CN116696596A - Tandem double-combustion-chamber rotary knocking ramjet engine and working method - Google Patents
Tandem double-combustion-chamber rotary knocking ramjet engine and working method Download PDFInfo
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- 238000000034 method Methods 0.000 title claims abstract description 15
- 238000002485 combustion reaction Methods 0.000 claims abstract description 147
- 238000002347 injection Methods 0.000 claims abstract description 107
- 239000007924 injection Substances 0.000 claims abstract description 107
- 239000000446 fuel Substances 0.000 claims abstract description 65
- 238000005474 detonation Methods 0.000 claims abstract description 62
- 238000002955 isolation Methods 0.000 claims abstract description 35
- 230000008602 contraction Effects 0.000 claims abstract description 32
- 239000007921 spray Substances 0.000 claims abstract description 17
- 238000013461 design Methods 0.000 claims description 16
- 230000009977 dual effect Effects 0.000 claims description 9
- 239000007788 liquid Substances 0.000 claims description 9
- 239000007789 gas Substances 0.000 claims description 7
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- 239000011152 fibreglass Substances 0.000 claims description 3
- 239000002737 fuel gas Substances 0.000 claims description 3
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- 238000002156 mixing Methods 0.000 description 4
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- 239000002828 fuel tank Substances 0.000 description 3
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- 238000009423 ventilation Methods 0.000 description 2
- 238000006243 chemical reaction Methods 0.000 description 1
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Abstract
The invention provides a tandem double-combustion-chamber rotary detonation ramjet engine and a working method thereof, wherein the tandem double-combustion-chamber rotary detonation ramjet engine comprises a combustion chamber outer shell, an air inlet channel, an isolation section, a first-stage combustion chamber, a first-stage contraction section, a second-stage combustion chamber, a second-stage contraction section, an outlet spray pipe, a combustion chamber central body, a first-stage fuel injection cavity, a second-stage fuel injection cavity, a first-stage fuel spray hole, a second-stage fuel spray hole, a central body front support and a central body rear support; considering the situation that the combustion back pressure forwarding distance is large in the high/low incoming flow Mach number working state, the problem of combustion back pressure forwarding under the wide incoming flow Mach number condition can be solved by arranging two combustion chambers in series and reasonably arranging the combustion organization working modes.
Description
Technical Field
The invention belongs to the technical field of rotary detonation ramjet engines, and particularly relates to a tandem double-combustion-chamber rotary detonation ramjet engine and a working method thereof.
Background
The rotary detonation combustion is used as a novel combustion mode, has the advantages of high heat release speed, continuous operation after one ignition, higher heat efficiency and the like, has unique advantages in the aspects of shortening the length of a combustion chamber, lightening the weight of a structure, improving the propulsion performance and the like, can be applied to an aero-engine, a rocket engine and a ramjet engine to replace the existing detonation combustion mode, and has wide application prospect in the field of aerospace propulsion. The rotary knock ramjet engine is a novel ramjet engine type which burns by rotary knock tissue, and the main working process is as follows: the high-speed air flow is captured by the air inlet channel and decelerated, enters the annular combustion chamber through the isolation section, and is fully mixed with fuel injected by the spray holes to form combustible mixed gas. Then, by ignition and detonation, detonation waves are formed in the annular combustion chamber and are propagated at high speed and high frequency along the circumferential direction of the combustion chamber, and rapid heat release of fuel is realized through the detonation waves. The detonation wave is followed by high temperature combustion products which are spread downwards along the flow direction and are accelerated and discharged through the spray pipe so as to realize the thrust conversion of the engine.
The detonation wave has good incoming flow adaptability, earlier research work has shown that rotary detonation in a stamping mode can realize efficient and stable combustion under different incoming flow Mach numbers, but in different Mach number ranges, the combustion chamber configuration when the rotary detonation stable combustion structure is different, and the rotary detonation stable and efficient combustion in a wide Mach number range is difficult to realize in a single combustion chamber. In addition, under different flight Mach numbers, the incoming flow state of high-speed air is changed drastically, and the capability of the isolation section for resisting the combustion back pressure is greatly different. Under the condition of higher Mach number, the laser front transmission distance of the isolation section is shorter, and the length of the isolation section is not required to be too long; and at lower Mach numbers, the shock wave forward distance of the isolation section is obviously prolonged, and a longer isolation section is needed to avoid the combustion back pressure forward. The length of the isolation section in the high or low incoming flow Mach number range is required to be greatly different, and if stable operation in the wide incoming flow Mach number range of the rotary detonation ramjet engine is required to be realized, the design of the isolation section is required to be reasonably considered.
Therefore, it is important to design a novel rotary detonation combustor configuration with controllable back pressure forward transmission and high combustion performance in a wide Mach number range under the same flow channel to solve the above problems. The invention provides a combustion chamber configuration of a tandem double-combustion-chamber rotary detonation ramjet engine, which can solve the problems.
Disclosure of Invention
In order to achieve the above purpose, the technical scheme of the invention is as follows:
the rotary knocking ramjet engine with the tandem double combustion chambers comprises an outer shell 1, wherein a central body 2 is arranged in the outer shell 1, and the central body 2 is connected with the outer shell 1 through a front central body support 3 and a rear central body support 4;
an annular flow passage is formed between the outer shell 1 and the central body 2, and sequentially comprises an air inlet passage 5, an isolation section 6, a first-stage combustion chamber 10, a first-stage contraction section 11, a second-stage combustion chamber 14, a second-stage contraction section 15 and an outlet spray pipe 16 along the airflow propagation direction;
a fuel storage tank 9 is arranged in the center body 2; the central body front support 3 is positioned at the isolation section 6 between the air inlet channel 5 and the first-stage combustion chamber 10, and the central body rear support 4 is positioned at the second-stage contraction section 15 between the second-stage combustion chamber 14 and the outlet spray pipe 16; in order to ensure the ventilation flow, the front support 3 of the central body and the rear support 4 of the central body are uniformly distributed in the circumferential direction;
the first-stage injection cavity 7 and the second-stage injection cavity 12 are of a double-injection cavity structure, and the double-injection cavities are respectively arranged on the outer shell 1 and the central body 2; along the airflow propagation direction, the first-stage injection cavity 7 is positioned between the tail end of the isolation section 6 far away from the incoming flow and the front end of the first-stage combustion chamber 10 near the incoming flow; along the air flow propagation direction, the second-stage injection cavity 12 is positioned between the tail end of the first-stage contraction section 11, which is far away from the incoming flow, and the front end of the second-stage combustion chamber 14, which is close to the incoming flow;
the outer shell 1 and the central body 2 corresponding to the first-stage injection cavity 7 are provided with first-stage injection holes 8, and the first-stage injection holes 8 are communicated to the first-stage injection cavity 7 through a runner; the outer shell 1 and the central body 2 corresponding to the second-stage injection cavity 12 are provided with second-stage injection holes 13, and the second-stage injection holes 13 are communicated to the second-stage injection cavity 12 through flow channels; the first-stage injection holes 8 and the second-stage injection holes 13 are uniformly distributed along the circumferential direction.
Wherein the outer shell 1 is used for supporting the central body 2;
an annular flow passage is formed between the central body 2 and the outer shell 1;
the front center body support 3 and the rear center body support 4 are used for supporting the center body;
the air inlet channel 5 is used for capturing incoming flow and forming annular air flow;
the fuel tank 9 is used for storing fuel;
the first-stage injection cavity 7 and the second-stage injection cavity 12 are used for uniformly filling fuel along the circumferential direction of the engine;
the first-stage injection holes 8 and the second-stage injection holes 13 are used for injecting fuel;
the first stage combustion chamber 10 and the second stage combustion chamber 14 are used for ignition initiation and form rotary detonation waves in the combustion chambers;
the outlet nozzle 16 is used for converting the high-temperature air flow into high-speed air flow so as to generate thrust;
under the low Mach number incoming flow condition, the second-stage combustion chamber 14 ignites and detonates, and the isolation section 6, the first-stage combustion chamber 10 and the first-stage contraction section 11 are used for isolating the back pressure of the second-stage combustion chamber 14 from entering the air inlet channel 5;
the isolation section 6 is used to isolate the back pressure of the first stage combustion chamber 10 from entering the inlet 5 under high Mach number inflow conditions.
Preferably, the first stage combustion chamber width delta 1 Second stage combustor width delta 2 The following formula is satisfied: delta is more than or equal to 0.5 lambda, and when liquid fuel is adopted, delta is more than or equal to d; first stage combustor lengthL 1 Second stage combustor lengthL 2 The following formula needs to be satisfied:Lnot less than 2 (12+/-5) lambda, wherein lambda is the size of detonation wave cell corresponding to the mixed gas under the current combustion chamber pressure, and d is the minimum diameter of fuel liquid drop;
on the basis of meeting the requirements, the profile structural design of the first-stage combustion chamber 10 takes Mach number Ma=5 incoming flow conditions as standard design points, and rotary detonation combustion in the Mach number Ma=4-6 incoming flow range is realized; the profile structural design of the second stage combustion chamber 14 takes Mach number Ma=3 incoming flow conditions as standard design points, and rotary detonation combustion in the Mach number Ma=2-4 incoming flow range is realized.
Preferably, the two-stage injection cavity structure of the first-stage injection cavity 7 and the second-stage injection cavity 12 has a square or circular injection cavity section, and the injection cavity section area is 4-25 mm 2 The fuel pressure of the injection cavity is 1-5 MPa.
As a preferable mode, the first-stage injection holes 8 and the second-stage injection holes 13 are uniformly distributed along the circumference, and the diameter of the spray holes is 0.2-0.6 mm for ensuring the mixing effect of fuel and air; for liquid fuel, the number of injection holes n satisfies the following formula:whereinmFor the fuel flow rate,C d for the injection orifice flow coefficient,Din order to inject the hole diameter,pin order to inject the cavity fuel pressure,ρis the fuel density; for gaseous fuels, the number of injection holes n satisfies +.>Wherein m is the fuel flow rate,T t is used for controlling the total temperature of the fuel,pin order to inject the cavity fuel pressure,C d for the injection orifice flow coefficient,Dfor the injection orifice diameter, K is a constant related to the fuel, and is derived from the following equation: />In which, in the process,γis the specific heat ratio of the glass fiber reinforced plastic material,Rfor the fuel gas constant, the diameter D of the injection hole is 0.2-0.6 mm to ensure the mixing effect of the fuel and the air.
Preferably, the flow area of the first stage constriction 11A th Smaller than the flow area of the first stage combustion chamber 10A c And the flow area of the second stage constriction 15A th Smaller than the flow area of the second stage combustion chamber 14A c The method comprises the steps of carrying out a first treatment on the surface of the The area contraction ratio AR is in the range of 1.2 to 2.5, and is the flow area of the first stage combustion chamber 10A c And the flow area of the first stage constriction 11A th Is the ratio of, or the flow area of, the second stage combustion chamber 14A c And the flow area of the second stage constriction 15A th Is a ratio of (2).
Preferably, the central body 2 is fixed to the inner center of the outer casing 1 by a central body front support 3 and a rear support 4.
Preferably, the high Mach number incoming flow condition is 4 < Ma.ltoreq.6;
the low Mach number incoming flow condition is that Ma is more than or equal to 2 and less than or equal to 4;
the high-temperature air flow means that the total temperature of the air flow is more than 2000K;
the high-speed air flow means that the air flow speed is more than 2 times of the sound speed.
The invention also provides a working method of the tandem double-combustion-chamber rotary detonation ramjet engine, which comprises the following steps:
under the low Mach number incoming flow condition, the incoming flow air is captured by the air inlet passage 5, enters the second-stage combustion chamber 14 through the isolation section 6, the first-stage combustion chamber 10 and the first-stage contraction section 11, is mixed with fuel injected into the second-stage injection hole 13, and is ignited and initiated in the second-stage combustion chamber 14 to form rotary detonation waves; the combustion products generated by the rotary detonation combustion are exhausted to the atmosphere through the second stage constriction 15 and the outlet nozzle 16, and thrust is generated; under the low Mach number incoming flow condition, the forward shock wave induced by the rotary detonation wave is isolated by a combined channel formed by the isolation section 6, the first-stage combustion chamber 10 and the first-stage contraction section 11, so that the combustion back pressure is not forwarded to the air inlet channel 5;
under the condition of high Mach number incoming flow, incoming flow air is captured by an air inlet passage 5 and enters a first-stage combustion chamber 10 through an isolation section 6, combustible mixed gas is formed after the incoming flow air is mixed with fuel injected by a first-stage injection hole 8, then an engine ignition device detonates a flow field in the first-stage combustion chamber 10 and forms rotary detonation waves in the combustion chamber, the detonation waves propagate along the circumferential direction and realize fuel consumption, and generated combustion products are discharged out of the atmosphere after passing through a first-stage contraction section 11, a second-stage combustion chamber 14, a second-stage contraction section 15 and an outlet spray pipe 16, so that thrust is generated; under the condition of high Mach number incoming flow, the forward shock wave induced by the rotary detonation wave is isolated by the isolation section 6, so that the combustion back pressure is not forwarded to the air inlet channel 5.
The beneficial effects of the invention are as follows: 1. according to the invention, aiming at the flow characteristics of the rotary detonation combustion structure in the high/low Mach number working state, the design can be respectively carried out aiming at the configurations of the first-stage combustion chamber and the second-stage combustion chamber, so that the design difficulty of the combustion chamber is reduced, the compatibility of the rotary detonation combustion working state in the same flow channel under the high/low Mach number incoming flow condition is realized, the high-efficiency combustion of the rotary detonation in the wide Mach number incoming flow condition is realized, the working Mach number range of the rotary detonation ramjet engine is widened, and the performance is improved; 2. considering the situation that the combustion back pressure forwarding distance is larger in the high/low incoming flow Mach number working state, the problem of combustion back pressure forwarding under the wide incoming flow Mach number condition can be solved by arranging two combustion chambers in series and reasonably arranging the combustion organization working modes; 3. by arranging the dual combustion chambers in the same flow passage, the flow passage can be fully utilized under the condition of high/low incoming flow Mach number, thereby being beneficial to reducing the volume of the engine, reducing the structural weight and reducing the structural complexity of the engine.
Drawings
FIG. 1 is a cross-sectional view of a flow path configuration of a tandem dual combustion chamber rotary detonation ramjet engine in accordance with the present invention.
FIG. 2 is a cross-sectional view of the flow path configuration of the rotary knock ramjet engine of section A-A of FIG. 1.
FIG. 3 is an enlarged view of a portion of a fuel injection chamber and an injection orifice according to the present invention.
Fig. 4 is a schematic diagram of an operation mode at a low incoming stream mach number according to the present invention.
Fig. 5 is a schematic diagram of an operation mode at a high incoming stream mach number according to the present invention.
Reference numerals
1. An outer housing; 2. a central body; 3. a center body front support; 4. a center body rear support; 5. an air inlet channel; 6. an isolation section; 7. a first stage injection chamber; 8. a first stage injection orifice; 9. a fuel tank; 10. a first stage combustion chamber; 11. a first stage constriction section; 12. a second stage injection chamber; 13. a second stage injection orifice; 14. a second stage combustion chamber; 15. a second stage constriction section; 16. an outlet nozzle.
Description of the embodiments
Other advantages and effects of the present invention will become apparent to those skilled in the art from the following disclosure, which describes the embodiments of the present invention with reference to specific examples. The invention may be practiced or carried out in other embodiments that depart from the specific details, and the details of the present description may be modified or varied from the spirit and scope of the present invention.
The embodiment provides a tandem double-combustion-chamber rotary detonation ramjet engine, the main structure of which is shown in fig. 1, comprising an outer shell 1, wherein a central body 2 is arranged in the outer shell 1, and the central body 2 is connected with the outer shell 1 through a central body front support 3 and a central body rear support 4;
an annular flow passage is formed between the outer shell 1 and the central body 2, and comprises an air inlet passage 5, an isolation section 6, a first-stage combustion chamber 10, a first-stage contraction section 11, a second-stage combustion chamber 14, a second-stage contraction section 15 and an outlet spray pipe 16 along the airflow propagation direction;
a fuel storage tank 9 is arranged in the center body 2; the central body front support 3 is positioned at the isolation section 6 between the air inlet channel 5 and the first-stage combustion chamber 10, and the central body rear support 4 is positioned at the second-stage contraction section 15 between the second-stage combustion chamber 14 and the outlet spray pipe 16; in order to ensure the ventilation flow, the front support 3 of the central body and the rear support 4 of the central body are uniformly distributed in the circumferential direction, and the specific arrangement scheme is shown in fig. 2;
fig. 3 shows an enlarged view of a portion of the structure of the first stage 7 or second stage 12 of the present invention. As can be seen from the figure, the first-stage injection cavity 7 and the second-stage injection cavity 12 are both of a double-injection cavity structure, and the double-injection cavities are respectively arranged on the outer shell 1 and the central body 2; along the airflow propagation direction, the first-stage injection cavity 7 is positioned between the tail end of the isolation section 6 far away from the incoming flow and the front end of the first-stage combustion chamber 10 near the incoming flow; along the air flow propagation direction, the second-stage injection cavity 12 is positioned between the tail end of the first-stage contraction section 11, which is far away from the incoming flow, and the front end of the second-stage combustion chamber 14, which is close to the incoming flow;
the outer shell 1 and the central body 2 corresponding to the first-stage injection cavity 7 are provided with first-stage injection holes 8, and the first-stage injection holes 8 are communicated to the first-stage injection cavity 7 through a runner; the outer shell 1 and the central body 2 corresponding to the second-stage injection cavity 12 are provided with second-stage injection holes 13, and the second-stage injection holes 13 are communicated to the second-stage injection cavity 12 through flow channels; the first-stage injection holes 8 and the second-stage injection holes 13 are uniformly distributed along the circumferential direction.
The outer shell 1 is used for supporting the central body 2;
an annular flow passage is formed between the central body 2 and the outer shell 1;
the front center body support 3 and the rear center body support 4 are used for supporting the center body;
the air inlet channel 5 is used for capturing incoming flow and forming annular air flow;
the fuel tank 9 is used for storing fuel;
the first-stage injection cavity 7 and the second-stage injection cavity 12 are used for uniformly filling fuel along the circumferential direction of the engine;
the first-stage injection holes 8 and the second-stage injection holes 13 are used for injecting fuel;
the first stage combustion chamber 10 and the second stage combustion chamber 14 are used for ignition initiation and form rotary detonation waves in the combustion chambers;
the outlet nozzle 16 is used for converting the high-temperature air flow into high-speed air flow so as to generate thrust;
under the low Mach number incoming flow condition, the second-stage combustion chamber 14 ignites and detonates, and the isolation section 6, the first-stage combustion chamber 10 and the first-stage contraction section 11 are used for isolating the back pressure of the second-stage combustion chamber 14 from entering the air inlet channel 5;
the isolation section 6 is used to isolate the back pressure of the first stage combustion chamber 10 from entering the inlet 5 under high Mach number inflow conditions.
In some embodiments, the first stage combustor width δ 1 Second stage combustor width delta 2 The following formula is satisfied: delta is more than or equal to 0.5 lambda, and when liquid fuel is adopted, delta is more than or equal to d; first stage combustor lengthL 1 Second stage combustor lengthL 2 The following formula needs to be satisfied:Lnot less than 2 (12+/-5) lambda, wherein lambda is the size of detonation wave cell corresponding to the mixed gas under the current combustion chamber pressure, and d is the minimum diameter of fuel liquid drop;
on the basis of meeting the requirements, the profile structural design of the first-stage combustion chamber 10 takes Mach number Ma=5 incoming flow conditions as standard design points, and rotary detonation combustion in the Mach number Ma=4-6 incoming flow range is realized; the profile structural design of the second stage combustion chamber 14 takes Mach number Ma=3 incoming flow conditions as standard design points, and rotary detonation combustion in the Mach number Ma=2-4 incoming flow range is realized.
In some embodiments, the dual-stage 7 and second-stage 12 injection chambers have square or circular cross-section during processing, with a cross-sectional area of 4-25 mm 2 The fuel pressure of the injection cavity is 1-5 MPa.
In some embodiments, the first-stage injection holes 8 and the second-stage injection holes 13 are uniformly distributed along the circumference, and the diameter of the spray holes is 0.2-0.6 mm to ensure the mixing effect of fuel and air; for liquid fuels, the number of injection holes n satisfies the following formula:whereinmFor the fuel flow rate,C d for the injection orifice flow coefficient,Din order to inject the hole diameter,pin order to inject the cavity fuel pressure,ρis the fuel density; for gaseous fuels, the number of injection holes n satisfies +.>In the followingmFor the fuel flow rate,T t is used for controlling the total temperature of the fuel,pin order to inject the cavity fuel pressure,C d for the injection orifice flow coefficient,Dfor the orifice diameter, K is a constant related to fuel, and is derived from the following equation: />In which, in the process,γis the specific heat ratio of the glass fiber reinforced plastic material,Rfor the fuel gas constant, the diameter D of the injection hole is 0.2-0.6 mm to ensure the mixing effect of the fuel and the air.
In some embodiments, the flow area of the first stage constrictor 11A th Smaller than the flow area of the first stage combustion chamber 10A c And the flow area of the second stage constriction 15A th Smaller than the flow area of the second stage combustion chamber 14A c The method comprises the steps of carrying out a first treatment on the surface of the Area shrinkage ratioThe AR range is 1.2-2.5, and the area contraction ratio AR is the flow area of the first stage combustion chamber 10A c And the flow area of the first stage constriction 11A th Is the ratio of, or the flow area of, the second stage combustion chamber 14A c And the flow area of the second stage constriction 15A th Is a ratio of (2).
In some embodiments, the central body 2 is secured to the inner center of the outer housing 1 by a central body front support 3 and a central body rear support 4.
In some embodiments, the high Mach number inflow condition refers to 4 < Ma.ltoreq.6;
the low Mach number incoming flow condition is that Ma is more than or equal to 2 and less than or equal to 4;
the high-temperature air flow means that the total temperature of the air flow is more than 2000K;
the high-speed air flow means that the air flow speed is more than 2 times of the sound speed.
The embodiment also provides a working method of the tandem double-combustion-chamber rotary detonation ramjet engine, which comprises the following steps:
the working diagram of the low Mach number incoming flow condition is shown in figure 4. The incoming air is captured by the air inlet channel 5, enters the second-stage combustion chamber 14 through the isolation section 6, the first-stage combustion chamber 10 and the first-stage contraction section 11, is mixed with fuel injected into the second-stage injection hole 13, and is ignited and detonated in the second-stage combustion chamber 14 to form rotary detonation waves; the combustion products generated by the rotary detonation combustion are exhausted to the atmosphere through the second stage constriction 15 and the outlet nozzle 16, and thrust is generated; under the low Mach number incoming flow condition, the forward shock wave induced by the rotary detonation wave is isolated by a combined channel formed by the isolation section 6, the first-stage combustion chamber 10 and the first-stage contraction section 11, so that the combustion back pressure is not forwarded to the air inlet channel 5;
the operation of the system under the high Mach number incoming flow condition is shown in figure 5. The incoming air is captured by the air inlet channel 5 and enters the first-stage combustion chamber 10 through the isolation section 6, combustible mixed gas is formed after the incoming air is mixed with fuel injected by the first-stage injection hole 8, then in the first-stage combustion chamber 10, an engine ignition device detonates a flow field and forms rotary detonation waves in the combustion chamber, the detonation waves propagate along the circumferential direction and realize fuel consumption, and generated combustion products are discharged out of the atmosphere after passing through the first-stage contraction section 11, the second-stage combustion chamber 14, the second-stage contraction section 15 and the outlet spray pipe 16, so that thrust is generated; under the condition of high Mach number incoming flow, the forward shock wave induced by the rotary detonation wave is isolated by the isolation section 6, so that the combustion back pressure is not forwarded to the air inlet channel 5.
The above embodiments are merely illustrative of the principles of the present invention and its effectiveness, and are not intended to limit the invention. Modifications and variations may be made to the above-described embodiments by those skilled in the art without departing from the spirit and scope of the invention. Accordingly, it is intended that all equivalent modifications and variations of the invention be covered by the claims of this invention, which are within the skill of those skilled in the art, can be made without departing from the spirit and scope of the invention disclosed herein.
Claims (8)
1. The utility model provides a rotatory detonation ramjet engine of tandem type double combustion chamber which characterized in that: the device comprises an outer shell (1), wherein a central body (2) is arranged in the outer shell (1), and the central body (2) is connected with the outer shell (1) through a front central body support (3) and a rear central body support (4);
an annular flow passage is formed between the outer shell (1) and the central body (2), and sequentially comprises an air inlet passage (5), an isolation section (6), a first-stage combustion chamber (10), a first-stage contraction section (11), a second-stage combustion chamber (14), a second-stage contraction section (15) and an outlet spray pipe (16) along the airflow propagation direction of the annular flow passage;
a fuel storage tank (9) is arranged in the central body (2); the central body front support (3) is positioned at an isolation section (6) between the air inlet channel (5) and the first-stage combustion chamber (10), and the central body rear support (4) is positioned at a second-stage contraction section (15) between the second-stage combustion chamber (14) and the outlet spray pipe (16); the front support (3) and the rear support (4) of the central body are uniformly distributed in the circumferential direction;
the first-stage injection cavity (7) and the second-stage injection cavity (12) are of a double-injection cavity structure, and the double-injection cavities are respectively arranged on the outer shell (1) and the central body (2); along the airflow propagation direction, a first-stage injection cavity (7) is positioned between the tail end of the isolation section (6) far away from the incoming flow and the front end of the first-stage combustion chamber (10) close to the incoming flow; along the air flow propagation direction, the second-stage injection cavity (12) is positioned between the tail end of the first-stage contraction section (11) far away from the incoming flow and the front end of the second-stage combustion chamber (14) close to the incoming flow;
the outer shell (1) and the central body (2) corresponding to the first-stage injection cavity (7) are provided with first-stage injection holes (8), and the first-stage injection holes (8) are communicated to the first-stage injection cavity (7) through a flow channel; the outer shell (1) and the central body (2) corresponding to the second-stage injection cavity (12) are provided with second-stage injection holes (13), and the second-stage injection holes (13) are communicated to the second-stage injection cavity (12) through a flow passage; the first-stage injection holes (8) and the second-stage injection holes (13) are uniformly distributed along the circumferential direction.
2. The tandem dual combustion chamber rotary detonation ramjet engine of claim 1, wherein: first stage combustor width delta 1 Second stage combustor width delta 2 The following formula is satisfied: delta is more than or equal to 0.5 lambda, and when liquid fuel is adopted, delta is more than or equal to d; first stage combustor lengthL 1 Second stage combustor lengthL 2 The following formula is satisfied:Lnot less than 2 (12+/-5) lambda, wherein lambda is the size of detonation wave cell corresponding to the mixed gas under the current combustion chamber pressure, and d is the minimum diameter of fuel liquid drop;
on the basis of meeting the requirements, the profile structural design of the first-stage combustion chamber (10) takes Mach number Ma=5 incoming flow conditions as standard design points, and rotary detonation combustion in the Mach number Ma=4-6 incoming flow range is realized; the profile structural design of the second-stage combustion chamber (14) takes Mach number Ma=3 incoming flow conditions as standard design points, and rotary detonation combustion in Mach number Ma=2-4 incoming flow ranges is realized.
3. The tandem dual combustion chamber rotary detonation ramjet engine of claim 1, wherein: the two-stage injection cavity structure of the first-stage injection cavity (7) and the second-stage injection cavity (12) has a square or round injection cavity section and a 4-stage injection cavity section area~25mm 2 The fuel pressure of the injection cavity is 1-5 MPa.
4. The tandem dual combustion chamber rotary detonation ramjet engine of claim 1, wherein: the first-stage injection holes (8) and the second-stage injection holes (13) are uniformly distributed along the circumference, and the diameters of the injection holes are 0.2-0.6 mm; for liquid fuels, the number of injection holes n satisfies the following formula:whereinmFor the fuel flow rate,C d for the injection orifice flow coefficient,Din order to inject the hole diameter,pin order to inject the cavity fuel pressure,ρis the fuel density; for gaseous fuels, the number of injection holes n satisfies +.>In the followingmFor the fuel flow rate,T t is used for controlling the total temperature of the fuel,Pin order to inject the cavity fuel pressure,C d for the injection orifice flow coefficient,Din order to inject the hole diameter,Kthe fuel-related constant is obtained from the following formula:in which, in the process,γis the specific heat ratio of the glass fiber reinforced plastic material,Rinjection hole diameter for fuel gas constantDIs 0.2-0.6 mm.
5. The tandem dual combustion chamber rotary detonation ramjet engine of claim 1, wherein: flow area of the first stage constriction (11)A th Is smaller than the flow area of the first stage combustion chamber (10)A c And the flow area of the second stage constriction (15)A th Is smaller than the flow area of the second stage combustion chamber (14)A c The method comprises the steps of carrying out a first treatment on the surface of the The area contraction ratio AR is in the range of 1.2 to 2.5, and is the flow area of the first stage combustion chamber (10)A c And the flow area of the first stage constriction (11)A th Is the ratio of, or the flow area of, the second stage combustion chamber (14)A c And the flow area of the second stage constriction (15)A th Is a ratio of (2).
6. The tandem dual combustion chamber rotary detonation ramjet engine of claim 1, wherein: the center body (2) is fixed at the inner center of the outer shell (1) through a front center body support (3) and a rear center body support (4).
7. The method of operating a tandem dual combustion chamber rotary detonation ramjet engine of any of claims 1 to 6, wherein:
under the low Mach number incoming flow condition, the incoming flow air is captured by an air inlet channel (5) and enters a second-stage combustion chamber (14) through an isolation section (6), a first-stage combustion chamber (10) and a first-stage contraction section (11), and then is mixed with fuel injected into a second-stage injection hole (13), and then is ignited and detonated in the second-stage combustion chamber (14) to form rotary detonation waves; the combustion products generated by the rotary detonation combustion are discharged out of the atmosphere through a second stage contraction section (15) and an outlet spray pipe (16) and generate thrust; under the low Mach number incoming flow condition, the forward shock wave induced by the rotary detonation wave is isolated by a combined channel formed by the isolation section (6), the first-stage combustion chamber (10) and the first-stage contraction section (11), so that the combustion back pressure is not forwarded to the air inlet channel (5);
under the condition of high Mach number incoming flow, incoming flow air is captured by an air inlet channel (5) and enters a first-stage combustion chamber (10) through an isolation section (6), combustible mixed gas is formed after the incoming flow air is mixed with fuel injected by a first-stage injection hole (8), then an engine ignition device detonates a flow field in the first-stage combustion chamber (10) and forms rotary detonation waves in the combustion chamber, the detonation waves propagate along the circumferential direction and realize fuel consumption, and generated combustion products are discharged out of the atmosphere after passing through a first-stage contraction section (11), a second-stage combustion chamber (14), a second-stage contraction section (15) and an outlet spray pipe (16), so that thrust is generated; under the condition of high Mach number incoming flow, the forward shock wave induced by the rotary detonation wave is isolated by the isolation section (6), so that the combustion back pressure is not forwarded to the air inlet channel (5).
8. The method of operating a tandem dual combustion chamber rotary detonation ramjet engine of claim 7, wherein: the high Mach number incoming flow condition is that Ma is more than 4 and less than or equal to 6; the low Mach number incoming flow condition is 2-4.
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Cited By (2)
| Publication number | Priority date | Publication date | Assignee | Title |
|---|---|---|---|---|
| CN118395640A (en) * | 2024-06-20 | 2024-07-26 | 中国人民解放军空军工程大学 | Design method of isolation section of rotary detonation engine |
| CN118669236A (en) * | 2024-08-01 | 2024-09-20 | 清华大学 | Detonation engine and aircraft |
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Cited By (3)
| Publication number | Priority date | Publication date | Assignee | Title |
|---|---|---|---|---|
| CN118395640A (en) * | 2024-06-20 | 2024-07-26 | 中国人民解放军空军工程大学 | Design method of isolation section of rotary detonation engine |
| CN118395640B (en) * | 2024-06-20 | 2024-08-30 | 中国人民解放军空军工程大学 | Design method of isolation section of rotary detonation engine |
| CN118669236A (en) * | 2024-08-01 | 2024-09-20 | 清华大学 | Detonation engine and aircraft |
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