CN112377325B - Hypersonic strong precooling turbine-based stamping combined engine - Google Patents
Hypersonic strong precooling turbine-based stamping combined engine Download PDFInfo
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- CN112377325B CN112377325B CN202011238937.4A CN202011238937A CN112377325B CN 112377325 B CN112377325 B CN 112377325B CN 202011238937 A CN202011238937 A CN 202011238937A CN 112377325 B CN112377325 B CN 112377325B
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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
- F02K7/00—Plants in which the working fluid is used in a jet only, i.e. the plants not having a turbine or other engine driving a compressor or a ducted fan; Control thereof
- F02K7/10—Plants in which the working fluid is used in a jet only, i.e. the plants not having a turbine or other engine driving a compressor or a ducted fan; Control thereof characterised by having ram-action compression, i.e. aero-thermo-dynamic-ducts or ram-jet engines
- F02K7/16—Composite ram-jet/turbo-jet engines
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- F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
- F02—COMBUSTION ENGINES; HOT-GAS OR COMBUSTION-PRODUCT ENGINE PLANTS
- F02C—GAS-TURBINE PLANTS; AIR INTAKES FOR JET-PROPULSION PLANTS; CONTROLLING FUEL SUPPLY IN AIR-BREATHING JET-PROPULSION PLANTS
- F02C7/00—Features, components parts, details or accessories, not provided for in, or of interest apart form groups F02C1/00 - F02C6/00; Air intakes for jet-propulsion plants
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- F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
- F02—COMBUSTION ENGINES; HOT-GAS OR COMBUSTION-PRODUCT ENGINE PLANTS
- F02C—GAS-TURBINE PLANTS; AIR INTAKES FOR JET-PROPULSION PLANTS; CONTROLLING FUEL SUPPLY IN AIR-BREATHING JET-PROPULSION PLANTS
- F02C7/00—Features, components parts, details or accessories, not provided for in, or of interest apart form groups F02C1/00 - F02C6/00; Air intakes for jet-propulsion plants
- F02C7/04—Air intakes for gas-turbine plants or jet-propulsion plants
- F02C7/057—Control or regulation
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- F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
- F02—COMBUSTION ENGINES; HOT-GAS OR COMBUSTION-PRODUCT ENGINE PLANTS
- F02C—GAS-TURBINE PLANTS; AIR INTAKES FOR JET-PROPULSION PLANTS; CONTROLLING FUEL SUPPLY IN AIR-BREATHING JET-PROPULSION PLANTS
- F02C7/00—Features, components parts, details or accessories, not provided for in, or of interest apart form groups F02C1/00 - F02C6/00; Air intakes for jet-propulsion plants
- F02C7/12—Cooling of plants
- F02C7/14—Cooling of plants of fluids in the plant, e.g. lubricant or fuel
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- F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
- F02—COMBUSTION ENGINES; HOT-GAS OR COMBUSTION-PRODUCT ENGINE PLANTS
- F02C—GAS-TURBINE PLANTS; AIR INTAKES FOR JET-PROPULSION PLANTS; CONTROLLING FUEL SUPPLY IN AIR-BREATHING JET-PROPULSION PLANTS
- F02C7/00—Features, components parts, details or accessories, not provided for in, or of interest apart form groups F02C1/00 - F02C6/00; Air intakes for jet-propulsion plants
- F02C7/12—Cooling of plants
- F02C7/16—Cooling of plants characterised by cooling medium
- F02C7/18—Cooling of plants characterised by cooling medium the medium being gaseous, e.g. air
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- Y—GENERAL TAGGING OF NEW TECHNOLOGICAL DEVELOPMENTS; GENERAL TAGGING OF CROSS-SECTIONAL TECHNOLOGIES SPANNING OVER SEVERAL SECTIONS OF THE IPC; TECHNICAL SUBJECTS COVERED BY FORMER USPC CROSS-REFERENCE ART COLLECTIONS [XRACs] AND DIGESTS
- Y02—TECHNOLOGIES OR APPLICATIONS FOR MITIGATION OR ADAPTATION AGAINST CLIMATE CHANGE
- Y02T—CLIMATE CHANGE MITIGATION TECHNOLOGIES RELATED TO TRANSPORTATION
- Y02T50/00—Aeronautics or air transport
- Y02T50/60—Efficient propulsion technologies, e.g. for aircraft
Abstract
The invention discloses a hypersonic strong precooling turbine-based stamping combined engine which comprises an air inlet device, an aviation gas turbine engine, a turbine engine exhaust device, a coolant system and a strong precooling system. When the forced precooling system is used, the first valve is opened, the second valve and the third valve are closed, the forced precooling system works, incoming flow air enters the engine from the air inlet device, is cooled by the forced precooling system, enters the aviation gas turbine engine and then enters the exhaust device of the turbine engine to generate thrust; after being cooled by the strong precooling system, the cooling working medium enters the aviation gas turbine engine to absorb the heat of the aviation gas turbine engine, so that the problem that the thrust is seriously reduced due to the limitation of the self pneumatic thermodynamic cycle of the turbine engine and the problem that a cold end part of the turbine engine faces high temperature due to high-temperature incoming flow are avoided, and the problems that the hot end part is difficult to cool and the turbojet engine is difficult to work normally and reliably due to high temperature of a cooling air source are avoided.
Description
Technical Field
The invention relates to the technical field of hypersonic aircrafts, in particular to a hypersonic strong precooling turbine-based stamping combined engine.
Background
Modern scientific and technological development has shown that the future scientific and technological development direction must shift to high altitude and near space, and as high-efficient transport means of high altitude and near space, military and civilian fields are urgent to can take off and land, used supersonic velocity, hypersonic velocity aircraft demand horizontally. In the military field, the aircraft has the characteristic of high flying speed, and also has the technical advantages of capability of horizontally taking off and landing on a conventional airport, flexible use, short take-off preparation period, capability of being repeatedly used and the like. In the civil field, the hypersonic aircraft capable of being horizontally lifted and repeatedly used has the capability of realizing double-stage or even single-stage rail entering, can effectively reduce unit payload cost, and can complete low-cost quick round-trip transportation tasks. Therefore, horizontal take-off and landing and repeated use of supersonic/hypersonic aircrafts have become one of the research hotspots in the aerospace field.
The origin of the excellent performance of the supersonic/hypersonic aircraft is the revolutionary power system of the supersonic/hypersonic aircraft, and as the flying speed of the supersonic/hypersonic aircraft is far higher than that of the traditional aircraft, the supersonic/hypersonic aircraft puts high requirements on the power of the aircraft. The supersonic/hypersonic aircraft capable of horizontal take-off and landing and being reused has higher requirements on power, and not only needs a power system to have excellent performance at high speed, but also needs the power system to be capable of starting at zero Mach, and has high fuel use efficiency and high reliability in the whole flight speed range including a low-speed section.
Currently, the types of power suitable for supersonic/hypersonic flight mainly include: turbojet, scramjet, rocket engine, rocket based Ramjet (RBCC), turbo ramjet (TBCC), Tri-Jet (Tri-Jet). The range of the operating Mach number of the turbojet engine is low, and the turbojet engine is not suitable for supersonic speed/hypersonic speed flight with high Mach number; the starting Mach number of the scramjet engine with the sub-combustion/super-combustion is higher, and the scramjet engine cannot independently start from the Mach number 0; the rocket engine technology is relatively mature, but the fuel utilization effectiveness is extremely low, the launching cost is extremely high, the horizontal take-off and landing and the repeated use are difficult, and the rocket engine is not suitable for being used as a supersonic/hypersonic aircraft power system capable of horizontal take-off and landing and repeated use; the RBCC engine can exert the advantages of a ramjet engine to make up the defects of a rocket, can improve the effectiveness of fuel utilization to a certain extent, but the low-speed performance of the power is poor, the power is not suitable for horizontal take-off and landing work on the ground, and the reusability and the use maneuverability are relatively poor.
The TBCC engine integrates the advantages of a turbojet engine and a ramjet engine in respective applicable flight ranges, has the advantages of good economy and higher specific impulse in a wide Mach number range, and is a potential power system suitable for horizontal take-off and landing and repeatedly using an ultrasonic/hypersonic aircraft. However, the power of the type has great technical difficulty, and mainly has a thrust gap: the effective working range of the existing turbojet engine is Mach number 0-2 +, and when the flying Mach number exceeds the range, the turbojet engine is limited by self pneumatic thermodynamic cycle due to the fact that the stagnation temperature of incoming high-speed airflow is increased, and thrust is seriously reduced; meanwhile, high-temperature incoming flow causes the problem that a cold end part of the turbojet engine faces high temperature, and the hot end part is difficult to cool due to the high temperature of a cooling air source, so that the turbojet engine is difficult to work normally and reliably. And when the effective working range of the ramjet is Mach 3+, the ramjet cannot generate enough thrust because the performance of the ramjet is deteriorated due to the fact that the flying Mach is too low when the Mach is lower than 3, and therefore the TBCC cannot generate enough total thrust within the range of Mach 2.5-3.5.
Therefore, the technical problem to be solved by those skilled in the art is how to solve the problem that the turbojet cannot work normally and reliably to generate sufficient thrust when flying at high mach number due to the limitations of the working principle and materials of the turbojet.
Disclosure of Invention
In view of the above, an object of the present invention is to provide a hypersonic strong precooling turbo-based ramjet combined engine, so as to solve the problem that the turbojet cannot work normally and reliably to generate sufficient thrust due to the limitations of the operating principle and materials of the engine itself when flying at a high mach number.
In order to achieve the above object, the present invention provides the following solutions:
a hypersonic strong precooling turbine-based stamping combined engine comprises an air inlet device, an aviation gas turbine engine and a turbine engine exhaust device, and further comprises a coolant system and a strong precooling system;
an outlet of the air inlet device is communicated with an air inlet of the strong precooling system, an air outlet of the strong precooling system is communicated with an air inlet of the aviation gas turbine engine, and an outlet of the aviation gas turbine engine is communicated with an inlet of the turbine engine exhaust device;
the working medium outlet of the coolant system is in conductive connection with the working medium inlet of the strong precooling system through a first valve, the working medium outlet of the coolant system is in conductive connection with the first fuel inlet of the aviation gas turbine engine through a second valve, and the working medium outlet of the strong precooling system is in conductive connection with the second fuel inlet of the aviation gas turbine engine through a third valve.
Preferably, in the hypersonic strong precooling turbine-based ramjet combined engine, the hypersonic strong precooling turbine-based ramjet combined engine further comprises a ramjet engine and a ramjet engine exhaust device;
the first fuel inlet of the ramjet is in conductive connection with the working medium outlet of the coolant system through a fourth valve;
an air inlet of the ramjet is communicated with an outlet of the air inlet device;
an inlet of the ramjet exhaust is in communication with an outlet of the ramjet;
a second fuel inlet of the ramjet is connected with a working medium outlet of the strong precooling system through a fifth valve;
when the strong precooling system does not work and the aviation gas turbine engine starts to work, the second valve is opened, the first valve, the third valve, the fourth valve and the fifth valve are all closed, and the working Mach number is in a first preset range value;
when the strong precooling system starts to work, the aviation gas turbine engine works, and the ramjet does not work, the first valve and the third valve are opened, the second valve, the fourth valve and the fifth valve are closed, and the working Mach number is in a second preset range value;
when the strong precooling system works, the aviation gas turbine engine works, and the ramjet engine starts to work, the first valve, the third valve and the fifth valve are opened, the second valve and the fourth valve are closed, and the working Mach number reaches a third preset range value;
when the strong precooling system and the aviation gas turbine engine both stop working and the ramjet engine works independently, the first valve, the second valve, the third valve and the fifth valve are closed, the fourth valve is opened, and the working Mach number reaches a fourth preset range value.
Preferably, in the turbo-based ramjet combined engine with super-supersonic speed pre-cooling, the first preset range value is greater than or equal to 0 and less than or equal to 2.5;
the second preset range value is greater than or equal to 0 and less than or equal to 2.5;
the third preset range value is greater than or equal to 2.0 and less than or equal to 4.0;
the fourth predetermined range value is greater than or equal to 3.0.
Preferably, in the above-mentioned high supersonic strong precooling turbine-based ramjet combined engine, the strong precooling system includes at least one set of basic heat exchange-expansion units;
the basic heat exchange-expansion unit comprises a first heat exchanger, a sixth valve, a medium turbine and a second heat exchanger;
a cold side inlet of the first heat exchanger is a working medium inlet of the strong precooling system, and a cold side outlet of the first heat exchanger is communicated with an inlet of the medium turbine;
a cold side inlet of the second heat exchanger is communicated with an outlet of the medium turbine, and a cold side outlet of the second heat exchanger is a working medium outlet of the strong precooling system;
the sixth valve is arranged in parallel with the medium turbine;
and a hot side inlet of the first heat exchanger is an air inlet of the strong precooling system, a hot side outlet of the first heat exchanger is communicated with a hot side inlet of the second heat exchanger, and a hot side outlet of the second heat exchanger is an air outlet of the strong precooling system.
Preferably, in the above mentioned high supersonic velocity strong precooling turbine-based ramjet combined engine, the aviation gas turbine engine comprises a gas compressor, a gas combustion chamber, a gas turbine and an afterburner;
an air inlet of the air compressor is an air inlet of the aviation gas turbine engine, an air outlet of the air compressor is communicated with an air inlet of the gas combustion chamber, and the air compressor is in power connection with the gas turbine through a rotating shaft;
a gas inlet of the gas turbine is communicated with a gas outlet of the gas combustion chamber;
the second valve comprises a first gas regulating valve and a first stress application regulating valve, the first gas regulating valve and the first stress application regulating valve are arranged in parallel, two ends of the first gas regulating valve are respectively communicated with a working medium outlet of the coolant system and a first gas inlet of the gas combustion chamber, and two ends of the first stress application regulating valve are respectively communicated with the working medium outlet of the coolant system and the first gas inlet of the stress application combustion chamber;
the third valve comprises a second fuel regulating valve and a second afterburning regulating valve, the second fuel regulating valve and the second afterburning regulating valve are arranged in parallel, two ends of the second fuel regulating valve are respectively communicated with a working medium outlet of the strong precooling system and a second fuel gas inlet of the gas combustion chamber, and two ends of the second afterburning regulating valve are respectively communicated with a working medium outlet of the strong precooling system and a second fuel gas inlet of the afterburning chamber;
the outlet of the afterburner is the outlet of the aero gas turbine engine.
Preferably, in the above mentioned super-supersonic strong precooling turbo-based ramjet combined engine, the ramjet engine comprises a ramjet combustion chamber; and/or
The coolant system includes a coolant pump.
Preferably, in the above mentioned super-supersonic strong precooling turbine-based ramjet combined engine, the turbine engine exhaust device is a turbine engine tail nozzle; and/or
The ramjet exhaust device is a ramjet exhaust nozzle.
Preferably, in the above mentioned high supersonic speed strong precooling turbo-based ramjet combined engine, the aviation gas turbine engine is a turbofan engine or a turbojet engine; and/or
The ramjet is a sub-combustion ramjet, a super-combustion ramjet or a bimodal ramjet.
Preferably, in the above mentioned hypersonic strong precooling turbine-based ramjet combined engine, the fuel used by the hypersonic strong precooling turbine-based ramjet combined engine includes: aviation kerosene; and/or
Liquid methane; and/or
Liquid hydrogen.
According to the technical scheme, when the hypersonic strong precooling turbine-based stamping combined engine is used, the hypersonic strong precooling turbine-based stamping combined engine starts to work, at the moment, the first valve is opened, the second valve and the third valve are closed, the strong precooling system starts to work, incoming flow air enters the aircraft gas turbine engine after entering the engine from the air inlet device and being cooled by the strong precooling system, and then enters the turbine engine exhaust device to generate thrust; after the cooling working medium is cooled by the strong precooling system, the cooling working medium enters the aviation gas turbine engine, absorbs the heat of the aviation gas turbine engine, avoids the turbine engine from being limited by self pneumatic thermodynamic cycle, leads to the problem of serious thrust reduction and avoids the problem that cold end parts of the turbine engine face high temperature due to high-temperature incoming flow, and because the air conveyed by the air inlet device meets cold through the strong precooling system, the problem that the hot end parts are difficult to cool due to high temperature of a cooling air source is avoided, and the turbojet engine is difficult to work normally and reliably. In conclusion, the invention solves the problem that the turbojet engine cannot normally and reliably work to generate enough thrust due to the limitation of the working principle, materials and the like of the engine when flying at a high Mach number.
Drawings
In order to more clearly illustrate the embodiments of the present invention or the technical solutions in the prior art, the drawings used in the embodiments or the prior art descriptions will be briefly described below, it is obvious that the drawings in the following description are only some embodiments of the present invention, and other drawings can be obtained by those skilled in the art without creative efforts.
Fig. 1 is a schematic structural diagram of a hypersonic strong precooling turbine-based ramjet combined engine according to an embodiment of the invention;
fig. 2 is a schematic structural diagram of a hypersonic strong precooling turbine-based ramjet combined engine provided in the second embodiment of the invention;
FIG. 3 is another schematic structural diagram of a hypersonic strong precooling turbine-based ramjet combined engine according to a second embodiment of the invention;
fig. 4 is a schematic structural diagram of a strong precooling system of a hypersonic strong precooling turbine-based ramjet combined engine provided by the invention.
Wherein, in fig. 1-4:
the system comprises an air inlet device 1, an aviation gas turbine engine 2, a turbine engine exhaust device 3, a coolant system 4, a strong precooling system 5, a first valve 6, a second valve 7, a third valve 8, a ramjet engine 9, a ramjet engine exhaust device 10, a fourth valve 11, a fifth valve 12, a first heat exchanger 501, a sixth valve 502, a medium turbine 503, a second heat exchanger 504, a compressor 201, a gas combustion chamber 202, a gas turbine 203, an afterburner 204, a first gas regulating valve 701, a first boost regulating valve 702, a second gas regulating valve 801 and a second boost regulating valve 802.
Detailed Description
In order to make the technical solutions of the present invention better understood, the present invention will be further described in detail with reference to the accompanying drawings and specific embodiments.
Example one
As shown in fig. 1-4, the invention discloses a hypersonic strong precooling turbine-based ramjet combined engine. The hypersonic strong precooling turbine-based stamping combined engine comprises an air inlet device 1, an aviation gas turbine engine 2, a turbine engine exhaust device 3, a coolant system 4 and a strong precooling system 5.
An outlet of the air inlet device 1 is communicated with an air inlet of the strong precooling system 5, an air outlet of the strong precooling system 5 is communicated with an air inlet of the aviation gas turbine engine 2, and an outlet of the aviation gas turbine engine 2 is communicated with an inlet of the turbine engine exhaust device 3.
The working medium outlet of the coolant system 4 and the working medium inlet of the strong precooling system 5 are in conductive connection through the first valve 6, namely the first valve 6 controls the on-off of the working medium outlet of the coolant system 4 and the working medium inlet of the strong precooling system 5. The working medium outlet of the coolant system 4 and the first fuel inlet of the aircraft gas turbine engine 2 are in conductive connection through a second valve 7, namely the second valve 7 controls the on-off of the working medium outlet of the coolant system 4 and the first fuel inlet of the aircraft gas turbine engine 2. The working medium outlet of the strong precooling system 5 and the second fuel inlet of the aircraft gas turbine engine 2 are in conductive connection through a third valve 8, namely the third valve 8 controls the connection and disconnection between the working medium outlet of the strong precooling system 5 and the second fuel inlet of the aircraft gas turbine engine 2.
Further, the invention specifically discloses that the aviation gas turbine engine 2 comprises a compressor 201, a gas combustion chamber 202, a gas turbine 203 and an afterburner 204. An air inlet of the compressor 201 is an air inlet of the aircraft gas turbine engine 2, an air outlet of the compressor 201 is communicated with an air inlet of the gas combustion chamber 202, and the compressor 201 is in power connection with the gas turbine 203 through a rotating shaft. The gas inlet of the gas turbine 203 communicates with the gas outlet of the gas combustor 202. The second valve 7 comprises a first gas regulating valve 701 and a first force-adding regulating valve 702, the first gas regulating valve 701 and the first force-adding regulating valve 702 are arranged in parallel, two ends of the first gas regulating valve 701 are respectively communicated with a working medium outlet of the coolant system 4 and a first gas inlet of the gas combustion chamber 202, and two ends of the first force-adding regulating valve 702 are respectively communicated with a working medium outlet of the coolant system 4 and a first gas inlet of the force-adding combustion chamber 204. The third valve 8 comprises a second fuel regulating valve 801 and a second afterburning valve 802, the second fuel regulating valve 801 and the second afterburning valve 802 are arranged in parallel, two ends of the second fuel regulating valve 801 are respectively communicated with a working medium outlet of the strong precooling system 5 and a second fuel gas inlet of the gas combustion chamber 202, and two ends of the second afterburning valve 802 are respectively communicated with a working medium outlet of the strong precooling system 5 and a second fuel gas inlet of the afterburning chamber 204. The outlet of the afterburner 204 is the outlet of the aircraft gas turbine engine 2.
Further, the present invention discloses that the strong precooling system 5 comprises at least one set of basic heat exchange-expansion units. The basic heat exchange-expansion unit comprises a first heat exchanger 501, a sixth valve 502, a medium turbine 503 and a second heat exchanger 504; the cold side inlet of the first heat exchanger 501 is a working medium inlet of the strong precooling system 5, and the cold side outlet of the first heat exchanger 501 is communicated with the inlet of the medium turbine 503. The cold side inlet of the second heat exchanger 504 is communicated with the outlet of the medium turbine 503, and the cold side outlet of the second heat exchanger 504 is a working medium outlet of the strong precooling system 5. The sixth valve 502 is arranged in parallel with the media turbine 503; a hot-side inlet of the first heat exchanger 501 is an air inlet of the strong precooling system 5, a hot-side outlet of the first heat exchanger 501 is communicated with a hot-side inlet of the second heat exchanger 504, and a hot-side outlet of the second heat exchanger 504 is an air outlet of the strong precooling system 5.
In the basic heat exchange-expansion unit, the number of the first heat exchanger 501, the sixth valve 502, the medium turbine 503, and the second heat exchanger 504 may be different, and may be set as needed.
In this embodiment, the strong precooling system 5 includes a set of basic heat exchange-expansion units as an example.
Further, the invention discloses that the turbine engine exhaust 3 is a turbine engine exhaust nozzle, the aircraft gas turbine engine 2 is a turbofan engine or a turbojet engine, and the ramjet engine 9 comprises a ramjet combustion chamber.
Further, the invention discloses a fuel used by the hypersonic strong precooling turbine-based stamping combined engine, which comprises any one or more working media of aviation kerosene, liquid methane, liquid hydrogen or other low-temperature fuels for aerospace. It should be noted that the fuel used by the hypersonic strong precooling turbo-based ramjet combined engine is liquid methane, which is taken as an example in the embodiment.
The hypersonic strong precooling turbine-based stamping combined engine in the embodiment is mainly used as a power system of a hypersonic aircraft with the working Mach number of 0-3. When the ground flight mach number is 0, the aircraft gas turbine engine 2 is started to work, at the moment, the second gas regulating valve 801, the second boost regulating valve 802 and the first valve 6 are opened, the first gas regulating valve 701, the first boost regulating valve 702 and the sixth valve 502 are closed, the strong precooling system 5 starts to work, incoming flow air enters from the air inlet device 1 and passes through the first heat exchanger 501 and the second heat exchanger 504, the temperature is reduced to 248K, then the incoming flow air enters the air compressor 201 for compression, the compressed air enters the gas combustion chamber 202 to be mixed with fuel and ignited to generate high-temperature gas, then the high-temperature gas enters the gas turbine 203 to be expanded to generate shaft work to drive the air compressor 201, then the high-temperature gas enters the boost combustion chamber 204 to be further increased in temperature and then enters the aircraft gas turbine engine 2 to be expanded and accelerated to generate thrust. In this embodiment, the coolant is liquid methane as engine fuel, and after being increased by a coolant pump, the coolant enters the strong precooling system 5 through the first valve 6, flows through the first heat exchanger 501 to absorb air heat, then rises in temperature, enters the gas turbine 203 to expand and do work, then falls in temperature and pressure, enters the second heat exchanger 504 to absorb air heat, further rises in temperature, and then flows into the gas combustion chamber 202 and the afterburner 204 through the second fuel regulating valve 801 and the second afterburner 802, respectively. When the ground flight Mach number is 0, the air flow of the engine is 14.8kg/s, the liquid methane flow is 0.7kg/s, the net thrust of the engine is 14.4kN, and the specific impulse is 2099 s. The flight Mach number is 0-3, and the strong precooling system 5 continuously works. At the flight Mach number of 3, the temperature of airflow behind the air inlet device 1 is 600K, the temperature of the air is reduced to 460K through the strong precooling system 5, the air flow of the engine is 8.4kg/s, the liquid methane flow is 0.4kg/s, the net thrust of the engine is 6.2kN, and the specific impulse is 1581 s.
Example two
As shown in fig. 2 to 3, in a second embodiment provided by the present invention, the structure of the hypersonic strong precooling turbine-based ramjet combined engine in this embodiment is similar to that of the hypersonic strong precooling turbine-based ramjet combined engine in the first embodiment, and the same parts are not repeated and only the differences are introduced.
In the present embodiment, it is specifically disclosed that the hypersonic strong precooling turbo-based ramjet combined engine further comprises a ramjet engine 9 and a ramjet engine exhaust 10.
The first fuel inlet of the ramjet 9 and the working medium outlet of the coolant system 4 are in conductive connection through a fourth valve 11, that is, the fourth valve 11 controls the on-off of the first fuel inlet of the ramjet 9 and the working medium outlet of the coolant system 4.
The air inlet of the ramjet 9 is in communication with the outlet of the air inlet device 1 and the inlet of the ramjet exhaust 10 is in communication with the outlet of the ramjet 9.
The second fuel inlet of the ramjet 9 is connected with the working medium outlet of the strong precooling system 5 through a fifth valve 12, that is, the fifth valve 12 controls the on-off of the second fuel inlet of the ramjet 9 and the working medium outlet of the strong precooling system 5.
When the strong precooling system 5 does not work and the aircraft gas turbine engine 2 starts to work, the second valve 7 is opened, the first valve 6, the third valve 8, the fourth valve 11 and the fifth valve 12 are all closed, and the working Mach number is within a first preset range value.
When the strong precooling system 5 starts to work, the aircraft gas turbine engine 2 works and the ramjet 9 does not work, the first valve 6 and the third valve 8 are opened, the second valve 7, the fourth valve 11 and the fifth valve 12 are closed, and the working Mach number is within a second preset range value.
When the strong precooling system 5 works, the aircraft gas turbine engine 2 works and the ramjet 9 starts to work, the first valve 6, the third valve 8 and the fifth valve 12 are opened, the second valve 7 and the fourth valve 11 are closed, and the working Mach number reaches a third preset range value.
When the strong precooling system 5 and the aviation gas turbine engine 2 both stop working and the ramjet engine 9 works alone, the first valve 6, the second valve 7, the third valve 8 and the fifth valve 12 are closed, the fourth valve 11 is opened, and the working mach number reaches a fourth preset range value.
Further, the invention specifically discloses that the first preset range value is greater than or equal to 0 and less than or equal to 2.5; the second preset range value is greater than or equal to 0 and less than or equal to 2.5; the third preset range value is greater than or equal to 2.0 and less than or equal to 4.0; the fourth predetermined range value is greater than or equal to 3.0.
It should be noted that the preset range values have overlapping range values, and when the overlapping range values, that is, when one operating state is switched to another operating state, the mach number reaches the next preset range value.
Further, the invention discloses in particular that the ramjet engine 9 comprises a ramjet combustion chamber and that the coolant system 4 comprises a coolant pump.
Further, the invention discloses that the ramjet exhaust device 10 is a ramjet 9 exhaust nozzle.
Further, the invention discloses that the ramjet 9 is a sub-combustion ramjet 9, a hyper-combustion ramjet 9 or a bimodal ramjet 9.
It should be noted that, depending on the specific implementation, the aero gas turbine engine 2 and the ramjet engine 9 may share an air intake passage in the air intake device 1 or may use separate air intake passages independently; the aero gas turbine engine 2 and the ramjet engine 9 may share a combustion chamber or may use separate combustion chambers independently according to specific implementations; the aero gas turbine engine 2 and ramjet engine 9 described above may share exhaust gas passages or may use separate exhaust gas passages alone in the turbine engine exhaust 3 and ramjet exhaust 10 depending on the implementation.
The strong precooling turbine engine is mainly used as a power system of a hypersonic aircraft with the working Mach number of 0-8. When the ground flight mach number is 0, the aircraft gas turbine engine 2 is started up and the ramjet 9 is not operated. At the moment, the second gas regulating valve 801, the second boost regulating valve 802 and the first valve 6 are opened, the first gas regulating valve 701, the first boost regulating valve 702, the sixth valve 502, the fourth valve 11 and the fifth valve 12 are closed, the strong precooling system 5 starts to work, incoming flow air enters from the air inlet device 1 and then passes through the first heat exchanger 501 and the second heat exchanger 504, the temperature is reduced to 248K, then the incoming flow air enters the compressor 201 for compression, the compressed air enters the gas combustion chamber 202 to be mixed with fuel and ignited to generate high-temperature gas, then enters the gas turbine 203 for expansion to generate shaft work to drive the compressor 201, and then enters the boost combustion chamber 204 for further temperature increase and then enters the tail nozzle of the aircraft gas turbine engine 2 for expansion and acceleration to generate thrust; in this embodiment, the coolant is liquid methane as engine fuel, and the coolant is increased by a coolant pump, enters the strong precooling system 5, flows through the first heat exchanger 501 to absorb heat of air, increases in temperature, then enters the medium turbine 503 to expand and work, decreases in temperature and pressure, then enters the second heat exchanger 504 to absorb heat of air, further increases in temperature, and then flows into the gas combustion chamber 202 and the afterburner 204 through the second fuel regulating valve 801 and the second afterburner 802 respectively. When the ground flight Mach number is 0, the air flow of the engine is 14.8kg/s, the liquid methane flow is 0.7kg/s, the net thrust of the engine is 14.4kN, and the specific impulse is 2099 s.
When the flight Mach number reaches 2.5, incoming air enters from the air inlet device 1, the rear part of the incoming air is shunted to the ramjet engine 9, the fourth valve 11 is closed, a ramjet combustion chamber of the ramjet engine 9 is ignited to work, the dual-mode ramjet engine works according to a sub-combustion ramjet mode, the temperature of air flow is increased after the air flow is combusted in the ramjet combustion chamber of the ramjet engine 9, and then the air flow enters a tail nozzle of the ramjet engine to expand and accelerate to generate thrust; the gas turbine 203 engine operates continuously in the above-described mode.
When the flight Mach number reaches 4.0, the first valve 6, the second valve 7, the third valve 8 and the fifth valve 12 are all closed, the fourth valve 11 is opened, the strong precooling system 5 is closed, the aviation gas turbine engine 2 is closed, the hypersonic strong precooling turbine-based ramjet combined engine works independently according to a sub-combustion ramjet mode, and the engine fuel is converted into aviation kerosene; after the flight Mach number is 5.0, the dual-mode ramjet engine is transited from the sub-combustion mode to the ram mode until the operation reaches Mach number 8.0. At Mach number 8.0, the engine thrust is 14kN, and the specific impulse is 780 s.
The invention has the following advantages:
(1) the power system is mainly used for supersonic/hypersonic aircrafts with flight Mach numbers of 0-3 and above. When the flight mach number reaches a second preset range, the strong precooling system 5 starts to work, and air entering the aviation gas turbine engine 2 is cooled in advance at the front section of the aviation gas turbine engine 2, so that the aviation gas turbine engine 2 can normally work in a high mach number range and can generate enough thrust; the invention effectively solves the problem of thrust gap existing in the TBCC engine, realizes the purpose of providing enough thrust for the supersonic/hypersonic aircraft when the TBCC engine flies from Mach 0 to high Mach, and simultaneously can be independently used as a strong precooling turbine engine after deleting the ramjet engine 9, thereby providing high-efficiency and reliable power for the supersonic aircraft.
(2) The strong precooling system 5 consists of a basic heat exchange-expansion unit, can effectively improve the available heat sink of the coolant by expanding the cooling working medium through the medium turbine 503, further improves the cooling capacity of the coolant with unit flow, and reduces the flow of the coolant required by cooling.
(3) The fuel type of the invention comprises single or multiple working media, and the fuel can be selected to be singly used or used in combination of two or more of aviation kerosene, liquid methane, liquid hydrogen and other low-temperature fuels for aviation and aerospace according to requirements, so that the problem that the heat sink carried by an aircraft is insufficient during high-speed flight and the problem that the fuel quality is low due to the fact that low-density fuel is carried in a limited space can be effectively solved.
The terms "first", "second", and the like in the present invention are used for descriptive distinction and have no other special meaning.
The previous description of the disclosed embodiments is provided to enable any person skilled in the art to make or use the present invention. Various modifications to these embodiments will be readily apparent to those skilled in the art, and the generic principles defined herein may be applied to other embodiments without departing from the spirit or scope of the invention. Thus, the present invention is not intended to be limited to the embodiments shown herein but is to be accorded the widest scope consistent with the principles and inventive features disclosed herein.
Claims (8)
1. A hypersonic strong precooling turbine-based ramjet combined engine comprises an air inlet device (1), an aviation gas turbine engine (2) and a turbine engine exhaust device (3), and is characterized by further comprising a coolant system (4) and a strong precooling system (5);
an outlet of the air inlet device (1) is communicated with an air inlet of the strong precooling system (5), an air outlet of the strong precooling system (5) is communicated with an air inlet of the aviation gas turbine engine (2), and an outlet of the aviation gas turbine engine (2) is communicated with an inlet of the turbine engine exhaust device (3);
a working medium outlet of the coolant system (4) is in conductive connection with a working medium inlet of the strong precooling system (5) through a first valve (6), a working medium outlet of the coolant system (4) is in conductive connection with a first fuel inlet of the aviation gas turbine engine (2) through a second valve (7), and a working medium outlet of the strong precooling system (5) is in conductive connection with a second fuel inlet of the aviation gas turbine engine (2) through a third valve (8);
the hypersonic strong precooling turbine-based stamping combined engine further comprises a stamping engine (9) and a stamping engine exhaust device (10);
a first fuel inlet of the ramjet (9) and a working medium outlet of the coolant system (4) are in conductive connection through a fourth valve (11);
an air inlet of the ram engine (9) is communicated with an outlet of the air inlet device (1);
the inlet of the ramjet exhaust (10) is in communication with the outlet of the ramjet (9);
a second fuel inlet of the ramjet (9) is connected with a working medium outlet of the strong precooling system (5) through a fifth valve (12);
when the strong precooling system (5) does not work and the aircraft gas turbine engine (2) starts to work, the second valve (7) is opened, the first valve (6), the third valve (8), the fourth valve (11) and the fifth valve (12) are all closed, and the working Mach number is in a first preset range value;
when the strong precooling system (5) starts to work, the aircraft gas turbine engine (2) works and the ramjet engine (9) does not work, the first valve (6) and the third valve (8) are opened, the second valve (7), the fourth valve (11) and the fifth valve (12) are closed, and the working Mach number is in a second preset range value;
when the strong precooling system (5) works, the aircraft gas turbine engine (2) works and the ramjet engine (9) starts to work, the first valve (6), the third valve (8) and the fifth valve (12) are opened, the second valve (7) and the fourth valve (11) are closed, and the working Mach number reaches a third preset range value;
when the strong precooling system (5) and the aviation gas turbine engine (2) both quit working, and the ramjet engine (9) works independently, the first valve (6), the second valve (7), the third valve (8) and the fifth valve (12) are closed, the fourth valve (11) is opened, and the working Mach number reaches a fourth preset range value.
2. The hypersonic strong precooling turbine-based ramjet combined engine of claim 1, wherein the first preset range value is greater than or equal to 0 and less than or equal to 2.5;
the second preset range value is greater than or equal to 0 and less than or equal to 2.5;
the third preset range value is greater than or equal to 2.0 and less than or equal to 4.0;
the fourth predetermined range value is greater than or equal to 3.0.
3. The hypersonic strong precooling turbine-based ramjet combined engine according to claim 1, characterized in that the strong precooling system (5) comprises at least one set of basic heat exchange-expansion units;
the basic heat exchange-expansion unit comprises a first heat exchanger (501), a sixth valve (502), a medium turbine (503) and a second heat exchanger (504);
a cold side inlet of the first heat exchanger (501) is a working medium inlet of the strong precooling system (5), and a cold side outlet of the first heat exchanger (501) is communicated with an inlet of the medium turbine (503);
a cold side inlet of the second heat exchanger (504) is communicated with an outlet of the medium turbine (503), and a cold side outlet of the second heat exchanger (504) is a working medium outlet of the strong precooling system (5);
the sixth valve (502) is arranged in parallel with the medium turbine (503);
the hot side inlet of the first heat exchanger (501) is an air inlet of the strong precooling system (5), the hot side outlet of the first heat exchanger (501) is communicated with the hot side inlet of the second heat exchanger (504), and the hot side outlet of the second heat exchanger (504) is an air outlet of the strong precooling system (5).
4. The hypersonic strong precooling turbine-based ramjet combined engine of claim 1, wherein the aero gas turbine engine (2) comprises a compressor (201), a gas combustion chamber (202), a gas turbine (203) and an afterburner (204);
an air inlet of the compressor (201) is an air inlet of the aviation gas turbine engine (2), an air outlet of the compressor (201) is communicated with an air inlet of the gas combustion chamber (202), and the compressor (201) is in power connection with the gas turbine (203) through a rotating shaft;
a gas inlet of the gas turbine (203) is in communication with a gas outlet of the gas combustor (202);
the second valve (7) comprises a first gas regulating valve (701) and a first force-adding regulating valve (702), the first gas regulating valve (701) and the first force-adding regulating valve (702) are arranged in parallel, two ends of the first gas regulating valve (701) are respectively communicated with a working medium outlet of the coolant system (4) and a first gas inlet of the gas combustion chamber (202), and two ends of the first force-adding regulating valve (702) are respectively communicated with a working medium outlet of the coolant system (4) and a first gas inlet of the force-adding combustion chamber (204);
the third valve (8) comprises a second fuel regulating valve (801) and a second boost regulating valve (802), the second fuel regulating valve (801) and the second boost regulating valve (802) are arranged in parallel, two ends of the second fuel regulating valve (801) are respectively communicated with a working medium outlet of the strong precooling system (5) and a second fuel gas inlet of the gas combustion chamber (202), and two ends of the second boost regulating valve (802) are respectively communicated with a working medium outlet of the strong precooling system (5) and a second fuel gas inlet of the boost combustion chamber (204);
the outlet of the afterburner (204) is the outlet of the aircraft gas turbine engine (2).
5. The hypersonic strong precooling turbine-based ramjet combined engine of claim 1, characterized in that the ramjet engine (9) comprises a ramjet combustion chamber; and/or
The coolant system (4) comprises a coolant pump.
6. The hypersonic strong precooling turbo-based ramjet combined engine as claimed in claim 1, characterized in that the turbo-engine exhaust means (3) is a turbo-engine exhaust nozzle; and/or
The ramjet exhaust device (10) is a ramjet exhaust nozzle.
7. The hypersonic strong precooling turbo-based ramjet combined engine according to claim 1, characterized in that the aero gas turbine engine (2) is a turbofan engine or a turbojet engine; and/or
The ramjet (9) is a sub-combustion ramjet, a super-combustion ramjet or a bimodal ramjet.
8. The hypersonic strong precooling turbo-based ramjet combined engine according to any one of claims 1 to 7, wherein the fuel used by the hypersonic strong precooling turbo-based ramjet combined engine comprises: aviation kerosene; and/or
Liquid methane; and/or
Liquid hydrogen.
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| CN113006947B (en) * | 2021-03-13 | 2023-01-03 | 西北工业大学 | Precooling engine of dual-fuel system |
| CN114151230B (en) * | 2021-10-20 | 2023-06-30 | 中国航发四川燃气涡轮研究院 | Cooling method and cooling structure of turbine punching combined engine |
| CN114396341A (en) * | 2022-01-05 | 2022-04-26 | 北京航空航天大学 | Intermediate circulation heat dissipation system for combined engine |
| CN114776474B (en) * | 2022-04-13 | 2024-04-12 | 清华大学 | Power system, aircraft and control method thereof |
| CN114658563A (en) * | 2022-04-15 | 2022-06-24 | 中国航空发动机研究院 | Turbine-based combined cycle engine, aircraft and control method |
| CN114704384A (en) * | 2022-04-24 | 2022-07-05 | 中国航发沈阳发动机研究所 | Method and structure for cooling turbine engine cabin of hypersonic vehicle |
| CN115992777B (en) * | 2023-02-15 | 2024-01-30 | 南京航空航天大学 | Dual-fuel precooling variable-cycle engine |
| CN116677498B (en) * | 2023-08-03 | 2023-10-17 | 中国航发四川燃气涡轮研究院 | Novel hypersonic combined engine based on hydrogen energy |
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