CN112879178A - Solid rocket ramjet based on detonation combustion - Google Patents
Solid rocket ramjet based on detonation combustion Download PDFInfo
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- CN112879178A CN112879178A CN202110090134.7A CN202110090134A CN112879178A CN 112879178 A CN112879178 A CN 112879178A CN 202110090134 A CN202110090134 A CN 202110090134A CN 112879178 A CN112879178 A CN 112879178A
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- 238000005474 detonation Methods 0.000 title claims abstract description 114
- 238000002485 combustion reaction Methods 0.000 title claims abstract description 106
- 239000007787 solid Substances 0.000 title claims abstract description 28
- 239000000446 fuel Substances 0.000 claims abstract description 39
- 238000002955 isolation Methods 0.000 claims abstract description 22
- 238000002347 injection Methods 0.000 claims abstract description 15
- 239000007924 injection Substances 0.000 claims abstract description 15
- 239000007789 gas Substances 0.000 claims description 40
- 239000002737 fuel gas Substances 0.000 claims description 18
- 239000003380 propellant Substances 0.000 claims description 7
- 230000000087 stabilizing effect Effects 0.000 claims description 6
- 239000003381 stabilizer Substances 0.000 claims description 5
- 239000000567 combustion gas Substances 0.000 claims description 4
- 230000007423 decrease Effects 0.000 claims 1
- 238000009413 insulation Methods 0.000 claims 1
- 239000000126 substance Substances 0.000 abstract description 2
- 239000004449 solid propellant Substances 0.000 description 12
- 238000006243 chemical reaction Methods 0.000 description 4
- 238000011161 development Methods 0.000 description 4
- 230000009467 reduction Effects 0.000 description 4
- QVGXLLKOCUKJST-UHFFFAOYSA-N atomic oxygen Chemical compound [O] QVGXLLKOCUKJST-UHFFFAOYSA-N 0.000 description 3
- 230000000694 effects Effects 0.000 description 3
- 238000002474 experimental method Methods 0.000 description 3
- 239000001301 oxygen Substances 0.000 description 3
- 229910052760 oxygen Inorganic materials 0.000 description 3
- 239000007788 liquid Substances 0.000 description 2
- 239000000203 mixture Substances 0.000 description 2
- 239000007800 oxidant agent Substances 0.000 description 2
- 239000000376 reactant Substances 0.000 description 2
- 238000011160 research Methods 0.000 description 2
- UFHFLCQGNIYNRP-UHFFFAOYSA-N Hydrogen Chemical compound [H][H] UFHFLCQGNIYNRP-UHFFFAOYSA-N 0.000 description 1
- 238000013461 design Methods 0.000 description 1
- 230000002349 favourable effect Effects 0.000 description 1
- 239000001257 hydrogen Substances 0.000 description 1
- 229910052739 hydrogen Inorganic materials 0.000 description 1
- 238000000034 method Methods 0.000 description 1
- 238000012986 modification Methods 0.000 description 1
- 230000004048 modification Effects 0.000 description 1
- 230000001590 oxidative effect Effects 0.000 description 1
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- 230000029058 respiratory gaseous exchange Effects 0.000 description 1
- 239000000243 solution Substances 0.000 description 1
- 238000012360 testing method Methods 0.000 description 1
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Classifications
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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
- F02K7/18—Composite ram-jet/rocket 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
- F02C3/00—Gas-turbine plants characterised by the use of combustion products as the working fluid
- F02C3/20—Gas-turbine plants characterised by the use of combustion products as the working fluid using a special fuel, oxidant, or dilution fluid to generate the combustion products
- F02C3/26—Gas-turbine plants characterised by the use of combustion products as the working fluid using a special fuel, oxidant, or dilution fluid to generate the combustion products the fuel or oxidant being solid or pulverulent, e.g. in slurry or suspension
- F02C3/28—Gas-turbine plants characterised by the use of combustion products as the working fluid using a special fuel, oxidant, or dilution fluid to generate the combustion products the fuel or oxidant being solid or pulverulent, e.g. in slurry or suspension using a separate gas producer for gasifying the fuel before combustion
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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
- F02K7/02—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 the jet being intermittent, i.e. pulse-jet
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- F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
- F23—COMBUSTION APPARATUS; COMBUSTION PROCESSES
- F23R—GENERATING COMBUSTION PRODUCTS OF HIGH PRESSURE OR HIGH VELOCITY, e.g. GAS-TURBINE COMBUSTION CHAMBERS
- F23R7/00—Intermittent or explosive combustion chambers
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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
- Y02E—REDUCTION OF GREENHOUSE GAS [GHG] EMISSIONS, RELATED TO ENERGY GENERATION, TRANSMISSION OR DISTRIBUTION
- Y02E20/00—Combustion technologies with mitigation potential
- Y02E20/34—Indirect CO2mitigation, i.e. by acting on non CO2directly related matters of the process, e.g. pre-heating or heat recovery
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- Engineering & Computer Science (AREA)
- Chemical & Material Sciences (AREA)
- Combustion & Propulsion (AREA)
- Mechanical Engineering (AREA)
- General Engineering & Computer Science (AREA)
- Fluidized-Bed Combustion And Resonant Combustion (AREA)
- Testing Of Engines (AREA)
Abstract
本发明公开了一种基于爆震燃烧的固体火箭冲压发动机，包括外壳体、内芯体、进气道、隔离段通道、爆震燃烧通道、燃气发生器、稳流环腔、主燃料喷注系统、旋转爆震燃烧室、尾喷管等结构；外壳体内壁和内芯体外壁之间形成进气通道、隔离段通道和爆震燃烧通道；由进气道捕获的空气在隔离段内完成减速增压，之后进入旋转爆震燃烧室；内芯体前部呈中空结构，并在其内布置燃气发生器，贫氧固体推进剂在燃气发生器内燃烧并形成富燃燃气，通过主燃料喷注器进入隔离段通道的扩张段，并与新鲜空气充分掺混后喷入旋转爆震燃烧室的环腔；旋转爆震燃烧室设有点火装置，出口安装尾喷管，将化学能转变为动能并产生推力。本发明具有更高的燃烧效率和工作性能。
The invention discloses a solid rocket ramjet engine based on detonation combustion, comprising an outer casing, an inner core body, an air inlet channel, an isolation section channel, a detonation combustion channel, a gas generator, a steady flow annular cavity, and a main fuel injection System, rotary detonation combustion chamber, tail nozzle and other structures; an intake passage, an isolation section passage and a detonation combustion passage are formed between the inner wall of the outer casing and the outer wall of the inner core; the air captured by the intake passage is completed in the isolation section Decelerate and pressurize, and then enter the rotary detonation combustion chamber; the front part of the inner core body is hollow, and a gas generator is arranged in it. The injector enters the expansion section of the isolation section channel, and is fully mixed with fresh air and injected into the annular cavity of the rotary detonation combustion chamber; the rotary detonation combustion chamber is equipped with an ignition device, and a tail nozzle is installed at the outlet to convert chemical energy into It is kinetic energy and produces thrust. The present invention has higher combustion efficiency and working performance.
Description
Technical Field
The invention belongs to the technical field of aerospace power, and particularly relates to a solid rocket ramjet engine based on detonation combustion.
Background
The chemical energy engine adopted in the current aerospace propulsion system has the internal combustion heat release process belonging to isobaric combustion. Through decades of development, power devices based on an isobaric combustion mode are improved day by day, the propulsion efficiency of the power devices is difficult to improve greatly, and a novel power propulsion device which meets the requirements of high speed and high efficiency is urgently needed to be developed. Compared with isobaric combustion, detonation combustion has the advantages of high energy release rate, high thermodynamic cycle efficiency, self-pressurization, short reaction zone and the like, and is close to constant volume combustion, so that a power system based on detonation combustion has a large promotion space in the aspect of propulsion performance, and has a wide application prospect.
The rotary detonation engine is a novel power propulsion device based on detonation combustion, can realize continuous rotary detonation by only one-time ignition, has the advantages of simple structure, high working frequency, self-pressurization, stable thrust and the like, can stably work under the condition of wide incoming flow speed, and can effectively work under a rocket mode and a stamping mode. In recent years, scientists in various countries have conducted research work on direct connection type experiments and free jet experiments of rotary detonation ramjet engines. Driscoll et al investigated the effect of geometry of an air breathing rotary detonation engine on engine operating mode and operating range. ([1] Driscoll R, Anand V, George A S, et al. investment on RDE operation by geometrical variation of the compressor and non-zle exit [ C ]. In:9th U.S. national compressor Meeting, Cincinnati, Ohio,2015:1-10.) Frolov et al performed wind tunnel experiments of the rotary detonation engine under conditions of different incoming air Mach numbers and total temperatures and different equivalence ratios, and obtained thrust and specific impulse. ([2] Frolov S M, Zvegientsev V I, Ivanov V S, et al. Hydrogen-fused destination ramjet model: Wind channel tests at advanced air stream machine number 5.7and registration temperature 1500K [ J ]. International Journal of Hydrogen Energy,2018,43(15):7515-7524.) the above experimental study further verified the feasibility of impact rotary detonation.
For a traditional solid rocket ramjet, the combustion mode in the afterburning chamber of the ramjet is isobaric combustion, and the space for further improving the propelling performance is limited. In addition, as the detonation combustion is supersonic combustion, the propagation speed of the detonation wave is in the order of kilometers per second, and the requirement on the mixing efficiency of reactants is high. Therefore, research on the rotary knocking engine is currently mostly limited to gaseous fuel, but the gaseous fuel has problems of low density and inconvenience in carrying. And the direct use of solid fuel, the efficient mixing of the solid fuel and air will be a key problem limiting the detonation and stable operation of the rotary detonation engine, and is also an urgent problem to be solved.
Disclosure of Invention
The invention aims to provide a solid rocket ramjet based on detonation combustion, which replaces an isobaric combustion mode in a afterburning chamber of the traditional solid rocket ramjet with a detonation combustion mode similar to isovolumetric combustion so as to improve the combustion efficiency and the working performance of the solid rocket ramjet. Meanwhile, the gas generator is used for pre-burning and gasifying the solid fuel so as to solve the problem of low mixing efficiency of the solid fuel and air.
The technical solution for realizing the purpose of the invention is as follows:
a solid rocket ramjet based on detonation combustion comprises an outer cylinder body which is cylindrical and continuously extends along the axial direction, and an inner core body which is arranged along the length direction of the outer cylinder body, wherein an annular channel formed between the inner wall of the outer cylinder body and the outer wall of the inner core body forms an air inlet channel, an isolation section channel and a detonation combustion channel;
the outer cylinder body and the inner core body extend along the axial direction and form an annular channel for air flow and rich fuel gas to flow, mix and detonate;
the air inlet channel, the isolation section channel and the detonation combustion channel are sequentially arranged along the axial direction and are continuously communicated;
the expansion section of the isolation section channel is connected with the main fuel injection channel;
the inner core body is internally provided with a fuel gas generator, the shell of the fuel gas generator is of an empty cylindrical structure, and oxygen-poor propellant filled in the fuel gas generator completes combustion in the shell of the fuel gas generator and forms fuel-rich gas;
the outlet of the fuel gas generator is connected with a steady flow ring cavity, and the diameter of the cross section of the steady flow ring cavity is larger than that of the outlet of the fuel gas generator;
the rear end of the flow stabilizing ring cavity is connected with a main fuel injector, and the fuel-rich gas passes through the flow stabilizing ring cavity and is injected into the expansion section of the isolation section channel through the main fuel injector and is mixed with air flow;
the rotary detonation combustor is connected with the expansion section of the isolation section channel, and the rich fuel gas is mixed with the air flow and then injected into the rotary detonation combustor;
an ignition device is arranged on the outer wall surface of the rotary detonation combustion chamber and is arranged near the head of the rotary detonation combustion chamber;
and an outlet of the rotary detonation combustion chamber is connected with a tail nozzle, and detonation products flow through the tail nozzle and are discharged out of the engine after being expanded.
Compared with the prior art, the invention has the following remarkable advantages:
the invention relates to a solid rocket ramjet based on detonation combustion, which is mainly divided into a primary isobaric combustion process of an oxygen-poor solid propellant in a gas generator and a secondary detonation combustion process in a rotary detonation combustion chamber, wherein the oxygen-poor solid propellant generates rich combustion gas through primary isobaric combustion, the rich combustion gas is sprayed into an expansion section of a channel of an isolation section through a main fuel injector arranged on the outer wall surface of an inner core body, and is fully mixed with fresh air captured by an air inlet channel and subjected to speed reduction and pressure boost through the isolation section, then the fresh air is sprayed into the rotary detonation combustion chamber, and the fresh air is expanded and discharged through a tail nozzle after undergoing the rotary detonation combustion process to generate thrust. Because the thermal cycle efficiency of detonation combustion is high, the detonation combustion is self-supercharging and the reaction area is short, the rotary detonation combustion chamber is used for replacing an isobaric combustion afterburning chamber of the traditional solid rocket ramjet engine, the working efficiency and the propelling performance of the engine can be effectively improved, the structure of the engine is simplified, the length of the afterburning chamber is shortened, the thrust-weight ratio of the engine is improved, and the engineering development of a novel power propelling system and a combined engine based on detonation combustion is promoted.
Drawings
FIG. 1 is a front cross-sectional view of an embodiment of a solid rocket ramjet engine based on detonation combustion.
FIG. 2 is an enlarged partial cross-sectional view of an embodiment of a solid rocket ramjet engine main fuel injector based on detonation combustion.
10-an inner core body; 101-core body front part; 102-a rear core body; 103-main fuel injectors; 20-outer cylinder; 30-a gas generator; 301-gas generator housing; 302-an oxygen-depleted solid propellant; 40-an ignition device; 50-a rotary detonation combustor; 60-tail nozzle; 90-a steady flow loop cavity; 110-a second insufflating system; 70-an air inlet channel; 80-isolating the segment channel; 100-a detonation combustion channel; 11-a flow of air; 22-a rich combustion gas stream; 33-detonation product stream;
Detailed Description
The invention is further described with reference to the following figures and embodiments.
Referring to fig. 1, the present invention provides a solid rocket ramjet engine based on detonation combustion for engineering development of novel power propulsion systems and combined engines. A solid rocket ramjet based on detonation combustion comprising: the device comprises an outer cylinder body 20 which is cylindrical and continuously extends along the axial direction, an inner core body 10 which is arranged along the length direction of the outer cylinder body 20, an annular channel which is formed between the inner wall of the outer cylinder body 20 and the outer wall of the inner core body 10 forms an air inlet channel 70 for capturing fresh air, an isolation section channel 80 for decelerating and pressurizing and a detonation combustion channel 100 for chemical reaction of fuel and oxidant, and the expansion section of the isolation section channel 80 is connected with a main fuel injector 103. The gas generator 30 is arranged in the inner core body 10, the shell 301 of the gas generator 30 is in an empty cylinder structure, the inner part of the shell is filled with the oxygen-poor solid propellant 302, an ignition device (not shown in the figure) of the gas generator 30 ignites at a head position, and the oxygen-poor solid propellant 302 combusts isobarically in the gas generator 30 and forms the high-temperature and high-pressure rich gas 22 rapidly. The outlet of the gas generator 30 is connected with the flow stabilizing ring cavity 90, the main fuel injector 103 is arranged on the outer wall surface of the inner core body 10, and the fuel-rich gas 22 is rectified by the flow stabilizing ring cavity 90, is injected into the expansion section of the isolation section channel 80 through the main fuel injector 103 and is mixed with the fresh air flow 11. The rotary detonation combustor 50 is connected with the expanding section of the isolating section channel 80, and the rich fuel gas 22 is fully mixed with the fresh air flow 11 and then is injected into the rotary detonation combustor 50. The ignition device 40 is arranged on the outer wall surface of the rotary detonation combustion chamber 50 and is installed near the head of the rotary detonation combustion chamber 50, the outlet of the rotary detonation combustion chamber 50 is connected with the tail nozzle 60, after the rich fuel gas 22 and the fresh air flow 11 are subjected to continuous detonation combustion, the detonation product flow 33 is expanded through the tail nozzle 60 and then is discharged out of the engine.
According to the solid rocket ramjet engine based on detonation combustion, incoming air 11 is captured by the air inlet 70, then the incoming air 11 is subjected to speed reduction and pressure increase through the isolation section channel 80, and is mixed with rich fuel gas 22 in the expansion section of the isolation section channel 80, and then the mixture is injected into the rotary detonation combustion chamber 50 together.
The isolation section passage 80 of the embodiment shown in fig. 1 sequentially comprises an equal straight section, a convergent section and an expansion section from front to back in the flow direction, incoming air 11 captured by the air inlet duct 70 is subjected to speed reduction and pressure boost through the isolation section, and then is subjected to secondary speed reduction and pressure boost through a structure of first convergence and then expansion, so that the incoming flow speed is further reduced, and the mixing efficiency of the rich fuel gas 22 and the fresh air 11 is improved. However, the present invention is not limited to the structure of the isolation section passage 80, and other flow passage structures, such as tapering, converging before diverging, etc., can be implemented.
The gas generator 30 is disposed in a cavity in the front core 101 and is comprised of a gas generator housing 301 and an oxygen-depleted solid propellant 302. The gas generator 30 in this embodiment is in an empty-barrel configuration, and the propellant is an oxygen-lean solid propellant, but the gas generator 30 and the propellant 302 in the present invention are not limited to the configuration and kind shown in this embodiment, and the propellant is not limited to solid matter, and can be a liquid or gaseous propellant, but the proportions of fuel and oxidizer in the propellant are both rich and lean in oxygen.
The oxygen-poor solid propellant in the gas generator 30 in the embodiment is subjected to isobaric combustion to generate high-temperature and high-pressure rich fuel gas 22, the gas generator 30 is connected with a steady flow ring cavity 90 at the rear, the diameter of the ring cavity is larger than the diameter of an outlet of the gas generator, the rich fuel gas flows through the steady flow ring cavity 90 from the gas generator, is rectified by the steady flow ring cavity 90, is sprayed into an expansion section of a channel of an isolation section through a main fuel injector 103, is mixed with fresh air, and then enters a combustion chamber. The main fuel injectors in the embodiment of the present invention shown in fig. 1 are circumferentially distributed small holes, and the diameter, the distance between the holes, the number of the small holes, and the injection angle of the small holes all affect the mixing state of the reactants entering the rotary detonation combustor 50. Wherein, the diameter and the number of the small holes are designed according to the incoming flow mass flow rate, the fuel-rich gas component and the variation range of the reaction equivalence ratio. The hole spacing is influenced by the circumferential size of the engine and the number of the small holes, and in order to ensure the mixing efficiency of the rich fuel gas 22 and the incoming flow air 11, the ratio of the hole spacing to the diameter of the small holes is less than or equal to 10. The injection angle of the fuel holes, namely the included angle between the injection direction of the fuel-rich gas 22 and the direction of the incoming air 11, is 30-90 degrees. The orifice configuration and injection angle in the example shown in fig. 2 is 60 deg.. In particular, the main fuel injector structure of the present invention is not limited to the structure shown in fig. 2.
The design of the flow stabilizer ring cavity 90 is related to the outlet geometric parameters of the gas generator 30, and is generally designed according to the outlet area of the gas generator 30, if the sectional area of the flow stabilizer ring cavity 90 is too small, the rectification effect is poor, and the axial speed of the gas generator 30 is too large, which is not favorable for the injection of the fuel-rich gas 22; if the cross-sectional area of the annular cavity is too large, the volume of the annular cavity is too large, the pressure drop of the fuel-rich gas 22 is too large, the effective injection and penetration depth of the fuel gas are affected, and the mixing effect of the fuel gas is further affected.
The rotary knocking combustor 50 is connected to the expanding section of the isolating section passage 80, and is an annular combustion passage composed of the inner wall surface of the outer cylinder 20 and the outer wall surface of the core body rear portion 102. The configuration of the rotary knocking combustion chamber 50 shown in the present embodiment is a circular ring structure. The ignition device 40 is arranged on the outer wall surface of the rotary detonation combustion chamber 50, the ignition device 40 is arranged near the head of the rotary detonation combustion chamber 50 in the embodiment, the outlet of the combustion chamber is connected with the tail nozzle 60, the detonation product flow 33 expands through the tail nozzle 60, the heat energy is converted into kinetic energy and then is discharged out of the engine, and thrust is generated.
The gas generator 30 of the embodiment shown in fig. 1 is a solid rocket engine, but the gas generator 30 of the present invention is not limited to the structure of the embodiment shown, and other types of fuel or structures of gas generators, such as liquid rocket engines, etc., can be implemented in combination with the rotary detonation combustor 50.
The rotary detonation combustor 50 of the embodiment shown in fig. 1 is in a circular ring structure, and the flow channel profile is in an equal straight shape, but the rotary detonation combustor 50 in the present invention is not limited to the structure shown in the embodiment, the rotary detonation combustor 50 may be in a cylindrical shape without an inner cylinder, a disc shape, and other combustor structures, and the flow channel profile may be in a divergent shape, a convergent shape, and any other profile.
The configuration and exit area of the aft nozzle 60 of the rotary detonation combustor 50, which may be a laval nozzle (as shown in fig. 1), a convergent nozzle, or a divergent nozzle, also affects the operating characteristics of the engine. The inner and outer wall surfaces of the detonation combustion channel 100 of the equal straight type rotary detonation combustion chamber 50 are parallel, and the cross-sectional area of the channel is not changed along the axial direction. In the divergent rotary detonation combustor 50, the outer wall of the detonation combustion channel 100 is expanded towards the inside of the outer cylinder 20, the inner wall of the detonation combustion channel is expanded towards the inside of the inner core body rear part 102, and the cross-sectional area of the channel is gradually increased along the axial direction. In the tapered rotary knocking combustion chamber 50, the inner and outer wall surfaces of the knocking combustion channel 100 simultaneously expand into the knocking combustion channel 100, and the channel cross-sectional area is gradually reduced along the axial direction.
The ignition device 40 for detonating the rotary detonation combustor 50 is mounted near the head of the rotary detonation combustor 50, but not limited to, the ignition device 40 in the embodiment shown in fig. 1 is mounted near the head of the rotary detonation combustor 50, the mounting position can be arranged along the axial position of the rotary detonation combustor 50, the mounting direction can be vertical to the wall surface of the outer cylinder 20 or tangential to the outer wall surface of the outer cylinder 20, and the igniter can be flush mounting, embedded mounting and protruding mounting. The requirement for the discharge energy of the ignition device is determined according to the components of the fuel-rich gas, and the ignition devices such as a high-energy spark plug, a pre-detonation pipe, a thermal jet pipe and the like can be selected.
As shown in fig. 1, a second injection system 110, consisting of an annular cavity and injectors, is provided within the inner core back 102, the second injection system 110 communicating with the detonation combustion channel 100 and injecting fuel into the rotating detonation combustor 50. When the main fuel injector 103 alone is insufficient to generate the spin detonation wave, or when the fuel is packed in a split-type manner to improve combustion efficiency, the fuel may be injected into the spin detonation combustor 50 through the second injection system 110, but the flow rate of the fuel injected through the second injection system 110 is generally no more than 20% of the total flow rate of the fuel.
The solid rocket ramjet based on detonation combustion uses the rotary detonation combustion chamber to replace an isobaric combustion afterburning chamber of the traditional solid rocket ramjet, can effectively improve the working efficiency and the propulsion performance of the engine, simplifies the structure of the engine, shortens the length of the afterburning chamber, improves the thrust-weight ratio of the engine, and promotes the engineering development of a novel power propulsion system and a combined engine based on detonation combustion.
The above description is only for the purpose of illustrating the preferred embodiments of the present invention and is not to be construed as limiting the invention, and any modifications, equivalents or improvements made within the spirit and principle of the present invention should be included in the scope of the present invention.
Claims (9)
1. A solid rocket ramjet based on detonation combustion is characterized by comprising an outer cylinder (20) which is cylindrical and continuously extends along the axial direction, and an inner core (10) which is arranged along the length direction of the outer cylinder (20), wherein an annular channel formed between the inner wall of the outer cylinder (20) and the outer wall of the inner core (10) forms an air inlet channel (70), an isolation section channel (80) and a detonation combustion channel (100);
the outer cylinder (20) and the inner core body (10) extend along the axial direction, and the outer cylinder and the inner core body are separated to form an annular channel for flowing, mixing and knocking combustion of the air flow (11) and the rich combustion gas (22);
the air inlet channel (70), the isolation section channel (80) and the detonation combustion channel (100) are sequentially arranged along the axial direction and are communicated with each other continuously;
the expanding section of the isolating section channel (80) is connected with the main fuel injection channel;
a gas generator (30) is arranged in the inner core body (10), a shell (301) of the gas generator (30) is of a hollow cylindrical structure, and oxygen-poor propellants (302) filled in the shell of the gas generator complete combustion and form rich fuel gas (22);
the outlet of the gas generator (30) is connected with a steady flow ring cavity (90), and the section diameter of the steady flow ring cavity (90) is larger than the outlet diameter of the gas generator (30);
the rear end of the flow stabilizing ring cavity (90) is connected with a main fuel injector (103), and the fuel-rich gas (22) passes through the flow stabilizing ring cavity (90) and is injected into the expansion section of the isolating section channel (80) through the main fuel injector (103) and is mixed with the air flow (11);
the rotary detonation combustor (50) is connected with the expanding section of the isolating section channel (80), and the fuel-rich gas (22) is mixed with the air flow (11) and then is injected into the rotary detonation combustor (50);
an ignition device (40) is arranged on the outer wall surface of the rotary detonation combustion chamber (50), and the ignition device is installed near the head of the rotary detonation combustion chamber (50);
the outlet of the rotary detonation combustion chamber (50) is connected with a tail pipe (60), and the detonation product flow (33) is expanded through the tail pipe (60) and then is discharged out of the engine.
2. The detonation combustion-based solid rocket ramjet engine according to claim 1, characterized in that said inner core (10) is divided into two parts, the inner core front part (101) and said outer barrel (20) constituting an air inlet channel (70) and an insulation section channel (80), and the inner core rear part (102) and said outer barrel (20) constituting a detonation combustion channel (100).
3. The detonation combustion-based solid rocket ramjet engine according to claim 2, characterized in that said flow stabilizer ring cavity (90) exit plane divides said inner core (10), said flow stabilizer ring cavity (90) being preceded by said inner core front portion (101) and said flow stabilizer ring cavity (90) being followed by said inner core rear portion (102).
4. The detonation combustion-based solid rocket ramjet engine according to claim 1, characterized in that the igniter of said gas generator (30) is located at the head position of the core (10).
5. The detonation combustion-based solid-rocket ramjet engine according to claim 1, characterized in that the configuration of the rotary detonation combustion chamber (50) is of the iso-straight type, the divergent type or the convergent type.
6. The detonation combustion-based solid rocket ramjet engine according to claim 1, characterized in that said rotary detonation combustion chamber (50) is of a straight configuration with parallel inner and outer wall surfaces of the detonation combustion channel (100) and a channel cross-sectional area that is axially constant.
7. The detonation combustion-based solid rocket ramjet engine according to claim 1, characterized in that said rotary detonation combustion chamber (50) is of a divergent configuration, the outer wall of said detonation combustion channel (100) expanding inwardly towards said outer cylinder (20), the inner wall expanding inwardly towards said inner core rear portion (102), the channel cross-sectional area increasing progressively in the axial direction.
8. The detonation combustion-based solid rocket ramjet engine according to claim 1, characterized in that said rotary detonation combustion chamber (50) is tapered in configuration, the inner and outer wall surfaces of said detonation combustion channel (100) expand simultaneously into said detonation combustion channel (100), and the channel cross-sectional area decreases gradually in the axial direction.
9. The detonation combustion-based solid rocket ramjet engine according to claim 1, characterized in that a second injection system (110) is provided inside said core aft portion (102), said second injection system (110) communicating with said detonation combustion channel (100) for injecting fuel into said detonation combustion channel (100).
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