CN109322761B - high-Mach-number in-flight engine annular combustion chamber and spiral oblique detonation combustion method - Google Patents
high-Mach-number in-flight engine annular combustion chamber and spiral oblique detonation combustion method Download PDFInfo
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- CN109322761B CN109322761B CN201811194017.XA CN201811194017A CN109322761B CN 109322761 B CN109322761 B CN 109322761B CN 201811194017 A CN201811194017 A CN 201811194017A CN 109322761 B CN109322761 B CN 109322761B
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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
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Abstract
The invention provides a high-Mach number in-flight engine annular combustion chamber and a spiral oblique detonation combustion method, which comprises the following steps: the invention relates to a fuel combustion device, which comprises an inner column, a shell and a shock wave generating device, wherein the shell surrounds the outer side of the inner column, an annular combustion chamber is formed between the inner wall of the shell and the outer wall of the inner column, supersonic combustible gas flows along the axial direction of the annular combustion chamber, the shock wave generating device is positioned in the annular combustion chamber and faces the flowing direction of air flow, the supersonic combustible gas generates shock waves when passing through the shock wave generating device as an obstacle, the shock waves ignite and induce the combustible gas to generate oblique detonation waves, the oblique detonation waves develop spirally in the annular combustion chamber, and finally the combustible gas in the whole annular combustion chamber is ignited.
Description
Technical Field
The invention belongs to the technical field of air-breathing hypersonic aircrafts, and particularly relates to an annular combustion chamber of an engine in flight with high Mach number and a method for carrying out spiral oblique detonation combustion by using the combustion chamber.
Background
The scramjet engine is a main power generation scheme of the existing air-breathing hypersonic aircraft and is suitable for flight conditions with the flight Mach number larger than 4. In high velocity gas flows, both fuel mixing and flame holding are relatively difficult, and at high mach numbers this is more severe. Therefore, conventional scramjet engines all employ hybrid enhancement and flame stabilization devices such as struts, ramp jets, cavities, and the like. Although these devices work well over a range of operating conditions, under high Mach number flight conditions (Ma ≧ 6), these devices have limited performance relative to existing engine configurations and can introduce significant drag. This limits the ability of scramjet engines to break through higher flight mach numbers. Therefore, there is a need to find a viable alternative to scramjet engines that can still operate stably in high mach number flight. At present, a skewed knock engine is a theoretically possible solution.
According to the scheme, oblique detonation combustion is generated by coupling shock waves and combustion waves through a shock wave generating device in a combustion chamber of the scramjet engine to generate thrust, and a schematic diagram of the scheme is shown in figure 1 and is a schematic diagram of the configuration of the scramjet engine based on the oblique detonation combustion scheme. The combustion scheme has good adaptability to high Mach number in theory, is a natural flame stabilizer under the condition of high Mach number, has higher theoretical thermodynamic cycle efficiency than that of slow combustion, and has remarkably shortened length of a combustion chamber due to rapid heat release of detonation, thereby reducing the structural weight of the aircraft and lightening the cooling burden of an engine. However, there are still serious challenges for the existing oblique knock combustion scheme, which results in that the oblique knock engine is still in the conceptual stage so far. These challenges include three main aspects:
(1) combustion schemes based on oblique knock premixing create a problem of fuel and air mixing at high mach numbers: at high mach numbers, both the fuel jet penetration depth and the rate of mixed layer growth are greatly inhibited, resulting in inadequate mixing of the fuel and air.
(2) Controllability of oblique knocking. At present, preliminary experiments and numerical studies show that the reason that the stability of the inclined detonation in the combustion chamber is poor is that the inclined detonation generates Mach reflection in a limited space to cause flow choking, so that the stability of the inclined detonation is unstable and the inclined detonation propagates upstream. And is more difficult to control if a scheme is adopted in which oblique knocking is generated by multiple reflections of the shock wave.
(3) The shock wave generating device brings about resistance to the aircraft: the shock wave generating device generates shock waves with enough strength to induce and generate oblique detonation, and the shock wave generating device needs to have enough windward area to realize detonation of the oblique detonation, so that the windward area of the shock wave generating device in the existing configuration is larger relative to the cross-sectional area of a flow passage, namely the blockage is larger, extra larger resistance is brought while the shock waves are generated, and the performance of an engine is weakened.
Aiming at the problems, the technical scheme of the invention adopts the annular combustion chamber of the high-Mach-number in-flight engine and the method for carrying out spiral oblique detonation combustion by using the combustion chamber, so that the three problems can be effectively improved, and the performance of the engine adopting the oblique detonation combustion scheme is obviously improved.
Disclosure of Invention
In view of the above-mentioned shortcomings of the prior art, it is an object of the present invention to provide a high mach number in-flight annular engine combustor and a method for helical oblique detonation combustion using the combustor.
In order to achieve the purpose, the technical scheme of the invention is as follows:
a high mach number in-flight engine annular combustor comprising: the shock wave generating device comprises an inner column, a shell and a shock wave generating device, wherein the shell surrounds the outer side of the inner column at a certain interval, an annular combustion chamber is formed between the inner wall of the shell and the outer wall of the inner column, supersonic combustible gas flows along the axial direction of the annular combustion chamber, the shock wave generating device is positioned in the annular combustion chamber and is placed in a direction facing the flowing direction of air flow to generate shock waves, the windward area of the shock wave generating device can realize detonation of oblique shock waves, the supersonic combustible gas generates shock waves when flowing through the shock wave generating device serving as an obstacle, the shock waves ignite the combustible gas and induce the oblique shock waves, the oblique shock waves develop spirally in the annular combustion chamber and finally ignite the combustible gas in the whole annular combustion chamber, the shock wave generating device adopts a unidirectional type or a bidirectional type, the unidirectional type generates the shock waves on one side of the shock wave generating device, and the bidirectional type simultaneously generates the shock waves on both sides of the shock wave generating device.
Preferably, the cross section of the inner column is circular, oval or racetrack shaped, and the inner column is solid or hollow column body, or solid or hollow cone body.
Preferably, the annular combustion chamber has an annular gap width of no more than 100mm between the inner wall of the outer casing and the outer wall of the inner column to facilitate mixing of the wall injected fuel with the incoming air.
Preferably, the shock wave generating device is one or more of a wedge, a boss, a sphere or a blunt body.
Preferably, the transverse width of the shock wave generating device extends from the outer wall of the inner column to the inner wall of the shell, or the shock wave generating device is suspended and fixed between the outer wall of the inner column and the inner wall of the shell and is not in contact with the outer wall of the inner column and the inner wall of the shell.
Preferably, the number of the shock wave generating devices is not more than 4 so as to reduce the total resistance of the shock wave generating devices, and the shock wave generating devices are equidistantly arranged in the annular combustion chamber so as to make the thrust of the engine uniform.
Preferably, the high mach number flight is: flight conditions at a flight mach number greater than or equal to 6. At this time, the air velocity is still very high even after the deceleration compression action of the air inlet of the hypersonic aircraft, which makes the traditional supersonic combustion method (for example, by means of a mixing enhancement and flame stabilization device such as a concave cavity, a support plate and the like) very difficult in terms of fuel mixing, engine drag reduction and flame stabilization, and the difficulty is increasingly serious along with the increase of the flight mach number.
Preferably, the material of the shock wave generator is a high temperature resistant alloy or a ceramic composite material.
In order to achieve the above object, the present invention further provides a helical oblique detonation combustion method for the annular combustion chamber of the high mach number in-flight engine, which comprises: supersonic combustible gas flows along the axial direction of the annular combustion chamber, when air flow passes through the shock wave generating device, shock waves are generated, shock wave induction generates oblique detonation waves, the combustible gas is completely combusted after passing through an oblique detonation wave surface, deflection occurs in the direction of the air flow, the oblique detonation waves are developed in a spiral shape along the annular combustion chamber, and finally the combustible gas of the whole annular combustion chamber is ignited.
The invention has the beneficial effects that: the invention provides an annular combustion chamber of an engine in flight with high Mach number and a spiral oblique detonation combustion method based on the annular combustion chamber, which can obviously improve fuel mixing under the condition of high Mach number, inhibit Mach reflection of oblique detonation, reduce total pressure loss caused by oblique detonation combustion, enhance controllability of oblique detonation and reduce relative resistance of a shock wave generating device. Because the annular seam width of the annular combustion chamber is small, when the scheme of injecting fuel on the wall surface of the annular combustion chamber is adopted under the high Mach number, the fuel can be completely mixed with air within a short distance, and therefore, a support plate, a support rod and other resistance components are not required to be adopted to enhance mixing like in a large-size combustion chamber. On the other hand, according to the placement scheme of the shock wave generating device, the windward area of the shock wave generating device is smaller (namely the blockage is smaller) than the circular seam cross-sectional area of the annular combustion chamber, so that the relative resistance of the shock wave generating device is also smaller, the Mach reflection of the spiral oblique detonation in a limited space is effectively inhibited, the flow choking is not easy to cause, and the controllability of the oblique detonation is enhanced.
Drawings
Fig. 1 is a schematic diagram of a conventional oblique knock engine.
FIG. 2 is a schematic representation of the engine annular combustor configuration of the present invention for high Mach number flight conditions.
Fig. 3 is a sectional view of the shock wave generating apparatus shown in fig. 2.
Fig. 4 is a sectional view of fig. 2 for showing an inner column and an outer shell.
FIG. 5 is a graph of the results of computer numerical simulation of the helical squib wave in the annular combustion chamber. The annular combustion chamber in the figure is in a runway shape, the shock wave generating devices are provided with one-way oblique wedges, the number of the shock wave generating devices is two, and the shock wave generating devices are arranged in a reverse symmetrical mode.
1 is an annular combustion chamber, 2 is a shell, 3 is an inner column, 4 is a shock wave generating device, and 5 is spiral oblique detonation wave.
Detailed Description
The embodiments of the present invention are described below with reference to specific embodiments, and other advantages and effects of the present invention will be easily understood by those skilled in the art from the disclosure of the present specification. The invention is capable of other and different embodiments and of being practiced or of being carried out in various ways, and its several details are capable of modification in various respects, all without departing from the spirit and scope of the present invention.
A high mach number in-flight engine annular combustor comprising: the shock wave generating device comprises an inner column 3, a shell 2 and a shock wave generating device 4, wherein the shell 2 surrounds the outer side of the inner column 3 at a certain interval, an annular combustion chamber 1 is formed between the inner wall of the shell 2 and the outer wall of the inner column 3, the annular gap width between the inner wall of the shell and the outer wall of the inner column of the annular combustion chamber is not more than 100mm, supersonic combustible gas flows along the axial direction of the annular combustion chamber, the shock wave generating device is positioned in the annular combustion chamber and faces the airflow flowing direction to generate shock waves, the windward area of the shock wave generating device can realize the detonation of oblique detonation, the supersonic combustible gas generates shock waves when flowing through the shock wave generating device serving as an obstacle, the shock waves ignite the combustible gas and induce the oblique detonation waves to generate oblique detonation waves, the oblique detonation waves spirally develop in the annular combustion chamber and finally ignite the combustible gas in the whole annular combustion chamber, the shock wave generating device adopts a unidirectional type or a bidirectional type, the unidirectional type generates shock waves on one side of the shock wave generating device, the bidirectional type generates shock waves simultaneously on both sides of the shock wave generating device.
The cross section of the inner column is circular, or elliptical, or racetrack-shaped, or other axisymmetric shape. The inner column is a solid or hollow column body or a solid or hollow cone body.
The annular combustion chamber has an annular seam width between the inner wall of the outer shell and the outer wall of the inner column of not more than 100 mm.
The shock wave generating device is one or more of a wedge, a boss, a sphere or a blunt body.
The transverse width of the shock wave generating device extends to the inner wall of the shell from the outer wall of the inner column, or the shock wave generating device is suspended and fixed between the outer wall of the inner column and the inner wall of the shell and is not in contact with the outer wall of the inner column and the inner wall of the shell.
The number of the shock wave generating devices is not more than 4, and the shock wave generating devices are equidistantly arranged in the annular combustion chamber.
The high mach number flight refers to: flight conditions at a flight mach number greater than or equal to 6. At this time, the air velocity is still very high even after the deceleration compression action of the air inlet of the hypersonic aircraft, which makes the traditional supersonic combustion method (for example, by means of a mixing enhancement and flame stabilization device such as a concave cavity, a support plate and the like) very difficult in terms of fuel mixing, engine drag reduction and flame stabilization, and the difficulty is increasingly serious along with the increase of the flight mach number.
The shock wave generating device is made of high-temperature-resistant alloy or ceramic composite material.
The embodiment also provides a helical oblique detonation combustion method of the annular combustion chamber of the high-mach-number in-flight engine, which comprises the following steps: supersonic combustible gas flows along the axial direction of the annular combustion chamber, when air flow passes through the shock wave generating device as an obstacle, shock waves are generated, shock wave induction generates oblique detonation waves, combustible gas is completely combusted after passing through an oblique detonation wave surface, deflection occurs in the direction of the air flow, the oblique detonation waves are developed spirally along the annular combustion chamber, and finally the combustible gas in the whole annular combustion chamber is ignited.
The foregoing embodiments are merely illustrative of the principles and utilities of the present invention and are not intended to limit the invention. Any person skilled in the art can modify or change the above-mentioned embodiments without departing from the spirit and scope of the present invention. Accordingly, it is intended that all equivalent modifications or changes which can be made by those skilled in the art without departing from the spirit and technical spirit of the present invention be covered by the claims of the present invention.
Claims (9)
1. A high mach number in-flight engine annular combustor, comprising: the shock wave generating device comprises an inner column (3), a shell (2) and a shock wave generating device (4), wherein the shell (2) surrounds the outer side of the inner column (3) at a certain interval, an annular combustion chamber (1) is formed between the inner wall of the shell (2) and the outer wall of the inner column (3), the shock wave generating device is positioned in the annular combustion chamber and is placed in a direction facing the flowing direction of airflow to generate shock waves, supersonic combustible gas flows along the axial direction of the annular combustion chamber and generates shock waves when flowing through the shock wave generating device serving as an obstacle, the shock waves ignite the combustible gas and induce the combustible gas to generate oblique shock waves, the acting surface of the shock wave generating device is perpendicular to the inner wall surface and the outer wall surface of the annular combustion chamber, the oblique shock waves develop spirally in the annular combustion chamber under the spatial limiting effect of the inner wall of the shell (2) and the outer wall of the inner column (3), and finally the combustible gas in the whole annular combustion chamber is ignited, and the shock wave generating device is of a unidirectional type or a bidirectional type, the unidirectional type generates shock waves at one side of the shock wave generating device, and the bidirectional type generates shock waves at two sides of the shock wave generating device simultaneously.
2. A high mach number in-flight engine annular combustor according to claim 1, wherein: the section of the inner column is round, oval or racetrack; the inner column is a solid or hollow column body or a solid or hollow cone body.
3. A high mach number in-flight engine annular combustor according to claim 1, wherein: the annular combustion chamber has an annular seam width between the inner wall of the outer shell and the outer wall of the inner column of not more than 100 mm.
4. A high mach number in-flight engine annular combustor according to claim 1, wherein: the shock wave generating device is one or more of a wedge, a boss, a sphere or a blunt body.
5. A high mach number in-flight engine annular combustor according to claim 1, wherein: the transverse width of the shock wave generating device extends to the inner wall of the shell from the outer wall of the inner column, or the shock wave generating device is suspended and fixed between the outer wall of the inner column and the inner wall of the shell and is not in contact with the outer wall of the inner column and the inner wall of the shell.
6. A high mach number in-flight engine annular combustor according to claim 1, wherein: the number of the shock wave generating devices is not more than 4, and the shock wave generating devices are equidistantly arranged in the annular combustion chamber.
7. A high mach number in-flight engine annular combustor according to claim 1, wherein: the high mach number flight refers to: flight conditions at a flight mach number greater than or equal to 6.
8. A high mach number in-flight engine annular combustor according to claim 1, wherein: the shock wave generating device is made of high-temperature-resistant alloy or ceramic composite material.
9. A high mach number in-flight helical detonation combustion method in an annular combustor of an engine according to any one of claims 1 to 8, characterised in that: supersonic combustible gas flows along the axial direction of the annular combustion chamber, when air flow passes through the shock wave generating device, shock waves are generated, shock wave induction generates oblique detonation waves, the combustible gas is completely combusted after passing through an oblique detonation wave surface, deflection occurs in the direction of the air flow, the oblique detonation waves are developed in a spiral shape along the annular combustion chamber, and finally the combustible gas of the whole annular combustion chamber is ignited.
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| CN111207007A (en) * | 2019-12-26 | 2020-05-29 | 中国空气动力研究与发展中心 | Method for enhancing stability of fixation of oblique detonation wave in closed space |
| CN111207009B (en) * | 2019-12-26 | 2023-01-13 | 中国空气动力研究与发展中心 | Method for initiating oblique detonation wave in supersonic velocity airflow by using external instantaneous energy source |
| CN112879177B (en) * | 2021-02-05 | 2022-07-08 | 中国空气动力研究与发展中心空天技术研究所 | Hypersonic mechanical transmission type gas-liquid dual-purpose pulse injection device and method |
| CN112922744B (en) * | 2021-03-05 | 2023-01-06 | 中国空气动力研究与发展中心空天技术研究所 | Wall-embedded aircraft fuel conveying device |
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| US6857261B2 (en) * | 2003-01-07 | 2005-02-22 | Board Of Regents, The University Of Texas System | Multi-mode pulsed detonation propulsion system |
| US8393158B2 (en) * | 2007-10-24 | 2013-03-12 | Gulfstream Aerospace Corporation | Low shock strength inlet |
| US7954754B2 (en) * | 2008-06-02 | 2011-06-07 | The United States Of America As Represented By The Secretary Of The Army | Mechanical acoustic noise generator system for scramjet engine |
| CN106065830B (en) * | 2016-06-01 | 2017-11-24 | 南京航空航天大学 | A kind of pulse detonation combustor device combined based on rotary valve with pneumatic operated valve |
| CN106968833B (en) * | 2017-03-29 | 2019-02-05 | 中国人民解放军国防科学技术大学 | A kind of supersonic speed detonation engine and its propulsion system |
| CN108332233B (en) * | 2018-01-08 | 2020-04-21 | 南京航空航天大学 | Combustion chamber for adjusting combustion area based on inlet rotational flow rectification support plate and special support plate thereof |
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