spray condenser 5 for cooling, and the airflow cooled by the water spray condenser 5 is discharged to the outer side of the device;
the system comprises a ground test bed 1, wherein a rocket engine 4 is installed on the ground test bed 1, and the ground test bed 1 is communicated with a rapid cooler 2 and a water spray condenser 5;
the vacuumizing system 3 is communicated with the ground test bed 1, the rapid cooler 2 and the water spray condenser 5; the ground test bed 1, the rapid cooler 2 and the water spray condenser 5 are all in a corresponding high vacuum state.
The rapid cooler 2 includes an ice-cooling module, the material of which is ice; the ice cooling module is provided with a cooling channel, the flame generated by the rocket engine 4 flows through the cooling channel and contacts the ice around the cooling channel in real time, and the flame flows through the cooling channel to be cooled into airflow at 100 +/-20 ℃.
The cooling channels of the embodiments of the invention communicate with the nozzle of the rocket motor 4. It is to be noted that the nozzle is an important part of the rocket motor 4, which means a device for accelerating the air flow by changing the geometry of the inner wall of the tube section. In the rocket engine 4, the flow of the fuel gas is controlled by the size of the throat area of the spray pipe, so that the fuel gas in the combustion chamber keeps a preset pressure, and the normal combustion of the fuel charge is ensured; the combustion products of the propellant are expanded and accelerated through the spray pipe, and the heat energy of the combustion products is fully converted into the kinetic energy of the fuel gas, so that the engine obtains propelling power-thrust.
In this embodiment, since the rocket motor 4 is operated, the generated flame mixture with high temperature, high speed and high mass flow rate passes through the ice channel inside the rapid cooler 2, and under the action of high temperature difference and high speed gas, the gas rapidly exchanges heat with the inner surface of the ice channel, and ice "sublimation" and "liquefaction-gasification" occur simultaneously. On one hand, ice absorbs a large amount of heat energy through heat energy exchange, changes from a solid state to a liquid state and a gaseous state, increases the temperature and the speed, and moves backwards together with tail gas until the tail gas is discharged from a motive outlet; on the other hand, the tail gas is subjected to energy exchange between internal energy and kinetic energy, the temperature and the kinetic energy of the tail gas are greatly reduced, and the tail gas is finally discharged to the vacuumizing module in a low-temperature form. In order to eliminate harmful gas HCl in the flame of the rocket motor 4, alkaline materials are added into ice in the rapid cooler 2 during manufacturing, so that the flame is subjected to harmless treatment through neutralization reaction in the energy conversion process.
The ice of this embodiment is frozen with alkaline water, which includes water and an alkaline material, which may be NaOH or NaHCO3And the like, and the specific type of the alkaline substance in the embodiment of the present invention may not be limited.
Preferably, the rapid cooler 2 is communicated with the water spray condenser 5, and the airflow with the temperature of 100 +/-20 ℃ enters the water spray condenser 5 to be cooled for the second time and then is discharged; the rapid cooler 2 and the water spray condenser 5 are arranged between the ground test bed 1 and the vacuumizing system 3, the rapid cooler 2 is arranged close to the ground test bed 1, and the water spray condenser 5 is arranged close to the vacuumizing system 3; the vacuumizing system 3 is also communicated with the water spray condenser 5, and the water spray condenser 5 is also in a corresponding high-altitude vacuum state when the vacuumizing system 3 works.
Referring to fig. 2, fig. 2 is a schematic diagram of a device for rocket engine high altitude simulation test. The device for the rocket engine high-altitude simulation test of the embodiment of the invention comprises a ground test bed 1, a quick cooler 2, a vacuumizing system 3 and a water spray condenser 5, and also comprises the following components: and the pressurizing module 6 is arranged between the rapid condenser and the water spray condenser 5, and the air flow passes through the pressurizing module 6 to increase the pressure and then enters the water spray condenser 5 for cooling. The boost module 6 is used to increase the pressure of the airflow.
Generally, the specific impulse of the rocket engine 4 is 2900 m/s-3000 m/s, the specific impulse of the airflow flowing out of the rapid cooler 2 in the horizontal direction is 30 m/s-40 m/s, the speed is reduced by more than 50 times after the airflow is cooled by the rapid cooler 2, and the flame flow speed can be rapidly reduced by the rapid cooler 2.
In most embodiments, referring to fig. 3, fig. 3 is a schematic flow direction diagram of a rapid cooler 2 of an apparatus for rocket motor high altitude simulation test, wherein the ice cooling module comprises at least one first cooling section 21 and at least one second cooling section 22; the first cooling section 21 and the second cooling section 22 are connected in sequence along the flame spraying direction; the cooling passage includes: one first cooling channel 211 provided at the first cooling stage 21, at least one second cooling channel 221 provided at the second cooling stage 22, each second cooling channel 221 communicating with the first cooling channel 211; the first cooling section 21 is communicated with the rocket engine 4, the flame ejected from the rocket engine 4 is guided into the first cooling channel 211 for cooling treatment, and the flame subjected to cooling treatment from the first cooling channel 211 enters the second cooling channel 221 for continuous cooling treatment; the first cooling section 21 and the second cooling section 22 are both communicated with the vacuum-pumping system 3, and the vacuum-pumping system 3 works to enable the first cooling section 21 and the second cooling section 22 to be in a corresponding high-vacuum state. The number of the second cooling channels 221 shown in fig. 3 is 5, one is located at the center of the second cooling section 22, and the remaining 4 are located around the second cooling channel 221 at the center. The 5 second cooling channels 221 are uniformly distributed on the end face of the second cooling section 22.
In practical applications, it is preferable that the ice-cooling module is frozen using the case as a mold. The shell is internally provided with a big ice block, a channel is arranged on the big ice block, the temperature of flame is reduced to about 100 ℃ after energy is exchanged between the flame and the ice in the first cooling channel 211, and the inner surface and the outer surface of the shell are sprayed with high-temperature resistant coatings which play a role in protecting the shell. In most embodiments, an integrated ice block is fixed in the inner cavity of the housing, a first cooling channel 211 is formed in the ice block, the first cooling channel 211 penetrates through the ice block along the length direction of the housing, flame at about 3000 ℃ enters the first cooling channel 211 to be cooled, and the flame sequentially flows through the first cooling channel 211 corresponding to each ice block to be cooled to about 100 ℃ and then enters the vacuum-pumping system 3.
In order to facilitate the fixation of the first cooling section 21 and the second cooling section 22, the peripheral wall of each of the first cooling section 22 and the second cooling section 22 of the embodiment of the invention is provided with a housing, and the fixation of different cooling sections is realized between adjacent first cooling sections 21, between adjacent second cooling sections 22, and between adjacent first cooling sections 21 and adjacent second cooling sections 22 through the housing. The shell of the embodiment of the invention can be used as a freezing mould and can bear ice blocks to finish the temperature reduction treatment of flame.
In some embodiments, it is preferred that the rapid cooler 2 further comprises: a shell for protecting the ice cooling module, wherein the shell is fixedly arranged at the outer side of the ice cooling module; the ice cooling module includes: at least one first cooling section 21 and at least one second cooling section 22 which are communicated in sequence; the housing includes: the first shell is fixedly arranged on the outer side of the first cooling section 21, and the second shell is fixedly arranged on the outer side of the second cooling section 22; the fixing of different cooling sections is realized between adjacent first cooling sections 21, between adjacent second cooling sections 22, and between adjacent first cooling sections 21 and second cooling sections 22 through the first casing and/or the second casing.
Preferably, the first housing and the second housing are each of a stubby structure. A sealing ring is arranged between two adjacent flange plates and is used for preventing ice from seeping from a gap between the two flange plates after being melted into water; two adjacent flanges are fixedly connected through a plurality of bolts.
Referring to fig. 4, fig. 4 is a schematic diagram of an application example of the device for rocket engine high altitude simulation test, the vacuum pumping system 3 of the embodiment includes: at least one group of vacuumizing modules 31, wherein each group of vacuumizing modules 31 comprises first-stage to Nth-stage vacuum pumps, and N is more than or equal to 2; the first-stage to Nth-stage vacuum pumps are connected in series in sequence; different groups of vacuum-pumping modules 31 are arranged in parallel; all the vacuumizing modules 31 are communicated with the rapid cooler 2, the ground test bed 1 and the water spray condenser 5, and each vacuumizing module 31 can vacuumize the rapid cooler 2, the ground test bed 1 and the water spray condenser 5.
More specifically, the vacuum pumping system 3 of the present embodiment includes: at least one group of vacuumizing modules, wherein each group of vacuumizing modules comprises first-stage to Nth-stage vacuum pumps, and N is more than or equal to 2; the first-stage to Nth-stage vacuum pumps are connected in series in sequence; different groups of vacuum-pumping modules are arranged in parallel; all evacuation modules all communicate with rapid cooler 2 and ground test platform 1, and every evacuation module all can be to rapid cooler 2 and ground test platform 1 evacuation. The vacuum pumping system 3 shown in fig. 4 includes a first group of vacuum pumping modules, a second group of vacuum pumping modules, a third group of vacuum pumping modules and a fourth group of vacuum pumping modules, and the first group of vacuum pumping modules, the second group of vacuum pumping modules, the third group of vacuum pumping modules and the fourth group of vacuum pumping modules are all arranged in parallel. Each set of evacuation modules shown in FIG. 4 includes a primary vacuum pump 311, a secondary vacuum pump 312, and a tertiary vacuum pump 313, where the primary vacuum pump 311, the secondary vacuum pump 312, and the tertiary vacuum pump 313 are arranged in series.
Referring to fig. 4, fig. 4 is a schematic diagram of an application example of a device for rocket engine altitude simulation test, and the device for rocket engine altitude simulation test of the present embodiment includes, in addition to a ground test stand 1, a rapid cooler 2, an evacuation system 3, a water spray condenser 5 and a pressurization module 6: the drying module 8 is used for drying the gas flowing out of the water spray condenser 5, the drying module 8 comprises a drying agent, and the drying module 8 is installed between the water spray condenser 5 and the vacuum pumping system 3; the evacuation system 3 is operated such that the drying module 8 is also in a correspondingly high vacuum state.
The drying module 8 is used for drying gas flowing out of the water spray condenser 5, the drying module 8 comprises a drying agent, and the drying module 8 is installed between the water spray condenser 5 and the vacuum pumping system 3; the evacuation system 3 is operated such that the drying module 8 is also in a correspondingly high vacuum state.
Referring to fig. 4, the number of the drying modules 8 is the same as the number of the vacuum-pumping modules 31 of the vacuum-pumping system 3; a drying module 8 is arranged between the water spray condenser 5 and each vacuumizing module 31.
In some embodiments, referring to fig. 5, the apparatus for rocket engine high altitude simulation test of the present embodiment includes, in addition to a ground test stand 1, a rapid cooler 2, an evacuation system 3, a water spray condenser 5 and a pressurization module 6: the M sub-cooling systems are uniformly arranged around the rapid cooler 2, and are circularly distributed around the center line of the rapid cooler 2; each sub-cooling system is communicated with the rapid cooler 2, and a pressure limiting valve is arranged at the communication part of each sub-cooling system and the rapid cooler 2; when the flame pressure in the rapid cooler 2 exceeds the threshold pressure of the pressure limiting valve, the flame exceeding the threshold pressure enters the sub-cooling system for rapid cooling.
The preferred sub-cooling system comprises: a sub-cooler and an evacuation subsystem; the rapid cooler 2, the sub-cooler and the vacuumizing subsystem are communicated with each other, and the vacuumizing subsystem is used for vacuumizing the sub-cooler so that the sub-cooler is in a corresponding high-altitude vacuum state.
An embodiment of the present invention further provides a method for a rocket engine high altitude simulation test, please refer to fig. 6, which includes:
starting a vacuumizing system 3, vacuumizing the ground test bed 1, the rapid cooler 2 and the water spray condenser 5 to a corresponding high-altitude vacuum state, continuously extracting flames of the ground test bed 1, the rapid cooler 2 and the water spray condenser 5 by the vacuumizing system 3, and continuously maintaining the vacuum state by the vacuumizing system 3;
starting the rocket engine 4, and spraying flame at about 3000 ℃ from a spray pipe of the rocket engine 4;
the rapid cooler 2 cools the flame into airflow of 100 +/-20 ℃, and the rapid cooler 2 cools the flame and simultaneously carries out first harmless treatment on the flame;
the water spray condenser 5 continuously cools the air flow, and the water spray condenser 5 cools the air flow and simultaneously carries out secondary harmless treatment on the air flow;
the gas flow after the second harmless treatment is discharged from the vacuum-pumping system 3.
More specifically, the rocket engine 4 is installed behind a ground test bed 1 of the test bed, the solid propellant is loaded into the rocket engine 4, the rocket engine 4 is ignited, the spray pipe sprays the ultra-high temperature flame, the rapid cooler is installed right opposite to the spray pipe of the rocket engine 4, the ultra-high temperature flame completely enters a cooling channel of the rapid cooler, the ultra-high temperature flame and the cooling channel added with alkaline substances carry out energy exchange of internal energy and kinetic energy, the temperature and the kinetic energy of the ultra-high temperature flame are greatly reduced, and finally the ultra-high temperature flame enters the water spray condenser 5 from an outlet of the cooling channel in a low-temperature.
As HCl harmful gas with the mass ratio of more than 20% is contained in the flame in the test of the rocket engine 2, the ice structure is formed by freezing alkaline water, and alkaline materials are added in the ice during the manufacturing process, so that the flame is harmlessly treated in the energy conversion process.
The high-altitude simulation test is ground test equipment capable of simulating conditions such as speed, height and the like of the rocket motor 4 during flying in the air, and is an important ground test for examining the performance and reliability of the rocket motor 4 and simultaneously accurately measuring the ballistic performance in the rocket motor 4. The working principle of the rocket engine 4 is as follows: the solid propellant is ignited and then burnt in the combustion chamber to generate high-temperature and high-pressure fuel gas, namely chemical energy is converted into heat energy; the gas expands and accelerates through the jet pipe, the heat energy is converted into kinetic energy, and the kinetic energy is discharged from the jet pipe at a very high speed so as to generate thrust to push the rocket to fly forwards.
The high-altitude simulation test is that a backfire phenomenon exists at the initial ignition stage and in the flameout process of the rocket motor 4, and great danger is brought to a nozzle of the rocket motor 4. Meanwhile, the high-temperature fuel gas reflowing during tempering poses serious threats to test equipment and circuits arranged in the high-altitude cabin.
In order to prevent backfire, the rapid cooler 2 and the water spray condenser 5 are adopted to cool the flame in the embodiment of the disclosure, and the temperature of the flame after cooling is only below 100 ℃, so that the high-temperature flame can be prevented from damaging the rocket engine 4, particularly the spray pipe, the test equipment and the like.
In the above method, starting the vacuum-pumping system 3 to vacuum the ground test stand 1, the rapid cooler 2 and the water spray condenser 5 to a corresponding high vacuum state, including:
starting a plurality of groups of vacuumizing modules 31 of the vacuumizing system 3, and vacuumizing the ground test bed 1, the rapid cooler 2 and the water spray condenser 5 by the plurality of groups of vacuumizing modules 31 at the same time;
the multistage vacuum pumps of each group of vacuumizing modules 31 are connected in series, and simultaneously vacuumize the ground test bed 1, the rapid cooler 2 and the water spray condenser 5;
and the multistage vacuum pump simultaneously pumps out airflow for a high-altitude simulation test of the rocket engine 4.
The thrust of the rocket engine 2 is continuously increased along with the rise of the height until the maximum thrust is reached in a vacuum environment, the environmental pressure of the rocket engine 2 in the vacuum environment is very low, and the approximate vacuum environment is realized by the vacuumizing module 4; the last ice cooling barrel 31 finishes the flame cooling within 1min after the rocket engine 2 stops injecting the flame.
The approximate vacuum environment disclosed by the invention is characterized in that the flame exhausted from the nozzle of the rocket engine 4 is used as the airflow of each stage of vacuum pump, the multistage vacuum pumps sequentially expand pressure, the last stage vacuum pump exhausts the airflow into the atmosphere at the ambient pressure, and the counter pressure felt at the nozzle is not the ambient atmospheric pressure but the local pressure, so that the high-altitude vacuum environment is simulated.
In the above method, the rocket engine 4 is started, and the nozzle of the rocket engine 4 ejects flames of about 3000 ℃, including:
starting the rocket engine 4, and enabling the rocket engine 4 to swing on the ground test bed 1;
the jet pipe of the rocket engine 4 jets out flame at about 3000 ℃, and the flame swings along with the swing of the rocket engine 4;
and in the region exceeding the threshold pressure in the rapid cooler 2, the pressure limiting valve of the corresponding region is opened, and the flame exceeding the threshold pressure enters the sub-cooling system of the corresponding region for cooling and discharging.
The flame of the rocket engine 4 passes through the rapid cooler 2 and the water spray condenser 5, so that low-temperature and harmless emission of the flame can be realized. Because the ice is used as the energy conversion material, the mass of the ice required is far less than that of the water under the condition of the same amount of energy exchange, the efficiency is higher, and the drainage performance of the product is stronger. The main technical advantages are as follows: (1) ice is used as a transduction processing medium, so that the transduction efficiency is higher; (2) due to the addition of the alkaline material, harmful component HCl in the flame of the solid rocket engine 4 can be eliminated simultaneously, and harmless treatment is realized; (3) the rapid cooler 2 and the water spray condenser 5 adopt a modular structure, have strong mobility and can be rapidly applied to various test sites. The temperature of the fluid flowing into the vacuum-pumping system from the water spray condenser 5 of the rapid cooler 2 and the water spray condenser 5 of the embodiment of the present disclosure is below 100 ℃, preferably below 80 ℃; after being treated by the rapid cooler 2 and the water spray condenser 5, the flame is discharged into the fluid of the vacuum-pumping system from the water spray condenser 5, and the noise is reduced by 60 percent within the range of 50 meters when the flame is treated by the rapid cooler 2 and the water spray condenser 5; after being treated by the rapid cooler 2 and the water spray condenser 5, the discharge amount of HCl gas in the flame is reduced by 95 percent.
The above-listed detailed description is merely a specific description of possible embodiments of the present disclosure, and is not intended to limit the scope of the disclosure, which is intended to include within its scope equivalent embodiments or modifications that do not depart from the technical spirit of the present disclosure.
It will be evident to those skilled in the art that the disclosure is not limited to the details of the foregoing illustrative embodiments, and that the present disclosure may be embodied in other specific forms without departing from the spirit or essential characteristics thereof. The present embodiments are therefore to be considered in all respects as illustrative and not restrictive, the scope of the disclosure being indicated by the appended claims rather than by the foregoing description, and all changes which come within the meaning and range of equivalency of the claims are therefore intended to be embraced therein. Any reference sign in a claim should not be construed as limiting the claim concerned.
Furthermore, it should be understood that although the present description refers to embodiments, not every embodiment may contain only a single embodiment, and such description is for clarity only, and those skilled in the art should integrate the description, and the embodiments may be combined as appropriate to form other embodiments understood by those skilled in the art.
Claims (10)
1. A device for rocket engine high altitude simulation test is characterized by comprising:
a rapid cooler in communication with a nozzle of the rocket engine, the rapid cooler comprising an ice cooling module of a material of ice; the ice cooling module is provided with a cooling channel, flame generated by the rocket engine is sprayed out from the spray pipe and flows through the cooling channel to contact ice around the cooling channel in real time, and the flame flows through the cooling channel to be cooled into airflow at 100 +/-20 ℃;
the water spray condenser is communicated with the rapid cooler, 100 +/-20 ℃ of airflow enters the water spray condenser for cooling, and the airflow cooled by the water spray condenser is discharged to the outer side of the device;
the ground test bed is provided with a rocket engine and is communicated with the rapid cooler and the water spray condenser;
the vacuumizing system is communicated with the ground test bed, the rapid cooler and the water spray condenser; the ground test bed, the rapid cooler and the water spray condenser are all in corresponding high vacuum states.
2. A device for rocket engine high altitude simulation tests according to claim 1, further comprising:
and the pressurizing module is installed between the rapid condenser and the water spray condenser, and the air flow flows through the pressurizing module to increase the pressure and then enters the water spray condenser for cooling.
3. An apparatus for rocket engine altitude simulation tests according to claim 1 or 2, wherein said evacuation system comprises: at least one group of vacuumizing modules, wherein each group of vacuumizing modules comprises first-stage to Nth-stage vacuum pumps, and N is more than or equal to 2;
the first-stage to Nth-stage vacuum pumps are connected in series in sequence;
different groups of vacuum-pumping modules are arranged in parallel;
all evacuation modules all with quick cooler ground test bench with the water spray condenser intercommunication, every evacuation module all can be right quick cooler ground test bench with the water spray condenser evacuation.
4. A device for rocket engine high altitude simulation tests according to claim 1, further comprising:
the drying module is used for drying the gas flowing out of the water spray condenser and comprises a drying agent, and the drying module is installed between the water spray condenser and the vacuumizing system; the vacuum-pumping system works so that the drying module is also in a corresponding high-altitude vacuum state.
5. A device for rocket engine high altitude simulation tests according to claim 4, wherein the number of said drying modules is the same as the number of said evacuation modules of said evacuation system;
and one drying module is arranged between the water spray condenser and each vacuumizing module.
6. A device for rocket engine high altitude simulation tests according to claim 1, further comprising: the M sub-cooling systems are uniformly arranged around the rapid cooler and are circularly distributed around the central line of the rapid cooler;
each sub-cooling system is communicated with the rapid cooler, and a pressure limiting valve is arranged at the communication position of each sub-cooling system and the rapid cooler;
and when the flame pressure in the rapid cooler exceeds the threshold pressure of the pressure limiting valve, the flame exceeding the threshold pressure enters the sub-cooling system for rapid cooling.
7. An apparatus for rocket engine altitude simulation testing according to claim 6, wherein said sub-cooling system comprises: a sub-cooler and an evacuation subsystem;
the rapid cooler, the sub-cooler and the vacuumizing subsystem are communicated with each other, and the vacuumizing subsystem is used for vacuumizing the sub-cooler, so that the sub-cooler is in a corresponding high-altitude vacuum state.
8. A method for high altitude simulation test of rocket engine is characterized by comprising the following steps:
starting a vacuumizing system, vacuumizing the ground test bed, the rapid cooler and the water spray condenser to a corresponding high-altitude vacuum state, continuously pumping flames of the ground test bed, the rapid cooler and the water spray condenser by the vacuumizing system, and continuously maintaining the vacuum state by the vacuumizing system;
starting the rocket engine, and spraying flame at about 3000 ℃ from a spray pipe of the rocket engine;
the rapid cooler cools the flame into airflow of 100 +/-20 ℃, and the rapid cooler performs first harmless treatment on the flame while cooling the flame;
the water spray condenser continuously cools the air flow, and the water spray condenser cools the air flow and simultaneously carries out secondary harmless treatment on the air flow;
and discharging the airflow subjected to the second harmless treatment from the vacuum-pumping system.
9. A method for rocket engine altitude simulation testing according to claim 8, wherein said starting vacuum pumping system to vacuum the ground test stand, the quencher and the water-spray condenser to respective altitude vacuum conditions comprises:
starting a plurality of groups of vacuumizing modules of a vacuumizing system, and vacuumizing the ground test bed, the quick cooler and the water spray condenser by the plurality of groups of vacuumizing modules simultaneously;
the multistage vacuum pumps of each group of vacuumizing modules are connected in series, and simultaneously vacuumize the ground test bed, the rapid cooler and the water spray condenser;
the multistage vacuum pump simultaneously extracts airflow out of the rocket engine high-altitude simulation test device.
10. A method for rocket motor high altitude simulation test according to claim 8, wherein said starting rocket motor, the rocket motor nozzle spraying flame about 3000 ℃, comprises:
starting the rocket engine, and enabling the rocket engine to swing on the ground test bed;
the jet pipe of the rocket engine sprays flame at about 3000 ℃, and the flame swings along with the swing of the rocket engine;
and in the region exceeding the threshold pressure in the rapid cooler, opening a pressure limiting valve corresponding to the region, and allowing the flame exceeding the threshold pressure to enter a sub-cooling system corresponding to the region for cooling and discharging.
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