A dual-continuous detonation mode turbine-based combined cycle engine

Abstract

The present invention discloses a dual-continuous detonation mode turbine-based combined cycle engine. The engine comprises an engine casing, a center body, a cylindrical casing, an annular regulator, and a core engine. A variable-section air inlet is formed between the center body and the front end of the engine casing. The core engine comprises a compressor, a turbine-type continuous detonation engine combustion chamber, and a turbine, which are sequentially connected. The annular regulator divides the annular cavity between the engine casing and the center body into a ramjet detonation flow channel and a turbine detonation flow channel, which is used to adjust the bypass ratio to change the operating mode. A ramjet-type continuous detonation engine combustion chamber is formed between the cylindrical casing, the engine casing, and the annular regulator. A turbine-type continuous detonation engine combustion chamber is formed between the cylindrical casing and the rotating shaft. The above-mentioned engine combines a low-Mach-number turbine detonation engine with a high-Mach-number ramjet detonation engine, which can achieve different operating modes and broaden the engine's flight speed range.

Description

Technical Field

[0001] The present invention relates to the technical field of aerospace engine structure design, and in particular to a dual-continuous detonation mode turbine-based combined cycle engine. Background Art

[0002] Traditional chemical jet propulsion engines, whether gas turbines, rocket engines, or ramjets, all utilize a deflagration cycle to convert chemical energy into thermal energy. Deflagration is a diffusion-driven, exothermic chemical reaction process. Therefore, the diffusion and thermal diffusion rates of smaller components determine the slow combustion wave propagation velocity, typically on the order of meters per second. Furthermore, this isobaric exothermic process increases entropy and has low thermal efficiency. After years of development, significantly improving the propulsion performance of deflagration-based jet propulsion systems has faced bottlenecks. Detonation, a novel energy release method, holds promise for resolving this issue. Detonation waves typically propagate at kilometres per second, approaching isochoric exothermicity. This results in minimal entropy gain, high thermal efficiency, and the advantage of self-pressurization.

[0003] Continuous detonation engines boast simple structure, compact size, fast combustion speed, easy control, self-compression, high efficiency, high specific impulse, highly adjustable thrust, and the ability to repeatedly shut down and re-ignite. They can provide propulsion for greater speeds and distances using less fuel. Currently, continuous detonation rocket engines, continuous detonation ramjets, and continuous detonation turbine engines are attracting widespread attention. However, traditional detonation engines operate in a single mode and cannot achieve optimal performance over a wide range of flight Mach numbers. For example, while turbine / ramjet detonation engines offer high specific impulse, they have stringent requirements for incoming flow conditions, resulting in a relatively narrow operating speed range. Therefore, a prototype turbine-based combined cycle engine with dual continuous detonation modes has been proposed and designed. Summary of the Invention

[0004] In view of this, the present invention provides a dual-continuous detonation mode turbine-based combined cycle engine, which combines a turbine detonation engine at low Mach number and a ramjet detonation engine at high Mach number, and can achieve different working modes, which not only broadens the engine flight speed range (0~8Ma), but also enables the engine to work efficiently in the entire flow range, and can solve the "thrust trap" problem of traditional turbine-based combined cycle engines in the transition range (2~3Ma).

[0005] The present invention adopts the following specific technical solutions:

[0006] A dual-continuous detonation mode turbine-based combined cycle engine, the engine having an axisymmetric structure and comprising an engine casing, a center body, a cylindrical casing, an annular regulator and a core engine;

[0007] The center body is coaxially arranged with the engine casing, with a front end extending outside the engine casing and a rear end located inside the engine casing and fixedly connected to the engine casing; a variable-section air intake duct is formed between the center body and the front end of the engine casing;

[0008] The engine casing comprises an isolation section adjacent to the variable-section air inlet duct and an expansion section adjacent to the isolation section; along the axial direction of the engine casing, the inner diameter of the isolation section gradually decreases, and the inner diameter of the expansion section rapidly increases;

[0009] The cylindrical shell is coaxially arranged in the engine casing and is spaced apart from the center body in the axial direction; the core engine and the plug nozzle are coaxially arranged in the cylindrical shell; the core engine includes a compressor, a turbine-type continuous detonation engine combustion chamber, and a turbine connected in sequence from the center body toward the cylindrical shell;

[0010] The annular regulator is axially movably sleeved on the outer wall of the cylindrical shell, with its front end located between the center body and the engine casing; the annular regulator divides the annular cavity between the engine casing and the center body into a ram detonation flow channel and a turbine detonation flow channel, the ram detonation flow channel being formed between the annular regulator and the engine casing, and the turbine detonation flow channel being formed between the annular regulator, the center body, the compressor, and the cylindrical shell; the annular regulator is used to adjust the bypass ratio by axial movement to change the operating mode; the bypass ratio is the ratio of the air flow rates of the ram detonation flow channel to the turbine detonation flow channel;

[0011] A ramjet-type continuous detonation engine combustion chamber is formed between the cylindrical shell, the engine casing, and the annular regulator, and is communicated with the ramjet-type continuous detonation engine combustion chamber; a plurality of ramjet-type continuous detonation engine fuel injection holes are distributed circumferentially in the expansion section of the engine casing and are communicated with the ramjet-type continuous detonation engine combustion chamber;

[0012] A turbine-type continuous detonation engine combustion chamber is formed between the cylindrical shell and the rotating shaft; a plurality of turbine flow channel fuel injection holes are circumferentially arranged on the cylindrical shell and are located in the turbine-type continuous detonation engine combustion chamber; the turbine flow channel fuel injection holes are connected to the turbine-type continuous detonation engine fuel tank;

[0013] A ramjet detonation engine tail nozzle communicating with the ramjet continuous detonation engine combustion chamber is formed between the rear end of the cylindrical shell and the rear end of the engine casing;

[0014] A turbine detonation engine tail nozzle connected to the turbine continuous detonation engine combustion chamber is formed between the rear end portion of the cylindrical shell and the plug nozzle.

[0015] Furthermore, a retractable outer lip cover is included;

[0016] The outer lip cover is fixedly mounted on the front end portion of the engine housing and is used for adjusting the variable cross-section air intake duct by extending and retracting the outer lip cover.

[0017] Furthermore, the rear end of the engine casing is connected to a tail nozzle variable structure;

[0018] When the engine enters the scramjet operating mode, the variable structure of the tail nozzle is adjusted from the Laval nozzle to the expansion nozzle.

[0019] Furthermore, along the axial direction of the engine casing, a plurality of rows of the ram flow channel fuel injection holes are spaced apart and distributed in the expansion section of the engine casing, and a plurality of rows of the turbine flow channel fuel injection holes are spaced apart and distributed in the cylindrical shell.

[0020] Furthermore, the plurality of ram flow channel fuel injection holes are evenly distributed along the circumferential direction;

[0021] The plurality of turbine flow channel fuel injection holes are evenly distributed along the circumferential direction.

[0022] Furthermore, the front end of the central body is a conical structure, and the rear end is a truncated cone structure with a gradually decreasing radius.

[0023] Beneficial effects:

[0024] 1. The dual-continuous detonation mode turbine-based combined cycle engine of the present invention is provided with a ramjet continuous detonation engine combustion chamber and a turbine continuous detonation engine combustion chamber, integrating the advantages of the turbine continuous detonation engine and the ramjet continuous detonation engine. It can provide power through multiple working modes within the range of 0 to 8 Ma, can start from zero speed on the ground, and can achieve hypersonic flight, broadening the engine flight speed range and overcoming the defect of the relatively narrow working speed range of single-power aircraft. In addition, the dual-continuous detonation mode turbine-based combined cycle engine can achieve efficient and stable operation in the entire flow range; the above-mentioned engine is capable of performing aerospace flight missions in the range from static to hypersonic speed, and has the characteristics of horizontal take-off and landing, round-trip between the earth and the sky, and flexibility and convenience. It has high application value and can provide a technical route for the research and development of aerospace combined cycle propulsion systems.

[0025] 2. The dual-continuous detonation mode turbine-based combined cycle engine of the present invention adopts a variety of variable structures, including an axially movable annular regulator, a retractable outer lip cover and a variable tail nozzle structure, which ensures a smooth transition during the engine's mode conversion process; in addition, multiple rows of ramjet flow channel fuel injection holes are arranged in the ramjet detonation flow channel. By adopting different fuel injection strategies for each mode, the mixing distance between the incoming air and the liquid fuel can be adjusted, thereby ensuring the efficient operation of the ramjet continuous detonation engine.

[0026] 3. The present invention addresses the "thrust trap" problem faced by traditional turbine-based combined cycle engines in the 2-3Ma range. By jointly providing thrust through a turbine-type continuous detonation engine and a ramjet-type continuous detonation engine, and combining the advantages of double detonation, the present invention solves the "thrust trap" problem faced by traditional turbine-based combined cycle engines in the transition range (2 to 3Ma). BRIEF DESCRIPTION OF THE DRAWINGS

[0027] Figure 1 Schematic diagram of the overall structure of the dual-continuous detonation mode turbine-based combined cycle engine of the present invention;

[0028] Figure 2 for Figure 1 Cross-sectional view of a turbine-based combined cycle engine in double continuous detonation mode.

[0029] Among them, 1-center body; 2-variable cross-section air inlet; 3-isolation section; 4-annular regulator; 5-ramjet flow channel fuel injection hole; 6-ramjet continuous detonation engine combustion chamber; 7-engine casing; 8-ramjet detonation engine tail nozzle; 9-turbine detonation flow channel; 10-cylindrical shell; 11-turbine flow channel fuel injection hole; 12-turbine continuous detonation engine combustion chamber; 13-compressor; 14-rotating shaft; 15-turbine; 16-plug nozzle; 17-outer lip cover; 18-connecting shaft; 19-connecting part; 20-turbine continuous detonation engine fuel tank, 21-expansion section. DETAILED DESCRIPTION

[0030] The present invention is described in detail below with reference to the accompanying drawings and embodiments.

[0031] The embodiment of the present invention provides a dual continuous detonation mode turbine-based combined cycle engine, such as Figure 1 and Figure 2 As shown in the structure, the engine is an axisymmetric structure, and includes an engine casing 7, a center body 1, a cylindrical shell 10, an annular regulator 4 and a core engine; in this embodiment, Figure 2 The left side of the middle is the front end, and the right side is the back end;

[0032] The center body 1 is coaxially arranged with the engine housing 7, with its front end extending outside the engine housing 7 and its rear end located inside and fixedly connected to the engine housing 7. The center body 1 can be fixedly connected to the engine housing 7 via a connecting shaft 18 on its outer periphery. A variable-section air intake duct 2 is formed between the center body 1 and the front end of the engine housing 7. The front end of the center body 1 is a conical structure, while the rear end is a frustum-shaped structure with a gradually decreasing radius.

[0033] The engine housing 7 includes an isolation section 3 adjacent to the variable-section air inlet duct 2 and an expansion section 21 adjacent to the isolation section 3. Along the axial direction of the engine housing 7, from the front end to the rear end of the engine housing 7, the inner diameter of the isolation section 3 gradually decreases, while the inner diameter of the expansion section 21 rapidly increases.

[0034] The cylindrical shell 10 is coaxially arranged in the engine casing 7 and is axially spaced relative to the center body 1; the cylindrical shell 10 is fixedly connected to the engine casing 7 through a connecting portion 19 on the outer peripheral side, and a turbine-type continuous detonation engine fuel tank 20 is formed in the connecting portion 19; a core engine and a plug nozzle 16 are coaxially arranged in the cylindrical shell 10; the core engine includes a compressor 13, a turbine-type continuous detonation engine combustion chamber 12 and a turbine 15 connected in sequence from the center body 1 toward the cylindrical shell 10; the outer peripheral surface of the compressor 13 cooperates with the inner peripheral surface of the cylindrical shell 10; the outer diameter of the rotating shaft 14 is smaller than the inner diameter of the cylindrical shell 10, thereby forming a cavity between the outer peripheral surface of the rotating shaft 14 and the inner peripheral surface of the cylindrical shell 10, that is, the turbine-type continuous detonation engine combustion chamber 12;

[0035] The annular regulator 4 is axially movable and sleeved on the outer wall of the cylindrical shell 10, with its front end located between the center body 1 and the engine casing 7. The annular regulator 4 divides the annular cavity between the engine casing 7 and the center body 1 into a ram detonation flow channel and a turbine detonation flow channel 9. The ram detonation flow channel is formed between the annular regulator 4 and the engine casing 7, and the turbine detonation flow channel 9 is formed between the annular regulator 4, the center body 1, the compressor 13, and the cylindrical shell 10. The annular regulator 4 is used to adjust the bypass ratio by axial movement to change the operating mode. The bypass ratio is the ratio of the air flow rate of the ram detonation flow channel to the turbine detonation flow channel 9.

[0036] A ramjet continuous detonation engine combustion chamber 6 communicating with the ramjet detonation flow channel is formed between the cylindrical casing 10, the engine housing 7, and the annular regulator 4. A plurality of ramjet flow channel fuel injection holes 5 communicating with the ramjet continuous detonation engine combustion chamber 6 are circumferentially distributed in the expansion section 21 of the engine housing 7. A plurality of rows of ramjet flow channel fuel injection holes 5 are provided in the expansion section 21 of the engine housing 7 along the axial direction of the engine housing 7. The plurality of ramjet flow channel fuel injection holes 5 are evenly distributed circumferentially. The ramjet flow channel fuel injection holes 5 are communicated with an external fuel tank.

[0037] A turbine-type continuous detonation engine combustion chamber 12 is formed between the cylindrical casing 10 and the rotating shaft 14. A plurality of turbine flow path fuel injection holes 11 are circumferentially arranged on the cylindrical casing 10 and are located within the turbine-type continuous detonation engine combustion chamber 12. The turbine flow path fuel injection holes 11 are connected to a turbine-type continuous detonation engine fuel storage tank 20, which in turn is connected to an external fuel storage tank. A plurality of rows of turbine flow path fuel injection holes 11 are spaced apart and evenly distributed along the circumference of the cylindrical casing 10 along the axial direction of the engine casing 7.

[0038] A ramjet detonation engine tail nozzle 8 communicating with the ramjet continuous detonation engine combustion chamber 6 is formed between the rear end of the cylindrical shell 10 and the rear end of the engine casing 7;

[0039] A turbo detonation engine tail nozzle communicating with the turbo continuous detonation engine combustion chamber 12 is formed between the rear end portion of the cylindrical casing 10 and the plug nozzle 16 .

[0040] In a specific embodiment, Figure 2 As shown, the above-mentioned dual-continuous detonation mode turbine-based combined cycle engine also includes a retractable outer lip cover 17; the outer lip cover 17 is fixedly installed on the front end of the engine casing 7, and is used to adjust the variable-section air inlet duct 2 by retracting and extending the outer lip cover 17.

[0041] Furthermore, the rear end of the engine housing 7 is connected to a tail nozzle variable structure;

[0042] When the engine enters the scramjet operating mode, the variable structure of the tail nozzle is adjusted from the Laval nozzle to the expansion nozzle.

[0043] The above-mentioned dual-continuous detonation mode turbine-based combined cycle engine is provided with a ramjet continuous detonation engine combustion chamber 6 and a turbine continuous detonation engine combustion chamber 12, which combines the advantages of the turbine continuous detonation engine and the ramjet continuous detonation engine. It can provide power through a variety of working modes in the range of 0 to 8 Ma. It can not only start from zero speed on the ground, but also achieve high-speed flight, broadening the engine flight speed range and overcoming the defect of a relatively narrow working speed range of a single-power aircraft. In addition, the above-mentioned dual-continuous detonation mode turbine-based combined cycle engine can achieve efficient and stable operation in the entire flow range; the above-mentioned engine can perform aerospace flight missions in the range from static to hypersonic, and has the characteristics of horizontal take-off and landing, round-trip between the earth and the sky, and flexibility and convenience. It has high application value and can provide a technical route for the research and development of aerospace combined cycle propulsion systems.

[0044] The above-mentioned dual-continuous detonation mode turbine-based combined cycle engine adopts a variety of variable structures, including an axially movable annular regulator 4, a retractable outer lip cover 17 and a variable tail nozzle structure, which ensures a smooth transition during the engine's mode conversion process; in addition, multiple rows of ramjet flow channel fuel injection holes 5 are arranged in the ramjet detonation flow channel. By adopting different fuel injection strategies for each mode, the mixing distance between the incoming air and the liquid fuel can be adjusted, thereby ensuring the efficient operation of the ramjet continuous detonation engine.

[0045] In order to solve the "thrust trap" problem faced by traditional turbine-based combined cycle engines in the 2-3Ma range, the thrust is jointly provided by a turbine continuous detonation engine and a ramjet continuous detonation engine, and the advantages of double detonation are combined to solve the "thrust trap" problem of traditional turbine-based combined cycle engines in the transition range (2~3Ma).

[0046] The operating method of the dual continuous detonation mode turbine-based combined cycle engine comprises the following stages:

[0047] Phase 1 (0-2 Ma): When the dual-continuous detonation mode turbine-based combined cycle engine is started at zero speed, it operates in turbine mode. The incoming air is diverted by the annular regulator 4 and simultaneously passes through the ramjet detonation channel and the turbine detonation channel 9. At this time, the turbine channel fuel injection holes 11 arranged circumferentially within the turbine-type continuous detonation engine combustion chamber 12 begin to inject fuel, and after fully mixing with the incoming air, they detonate, forming a self-sustaining and propagating continuous rotating detonation wave. The high-temperature combustion gas is ejected through the plug nozzle 16, thereby generating thrust. Due to the low flight Mach number, the ramjet channel fuel injection holes 5 do not inject fuel, and the ramjet continuous detonation engine combustion chamber 6 does not operate. At this time, the air passing through the ramjet detonation channel can cool the cylindrical shell 10 forming the turbine-type continuous detonation engine combustion chamber 12. In this mode, the annular regulator 4 gradually moves backward, the bypass ratio decreases, the air flow in the turbine detonation flow channel 9 increases, and the fuel injection flow is changed at the same time, and the thrust generated by the turbine continuous detonation engine combustion chamber 12 gradually increases.

[0048] The second stage (2-3 Ma): After the aircraft reaches 2 Ma in the turbine mode, it enters the turbine-ramjet working mode. At this time, the ramjet flow channel fuel injection hole 5 begins to spray and detonate, and the liquid fuel and the incoming air form a self-sustaining continuous detonation wave in the ramjet continuous detonation engine combustion chamber 6. The high-temperature gas is ejected through the Laval nozzle, and the ramjet continuous detonation engine combustion chamber 6 begins to work. At this time, the thrust of the dual-continuous detonation mode turbine-based combined cycle engine is jointly provided by the turbine-based continuous detonation engine combustion chamber 12 and the ramjet continuous detonation engine combustion chamber 6, thus solving the "thrust trap" problem faced by traditional turbine-based combined cycle engines in the 2-3 Ma range. In this mode, the annular regulator 4 gradually moves forward, the bypass ratio increases, the air flow rate of the ramjet detonation flow channel increases, and the air flow rate of the turbine detonation flow channel 9 decreases. At the same time, the fuel injection flow rate of the dual combustion chamber is changed to ensure a suitable equivalence ratio; when the annular regulator 4 moves forward to match the center body 1, the turbine detonation flow channel 9 is closed, and the turbine continuous detonation engine combustion chamber 12 is closed. At this time, the ramjet continuous detonation engine combustion chamber 6 is completely used for power.

[0049] Phase 3 (3-5 Mach): After the flight Mach number reaches 3 Mach, the aircraft enters the sub-ramjet operating mode, generating a pre-combustion shock wave train within the isolation section 3 between the annular regulator 4 and the engine casing 7. In this mode, the outer lip 17 moves rearward to ensure that the variable-area inlet 2 structure meets the operating characteristics of a ramjet continuous detonation engine.

[0050] Phase 4 (5-8 Ma): After the flight Mach number reaches 5 Ma, the vehicle enters the scramjet operating mode. At this point, most of the flow channel is in a supersonic state, the outer lip cover 17 continues to move backward, and the variable nozzle structure of the ramjet detonation engine tail nozzle 8 is mechanically adjusted from a Laval nozzle to a diverging nozzle. The fuel injection position is replaced by the upstream injection hole from the downstream injection hole, extending the mixing distance between the liquid fuel and the supersonic incoming air. The engine gradually transitions to the high Mach number scramjet continuous detonation engine operating mode.

[0051] In summary, the above are only preferred embodiments of the present invention and are not intended to limit the scope of protection of the present invention. Any modifications, equivalent replacements, improvements, etc. made within the spirit and principles of the present invention should be included in the scope of protection of the present invention.

Claims

1. A dual continuous detonation mode turbine-based combined cycle engine, characterized in that: It has an axisymmetric structure and includes an engine casing, a center body, a cylindrical shell, an annular regulator and a core engine; The center body is coaxially arranged with the engine casing, with a front end extending outside the engine casing and a rear end located inside the engine casing and fixedly connected to the engine casing; a variable-section air intake duct is formed between the center body and the front end of the engine casing; The engine casing comprises an isolation section adjacent to the variable-section air inlet duct and an expansion section adjacent to the isolation section; along the axial direction of the engine casing, the inner diameter of the isolation section gradually decreases, and the inner diameter of the expansion section rapidly increases; The cylindrical shell is coaxially arranged in the engine casing and is spaced apart from the center body in the axial direction; the core engine and the plug nozzle are coaxially arranged in the cylindrical shell; the core engine includes a compressor, a turbine-type continuous detonation engine combustion chamber, and a turbine connected in sequence from the center body toward the cylindrical shell; The annular regulator is axially movably sleeved on the outer wall of the cylindrical shell, with its front end located between the center body and the engine casing; the annular regulator divides the annular cavity between the engine casing and the center body into a ram detonation flow channel and a turbine detonation flow channel, the ram detonation flow channel being formed between the annular regulator and the engine casing, and the turbine detonation flow channel being formed between the annular regulator, the center body, the compressor, and the cylindrical shell; the annular regulator is used to adjust the bypass ratio by axial movement to change the operating mode; the bypass ratio is the ratio of the air flow rates of the ram detonation flow channel to the turbine detonation flow channel; A ramjet-type continuous detonation engine combustion chamber is formed between the cylindrical shell, the engine casing, and the annular regulator, and is communicated with the ramjet-type continuous detonation engine combustion chamber; a plurality of ramjet-type continuous detonation engine fuel injection holes are distributed circumferentially in the expansion section of the engine casing and are communicated with the ramjet-type continuous detonation engine combustion chamber; A turbine-type continuous detonation engine combustion chamber is formed between the cylindrical shell and the rotating shaft; a plurality of turbine flow channel fuel injection holes are circumferentially arranged on the cylindrical shell and are located in the turbine-type continuous detonation engine combustion chamber; the turbine flow channel fuel injection holes are connected to the turbine-type continuous detonation engine fuel tank; A ramjet detonation engine tail nozzle communicating with the ramjet continuous detonation engine combustion chamber is formed between the rear end of the cylindrical shell and the rear end of the engine casing; A turbine detonation engine tail nozzle connected to the turbine continuous detonation engine combustion chamber is formed between the rear end portion of the cylindrical shell and the plug nozzle.

2. The engine according to claim 1, wherein: Also included is a retractable outer lip cover; The outer lip cover is fixedly mounted on the front end portion of the engine housing and is used for adjusting the variable cross-section air intake duct by extending and retracting the outer lip cover.

3. The engine according to claim 1, wherein: The rear end of the engine casing is connected to a tail nozzle variable structure; When the engine enters the scramjet operating mode, the variable structure of the tail nozzle is adjusted from the Laval nozzle to the expansion nozzle.

4. The engine according to claim 1, wherein Along the axial direction of the engine casing, a plurality of rows of spaced-apart ram flow channel fuel injection holes are provided in the expansion section of the engine casing, and a plurality of rows of spaced-apart turbine flow channel fuel injection holes are provided in the cylindrical shell.

5. The engine according to claim 4, characterized in that The plurality of ram flow channel fuel injection holes are evenly distributed along the circumferential direction; The plurality of turbine flow channel fuel injection holes are evenly distributed along the circumferential direction.

6. The engine according to any one of claims 1 to 5, characterized in that: The front end of the central body is a conical structure, and the rear end is a truncated cone structure with a gradually decreasing radius.

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