CN112747335B - Lobe type backward step structure, lobe type concave cavity device and afterburning chamber - Google Patents
Lobe type backward step structure, lobe type concave cavity device and afterburning chamber Download PDFInfo
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- CN112747335B CN112747335B CN202110010395.3A CN202110010395A CN112747335B CN 112747335 B CN112747335 B CN 112747335B CN 202110010395 A CN202110010395 A CN 202110010395A CN 112747335 B CN112747335 B CN 112747335B
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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
- F23R3/00—Continuous combustion chambers using liquid or gaseous fuel
- F23R3/42—Continuous combustion chambers using liquid or gaseous fuel characterised by the arrangement or form of the flame tubes or combustion chambers
- F23R3/58—Cyclone or vortex type combustion chambers
Abstract
The invention discloses a lobe type backward step structure, a lobe type cavity device and a afterburning chamber, wherein the lobe type cavity device is a component of a solid rocket scramjet engine wall surface; the wave petal type backward step structure comprises a plurality of lobe grooves and backward steps which are distributed along the wall surface of the solid rocket scramjet in a spread-out staggered manner; the lobe-type cavity device comprises a cavity, a forward step and a lobe-type backward step structure; the wave petal type backward step structure is arranged at one end, located at the upstream of the air flow in the engine, of the concave cavity, and the forward step is arranged at one end, located at the downstream of the air flow in the engine, of the concave cavity. Through adopting foretell lobe type cavity device, can effectual stable flame, reinforcing in the afterburning room rich fuel gas and the mixing burning of incoming flow air, compromise simple structure's the scramjet that is applicable to solid rocket scramjet simultaneously.
Description
Technical Field
The invention relates to the technical field of solid rocket scramjet engines, in particular to a lobe type backward step structure, a lobe type cavity device and a afterburning chamber.
Background
The solid rocket scramjet engine is characterized in that rich fuel gas generated by a fuel gas generator is mixed with air to combust to generate high-temperature fuel gas, and thrust is generated under the expansion action of a spray pipe. Compared with a liquid scramjet engine, the solid rocket scramjet engine has the advantages of simple structure, low cost, short operation reaction time, good maneuverability and safety, long storage time and the like, and compared with a solid fuel scramjet engine, the solid rocket scramjet engine has the advantages of easy flow adjustment, no ignition flame stabilization problem, small influence of incoming flow parameters on the working process of a combustion chamber, long working time and the like, so the solid rocket scramjet engine has good application prospect.
At present, the research on the solid rocket scramjet at home and abroad is still in a primary stage, two schemes of head and lateral air intake are designed in China, and the feasibility of the solid rocket scramjet is verified through value simulation and experimental research; the Lexuan and the like develop the influence of the configuration of the afterburning chamber of the solid rocket scramjet engine on the performance of the engine, and respectively research the influence of multi-stage small-angle expansion, a concave cavity flame stabilizer and a turbulent flow device on the mixing combustion performance of the afterburning chamber of the engine aiming at a pure gas phase combustion product; influence analysis of the solid rocket scramjet engine afterburning chamber configuration and the gas jet hole structure is carried out by Liu Zi, Chen Lin quan and the like, and the influence of the length, the expansion angle, the gas jet hole profile, the number, the jet angle and the like of the combustion chamber on the engine performance is analyzed aiming at pure gas phase combustion products; the influence of the central support plate on the performance of the solid rocket scramjet engine is researched by Liuyang, high brave steel and the like, and the central support plate with the tail part provided with a staggered framework is provided for turbulent flow of the solid rocket scramjet engine; researches on the mixing effect and the combustion efficiency of the solid rocket scramjet engine are developed in Qin Fei, Lian and the like, the influence of a combination mode of a turbulence device and a cavity flame stabilizer in a afterburning chamber on the performance of the engine is analyzed through numerical simulation and experiments, and meanwhile, a device for enhancing mixing, which is formed by combining an oblique wedge and a cavity, is provided; the influence of different turbulence devices in a combustion chamber of the solid rocket scramjet engine on the performance of the engine is developed by the mown gazania and the Licheng and the like, and the research on the ignition and flame stabilizing effects in a afterburning chamber by the inclined clefts, the support plates and the support plate-cavity flame stabilizer combination device is analyzed; the solid rocket scramjet engine with the concave cavity flame stabilizer, such as Marikun, develops experimental research, and analyzes the influence of different structures and positions of the concave cavity flame stabilizer on the performance of the engine; the Liujie et al proposed an organization mode combining a cavity flame stabilizer and wall injection for a solid rocket scramjet engine, and studied the influence of the length-depth ratio and the depth of a gas injection mode and the cavity flame stabilizer on the engine performance.
Through the existing research, the existing device for enhancing the mixing combustion of gas and solid phases in the afterburning chamber of the solid rocket scramjet engine mainly comprises a central support plate, a wedge and a cavity flame stabilizer.
The central support plate can enable the local area in the afterburner to generate vortexes and backflow areas, the fuel mixing process is improved, the mixed gas residence time is prolonged, the combustion efficiency is further improved, the backflow area plays a role in flame stabilization to a certain extent, the support plate causes the turbulence degree of airflow to be increased, and the propagation speed of flame is improved. However, the central support plate is directly exposed to high-speed airflow, so that the resistance and total pressure loss are large, and advanced composite materials are required for cooling and heat protection measures.
The slope is generally placed in parallel, is used for enhancing flow mixing near a wall surface, is flexible in placement position and quantity, and increases the flow direction vorticity of a flow field by disturbing incoming flow and generating phenomena of shock waves, expansion waves, flow direction vortexes and the like in the flow field so as to achieve the purpose of enhancing mixing. But the slope structure protrudes out of the wall surface of the afterburning chamber, so that small total pressure loss can be caused; the heat load of the front side is large, and the requirements on the heat protection and the material performance of the oblique wedge are high; the produced recirculation zone has a small range and insufficient flame stabilization capability, so the device is often used as a mixing enhancement device to be matched with other flame stabilization modes.
The concave cavity flame stabilizer is concerned by various scholars, has the outstanding advantages of small total pressure loss and good stable combustibility, is widely used as a flame stabilizing device for a liquid scramjet engine at present, and forms a low-speed high-temperature region to provide a stable ignition source. However, for the solid rocket scramjet engine, the performance of the solid rocket scramjet engine is greatly influenced by mixing, and due to the compressibility of supersonic airflow, the shear layer of the device is slowly developed, so that the traditional cavity flame stabilizer has the problems of poor airflow mixing effect and small flame propagation range.
Disclosure of Invention
Aiming at the defects of the enhanced mixing combustion device of the solid rocket scramjet in the prior art, the invention provides a lobe type backward step structure, a lobe type cavity device and a afterburning chamber, which can stabilize flame, enhance the mixing combustion of rich fuel gas and incoming air in the afterburning chamber, and are suitable for the solid rocket scramjet with simple structure.
In order to achieve the purpose, the invention provides a lobe type backward step structure which is a component of a solid rocket scramjet engine wall surface;
the wave petal type backward step structure comprises a plurality of lobe grooves and backward steps which are distributed along the wall surface of the solid rocket scramjet in a spread-out staggered mode.
In one embodiment, the lobe grooves include an inclined plane and side faces on both sides of the inclined plane, the rearward step includes a step face and side faces on both sides of the step face, and adjacent lobe grooves share one side face with the rearward step.
In one embodiment, one end of the inclined surface of the lobe groove located downstream of the airflow in the engine is inclined away from the supplementary combustion chamber of the engine, the step surface in the backward step comprises a first wall surface and a second wall surface, the first wall surface is parallel to the airflow in the engine, one end of the second wall surface is connected with one end of the first wall surface located downstream of the airflow in the engine, and the other end of the second wall surface extends away from the supplementary combustion chamber of the engine.
In one embodiment, the first wall surface and the second wall surface of the step surface in the same backward step are perpendicular to each other.
In order to achieve the purpose, the invention also provides a lobe type cavity device which is a component of the wall surface of the solid rocket scramjet engine;
the lobe-type cavity device comprises a cavity, a forward step and the lobe-type backward step structure;
the wave petal type backward step structure is arranged at one end, located at the upstream of the air flow in the engine, of the concave cavity, and the forward step is arranged at one end, located at the downstream of the air flow in the engine, of the concave cavity.
In one embodiment, the bottom wall of the cavity is provided with a gas injection port connected with the gas generator.
In order to achieve the purpose, the invention also provides a afterburning chamber of the solid rocket scramjet engine, which comprises an air inlet channel isolation section and an afterburning chamber which are sequentially connected from front to back along the flow direction, wherein the wall of the afterburning chamber is provided with the wave-petal-shaped cavity device.
In one embodiment, the afterburner is in a single-side multi-stage continuous expansion rectangular configuration, and the lobe type cavity devices are multiple in number and are arranged on the same wall surface of the afterburner from front to back in the flow direction.
In one embodiment, the afterburner is in a one-side multi-stage continuous expansion rectangular configuration, and the lobe type cavity devices are in a plurality and are respectively arranged on adjacent or opposite wall surfaces of the afterburner.
In one embodiment, the afterburner is in a single-side multi-stage continuous expansion rectangular configuration, and the lobe type cavity devices are multiple in number and are arranged on each wall surface of the afterburner from front to back in the flow direction.
Compared with the prior art, the lobe type backward step structure, the lobe type concave cavity device and the afterburning chamber provided by the invention have the beneficial technical effects that:
(1) the flow direction vortex is generated through the lobe structure, so that the residence time of the rich fuel gas and the incoming flow air in the afterburning chamber is prolonged, the mixing effect of the rich fuel gas and the incoming flow air in the afterburning chamber is enhanced, and the combustion efficiency is improved;
(2) the invention increases the turbulence degree of the inner region of the afterburner, increases the mass exchange rate between the gas inside and outside the concave cavity flame stabilizer and enlarges the flame propagation range;
(3) the height of backward steps, the geometric dimension of the lobes, the number of the lobes and the backward steps of the lobe type concave cavity device can be flexibly adjusted, so that the optimal combustion organization mode is realized;
(4) the invention has simple structure, easy realization, less influence on the flow field of the afterburning chamber of the solid rocket scramjet engine and small flow loss.
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 description of the embodiments or the prior art will be briefly described below, it is obvious that the drawings in the following description are only some embodiments of the present invention, and for those skilled in the art, other drawings can be obtained according to the structures shown in the drawings without creative efforts.
FIG. 1 is a schematic structural diagram of a lobe-type backward step structure in an embodiment of the present invention;
FIG. 2 is a schematic diagram of a lobe-type re-entrant device in accordance with an embodiment of the present invention;
FIG. 3 is a schematic diagram of a preferred arrangement of lobe-shaped re-entrant means in accordance with an embodiment of the invention;
FIG. 4 is a schematic structural diagram of an afterburner in an embodiment of the invention;
FIG. 5 is a graph showing the variation of the combustion efficiency of the rich fuel gas component in the afterburner of the solid rocket scramjet engine using the conventional concave-cavity flame stabilizer in the embodiment of the present invention;
FIG. 6 is a graph showing the combustion efficiency variation of the rich fuel gas component in the afterburner of the solid rocket scramjet engine using the lobe-type cavity device in the embodiment of the invention;
FIG. 7 is a graph of total pressure recovery in an afterburner using a conventional re-entrant flame holder in an embodiment of the invention and a lobed re-entrant device in an embodiment of the invention.
The reference numbers illustrate: lobe grooves 100, backward steps 110, inclined surfaces 101, first wall surfaces 102, second wall surfaces 103, and side surfaces 104; the gas injection device comprises a cavity 200, a forward step 210, a lobe type backward step structure 220 and a gas injection port 201; the inlet duct isolation section 300, the afterburner 310, the lobe cavity arrangement 320, the gas generator 330, and the gas conduit 340.
The implementation, functional features and advantages of the objects of the present invention will be further explained with reference to the accompanying drawings.
Detailed Description
The technical solutions in the embodiments of the present invention will be clearly and completely described below with reference to the drawings in the embodiments of the present invention, and it is obvious that the described embodiments are only a part of the embodiments of the present invention, and not all of the embodiments. All other embodiments, which can be derived by a person skilled in the art from the embodiments given herein without making any creative effort, shall fall within the protection scope of the present invention.
It should be noted that all the directional indicators (such as up, down, left, right, front, and rear … …) in the embodiment of the present invention are only used to explain the relative position relationship between the components, the movement situation, etc. in a specific posture (as shown in the drawing), and if the specific posture is changed, the directional indicator is changed accordingly.
In addition, the descriptions related to "first", "second", etc. in the present invention are only for descriptive purposes and are not to be construed as indicating or implying relative importance or implicitly indicating the number of technical features indicated. Thus, a feature defined as "first" or "second" may explicitly or implicitly include at least one such feature. In the description of the present invention, "a plurality" means at least two, e.g., two, three, etc., unless specifically limited otherwise.
In the present invention, unless otherwise expressly stated or limited, the terms "connected," "secured," and the like are to be construed broadly, and for example, "secured" may be a fixed connection, a removable connection, or an integral part; the connection can be mechanical connection, electrical connection, physical connection or wireless communication connection; they may be directly connected or indirectly connected through intervening media, or they may be connected internally or in any other suitable relationship, unless expressly stated otherwise. The specific meanings of the above terms in the present invention can be understood by those skilled in the art according to specific situations.
In addition, the technical solutions in the embodiments of the present invention may be combined with each other, but it must be based on the realization of those skilled in the art, and when the technical solutions are contradictory or cannot be realized, such a combination of technical solutions should not be considered to exist, and is not within the protection scope of the present invention.
Fig. 1 shows a lobe type backward step structure disclosed in the present embodiment, which is an integral part of a solid rocket scramjet engine wall surface. Specifically, the lobed backward step structure comprises a plurality of lobed slots 100 and backward steps 110 distributed alternately along the wall span of the solid rocket scramjet. The lobe grooves 100 comprise inclined planes 101 and side faces located on two sides of the inclined planes 101, the backward steps 110 comprise step faces and side faces located on two sides of the step faces, and adjacent lobe grooves 100 and the backward steps 110 share one side face 104.
More specifically, the end of the inclined surface 101 of the lobe groove 100 located upstream of the air flow in the engine is a fixed end, and the end of the inclined surface 101 of the lobe groove 100 located downstream of the air flow in the engine is an inclined end, wherein the fixed end is located on a conventional wall surface of the solid rocket scramjet engine, and the inclined end is inclined away from the supplementary combustion chamber of the engine. The step surface of the backward step 110 comprises a first wall surface 102 and a second wall surface 103, the first wall surface 102 is positioned on the conventional wall surface of the solid rocket scramjet engine and is parallel to the direction of air flow in the engine, the head end of the second wall surface 103 is connected with one end of the first wall surface 102 positioned at the downstream of the air flow in the engine, and the tail end of the second wall surface 103 extends in the direction away from the supplementary combustion chamber of the engine, wherein the inclined end on the inclined surface 101 can be connected to the middle position of the second wall surface 103 or the tail end position of the second wall surface 103. In the embodiment shown, the inclined end of the inclined surface 101 is connected to the rear end of the second wall 103, i.e. the side surface 104 shared between the lobe groove 100 and the backward step 110 has a triangular structure.
In a preferred embodiment, in the same backward step 110, the first wall surface 102 and the second wall surface 103 in the step surface are perpendicular to each other, the side surface is perpendicular to the first wall surface 102 and the second wall surface 103, and in the same lobe groove 100, the side surface is perpendicular to the inclined surface 101. Of course, in the same backward step 110, the angles between the first wall surface 102 and the second wall surface 103, the angles between the side surfaces and the inclined surface 101, and the sizes of the first wall surface 102, the second wall surface 103, the inclined surface 101, and the side surfaces in the step surface can be flexibly adjusted according to actual requirements, so as to realize the optimal combustion organization mode.
It should be noted that the present embodiment does not limit the number of the lobe grooves 100 and the backward steps 110 in the lobe-type backward step structure, and the lobe-type backward step structure shown in fig. 1 is composed of three backward steps 110 and two lobe grooves 100 alternately, wherein both ends of the entire lobe-type backward step structure are backward steps.
The working process of the wave petal type backward step structure in this embodiment is: the lobe type backward step structure realizes flame stabilization and simultaneously enhances gas-solid two-phase flow mixing combustion in the afterburning chamber of the solid rocket scramjet engine by alternately distributing backward steps 110 and lobe grooves 100, wherein under the action of the backward steps 110, the afterburning chamber forms a local high-temperature low-speed backflow area for ignition and flame stabilization; under the action of the lobe groove 100, flow direction vortexes are generated near the high-temperature area, the mixing process of fuel is improved, the retention time of mixed gas is prolonged, the combustion efficiency is further improved, meanwhile, airflow turbulence degree is increased due to a vortex structure, the propagation speed of flame is increased, and the range of the high-temperature area in the afterburner is enlarged.
Fig. 2 shows a lobe-type cavity device disclosed in this embodiment, which is an integral part of the wall surface of a solid rocket scramjet engine. Specifically, the lobe cavity device comprises a cavity 200, a forward step 210 and the lobe-shaped backward step structure 220, wherein the lobe-shaped backward step structure 220 is arranged at one end of the cavity 200 which is positioned at the upstream of the air flow in the engine, and the forward step 210 is arranged at one end of the cavity 200 which is positioned at the downstream of the air flow in the engine.
Referring to fig. 3, as a preferred embodiment, the bottom wall of the cavity 200 is provided with gas injection ports 201 connected with the gas generator, wherein the number of the gas injection ports 201 on the bottom wall of the cavity 200 can be set to be multiple, so that the fuel-rich gas is injected from the bottom wall of the cavity 200 in a wall injection manner, and the fuel distribution is improved.
The working process of the wave petal type cavity device in the embodiment is as follows: the lobe-type cavity device realizes flame stabilization and simultaneously enhances gas-solid two-phase flow mixing combustion in the afterburning chamber of the solid rocket scramjet engine by alternately distributing backward steps and lobe grooves, wherein the afterburning chamber forms a local high-temperature low-speed backflow area for ignition and flame stabilization under the action of the backward steps; under the action of the lobe grooves, flow direction vortexes are generated near the high-temperature area, the mixing process of fuel is improved, the retention time of mixed gas is prolonged, the combustion efficiency is further improved, and meanwhile, airflow turbulence degree is increased due to the vortex structure, so that the propagation speed of flame is increased, and the range of the high-temperature area in the afterburning chamber is enlarged.
Fig. 4 shows an afterburning chamber of a solid rocket scramjet engine disclosed in this embodiment, specifically, the afterburning chamber includes an intake duct isolation section 300 and an afterburning chamber 310 sequentially connected from front to back along a flow direction, the intake duct isolation section 300 is a rectangular with an equal cross section, and the afterburning chamber 310 is a one-side multi-stage continuous expansion rectangular configuration. The wall of the afterburning chamber 310 is provided with the wave-petal-shaped concave cavity device 320, so that flame is effectively stabilized, and the mixing and burning of rich fuel gas and incoming air in the afterburning chamber 310 are enhanced. Wherein the external gas generator 330 is communicated with a gas injection port on the lobed cavity apparatus 320 through a gas conduit 340.
The embodiment also discloses an arrangement mode of the lobe-shaped cavity devices 320 in the afterburning chamber of the solid rocket scramjet engine, wherein the number of the lobe-shaped cavity devices 320 is multiple, and the multiple lobe-shaped cavity devices 320 can be arranged on the same wall surface of the afterburning chamber 310 from front to back along the flow direction, can also be arranged on adjacent or opposite wall surfaces of the afterburning chamber 310 respectively, and can also be a combination of the two arrangement modes, namely the multiple lobe-shaped cavity devices 320 are arranged on each wall surface of the afterburning chamber 310 from front to back along the flow direction.
The afterburning chamber of the solid rocket scramjet engine is further explained by combining a specific simulation example.
Aiming at a fuel-rich solid propellant with a certain formula, a solid rocket scramjet engine using a traditional cavity flame stabilizer and a solid rocket scramjet engine adopting the lobe cavity device of the embodiment are obtained by a numerical simulation method, the combustion efficiency of each component of the fuel-rich gas in a afterburning chamber is shown in fig. 5 and 6, and the combustion efficiency value is shown in table 1.
TABLE 1 Combustion efficiency of rich-burn gas in afterburner
| Efficiency of combustion | C | H2 | CO | General of |
| Conventional |
22％ | 100％ | 100％ | 82.6％ |
| Wave petal type concave cavity device | 51％ | 100％ | 100％ | 89.1％ |
The total pressure of the flow channel in the afterburning chamber of the solid rocket scramjet engine in the embodiment is detected to obtain the change situation of the total pressure of the flow channel in the afterburning chamber along the flow direction, fig. 7 is a total pressure change curve of the traditional cavity flame stabilizer and the lobe-shaped cavity device in the embodiment in the afterburning chamber, and as can be seen from fig. 7, the total pressure change trends of the traditional cavity flame stabilizer and the lobe-shaped cavity device in the embodiment are consistent, and the total pressure recovery coefficients are basically consistent, so that the design expectation is reached.
The above description is only a preferred embodiment of the present invention, and is not intended to limit the scope of the present invention, and all modifications and equivalents of the present invention, which are made by the contents of the present specification and the accompanying drawings, or directly/indirectly applied to other related technical fields, are included in the scope of the present invention.
Claims (8)
1. The lobe type backward step structure is characterized by being a component of a solid rocket scramjet engine wall surface, and comprising a plurality of lobe grooves and backward steps which are distributed in a staggered mode along the spanwise direction of the solid rocket scramjet engine wall surface;
the lobe grooves comprise inclined planes and side faces positioned on two sides of the inclined planes, the backward steps comprise step faces and side faces positioned on two sides of the step faces, and the adjacent lobe grooves and the backward steps share one side face;
one end, located at the downstream of the engine internal airflow, of the inclined plane of the lobe groove inclines towards the direction far away from the engine supplementary combustion chamber, the step plane in the backward step comprises a first wall surface and a second wall surface, the first wall surface is parallel to the direction of the engine internal airflow, one end of the second wall surface is connected with one end, located at the downstream of the engine internal airflow, of the first wall surface, and the other end of the second wall surface extends towards the direction far away from the engine supplementary combustion chamber.
2. The petal-shaped backward step structure according to claim 1, wherein the first wall surface and the second wall surface of the step surface in the same backward step are perpendicular to each other.
3. A lobe-type cavity device is characterized in that the lobe-type cavity device is a component of a solid rocket scramjet engine wall surface;
the lobe cavity device comprises a cavity, a forward step and the lobe-shaped backward step structure of claim 1 or 2;
the wave petal type backward step structure is arranged at one end, located at the upstream of the air flow in the engine, of the concave cavity, and the forward step is arranged at one end, located at the downstream of the air flow in the engine, of the concave cavity.
4. The lobe-type pocket assembly of claim 3, wherein the bottom wall of the pocket is provided with a gas injection port associated with the gas generator.
5. A afterburning chamber of a solid rocket scramjet engine is characterized by comprising an air inlet channel isolation section and an afterburning chamber which are sequentially connected from front to back along the flow direction, wherein the wall of the afterburning chamber is provided with a lobe-shaped cavity device according to claim 3 or 4.
6. The afterburner of a solid-rocket scramjet engine as defined in claim 5, wherein said afterburner has a one-sided multi-stage continuous expanding rectangular configuration, and said plurality of lobe-shaped cavity means are arranged on the same wall surface of the afterburner from front to back in the flow direction.
7. The afterburner of a solid rocket scramjet engine as defined in claim 5, wherein said afterburner has a one-sided multi-stage continuous expanding rectangular configuration, and said plurality of lobe-shaped cavity means are disposed on adjacent or opposite wall surfaces of said afterburner.
8. The afterburner of a solid-rocket scramjet engine as defined in claim 5, wherein said afterburner has a one-sided multi-stage continuously expanding rectangular configuration, and said lobed cavity means are plural in number and arranged on each wall surface of the afterburner from front to rear in the flow direction.
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| CN113190932B (en) * | 2021-05-31 | 2022-04-19 | 中国人民解放军国防科技大学 | Method for improving combustion efficiency of scramjet engine based on pneumatic virtual concave cavity |
| CN113551259B (en) * | 2021-07-19 | 2022-09-30 | 南昌航空大学 | Wavy middle-slit type V-shaped flame stabilizer with lobe partition plate |
| CN113700574B (en) * | 2021-09-22 | 2022-09-27 | 西北工业大学 | A reinforcing mixing device for solid rocket ramjet |
| CN114352437A (en) * | 2022-01-07 | 2022-04-15 | 北京理工大学 | Solid fuel stamping combined engine suitable for wide Mach number flight |
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| GB758345A (en) * | 1952-10-15 | 1956-10-03 | Nat Res Dev | Improvements relating to combustion systems |
| EP0999367B1 (en) * | 1998-11-06 | 2003-02-12 | ALSTOM (Switzerland) Ltd | Flow conduit with cross-section discontinuity |
| CN101418956A (en) * | 2007-10-22 | 2009-04-29 | 周建兴 | A kind of supersonic speed combustion chamber burner scheme of novel injection structure |
| JP5529650B2 (en) * | 2010-07-01 | 2014-06-25 | 三菱重工業株式会社 | Supersonic combustor |
| CN102966974B (en) * | 2012-12-18 | 2015-01-21 | 中国人民解放军国防科学技术大学 | Supersonic combustor wall surface concave cavity structure and engine combustor comprising same |
| CN102980206B (en) * | 2012-12-18 | 2014-12-24 | 中国人民解放军国防科学技术大学 | Flame stabilizer and engine combustion chamber comprising same |
| CN107504517B (en) * | 2017-08-15 | 2023-09-29 | 南京航空航天大学 | Step type center cone with circumferential staggered terrace |
| CN107620653B (en) * | 2017-09-27 | 2018-11-02 | 西北工业大学 | A kind of disturbing flow device for solid-rocket combustion gas scramjet engine |
| CN109139297B (en) * | 2018-07-10 | 2019-06-07 | 西北工业大学 | A kind of device combining enhancing blending for solid-rocket scramjet engine |
| CN109882886B (en) * | 2018-12-21 | 2020-10-16 | 西安航天动力研究所 | RBCC engine combustion chamber with slope rocket layout mode and design method thereof |
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