CN118997950A - Rotary detonation rocket engine with staged combustion - Google Patents
Rotary detonation rocket engine with staged combustion Download PDFInfo
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- CN118997950A CN118997950A CN202411106481.4A CN202411106481A CN118997950A CN 118997950 A CN118997950 A CN 118997950A CN 202411106481 A CN202411106481 A CN 202411106481A CN 118997950 A CN118997950 A CN 118997950A
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- 238000002485 combustion reaction Methods 0.000 title claims abstract description 140
- 238000005474 detonation Methods 0.000 title claims abstract description 110
- 239000000446 fuel Substances 0.000 claims abstract description 156
- 238000002347 injection Methods 0.000 claims abstract description 79
- 239000007924 injection Substances 0.000 claims abstract description 79
- QVGXLLKOCUKJST-UHFFFAOYSA-N atomic oxygen Chemical compound [O] QVGXLLKOCUKJST-UHFFFAOYSA-N 0.000 claims abstract description 77
- 239000001301 oxygen Substances 0.000 claims abstract description 77
- 229910052760 oxygen Inorganic materials 0.000 claims abstract description 77
- 239000007800 oxidant agent Substances 0.000 claims abstract description 53
- 230000001590 oxidative effect Effects 0.000 claims abstract description 51
- 239000007789 gas Substances 0.000 claims abstract description 37
- 239000007788 liquid Substances 0.000 claims abstract description 26
- 230000007704 transition Effects 0.000 claims abstract description 16
- 230000000694 effects Effects 0.000 claims abstract description 13
- MYMOFIZGZYHOMD-UHFFFAOYSA-N Dioxygen Chemical compound O=O MYMOFIZGZYHOMD-UHFFFAOYSA-N 0.000 claims description 8
- 230000004323 axial length Effects 0.000 claims description 6
- 230000008602 contraction Effects 0.000 claims description 6
- 238000001816 cooling Methods 0.000 claims description 6
- 230000001105 regulatory effect Effects 0.000 claims description 4
- 239000004215 Carbon black (E152) Substances 0.000 abstract description 9
- 229930195733 hydrocarbon Natural products 0.000 abstract description 9
- 150000002430 hydrocarbons Chemical class 0.000 abstract description 9
- 238000005336 cracking Methods 0.000 abstract description 3
- 239000003380 propellant Substances 0.000 description 7
- 238000002156 mixing Methods 0.000 description 5
- 230000008901 benefit Effects 0.000 description 4
- 239000000945 filler Substances 0.000 description 4
- 238000000034 method Methods 0.000 description 4
- 238000004200 deflagration Methods 0.000 description 3
- 238000001704 evaporation Methods 0.000 description 3
- 230000008569 process Effects 0.000 description 3
- 238000000889 atomisation Methods 0.000 description 2
- 238000006243 chemical reaction Methods 0.000 description 2
- 230000008020 evaporation Effects 0.000 description 2
- 230000000977 initiatory effect Effects 0.000 description 2
- 239000000243 solution Substances 0.000 description 2
- UGFAIRIUMAVXCW-UHFFFAOYSA-N Carbon monoxide Chemical class [O+]#[C-] UGFAIRIUMAVXCW-UHFFFAOYSA-N 0.000 description 1
- 230000009286 beneficial effect Effects 0.000 description 1
- 229910002090 carbon oxide Inorganic materials 0.000 description 1
- 230000007547 defect Effects 0.000 description 1
- 230000018109 developmental process Effects 0.000 description 1
- 238000010438 heat treatment Methods 0.000 description 1
- 239000003350 kerosene Substances 0.000 description 1
- 239000007791 liquid phase Substances 0.000 description 1
- 230000007246 mechanism Effects 0.000 description 1
- 230000008520 organization Effects 0.000 description 1
- 230000035515 penetration Effects 0.000 description 1
- 239000012466 permeate Substances 0.000 description 1
- 239000012071 phase Substances 0.000 description 1
- 238000004321 preservation Methods 0.000 description 1
- 230000009257 reactivity Effects 0.000 description 1
- 239000007921 spray Substances 0.000 description 1
- 239000000126 substance Substances 0.000 description 1
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
- F02K9/00—Rocket-engine plants, i.e. plants carrying both fuel and oxidant therefor; Control thereof
- F02K9/42—Rocket-engine plants, i.e. plants carrying both fuel and oxidant therefor; Control thereof using liquid or gaseous propellants
-
- 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
- F02K9/42—Rocket-engine plants, i.e. plants carrying both fuel and oxidant therefor; Control thereof using liquid or gaseous propellants
- F02K9/44—Feeding propellants
- F02K9/52—Injectors
-
- 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
- F02K9/42—Rocket-engine plants, i.e. plants carrying both fuel and oxidant therefor; Control thereof using liquid or gaseous propellants
- F02K9/44—Feeding propellants
- F02K9/56—Control
-
- 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
- F02K9/42—Rocket-engine plants, i.e. plants carrying both fuel and oxidant therefor; Control thereof using liquid or gaseous propellants
- F02K9/60—Constructional parts; Details not otherwise provided for
- F02K9/62—Combustion or thrust chambers
-
- 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
- F02K9/42—Rocket-engine plants, i.e. plants carrying both fuel and oxidant therefor; Control thereof using liquid or gaseous propellants
- F02K9/60—Constructional parts; Details not otherwise provided for
- F02K9/62—Combustion or thrust chambers
- F02K9/66—Combustion or thrust chambers of the rotary type
-
- 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)
Abstract
The invention discloses a staged combustion rotary detonation rocket engine, which comprises a precombustion injection section, an oxygen-enriched precombustion combustion chamber, a combustion transition section and a rotary detonation combustion chamber which are coaxially arranged in sequence along the axial direction; the oxygen-enriched precombustion combustion chamber is used for carrying out oxygen-enriched combustion on the oxidant and the combustion A to form high-temperature high-pressure oxygen-enriched gas; the combustion transition section is provided with an oxygen-enriched flow passage and is used for injecting oxygen-enriched gas formed by the oxygen-enriched precombustion combustion chamber into the rotary detonation combustion chamber; the rotary detonation combustor is provided with an annular detonation chamber and a fuel flow channel; the fuel runner is used for filling, preheating and injecting fuel B into the annular knocking cavity; the annular detonation chamber is capable of rotary detonation combustion of oxygen-enriched gas and fuel B. The invention adopts precombustion staged combustion, can promote fuel cracking, improve the activity of fuel, optimize the combustion process, realize the rapid detonation of the rotary detonation of liquid hydrocarbon fuel and the stable propagation of rotary detonation waves, and can also realize the mass flow rate adjustment in a large range and the thrust adjustment.
Description
Technical Field
The invention relates to a rocket engine, in particular to a staged combustion rotary detonation rocket engine.
Background
With the continuous development of aerospace industry, the performance requirements of rocket engines are continuously improved. The current rocket engines all adopt the deflagration combustion mode, and under the continuous research of the last decades of the world, the performance of the rocket engines based on the deflagration combustion mode is nearly exerted to the greatest extent. As another novel combustion mode, knocking combustion has the advantages of low entropy, high thermodynamic cycle efficiency and the like. Detonation combustion is different from deflagration approximately equal-pressure combustion, has the characteristic of approximately equal-volume combustion, and therefore potential work of gas expansion loss can be avoided, and more effective work can be output by higher thermodynamic cycle efficiency. The rocket engine adopts a detonation combustion mode, so that the specific impact performance of the engine is greatly improved. The rocket engine adopting detonation combustion has the characteristics of simple and compact structure, self-pressurization, large thrust-weight ratio, high heat release rate and the like.
In theory, the liquid oxidant and the fuel are the trend of the detonation rocket engine, and the form of detonation combustion with high thermodynamic cycle efficiency is adopted, so that the fuel and the oxidant are fully combusted, the emission of carbon oxides is reduced, and the pollution to the environment is reduced. Meanwhile, the normal-temperature liquid hydrocarbon fuel and the storage and transportation cost thereof are low, and the normal-temperature liquid hydrocarbon fuel and the liquid oxidant also have the advantage of volume density in the space limited by the spacecraft.
The propellant used in current rotary knock research is mostly gaseous, and when it is also a gaseous oxidizer, the phase of the fuel directly affects the conditions, form and engine design of the combustion structure. The gaseous fuel has higher activity, faster blending and reaction rate, easier detonation initiation, easy realization of stable detonation wave propagation, is generally used for researching detonation combustion mechanism and theoretical working condition, and is applied to the carrying amount of the gaseous propellant in engine design. The normal temperature liquid hydrocarbon fuel, such as kerosene, has many advantages over gaseous fuel, generally has high energy density, is convenient to transport and store, and does not need special equipment and conditions for use and storage. The low-temperature liquid fuel needs pre-cooling and filling, and is not easy to store for a long time due to special heat preservation measures.
From the spray combustion organization process, the normal temperature liquid hydrocarbon fuel or oxidant needs to undergo the processes of crushing, atomizing, evaporating, blending and the like before being combusted. Due to the influences of the structural design of the injector and other configuration or flow working condition factors, the crushing effect and the mixing effect of fuel or oxidant liquid drops in the detonation combustion chamber are poor, the liquid drops can permeate out from the intersection point of the detonation wave and the shear layer to form local hot spots, and the stable propagation of the detonation wave is seriously influenced. In low temperature propellant liquid rocket detonation engine designs, the low temperature liquid fuel and oxidizer also undergo an atomization process prior to combustion. In addition, due to poor chemical reactivity caused by low temperature, detonation energy required by detonation combustion of a low-temperature liquid propellant tissue is higher than that required by normal-temperature liquid propellant, detonation is very difficult, and stable self-sustaining of detonation waves is not easy to maintain. In addition, direct tissue combustion of liquid propellants may suffer from low combustion efficiency, low energy conversion utilization, and the like.
The rotary detonation combustion chamber of the normal-temperature liquid hydrocarbon fuel which is studied more at present only realizes rotary detonation combustion in the mass flow of the oxidant and the fuel with small fluctuation range, and when the mass flow of the oxidant and the fuel is changed, detonation waves become unstable and even are decoupled and quenched. It is difficult to achieve a wide range of mass flow adjustment, i.e., a wide range of thrust adjustment, in the current rotary detonation combustors using normal temperature liquid hydrocarbon fuels.
Disclosure of Invention
The technical problem to be solved by the invention is to provide a staged combustion rotary detonation rocket engine design aiming at the defects of the prior art, wherein the staged combustion rotary detonation rocket engine adopts a pre-combustion staged combustion mode, can combust all liquid oxygen and part of fuel to generate high-temperature and high-pressure oxygen-enriched gas in an oxygen-enriched combustion chamber, actively cool the combustion chamber and preheat the fuel while cooling the combustion chamber, promote fuel cracking, improve the activity of the fuel, optimize the combustion process and realize stable propagation of rotary detonation waves. The problems that the current liquid hydrocarbon fuel is difficult to detonate by rotating detonation, the breaking effect of fuel droplets is poor, partial fuel is combusted in a slow combustion mode, detonation wave propagation is unstable, mass flow is regulated in a large range to detonate, and the equivalent ratio range of stable propagation of detonation wave is small are effectively solved.
In order to solve the technical problems, the invention adopts the following technical scheme:
A staged combustion rotary detonation rocket engine comprises a precombustion injection section, an oxygen-enriched precombustion combustion chamber, a combustion transition section and a rotary detonation combustion chamber which are coaxially arranged in sequence along the axial direction.
The precombustion injection section is used for injecting liquid oxidant and fuel A into the oxygen-enriched precombustion chamber; wherein the mass flow rate of the liquid oxidant injection is greater than the mass flow rate of the fuel A injection.
The oxygen-enriched precombustion chamber is used for carrying out oxygen-enriched combustion on the oxidant and the combustion A to form high-temperature high-pressure oxygen-enriched gas.
The combustion transition section is provided with an oxygen-enriched flow passage and is used for injecting oxygen-enriched gas formed by the oxygen-enriched precombustion combustion chamber into the rotary detonation combustion chamber.
The rotary detonation combustor has an annular detonation chamber and a fuel flow path.
The fuel flow passage is used for filling, preheating and injecting fuel B into the annular knocking cavity.
The annular detonation chamber can perform rotary detonation combustion on oxygen-enriched gas injected by the combustion transition section and fuel B injected by the fuel flow passage.
The oxygen-enriched flow passage comprises an annular flow passage, a shrinkage flow passage and an expansion flow passage which are sequentially arranged along the axial direction; the junction of the contracted flow channel and the expanded flow channel forms a throat part; the flow velocity of the oxygen-enriched gas forms sound velocity at the throat, reaches supersonic velocity in the expansion flow channel, and is sprayed to the annular detonation cavity at supersonic velocity.
The center of the combustion transition section is provided with a center connector; the central connector comprises a cylindrical section and a circular table section which are coaxially arranged in sequence along the axial direction.
The tail end of the round table section is provided with a chamfer to form an inverted round table part.
The cylindrical section and the inner wall of the oxygen-enriched precombustion combustion chamber form the annular flow passage.
The circular table section and the inner wall of the oxygen-enriched precombustion combustion chamber form the contraction flow passage.
The inverted circular truncated cone portion and the inner wall of the oxygen-enriched precombustion combustion chamber form the expansion flow passage.
The included angle between the generatrix of the circular bench section and the center axis of the circular bench section is smaller than 30 degrees, and the included angle between the generatrix of the inverted circular bench section and the center axis of the inverted circular bench section is larger than 45 degrees.
The axial length of the expanding runner is not more than one half of the axial length of the contracting runner.
The center of the rotary detonation combustion chamber is provided with a center cone; the central cone comprises a central large cylindrical section and a cone section which are sequentially arranged along the axial direction.
The outer diameter of the central large cylindrical section is larger than that of the large circular section, but not larger than the minimum outer diameter of the inverted circular table part; the annular detonation chamber is formed between the central large cylindrical segment and the inner wall of the rotary detonation combustion chamber.
The cylindrical section, the circular table section, the central large cylindrical section and the cone section are sequentially and coaxially sealed and detachably connected.
The fuel runner includes a fuel filler inlet, a fuel preheating runner, and a fuel injection hole.
The fuel preheating flow passage is axially distributed and is annular.
The fuel filler port is used for filling fuel into the fuel preheating flow passage.
The fuel injection holes are uniformly distributed along the circumferential direction of the front end of the fuel preheating flow channel, and each fuel injection hole is vertically distributed.
The precombustion injection section includes an oxidant injection plate and a precombustion fuel injection plate.
The precombustion fuel injection plate is coaxially and hermetically arranged on the front end face of the oxygen-enriched precombustion chamber.
The oxidant injection plate is coaxially and hermetically covered on the outer end surface of the pre-combustion fuel injection plate, and an oxidant gas collection cavity is formed between the oxidant injection plate and the pre-combustion fuel injection plate.
The center of the oxidant injection plate is provided with an oxidant injection port, and the oxidant injection port penetrates through the center of the precombustion fuel injection plate.
A plurality of precombustion fuel injection holes are distributed on the outer side of the precombustion fuel injection plate along the circumferential direction; each injection hole is communicated with the oxidant gas collection cavity and the oxygen-enriched precombustion combustion chamber, and the straight line of each injection hole is intersected with the straight line of the oxidant injection hole.
The fuel A and the fuel B are the same fuel, and the mass flow of the fuel B is determined according to the fuel specific impulse and the designed thrust range; the mass flow rate of the fuel A is not more than 1/4 of the mass flow rate of the fuel B; under the condition that the rotary detonation wave is not decoupled, the mass flow of the total fuel and the liquid oxygen is regulated simultaneously through the flow control valve, so that the thrust of the rotary detonation engine is controlled.
Oxygen-enriched gas entering the detonation combustion chamber is pre-burnt, so that the temperature is high; the fuel also has a higher temperature before flowing through the active cooling flow passage into the combustion chamber; therefore, the fuel and the oxidant have higher activity, and can realize rotary knocking when the equivalent ratio is 0.7-1.3, so that the range of the equivalent ratio for realizing stable propagation of the detonation wave is widened.
The invention has the following beneficial effects:
1. The invention can adopt the knocking combustion mode, has the advantages of simple structure, large thrust-weight ratio, high thermodynamic cycle efficiency and the like, optimizes the combustion process, and greatly improves the combustion efficiency and the energy utilization rate of the liquid-phase fuel.
2. The oxygen-enriched precombustion combustion chamber can realize oxygen-enriched staged circulating combustion, and partial liquid fuel and all liquid oxygen are combusted in the oxygen-enriched precombustion combustion chamber to generate a large amount of high-temperature and high-pressure oxygen-enriched gas, so that the liquid oxygen is quickly converted into gaseous oxygen, the activity of an oxidant is greatly improved, the rapid mixing and evaporation with the fuel can be realized in the detonation combustion chamber, the rotary detonation can be quickly initiated, and the problem of difficult initiation of a low-temperature propellant is solved.
3. The fuel flow passage of the detonation combustion chamber can cool the detonation combustion chamber and heat the fuel B, so that the fuel activity is improved, and the cracking of liquid hydrocarbon fuel and the like is accelerated.
4. The flow speed of the oxygen-enriched gas can be greatly improved through the contraction flow passage, the sound speed is achieved at the throat part of the contraction flow passage, and the supersonic speed is achieved at the flow speed of the oxygen-enriched gas in the expansion flow passage. The supersonic oxygen-enriched gas can greatly improve the breaking and atomizing effects of the fuel, accelerate the atomization and evaporation of the fuel and improve the mixing effect of the fuel and the oxygen-enriched gas. More fuel can be combusted in a knocking mode, the fuel combusted in a slow combustion mode is reduced, the penetration of fuel droplets from the intersection point of the detonation wave and the shear layer is avoided, local hot spots are formed, and stable propagation of the rotary detonation wave is realized.
5. Because the oxygen-enriched gas generated by the precombustion combustion chamber is used as the oxidant for detonation combustion, the invention can obviously improve the range of liquid oxygen and fuel mass flow for realizing stable propagation of detonation waves under the condition of stoichiometric ratio. Therefore, the thrust of the rotary detonation engine can be controlled by adjusting the mass flow of fuel and liquid oxygen under the condition that the rotary detonation wave is not decoupled. In addition, the range of equivalent ratio for realizing stable propagation of detonation waves can be expanded.
Drawings
FIG. 1 shows a three-dimensional schematic of a staged combustion rotary detonation rocket engine in accordance with the present invention.
FIG. 2 shows a longitudinal cross-sectional view of a staged combustion rotary detonation rocket engine in accordance with the present invention.
Fig. 3 shows an enlarged schematic view of the encircled area i of fig. 2.
The method comprises the following steps:
10. a precombustion injection section;
11. An oxidant injection plate; 111, an oxidant injection port;
12. a pre-combustion fuel injection plate; 121 pre-burning fuel injection holes;
13. an oxidant gas collection chamber;
20. An oxygen-enriched precombustion chamber; 21. a spark plug;
30. a fuel transition section;
31. A central connection body; 311. a cylindrical section; 312. a circular table section; 313. a reverse round platform part;
32. an annular flow passage; 33. contracting the flow channel; 34. expanding the flow channel; 35. a throat;
40. A rotary detonation combustor;
41. a central cone; 411. a central large cylindrical section; 412. a cone section;
42. An annular detonation chamber;
43. a fuel flow passage; 431. a fueling port; 432. a fuel preheating flow passage; 433. and a fuel injection hole.
Detailed Description
The invention will be described in further detail with reference to the accompanying drawings and specific preferred embodiments.
In the description of the present invention, it should be understood that the terms "left", "right", "upper", "lower", etc. indicate orientations or positional relationships based on the orientations or positional relationships shown in the drawings, are merely for convenience in describing the present invention and simplifying the description, and do not indicate or imply that the apparatus or element being referred to must have a specific orientation, be configured and operated in a specific orientation, and "first", "second", etc. do not indicate the importance of the components, and thus are not to be construed as limiting the present invention. The specific dimensions adopted in the present embodiment are only for illustrating the technical solution, and do not limit the protection scope of the present invention.
As shown in fig. 1 and 2, a staged combustion rotary detonation rocket engine includes a precombustion injection section 10, an oxygen-enriched precombustion chamber 20, a combustion transition section 30 and a rotary detonation combustion chamber 40 coaxially arranged in sequence in an axial direction.
In this embodiment, the pre-combustion injection section 10, the oxygen-enriched pre-combustion chamber 20, the combustion transition section 30 and the rotary detonation combustion chamber 40 are preferably detachably and sealingly connected by bolts and nuts.
The precombustion injection section is used for injecting liquid oxidant and fuel A into the oxygen-enriched precombustion chamber; wherein the mass flow rate of the liquid oxidant injection is greater than the mass flow rate of the fuel a injection.
The pre-combustion injection section preferably includes an oxidant injection plate 11 and a pre-combustion fuel injection plate 12.
The precombustion fuel injection plate is preferably coaxially and hermetically arranged on the front end surface of the oxygen-enriched precombustion chamber.
The oxidant injection plate is preferably coaxially and hermetically covered on the outer end surface of the precombustion fuel injection plate, and an oxidant gas collection cavity 13 is formed between the oxidant injection plate and the precombustion fuel injection plate.
The center of the oxidant injection plate is provided with an oxidant injection port 111, and it penetrates the center of the pre-combustion fuel injection plate, i.e. on the central axis.
A plurality of precombustion fuel injection holes 121 are distributed on the outer side of the precombustion fuel injection plate along the circumferential direction; each injection hole is communicated with the oxidant gas collection cavity and the oxygen-enriched precombustion combustion chamber, and the straight line of each injection hole is intersected with the straight line of the oxidant injection hole.
The spark plug 21 is arranged in the oxygen-enriched precombustion combustion chamber, so that the oxidant and the combustion A can be ignited, oxygen-enriched combustion is realized, and high-temperature and high-pressure oxygen-enriched gas is formed.
The center of the combustion transition section is provided with a center connector 31; the central connecting body preferably comprises a cylindrical section 311 and a truncated cone section 312 which are coaxially arranged in sequence along the axial direction.
Further, the tail end of the rounded bench section is preferably provided with a chamfer forming a rounded bench section 313.
An oxygen-enriched flow passage is formed between the central connector and the inner wall of the combustion transition section, and preferably comprises an annular flow passage 32, a contracted flow passage 33 and an expanded flow passage 34 which are sequentially arranged along the axial direction. Wherein the junction of the converging flow path and the diverging flow path forms a throat 35.
The cylindrical section and the inner wall of the oxygen-enriched precombustion combustion chamber form an annular flow passage.
The circular table section and the inner wall of the oxygen-enriched precombustion combustion chamber form a contracted flow passage.
The inverted circular truncated cone part and the inner wall of the oxygen-enriched precombustion combustion chamber form an expansion flow passage.
Further, the included angle between the generatrix of the circular table section and the center axis of the circular table section is smaller than 30 degrees, and the included angle between the generatrix of the inverted circular table section and the center axis of the inverted circular table section is larger than 45 degrees.
Still further, the axial length of the diverging flow passage preferably does not exceed one-half the axial length of the converging flow passage.
The oxygen-enriched flow passage is used for injecting oxygen-enriched gas formed by the oxygen-enriched precombustion combustion chamber into the rotary detonation combustion chamber.
In this embodiment, the oxygen-enriched flow channel adopts a contraction-expansion mode, the flow rate of the oxygen-enriched gas can be increased by reducing the flow channel area below the sound velocity of the gas flow rate, the flow rate is increased in the contraction section, the sound velocity is reached in the throat of the contraction flow channel, and then the flow rate is further increased to the sound velocity. The constriction flow passage functions to gradually increase the flow rate of the oxygen-enriched gas to achieve a high initial velocity of the oxygen-enriched gas in entering the detonation combustor. The flow speed of the oxygen-enriched gas forms sound velocity at the throat, reaches supersonic speed in the expansion flow channel, and is sprayed to the annular detonation cavity at supersonic speed.
The center of the rotary detonation combustor is provided with a center cone 41; the central cone preferably comprises a central large cylindrical section 411 and a cone section 412 arranged in sequence in the axial direction.
The outer diameter of the central large cylindrical section is larger than that of the large circular section, but not larger than the minimum outer diameter of the inverted circular table part; an annular detonation chamber 42 is formed between the central large cylindrical segment and the inner wall of the rotary detonation combustor.
The cylindrical section, the circular table section, the central large cylindrical section and the cone section are sequentially and coaxially sealed and detachably connected.
A fuel flow channel 43 is preferably provided in the inner wall of the rotary detonation combustor for filling, preheating and injecting fuel B into the annular detonation chamber.
The fuel flow passages preferably include a fuel filler neck 431, a fuel pre-heating flow passage 432, and a fuel injection hole 433.
The fuel preheating flow passage is axially distributed and annular, and can heat the fuel B while cooling the rotary detonation combustion chamber.
The fuel filler port is used for filling fuel into the fuel preheating flow passage.
The fuel injection holes are uniformly distributed along the circumferential direction of the front end of the fuel preheating flow channel, and each fuel injection hole is vertically distributed.
The annular detonation chamber can perform rotary detonation combustion on oxygen-enriched gas injected by the combustion transition section and fuel B injected by the fuel flow passage.
In the embodiment, the fuel A and the fuel B are the same fuel, and the mass flow of the fuel B is determined according to the fuel specific impulse and the designed thrust range; the mass flow rate of the fuel A is not more than 1/4 of the mass flow rate of the fuel B; under the condition that the rotary detonation wave is not decoupled, the mass flow of the total fuel and the liquid oxygen is regulated simultaneously through the flow control valve, and the report realizes the control of the thrust of the rotary detonation engine.
Oxygen-enriched gas entering the detonation combustion chamber is pre-burnt, so that the temperature is high; the fuel also has a higher temperature before flowing through the active cooling flow passage into the combustion chamber; therefore, the fuel and the oxidant have higher activity, and can realize rotary knocking when the equivalent ratio is 0.7-1.3, so that the range of the equivalent ratio for realizing stable propagation of the detonation wave is widened.
The preferred embodiments of the present invention have been described in detail above, but the present invention is not limited to the specific details of the above embodiments, and various equivalent changes can be made to the technical solution of the present invention within the scope of the technical concept of the present invention, and all the equivalent changes belong to the protection scope of the present invention.
Claims (10)
1. A staged combustion rotary detonation rocket engine, characterized by: comprises a precombustion injection section, an oxygen-enriched precombustion combustion chamber, a combustion transition section and a rotary knocking combustion chamber which are coaxially arranged in sequence along the axial direction;
The precombustion injection section is used for injecting liquid oxidant and fuel A into the oxygen-enriched precombustion chamber; wherein the mass flow rate of the liquid oxidant injection is greater than the mass flow rate of the fuel A injection;
The oxygen-enriched precombustion combustion chamber is used for carrying out oxygen-enriched combustion on the oxidant and the combustion A to form high-temperature high-pressure oxygen-enriched gas;
The combustion transition section is provided with an oxygen-enriched flow passage and is used for injecting oxygen-enriched gas formed by the oxygen-enriched precombustion combustion chamber into the rotary detonation combustion chamber; the rotary detonation combustor is provided with an annular detonation chamber and a fuel flow channel;
the fuel runner is used for filling, preheating and injecting fuel B into the annular knocking cavity;
The annular detonation chamber can perform rotary detonation combustion on oxygen-enriched gas injected by the combustion transition section and fuel B injected by the fuel flow passage.
2. The staged combustion rotary detonation rocket engine of claim 1, wherein: the oxygen-enriched flow passage comprises an annular flow passage, a shrinkage flow passage and an expansion flow passage which are sequentially arranged along the axial direction; the junction of the contracted flow channel and the expanded flow channel forms a throat part; the flow velocity of the oxygen-enriched gas forms sound velocity at the throat, reaches supersonic velocity in the expansion flow channel, and is sprayed to the annular detonation cavity at supersonic velocity.
3. The staged combustion rotary detonation rocket engine of claim 2, wherein: the center of the combustion transition section is provided with a center connector; the central connector comprises a cylindrical section and a circular table section which are coaxially arranged in sequence along the axial direction;
The tail end of the round table section is provided with a chamfer to form an inverted round table part;
the cylindrical section and the inner wall of the oxygen-enriched precombustion combustion chamber form the annular flow passage;
the circular table section and the inner wall of the oxygen-enriched precombustion combustion chamber form the contraction flow passage;
the inverted circular truncated cone portion and the inner wall of the oxygen-enriched precombustion combustion chamber form the expansion flow passage.
4. A staged combustion rotary detonation rocket engine as recited in claim 3, wherein: the included angle between the generatrix of the circular bench section and the center axis of the circular bench section is smaller than 30 degrees, and the included angle between the generatrix of the inverted circular bench section and the center axis of the inverted circular bench section is larger than 45 degrees.
5. A staged combustion rotary detonation rocket engine as recited in claim 3, wherein: the axial length of the expanding runner is not more than one half of the axial length of the contracting runner.
6. A staged combustion rotary detonation rocket engine as recited in claim 3, wherein: the center of the rotary detonation combustion chamber is provided with a center cone; the central cone comprises a central large cylindrical section and a cone section which are sequentially arranged along the axial direction;
the outer diameter of the central large cylindrical section is larger than that of the large circular section, but not larger than the minimum outer diameter of the inverted circular table part; the annular detonation cavity is formed between the central large cylindrical section and the inner wall of the rotary detonation combustion chamber;
The cylindrical section, the circular table section, the central large cylindrical section and the cone section are sequentially and coaxially sealed and detachably connected.
7. The staged combustion rotary detonation rocket engine of claim 1, wherein: the fuel runner comprises a fuel filling port, a fuel preheating runner and a fuel injection hole;
the fuel preheating flow passage is axially distributed and is annular;
the fuel filling port is used for filling fuel into the fuel preheating runner;
the fuel injection holes are uniformly distributed along the circumferential direction of the front end of the fuel preheating flow channel, and each fuel injection hole is vertically distributed.
8. The staged combustion rotary detonation rocket engine of claim 1, wherein: the precombustion injection section comprises an oxidant injection plate and a precombustion fuel injection plate;
the precombustion fuel injection plate is coaxially and hermetically arranged on the front end surface of the oxygen-enriched precombustion chamber;
The oxidant injection plate is coaxially and hermetically covered on the outer end surface of the pre-combustion fuel injection plate, and an oxidant gas collection cavity is formed between the oxidant injection plate and the pre-combustion fuel injection plate;
the center of the oxidant injection plate is provided with an oxidant injection port, and the oxidant injection port penetrates through the center of the precombustion fuel injection plate;
A plurality of precombustion fuel injection holes are distributed on the outer side of the precombustion fuel injection plate along the circumferential direction; each injection hole is communicated with the oxidant gas collection cavity and the oxygen-enriched precombustion combustion chamber, and the straight line of each injection hole is intersected with the straight line of the oxidant injection hole.
9. The staged combustion rotary detonation rocket engine of claim 1, wherein: the fuel A and the fuel B are the same fuel, and the mass flow of the fuel B is determined according to the fuel specific impulse and the designed thrust range; the mass flow rate of the fuel A is not more than 1/4 of the mass flow rate of the fuel B; under the condition that the rotary detonation wave is not decoupled, the mass flow of the total fuel and the liquid oxygen is regulated simultaneously through the flow control valve, so that the thrust of the rotary detonation engine is controlled.
10. The staged combustion rotary detonation rocket engine of claim 1, wherein: oxygen-enriched gas entering the detonation combustion chamber is pre-burnt, so that the temperature is high; the fuel also has a higher temperature before flowing through the active cooling flow passage into the combustion chamber; therefore, the fuel and the oxidant have higher activity, and can realize rotary knocking when the equivalent ratio is 0.7-1.3, so that the range of the equivalent ratio for realizing stable propagation of the detonation wave is widened.
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