CN115342380A - Nonlinear detonation combustion chamber - Google Patents
Nonlinear detonation combustion chamber Download PDFInfo
- Publication number
- CN115342380A CN115342380A CN202210826310.3A CN202210826310A CN115342380A CN 115342380 A CN115342380 A CN 115342380A CN 202210826310 A CN202210826310 A CN 202210826310A CN 115342380 A CN115342380 A CN 115342380A
- Authority
- CN
- China
- Prior art keywords
- combustor
- combustion chamber
- assembly
- oxidant
- fuel
- Prior art date
- Legal status (The legal status is an assumption and is not a legal conclusion. Google has not performed a legal analysis and makes no representation as to the accuracy of the status listed.)
- Pending
Links
Images
Classifications
-
- 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
- F23R3/28—Continuous combustion chambers using liquid or gaseous fuel characterised by the fuel supply
-
- F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
- F02—COMBUSTION ENGINES; HOT-GAS OR COMBUSTION-PRODUCT ENGINE PLANTS
- F02K—JET-PROPULSION PLANTS
- F02K7/00—Plants in which the working fluid is used in a jet only, i.e. the plants not having a turbine or other engine driving a compressor or a ducted fan; Control thereof
- F02K7/02—Plants in which the working fluid is used in a jet only, i.e. the plants not having a turbine or other engine driving a compressor or a ducted fan; Control thereof the jet being intermittent, i.e. pulse-jet
-
- 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
- F02K9/95—Rocket-engine plants, i.e. plants carrying both fuel and oxidant therefor; Control thereof characterised by starting or ignition means or arrangements
-
- 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
- F23R3/52—Toroidal combustion chambers
-
- F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
- F23—COMBUSTION APPARATUS; COMBUSTION PROCESSES
- F23R—GENERATING COMBUSTION PRODUCTS OF HIGH PRESSURE OR HIGH VELOCITY, e.g. GAS-TURBINE COMBUSTION CHAMBERS
- F23R7/00—Intermittent or explosive combustion chambers
-
- 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
Abstract
The invention provides a nonlinear detonation combustor, which comprises an oxidant and fuel injection assembly, a combustor assembly and a tail nozzle assembly, wherein the oxidant and fuel injection assembly, the combustor assembly and the tail nozzle assembly are sequentially communicated from an inlet end to an outlet end; the oxidant and fuel injection assembly comprises an external fuel inlet, an external oxidant inlet, an oxidant pressure stabilizing cavity and a fuel pressure stabilizing cavity, the external fuel inlet is communicated with the fuel pressure stabilizing cavity, the external oxidant inlet is communicated with the oxidant pressure stabilizing cavity, and the outlet end of the oxidant pressure stabilizing cavity and the outlet end of the fuel pressure stabilizing cavity are both communicated with the inlet end of the combustion chamber assembly; the combustion chamber assembly comprises a shell, a circumferential seam detonation combustion chamber and a combustion chamber inner ring body, wherein the combustion chamber outer tube shell is coaxially sleeved outside the combustion chamber inner ring body, and the circumferential seam detonation combustion chamber is a circumferential seam gap between the combustion chamber outer tube shell and the combustion chamber inner ring body; the tail nozzle assembly comprises a tail nozzle, and the inlet end of the tail nozzle is communicated with the outlet end of the combustion chamber assembly. Through nonlinear design, the problem of applying the continuous rotation detonation engine under different scenes can be solved.
Description
Technical Field
The present description relates to the field of continuous detonation combustors, and in particular to a non-linear detonation combustor.
Background
The slow combustion and detonation combustion modes exist in nature, the propagation rate of flame of the slow combustion is relatively low, and the combustion modes in power devices such as an internal combustion engine, an aircraft engine, a gas turbine and the like are all slow combustion; detonation combustion is characterized in that the upstream of a combustion area is of a shock wave structure, shock waves are coupled with the combustion area to propagate, and the flame propagation speed of detonation combustion is far higher than that of slow combustion and can reach thousands of meters per second generally.
The aerospace field is more and more competitive, and the research on key innovation technology in the aerospace field draws more and more attention from various countries. In recent years, with the continuous and deep research on hypersonic aircrafts and single-stage in-orbit power systems, the technology of novel continuous rotation detonation engines is rapidly developed. Researches show that the propelling technology based on detonation combustion can greatly reduce fuel consumption, greatly improve the specific impulse characteristic of a power device and has important significance for widening the working envelope of the air-breathing aircraft and improving the economy and performance of the conventional weaponry. As a leading technology capable of overtaking at a curve, comprehensive and deep research on the technology is more urgent.
The continuous rotation detonation engine is a power technology utilizing detonation combustion, and has the following characteristics and advantages in summary: 1. only one time of successful detonation is needed, and the detonation wave can be continuously transmitted along the circumferential direction of the combustion chamber; 2. the combustion rate is high, the heat release intensity is high, the structure of the combustion chamber is compact, and the length of the engine can be shortened; 3. the boosting characteristic is provided, the number of stages of a gas compressor of the turbine engine can be reduced or the total pressure loss of an air inlet passage of the ramjet engine can be reduced, the design of a propulsion system is facilitated, and the thrust-weight ratio of the engine is improved; 4. the device can work in an air suction mode or a rocket mode, and the working range can be changed from subsonic speed to high Mach number supersonic speed. Therefore, the research of the continuous rotation knocking engine gradually draws a great attention of the scientific field.
At present, a great deal of results and relatively many experiences are obtained and accumulated in the research of the continuous rotation knocking engine, but the research on the minimum section, the shortest length, a non-linear flow passage and the like of the knocking combustion chamber is relatively little. Much current research is directed to the cross-section of a detonation combustor, but the non-linear flow channels are relatively less researched than the cross-section's effect on the formation of pressure waves and the circumferential and axial propagation of the converted detonation waves.
Disclosure of Invention
In view of this, the embodiments of the present disclosure provide a non-linear detonation combustor to solve the problem of applying a continuous rotation detonation engine in different scenarios.
The embodiment of the specification provides the following technical scheme:
a nonlinear detonation combustor comprises an oxidant and fuel injection assembly, a combustor assembly and a tail nozzle assembly which are communicated in sequence from an inlet end to an outlet end;
the oxidant and fuel injection assembly comprises an external fuel inlet, an external oxidant inlet, an oxidant pressure stabilizing cavity and a fuel pressure stabilizing cavity, the external fuel inlet is communicated with the fuel pressure stabilizing cavity, the external oxidant inlet is communicated with the oxidant pressure stabilizing cavity, and the outlet end of the oxidant pressure stabilizing cavity and the outlet end of the fuel pressure stabilizing cavity are both communicated with the inlet end of the combustion chamber assembly;
the combustion chamber assembly comprises a combustion chamber outer tube shell, a circumferential seam detonation combustion chamber and a combustion chamber inner ring body, the combustion chamber outer tube shell is coaxially sleeved outside the combustion chamber inner ring body, and the circumferential seam detonation combustion chamber is a circumferential seam gap between the combustion chamber outer tube shell and the combustion chamber inner ring body;
the tail nozzle assembly comprises a tail nozzle, and the inlet end of the tail nozzle is communicated with the outlet end of the combustion chamber assembly.
Further, the combustor assembly has an axis, and the axis is curvilinear.
Furthermore, the combustion chamber component is one or a combination of a plurality of L-shaped, S-shaped, spiral-shaped, U-shaped and semi-circular.
Furthermore, the oxidant and fuel injection assembly further comprises a combustion chamber connecting and rotating device, the oxidant pressure stabilizing cavity and the fuel pressure stabilizing cavity are detachably connected to the combustion chamber assembly through the combustion chamber connecting and rotating device, and the combustion chamber assembly can axially rotate around the central axis of the oxidant and fuel injection assembly through the combustion chamber connecting and rotating device.
Further, the combustion chamber is connected with a rotating device, is driven by a motor and rotates according to a set angle.
Further, the jet nozzle assembly further comprises a combustor jet nozzle connection device, and the jet nozzle is detachably connected to the combustor assembly through the combustor jet nozzle connection device.
Furthermore, the tail nozzle is of a cone-shaped structure, the large-diameter end of the tail nozzle is connected with the combustion chamber tail nozzle connecting device, the small-diameter end of the tail nozzle is arranged on the outer side of the combustion chamber tail nozzle connecting device, and the tail nozzle is provided with an axially-through jet hole.
Further, the inner ring body of the combustion chamber is of a hollow ring body structure.
Furthermore, an automatic control system is arranged at the external fuel inlet and can control the injection time and flow of the fuel.
Furthermore, an automatic control system is arranged at the external oxidant inlet and can control the injection time and flow of the oxidant.
Compared with the prior art, the beneficial effects that can be achieved by the at least one technical scheme adopted by the embodiment of the specification at least comprise:
the problem that the fuel and the oxidant can be fully mixed only after being injected into the combustion chamber and needing a certain length of circulation under the condition that the axial arrangement space of the current non-premixed detonation combustion chamber is limited is solved. When the detonation combustion chamber is mainly used for creating a good environment for the work of the detonation engine, the space of the combustion chamber is filled with the explosive combustible, basic conditions are created for the continuous generation and transmission of detonation waves, the continuous work of the detonation engine is achieved, the nonlinear flow channel mainly can realize good mixing of the oxidant and the fuel under the condition that the combustion chamber meets the short-length design, vector control (engine transverse arrangement and rotatable design) of the engine or the spray pipe and an engine injection cavity are arranged side by side, and the length of the whole engine can be greatly shortened. Through the design of a nonlinear flow channel, the thrust direction of an engine is controlled, and the functions of horizontal installation of an aircraft engine, vertical take-off and landing of the aircraft, rapid avoidance of attack in the air, multi-dimensional direction change of a guided missile and the like are realized.
Drawings
In order to more clearly illustrate the technical solutions of the embodiments of the present application, the drawings needed to be used in the embodiments will be briefly described below, and it is obvious that the drawings in the following description are only some embodiments of the present application, and it is obvious for those skilled in the art to obtain other drawings based on these drawings without creative efforts.
FIG. 1 is a sectional view of a detonation combustor of an L-shaped combustor of a first embodiment of the present invention;
FIG. 2 is a perspective view of a detonation combustor of the first embodiment L-shaped combustor of the present invention;
FIG. 3 is a cross-sectional view of a detonation combustor of a second embodiment S-type combustor of the present invention;
FIG. 4 is a perspective view of a detonation combustor of a second embodiment S-type combustor of the present invention;
fig. 5 is a perspective view of a knocking combustion chamber of a helical combustion chamber of a third embodiment of the present invention.
Description of the reference numerals: 1. is externally connected with a fuel inlet; 2. is externally connected with an oxidant inlet; 3. an oxidant pressure stabilizing cavity; 4. a fuel plenum chamber; 5. the combustion chamber is connected with the rotating device; 6. a combustion chamber outer tube housing; 7. a circular seam detonation combustor; 8. a combustion chamber inner ring; 9. a tail nozzle; 10. an oxidant and fuel injector assembly; 11. a combustor tail nozzle connecting device; 12. a tail nozzle assembly.
Detailed Description
The embodiments of the present application will be described in detail below with reference to the accompanying drawings.
The following description of the embodiments of the present application is provided by way of specific examples, and other advantages and effects of the present application will be readily apparent to those skilled in the art from the disclosure herein. It is to be understood that the embodiments described are only a few embodiments of the present application and not all embodiments. The present application is capable of other and different embodiments and its several details are capable of modifications and/or changes in various respects, all without departing from the spirit of the present application. It should be noted that the features in the following embodiments and examples may be combined with each other without conflict. All other embodiments obtained by a person of ordinary skill in the art based on the embodiments in the present application without making any creative effort belong to the protection scope of the present application.
It is noted that various aspects of the embodiments are described below within the scope of the appended claims. It should be apparent that the aspects described herein may be embodied in a wide variety of forms and that any specific structure and/or function described herein is merely illustrative. Based on the present application, one skilled in the art should appreciate that one aspect described herein may be implemented independently of any other aspects and that two or more of these aspects may be combined in various ways. For example, an apparatus may be implemented and/or a method practiced using any number and aspects set forth herein. Additionally, such an apparatus may be implemented and/or such a method may be practiced using other structure and/or functionality in addition to one or more of the aspects set forth herein.
It should be noted that the drawings provided in the following embodiments are only for illustrating the basic idea of the present application, and the drawings only show the components related to the present application rather than the number, shape and size of the components in actual implementation, and the type, amount and ratio of the components in actual implementation may be changed arbitrarily, and the layout of the components may be more complicated.
In addition, in the following description, specific details are provided to facilitate a thorough understanding of the examples. However, it will be understood by those skilled in the art that the aspects may be practiced without these specific details.
The combustion chamber of the continuous rotation detonation engine in the prior art has the following defects:
1. the combustion chamber has a single flow channel, is a straight flow channel, has longer required axial length, is used as a core component of the detonation engine, and basically determines most length dimensions of the engine;
2. most of the mixing of the oxidant and the fuel is ensured by the axial length, and once the length space is insufficient, the mixing effect directly influences the efficient and stable operation of the detonation engine;
3. the engine is generally arranged in the same direction as the flow passage of a combustion chamber, so that the thrust direction of the engine can only be arranged backwards, or the non-coaxial arrangement of the engine and a nozzle is realized by a vectoring nozzle which is not economically applied.
The detonation combustion chamber mainly creates a good environment for the work of the detonation engine, so that the space of the combustion chamber is filled with explosive combustible, basic conditions are created for the continuous generation and transmission of detonation waves, and the continuous work of the detonation engine is achieved.
Terms used in the embodiments of the present invention are explained.
Detonation (Detonation): the combustion mode is a combustion mode coupling shock waves and flames (chemical reaction), the chemical reaction speed is high, the flame propagation speed is high and can reach 1000+ km/s, and extremely high pressure and temperature can be generated. The detonation wave generates extremely high gas pressure (1.5-5.5 MPa) and extremely high gas temperature (more than 2800K).
The technical solutions provided by the embodiments of the present application are described below with reference to the accompanying drawings.
Fig. 1 and 2 show an L-shaped detonation combustor according to a first embodiment of the present invention, which is composed of three parts, namely an oxidizer and fuel injection assembly 10, a combustor assembly and a tail nozzle assembly 11, which are connected with each other in the order of the oxidizer and fuel injection assembly 10, the combustor assembly and the tail nozzle assembly 11 from the air inlet side to the air outlet side. Wherein, the combustion chamber component is an L-shaped nonlinear runner combustion chamber.
The combustor assembly includes a combustor outer tube housing 6, a cyclic detonation combustor 7, and a combustor inner ring 8. The outer tube shell 6 of the combustion chamber is coaxially sleeved outside the inner ring body 8 of the combustion chamber, and the circumferential seam detonation combustion chamber 7 is a circumferential seam gap between the outer tube shell 6 of the combustion chamber and the inner ring body 8 of the combustion chamber. The circular seam detonation combustor 7 is a core part of a detonation engine, a main space for fuel mixed combustion, continuous rotation detonation is a power technology utilizing detonation combustion, and the continuous rotation detonation combustor is mostly in a circular seam form. The inner part of the inner ring body 8 of the combustion chamber is hollow, and the weight of the whole combustion chamber component can be reduced as much as possible due to the hollow structure. The whole combustion chamber assembly is L-shaped, and the inlet end and the outlet end of the combustion chamber assembly have a certain included angle.
The oxidant and fuel injection assembly 10 comprises an external fuel inlet 1, an external oxidant inlet 2, an oxidant pressure stabilizing cavity 3, a fuel pressure stabilizing cavity 4 and a combustion chamber connecting rotary module 5. The fuel enters the fuel pressure stabilizing cavity 4 through the external fuel inlet 1, and the oxidant enters the fuel pressure stabilizing cavity 4 through the external oxidant inlet 2. The oxidant pressure stabilizing cavity 3 and the fuel pressure stabilizing cavity 4 are connected with the rotary module 5 and the combustion chamber component through a combustion chamber. The combustion chamber is connected with the rotating module 5, the rotating device is arranged inside the combustion chamber connecting rotating module, the rotating device is located on one side of an interface of a combustion chamber component, the rotating device is of a gear structure, a mounting driving motor is installed on the outer portion of the rotating device, and the driving gear of the driving motor achieves the purpose of rotation. The combustor assembly and the jet nozzle 9 are rotatable relative to the oxidizer and fuel injection assembly 10 by the combustor attachment rotation module 5.
The tail nozzle assembly 11 comprises a tail nozzle 9 and a combustion chamber tail nozzle connecting device, and the tail nozzle 9 is detachably connected to the outlet end of the combustion chamber assembly through the combustion chamber tail nozzle connecting device.
The embodiment also comprises an automatic control system, wherein the automatic control system comprises an electromagnetic valve, a flowmeter and a power supply. The automatic control system can control the injection of fuel and oxidant according to the settings.
When the engine works, the fuel is timely, properly and orderly controlled to be injected by controlling the electromagnetic valve and the flowmeter which are connected with the fuel tank and fuel through the external fuel inlet 1, the fuel enters the fuel pressure stabilizing cavity 4 after being switched on, a fine circular or gap channel is designed between the fuel pressure stabilizing cavity 4 and the combustion chamber, and the fuel is sprayed into the circumferential seam detonation combustion chamber 7 in a good atomization state through the fine channels. Similarly, the engine controls the injection of proper and orderly oxidant in time by controlling an electromagnetic valve and a flowmeter which are connected with an oxidant box and externally connected with an oxidant inlet 2, the oxidant is firstly injected into an oxidant pressure stabilizing cavity 3, then is accelerated and sprayed into a circumferential seam detonation combustion chamber 7 through a contraction structure connected with the combustion chamber, and is fully mixed with the sprayed fuel, and an igniter is arranged on an outer tube shell 6 of the combustion chamber and is ignited in time. The rotating module 5 is connected with the combustion chamber to enable the outer tube shell 6 of the combustion chamber and the inner ring body 8 of the combustion chamber to rotate and drive the tail nozzle 9 to rotate, so that the aim of adjusting the direction of the nozzle of the tail nozzle 9 is fulfilled. The thrust direction of the engine can be controlled by changing the direction of the nozzle, so that the horizontal installation of the engine of the aircraft is realized, and the vertical take-off and landing of the aircraft are realized; the missile can quickly avoid attack in the air, and the missile can change direction in multiple dimensions. Through the design of the nonlinear flow channel, the size L of the detonation engine in the length direction can be shortened, and the detonation engine is suitable for an aircraft with strict length control. In addition, compared with a linear flow channel, the nonlinear flow channel has the same space length, the mixing path of the fuel and the oxidant is longer, the mixing of the fuel and the oxidant is more sufficient, and the combustion is easier to realize.
The L-shaped detonation combustor is characterized in that an independent nonlinear flow channel design is carried out on a main combustion area of the combustor, an oxidant and fuel injection assembly 10 is designed into a detachable module, the oxidant and fuel injection assembly 10, namely an upstream injection system of the combustor, is connected with the main combustion area of the combustor, namely the independent nonlinear flow channel combustor, through a combustor connecting rotary module 5, and a downstream tail nozzle 9 is connected with the combustor through a detachable combustor tail nozzle connecting device, so that the tail nozzle 9 is replaceable.
As shown in fig. 1, L is the radial distance of the fuel and oxidant injection inlets and the jet nozzle 9. The L-shaped non-linear flow path combustor can accommodate aircraft with short horizontal distances between the inlet and outlet ends, with the fuel and oxidant injection inlets and the jet ports of the jet nozzle 9 being the inlet and outlet ends.
Fig. 3 and 4 show a second embodiment S-type detonation combustor according to the invention, comprising three main assemblies of oxidizer and fuel injector assembly 10, combustor assembly and tailpipe assembly 11, which are interconnected in the order oxidizer and fuel injector assembly 10, combustor assembly and tailpipe assembly 11 from inlet to outlet. Wherein the shape of the combustion chamber component is an S-shaped nonlinear runner combustion chamber.
The S-shaped non-linear flow channel combustor accommodates aircraft having inlet and outlet ports that are horizontal and vertical, with one port being a short distance and the other port being a long distance, using the fuel and oxidant injection inlets and the jet ports of the jet nozzle 9 as the inlet and outlet ports.
Fig. 5 shows a third embodiment of the present invention of a helical detonation combustor comprising three main components, an oxidizer and fuel injector assembly 10, a combustor assembly and a tail nozzle assembly 11, interconnected in the order of oxidizer and fuel injector assembly 10, combustor assembly and tail nozzle assembly 11 from inlet to outlet. Wherein, the combustion chamber component is a spiral nonlinear runner combustion chamber.
The spiral, nonlinear flow channel combustor accommodates aircraft having relatively short horizontal and vertical distances between the inlet and outlet ends, with the inlet and outlet ends being substantially horizontal, using fuel and oxidant injection inlets and the jet ports of the jet nozzle 9 as the inlet and outlet ends.
In other embodiments, the combustor assembly may be a U-shaped, non-linear flow channel combustor that accommodates aircraft with closely spaced inlet and outlet ports, using fuel and oxidant injection inlets and nozzles from the jet nozzle 9 as the inlet and outlet ports.
In other embodiments, the combustor assembly may also be semi-circular, or other shapes having axes that are curvilinear.
The combustion space required by the continuous detonation combustion chamber can be ensured while the overall length of the combustion chamber with the nonlinear flow channel is shortened; due to the design of the nonlinear flow channel, the full mixing of the fuel and the oxidant is realized, and the combustion efficiency is improved; adapting different application scenes and engine thrust requirements by replacing the tail nozzles 9 with different expansion ratios; the outlet position of the tail nozzle 9 can be arranged according to the requirements of application scenes; the main body part of the combustion chamber can be rotated to adapt to arrangement or realize real-time adjustment of the nozzle direction by setting the rotating mechanism in the working process of the engine, so that vector adjustment of the spray pipe is realized. The shape of the combustor assembly is modified to accommodate aircraft of different configurations, depending on the specific requirements of the various aircraft shapes. The application field of the detonation engine is widened through the detonation engine with the non-linear flow passage, so that the aircraft which cannot use the detonation engine due to the problems of space and the like can also use the detonation engine.
The embodiments in the present specification are described in a progressive manner, and the same and similar parts among the embodiments can be referred to each other, and each embodiment focuses on the differences from the other embodiments. In particular, for the method embodiments described later, since they correspond to the system, the description is simple, and for the relevant points, reference may be made to the partial description of the system embodiments.
The above description is only for the specific embodiments of the present application, but the scope of the present application is not limited thereto, and any changes or substitutions that can be easily conceived by those skilled in the art within the technical scope of the present application should be covered within the scope of the present application. Therefore, the protection scope of the present application shall be subject to the protection scope of the claims.
Claims (10)
1. A non-linear detonation combustor comprising an oxidizer and fuel injector assembly (10), a combustor assembly and a tail nozzle assembly (12) in sequential communication from an inlet end to an outlet end;
the oxidant and fuel injection assembly (10) comprises an external fuel inlet (1), an external oxidant inlet (2), an oxidant pressure stabilizing cavity (3) and a fuel pressure stabilizing cavity (4), wherein the external fuel inlet (1) is communicated with the fuel pressure stabilizing cavity (4), the external oxidant inlet (2) is communicated with the oxidant pressure stabilizing cavity (3), and the outlet ends of the oxidant pressure stabilizing cavity (3) and the fuel pressure stabilizing cavity (4) are communicated with the inlet end of the combustion chamber assembly;
the combustion chamber component comprises a combustion chamber outer tube shell (6), a circumferential seam detonation combustion chamber (7) and a combustion chamber inner ring body (8), the combustion chamber outer tube shell (6) is coaxially sleeved outside the combustion chamber inner ring body (8), and the circumferential seam detonation combustion chamber (7) is a circumferential seam gap between the combustion chamber outer tube shell (6) and the combustion chamber inner ring body (8);
the tail nozzle assembly (12) comprises a tail nozzle (9), and the inlet end of the tail nozzle (9) is communicated with the outlet end of the combustion chamber assembly.
2. The non-linear detonation combustor of claim 1, wherein the combustor assembly has an axis and the axis is curvilinear.
3. The non-linear detonation combustor of claim 2, wherein the combustor assembly is one or a combination of L-shaped, S-shaped, spiral-shaped, U-shaped, and semi-circular shaped.
4. The non-linear detonation combustor of claim 1, characterized in that the oxidizer and fuel injection assembly (10) further comprises a combustor attachment rotating means (5), both the oxidizer plenum (3) and the fuel plenum (4) being detachably connected to the combustor assembly by the combustor attachment rotating means (5), the combustor assembly being axially rotatable about a central axis of the oxidizer and fuel injection assembly (10) by the combustor attachment rotating means (5).
5. The non-linear detonation combustor according to claim 4, characterised in that the combustor connection rotation device (5) is driven by a motor and rotates according to a set angle.
6. The non-linear detonation combustor of claim 1, characterized in that the tailpipe assembly (12) further includes a combustor tailpipe connection (11), the tailpipe (9) being detachably connected to the combustor assembly by the combustor tailpipe connection (11).
7. The nonlinear detonation combustor according to claim 6, characterized in that the jet nozzle (9) is of a cone-shaped structure, the large diameter end of the jet nozzle (9) is connected with the combustor jet nozzle connecting device (11), the small diameter end of the jet nozzle (9) is arranged outside the combustor jet nozzle connecting device (11), and the jet nozzle (9) is provided with an axially through jet hole.
8. The non-linear detonation combustor of claim 1, characterised in that the combustor inner annulus (8) is of internally hollow annulus construction.
9. The non-linear detonation combustor of claim 1, characterised in that the external fuel inlet (1) is provided with an automatic control system capable of controlling the timing and flow of fuel injection.
10. The non-linear detonation combustor of claim 1, characterised in that the external oxidant inlet (2) is provided with an automatic control system capable of controlling the time and flow of oxidant injection.
Priority Applications (1)
| Application Number | Priority Date | Filing Date | Title |
|---|---|---|---|
| CN202210826310.3A CN115342380A (en) | 2022-07-13 | 2022-07-13 | Nonlinear detonation combustion chamber |
Applications Claiming Priority (1)
| Application Number | Priority Date | Filing Date | Title |
|---|---|---|---|
| CN202210826310.3A CN115342380A (en) | 2022-07-13 | 2022-07-13 | Nonlinear detonation combustion chamber |
Publications (1)
| Publication Number | Publication Date |
|---|---|
| CN115342380A true CN115342380A (en) | 2022-11-15 |
Family
ID=83947576
Family Applications (1)
| Application Number | Title | Priority Date | Filing Date |
|---|---|---|---|
| CN202210826310.3A Pending CN115342380A (en) | 2022-07-13 | 2022-07-13 | Nonlinear detonation combustion chamber |
Country Status (1)
| Country | Link |
|---|---|
| CN (1) | CN115342380A (en) |
Citations (14)
| Publication number | Priority date | Publication date | Assignee | Title |
|---|---|---|---|---|
| EP0478881A1 (en) * | 1990-09-24 | 1992-04-08 | Franz Gähler | Process and device for processing fragmentary parts and/or flowable materials |
| US5758611A (en) * | 1997-04-15 | 1998-06-02 | Collins; Imack L. | Flex-rod |
| CN1252119A (en) * | 1997-04-17 | 2000-05-03 | 联合讯号公司 | Connector tube for linking cold plenum to hot chamber |
| CN101021307A (en) * | 2007-03-14 | 2007-08-22 | 梁兆志 | Efficient low-discharging heat accumulating gas boiler |
| CN101865459A (en) * | 2010-06-11 | 2010-10-20 | 杭州电子科技大学 | Coking-proof multistage ignition combustion apparatus with self-adaption function |
| CN205315112U (en) * | 2015-12-22 | 2016-06-15 | 重庆新颜达机电设备有限公司 | Energy -conserving gasoline engine |
| CN106762220A (en) * | 2016-12-22 | 2017-05-31 | 清华大学 | Turbogenerator |
| CN109458271A (en) * | 2018-11-07 | 2019-03-12 | 厦门大学 | A kind of rotation detonation engine air intake duct and jet pipe integrated design method |
| CN109642519A (en) * | 2016-08-25 | 2019-04-16 | 康宁股份有限公司 | Segmentation thermal barriers and its manufacturing method for internal combustion engine |
| CN109736935A (en) * | 2018-12-30 | 2019-05-10 | 北京工业大学 | Using purity oxygen as the control method of the nitrogen discharged engine of HC fuel spark ignition type zero of oxidant |
| CN212716929U (en) * | 2020-01-15 | 2021-03-16 | 清华大学 | Continuous rotation detonation rocket engine manufactured by additive |
| CN112879178A (en) * | 2021-01-22 | 2021-06-01 | 南京理工大学 | Solid rocket ramjet based on detonation combustion |
| CN113513429A (en) * | 2021-04-16 | 2021-10-19 | 中国人民解放军战略支援部队航天工程大学 | Engine and method capable of realizing tangential unstable combustion and continuous rotation detonation |
| CN113739206A (en) * | 2021-09-22 | 2021-12-03 | 西北工业大学 | Zoned combustion scheme for improving space utilization rate of rotary detonation combustion chamber |
-
2022
- 2022-07-13 CN CN202210826310.3A patent/CN115342380A/en active Pending
Patent Citations (14)
| Publication number | Priority date | Publication date | Assignee | Title |
|---|---|---|---|---|
| EP0478881A1 (en) * | 1990-09-24 | 1992-04-08 | Franz Gähler | Process and device for processing fragmentary parts and/or flowable materials |
| US5758611A (en) * | 1997-04-15 | 1998-06-02 | Collins; Imack L. | Flex-rod |
| CN1252119A (en) * | 1997-04-17 | 2000-05-03 | 联合讯号公司 | Connector tube for linking cold plenum to hot chamber |
| CN101021307A (en) * | 2007-03-14 | 2007-08-22 | 梁兆志 | Efficient low-discharging heat accumulating gas boiler |
| CN101865459A (en) * | 2010-06-11 | 2010-10-20 | 杭州电子科技大学 | Coking-proof multistage ignition combustion apparatus with self-adaption function |
| CN205315112U (en) * | 2015-12-22 | 2016-06-15 | 重庆新颜达机电设备有限公司 | Energy -conserving gasoline engine |
| CN109642519A (en) * | 2016-08-25 | 2019-04-16 | 康宁股份有限公司 | Segmentation thermal barriers and its manufacturing method for internal combustion engine |
| CN106762220A (en) * | 2016-12-22 | 2017-05-31 | 清华大学 | Turbogenerator |
| CN109458271A (en) * | 2018-11-07 | 2019-03-12 | 厦门大学 | A kind of rotation detonation engine air intake duct and jet pipe integrated design method |
| CN109736935A (en) * | 2018-12-30 | 2019-05-10 | 北京工业大学 | Using purity oxygen as the control method of the nitrogen discharged engine of HC fuel spark ignition type zero of oxidant |
| CN212716929U (en) * | 2020-01-15 | 2021-03-16 | 清华大学 | Continuous rotation detonation rocket engine manufactured by additive |
| CN112879178A (en) * | 2021-01-22 | 2021-06-01 | 南京理工大学 | Solid rocket ramjet based on detonation combustion |
| CN113513429A (en) * | 2021-04-16 | 2021-10-19 | 中国人民解放军战略支援部队航天工程大学 | Engine and method capable of realizing tangential unstable combustion and continuous rotation detonation |
| CN113739206A (en) * | 2021-09-22 | 2021-12-03 | 西北工业大学 | Zoned combustion scheme for improving space utilization rate of rotary detonation combustion chamber |
Similar Documents
| Publication | Publication Date | Title |
|---|---|---|
| CN113294264B (en) | Double-component variable-thrust rotary detonation rocket engine based on pintle injector | |
| US10641169B2 (en) | Hybrid combustor assembly and method of operation | |
| JP4555654B2 (en) | Two-stage pulse detonation system | |
| CN109028151B (en) | Multi-chamber rotary detonation combustor | |
| CN109028142B (en) | Propulsion system and method of operating the same | |
| US7669406B2 (en) | Compact, low pressure-drop shock-driven combustor and rocket booster, pulse detonation based supersonic propulsion system employing the same | |
| US8205433B2 (en) | Pulse detonation/deflagration apparatus and related methods for enhancing DDT wave production | |
| CN112902225B (en) | Multistage afterburning chamber with outer ring rotary detonation supercharged combustion chamber | |
| US20110047962A1 (en) | Pulse detonation combustor configuration for deflagration to detonation transition enhancement | |
| CN109028149A (en) | The rotation detonating combustion device and its operating method of geometry-variable | |
| US8438834B2 (en) | Helical cross flow (HCF) pulse detonation engine | |
| CN110131074B (en) | Bipropellant air turbine rocket propulsion system | |
| CN112728585B (en) | System for rotary detonation combustion | |
| CN113154458B (en) | Continuous rotation detonation combustion chamber and ramjet | |
| US10520195B2 (en) | Effervescent atomizing structure and method of operation for rotating detonation propulsion system | |
| CN115342380A (en) | Nonlinear detonation combustion chamber | |
| CN114877378A (en) | Inner ring detonation combustion chamber | |
| CN113154451B (en) | Guide spray pipe of rotary detonation combustion chamber | |
| CN114877377A (en) | Outer ring detonation combustion chamber | |
| CN115342382A (en) | Single-channel oxygen supply detonation combustion chamber module and detonation combustion chamber | |
| CN114877376A (en) | Double-channel detonation combustion chamber | |
| CN113864824B (en) | Pre-explosion tube applicable to rotary detonation combustion chamber with variable length | |
| CN115342381A (en) | Detonation combustion chamber module and detonation combustion chamber | |
| CN115263609A (en) | Modular secondary detonation engine | |
| Banga et al. | Design and Study of Swirl Injector of Pulse Detonation Engine |
Legal Events
| Date | Code | Title | Description |
|---|---|---|---|
| PB01 | Publication | ||
| PB01 | Publication | ||
| SE01 | Entry into force of request for substantive examination | ||
| SE01 | Entry into force of request for substantive examination |