A floating detonation turbine engine

Abstract

The present invention provides a floating detonation turbine engine, which relates to the field of turbine technology. It comprises: a detonation combustion chamber, which is composed of a plurality of outer ring sections of an arc-shaped surface, which are circumferentially spliced ​​together to form a cylindrical combustion chamber structure; a positioning groove is respectively provided on the outer wall surface of each outer ring section of the combustion chamber; an engine casing, which is coaxially arranged on the outer ring of the detonation combustion chamber, and mounting holes are respectively provided on the engine casing at positions corresponding to the positioning grooves, and positioning pins are provided in the mounting holes, and the ends of the pins are inserted into the positioning grooves; wherein, a first gap is provided between the detonation combustion chamber and the engine casing, and the first gap forms an annular fuel cavity, and the positioning pins connect the detonation combustion chamber and the engine casing radially, so that each outer ring section of the combustion chamber can move radially along the end of the positioning pin. The present invention can reduce the thermal deformation of the detonation combustion chamber, ensure the long-term stable operation of the detonation combustion chamber, extend the service life, and improve the overall reliability of the turbine.

Description

Technical Field

[0001] The present invention relates to the technical field of turbines, and in particular to a floating detonation turbine engine. Background Art

[0002] Turbine engines are a type of heat engine commonly used in power generation and aerospace. Their primary operating principle is that air enters a compressor, where it is pressurized and mixed with fuel. This generates high-temperature exhaust gases in the combustion chamber, which propel the turbine to produce work. The turbine then drives the compressor and the load. Common turbine engines include turbojets and turbofans, and gas turbines. Turbine engines primarily utilize the high-temperature, high-pressure exhaust gases at the turbine outlet to generate high-temperature, high-speed airflow in the exhaust nozzle, generating thrust. Gas turbines, on the other hand, use the turbine to drive a generator load, generating electricity.

[0003] In order to improve the combustion stability and combustion efficiency in the turbine engine combustion chamber, a detonation combustion chamber can be used instead of the traditional slow combustion combustion chamber, which can not only effectively reduce the pressure loss and greatly improve the combustion efficiency, but also reduce the nitrogen oxides (NO x ) emissions. Current detonation combustion chamber structures are all rigidly fixed within turbine engines. Due to the high combustion temperatures and rapid flame propagation within detonation combustion chambers, this rigid fixing method results in significant thermal stress when the combustion chamber expands due to heat. Long-term operation can easily lead to thermal deformation and even cracking of the combustion chamber, shortening its service life. Summary of the Invention

[0004] In view of this, an embodiment of the present application provides a floating detonation turbine to reduce thermal deformation of the combustion chamber caused by high temperature and increase the service life of the detonation combustion chamber.

[0005] The embodiment of the present application provides the following technical solution: a floating detonation turbine engine, comprising:

[0006] A detonation combustion chamber, wherein the detonation combustion chamber is composed of a plurality of arc-shaped outer ring sections of the combustion chamber, which are circumferentially spliced ​​to form a cylindrical detonation combustion chamber structure; a positioning groove is respectively provided on the outer wall surface of each section of the outer ring section of the combustion chamber;

[0007] An engine casing, wherein the detonation combustion chamber is coaxially arranged on the inner ring of the engine casing, and mounting holes are respectively provided on the engine casing at positions corresponding to the positioning grooves, and positioning pins are provided in the mounting holes so that the ends of the positioning pins are inserted into the positioning grooves;

[0008] In which, a first gap is provided between the detonation combustion chamber and the engine casing, the first gap forms an annular fuel cavity for accommodating fuel, and the positioning pin connects the detonation combustion chamber and the engine casing radially so that each outer ring segment of the combustion chamber can move radially along the end of the positioning pin.

[0009] According to an embodiment of the present application, elastic sealing structures are respectively provided on the splicing end faces of adjacent outer ring segments of the combustion chamber, and the elastic sealing structures are made of any one of graphite tape and metal elastic sealing gaskets.

[0010] According to one embodiment of the present application, a second gap is provided between the elastic sealing structures on the spliced ​​end faces of adjacent outer ring segments of the combustion chamber, and the second gap forms a fuel passage for fuel flow, so that the fuel in the annular fuel cavity enters the detonation combustion zone in the inner cavity of the detonation combustion chamber through the fuel passage.

[0011] According to an embodiment of the present application, the width of the second gap is 0.1 to 3 mm.

[0012] According to one embodiment of the present application, the combustion chamber outer ring section includes a metal base layer and a ceramic tile layer arranged in sequence from the outer ring of the combustion chamber toward the axial direction, and the ceramic tile layer wraps the inner wall surface, and the front and rear wall surfaces of the metal base layer to form the combustion chamber outer ring section.

[0013] According to one embodiment of the present application, the thickness of the metal base layer is at least three times the thickness of the ceramic tile layer.

[0014] According to one embodiment of the present application, a plurality of fuel inlets are provided on the engine casing, and the fuel inlets are connected to the annular fuel cavity.

[0015] According to one embodiment of the present application, it also includes a compressor and a turbine, and the compressor and the turbine are coaxially connected; the inlet of the compressor inhales air, and the outlet of the compressor is connected to the air inlet of the detonation combustion chamber, so that the pressurized air enters the detonation combustion chamber and is mixed with fuel to undergo detonation combustion; the flue gas outlet of the detonation combustion chamber is connected to the turbine, driving the turbine to do work.

[0016] According to one embodiment of the present application, the compressor includes multiple stages of moving blades and multiple stages of stationary blades, and the moving blades and the stationary blades are arranged alternately in sequence; wherein, the number of stages of the moving blades is one more than the number of stages of the stationary blades, so that the primary blades and the last stage blades of the compressor are both the moving blades.

[0017] According to one embodiment of the present application, the turbine is an axial flow turbine without guide vanes.

[0018] Compared with the prior art, the beneficial effects that can be achieved by at least one of the above-mentioned technical solutions adopted in the embodiments of this specification include at least the following: the detonation combustion chamber of the embodiment of the present invention adopts a floating connection structure, which can effectively reduce the thermal stress generated by high temperature, absorb circumferential and radial thermal expansion, reduce thermal deformation, ensure the long-term stable operation of the detonation combustion chamber, extend the service life, and improve the overall reliability of the turbine. BRIEF DESCRIPTION OF THE DRAWINGS

[0019] In order to more clearly illustrate the technical solutions of the embodiments of the present application, the following briefly introduces the drawings required for use in the embodiments. Obviously, the drawings described below are only some embodiments of the present application. For ordinary technicians in this field, other drawings can be obtained based on these drawings without any creative work.

[0020] Figure 1 is a partial structural diagram of a floating detonation turbine according to an embodiment of the present invention;

[0021] Figure 2 1 is a schematic structural diagram of a detonation combustion chamber of a turbine according to an embodiment of the present invention;

[0022] Figure 3 2. It is a schematic structural diagram of the outer ring section of the detonation combustion chamber of an embodiment of the present invention;

[0023] Figure 4 This is another structural schematic diagram of the outer ring section of the detonation combustion chamber of an embodiment of the present invention;

[0024] Figure 5 1 is a schematic diagram of the overall structure of a floating detonation turbine according to an embodiment of the present invention;

[0025] Among them, 1-engine casing, 2-combustion chamber outer ring section, 2-1-metal matrix layer, 2-2-ceramic tile layer, 3-annular fuel cavity, 4-fuel inlet, 5-detonation combustion zone, 6-locating pin, 7-locating groove, 8-engine rotor, 9-detonation combustion chamber, 10-turbine moving blade, 11-compressor. DETAILED DESCRIPTION

[0026] The embodiments of the present application are described in detail below with reference to the accompanying drawings.

[0027] The following describes the embodiments of the present application through specific examples, and those skilled in the art can easily understand other advantages and effects of the present application from the contents disclosed in this specification. Obviously, the described embodiments are only a part of the embodiments of the present application, rather than all the embodiments. The present application can also be implemented or applied through other different specific embodiments, and the details in this specification can also be modified or changed in various ways based on different viewpoints and applications without departing from the spirit of the present application. It should be noted that, in the absence of conflict, the features in the following embodiments and embodiments can be combined with each other. Based on the embodiments in the present application, all other embodiments obtained by those of ordinary skill in the art without making creative work are within the scope of protection of this application.

[0028] like Figure 1-Figure 5 As shown, an embodiment of the present invention provides a floating detonation turbine engine, comprising: a detonation combustion chamber 9, wherein the detonation combustion chamber 9 is composed of a plurality of sections of outer combustion chamber ring segments 2 with arc-shaped curved surfaces spliced ​​circumferentially to form a cylindrical detonation combustion chamber 9 structure; a positioning groove 7 is respectively provided on the outer wall surface of each section of the outer combustion chamber ring segment 2; an engine casing 1, wherein the detonation combustion chamber 9 is coaxially arranged on the inner ring of the engine casing 1, and mounting holes are respectively provided on the engine casing 1 at positions corresponding to the positioning grooves 7, and positioning pins 6 are provided in the mounting holes so that the ends of the positioning pins 6 are inserted into the positioning grooves 7; wherein a first gap is provided between the detonation combustion chamber 9 and the engine casing 1, and the first gap forms an annular fuel cavity 3 for accommodating fuel, and the positioning pins 6 movably connect the detonation combustion chamber 9 and the engine casing 1 in the radial direction, so that each section of the outer combustion chamber ring segment 2 can move radially along the end of the positioning pin 6.

[0029] The detonation combustion chamber 9 of the embodiment of the present invention is configured as a multi-section splicing structure, such as Figure 2 As shown, each section can float radially along the positioning pin 6. Compared with the traditional rigid fixed detonation combustion chamber 9 structure, when the floating detonation combustion chamber 9 of the embodiment of the present invention produces thermal expansion in a high-temperature working environment, the independent outer ring section 2 of the combustion chamber can float radially, effectively reducing the thermal stress generated by high temperature, ensuring the long-term stable operation of the detonation combustion chamber 9, extending the service life, and improving the overall reliability of the turbine.

[0030] In this embodiment, the positioning pins 6 position each section of the combustion chamber outer ring segment 2 in the circumferential direction, so that multiple sections of the combustion chamber outer ring segment 2 are circumferentially spliced ​​in the inner ring of the engine casing 1 to form the floating detonation combustion chamber 9. In addition, an annular fuel cavity 3 is formed between the engine casing 1 and the detonation combustion chamber 9, which is used to first introduce the fuel into the annular fuel cavity 3 and then enter the inner cavity of the detonation combustion chamber 9 from the annular fuel cavity 3. On the one hand, the fuel entering the annular fuel cavity 3 can cool the detonation combustion chamber 9 and further prevent the detonation combustion chamber 9 from thermal deformation due to excessive temperature; on the other hand, the liquid fuel is heated and vaporized in the annular fuel cavity 3, and then enters the detonation combustion chamber 9 to mix with the air. The gaseous fuel can be better mixed with the air, which is conducive to detonation combustion.

[0031] In order to further reduce the thermal stress generated by high temperature, in one embodiment, elastic sealing structures are respectively provided on the splicing end faces of the adjacent outer ring segments 2 of the combustion chamber, and the elastic sealing structures are made of any one of graphite tape and metal elastic sealing gaskets. In this embodiment, the splicing end faces of the adjacent outer ring segments 2 of the combustion chamber are all free end faces. The elastic sealing structure is provided on the splicing end faces. On the one hand, the splicing can be sealed. On the other hand, the elastic sealing structure can absorb the thermal expansion of the outer ring segment 2 in the circumferential direction, further reducing thermal stress. The structural setting of this embodiment can meet the thermal expansion of the detonation combustion chamber 9 in the radial and circumferential directions, ensure that the detonation combustion chamber 9 can operate stably for a long time in a high temperature environment, extend the service life of the detonation combustion chamber 9, and improve the reliability of the equipment. The present invention does not specifically limit the above-mentioned elastic sealing structure. Any elastic sealing member that can be applied to a high temperature environment can be used. In this embodiment, graphite tape or metal elastic sealing gaskets are preferably used.

[0032] In one embodiment, a second gap is provided between the elastic sealing structures on the spliced ​​end faces of adjacent outer ring sections 2 of the combustion chamber, and the second gap forms a fuel passage for fuel flow, so that the fuel in the annular fuel cavity 3 enters the detonation combustion zone 5 in the inner cavity of the detonation combustion chamber 9 through the fuel passage. In this embodiment, the adjacent elastic sealing structures are not in a completely squeezed state, and the second gap is reserved so that the vaporized fuel in the annular fuel cavity 3 enters the inner cavity of the detonation combustion chamber 9 through the gap. This structural setting is reasonable, which not only simplifies the equipment structure, but also optimizes the fuel flow path, so that the gaseous fuel can evenly enter the detonation combustion, ensuring uniform mixing of the fuel and air, and thus achieving fully stable combustion.

[0033] Furthermore, in order to ensure that the fuel passage is not completely compressed and blocked after the combustion chamber outer ring section 2 undergoes thermal expansion, and at the same time ensure that the fuel passage is not too wide to affect detonation combustion, the preferred width of the second gap in this embodiment is 0.1 to 3 mm.

[0034] In one embodiment, if Figure 3 As shown, the combustion chamber outer ring section 2 comprises a metal matrix layer 2-1 and a ceramic tile layer 2-2, arranged sequentially from the outer periphery of the combustion chamber toward the axial center. The ceramic tile layer 2-2 wraps around the inner wall surface, as well as the front and rear walls of the metal matrix layer 2-1, to form the combustion chamber outer ring section 2. The metal matrix layer 2-1 provides rigid support, ensuring the rigid structure of the detonation combustion chamber 9. The ceramic tile layer 2-2 is fixedly mounted on the inner layer. The ceramic tile layer 2-2 has excellent thermal insulation properties, reduces heat conduction, and can reduce the cooling requirements of the detonation combustion chamber 9, simplifying the equipment structure and reducing equipment costs.

[0035] The metal base layer 2-1 needs to ensure sufficient rigid support capability. In this embodiment, preferably, the thickness of the metal base layer 2-1 is at least three times the thickness of the ceramic tile layer 2-2.

[0036] In the preferred structure of this embodiment, a plurality of fuel inlets 4 are provided on the engine casing 1 , and the fuel inlets 4 are communicated with the annular fuel cavity 3 so as to allow fuel to flow into the annular fuel cavity 3 .

[0037] In this embodiment, continuous rotating detonation combustion is performed in the detonation combustion chamber. The engine rotor 8 is positioned at the center of the detonation combustion chamber, forming an annular cavity with the combustion chamber. This cavity serves as the detonation combustion zone 5, forming an annular channel for continuous rotating detonation combustion. A thermal insulation layer may be provided between the engine rotor 8 and the detonation combustion zone 5.

[0038] In this embodiment, a compressor 11 and a turbine are further included, and the compressor 11 and the turbine are coaxially connected; the inlet of the compressor 11 inhales air, and the outlet of the compressor 11 is connected to the air inlet of the detonation combustion chamber 9, so that the pressurized air enters the detonation combustion chamber 9 and mixes with the fuel to undergo detonation combustion; the flue gas outlet of the detonation combustion chamber 9 is connected to the turbine, driving the turbine to perform work.

[0039] In a traditional multi-stage axial-flow compressor, the multiple stages of rotor blades and stator blades are arranged in alternating intervals. After air is accelerated by the rotor blades, the total pressure rises. It is then decelerated and pressurized again by the stator blades, causing a slight decrease in total pressure and an increase in static pressure. The air then passes through two additional deceleration structures, the outlet guide vanes (OGVs) and the diffuser, before entering the combustion chamber. This deceleration of the air after entering the final stator blades, OGV, and diffuser results in additional total pressure and energy losses, sacrificing some efficiency and leading to a decrease in overall engine efficiency.

[0040] Therefore, in this embodiment, a floating detonation combustion chamber 9 is used. Since the flame propagation speed of the continuous rotating detonation combustion is extremely fast, ranging from 1000 to 2000 m / s depending on the fuel, temperature and equivalence ratio, there is no need for deceleration, which reduces the pressure loss during the deceleration and acceleration of the airflow. In addition, the compressor 11 in this embodiment includes multiple stages of moving blades and multiple stages of stationary blades, and the moving blades and the stationary blades are arranged alternately in sequence; wherein the number of stages of the moving blades is one more than the number of stages of the stationary blades, so that the primary blades and the final blades of the compressor 11 are both the moving blades. Figure 5 As shown, the compressor 11 includes multiple stages of moving blades R1, R2, R3, and R4. Except for the last stage moving blades, each stage of moving blades is followed by a stage of stationary blades S1, S2, and S3. The last stage is the moving blade R4. After passing through the last stage moving blade R4, the air enters the detonation combustion chamber 9 and mixes with the fuel.

[0041] The turbine of this embodiment has a smaller axial dimension than conventional designs, making it suitable for applications with stringent axial dimension requirements. Furthermore, the turbine has fewer components, reducing costs. More importantly, it reduces total pressure losses during airflow acceleration and deceleration, thereby improving efficiency. Furthermore, the turbine is preferably a vaneless axial-flow turbine, allowing the high-temperature, high-speed airflow from the continuously rotating detonation outlet to directly drive the turbine blades 10 to produce work, thus avoiding total pressure losses after the high-temperature flue gas passes through the turbine guide vanes, which in turn reduces efficiency.

[0042] The above description is merely a specific embodiment of the present application, but the scope of protection of the present application is not limited thereto. Any changes or substitutions that can be easily conceived by a person skilled in the art within the technical scope disclosed in the present application should be included in the scope of protection of the present application. Therefore, the scope of protection of the present application should be based on the scope of protection of the claims.

Claims

1. A floating detonation turbine engine, characterized in that: include: A detonation combustion chamber, wherein the detonation combustion chamber is composed of a plurality of arc-shaped combustion chamber outer ring sections spliced ​​circumferentially to form a cylindrical detonation combustion chamber structure; Positioning grooves are respectively provided on the outer wall surface of the outer ring section of each combustion chamber; An engine casing, wherein the detonation combustion chamber is coaxially arranged on the inner ring of the engine casing, and mounting holes are respectively provided on the engine casing at positions corresponding to the positioning grooves, and positioning pins are provided in the mounting holes so that the ends of the positioning pins are inserted into the positioning grooves; A first gap is provided between the detonation combustion chamber and the engine casing, the first gap forming an annular fuel cavity for accommodating fuel, and the positioning pin movably connects the detonation combustion chamber and the engine casing in the radial direction, so that each outer ring segment of the combustion chamber can move radially along the end of the positioning pin; Elastic sealing structures are respectively provided on the splicing end faces of the adjacent outer ring sections of the combustion chamber, and the elastic sealing structures are made of any one of graphite tape and metal elastic sealing gaskets; a plurality of fuel inlets are provided on the engine casing, and the fuel inlets are connected to the annular fuel cavity.

2. The floating detonation turbine engine according to claim 1, characterized in that: A second gap is provided between the elastic sealing structures on the spliced ​​end faces of adjacent outer ring segments of the combustion chamber, and the second gap forms a fuel passage for fuel flow, so that the fuel in the annular fuel cavity enters the detonation combustion zone in the inner cavity of the detonation combustion chamber through the fuel passage.

3. The floating detonation turbine engine according to claim 2, characterized in that: The width of the second gap is 0.1-3 mm.

4. The floating detonation turbine engine according to claim 1, characterized in that: The combustion chamber outer ring section includes a metal base layer and a ceramic tile layer arranged in sequence from the outer ring of the combustion chamber toward the axial direction. The ceramic tile layer wraps the inner wall surface, front and rear wall surfaces of the metal base layer to form the combustion chamber outer ring section.

5. The floating detonation turbine engine according to claim 4, characterized in that: The thickness of the metal matrix layer is at least three times the thickness of the ceramic tile layer.

6. The floating detonation turbine engine according to claim 1, characterized in that: It also includes a compressor and a turbine, which are coaxially connected; the inlet of the compressor inhales air, and the outlet of the compressor is connected to the air inlet of the detonation combustion chamber, so that the pressurized air enters the detonation combustion chamber and is mixed with fuel to undergo detonation combustion; the flue gas outlet of the detonation combustion chamber is connected to the turbine to drive the turbine to perform work.

7. The floating detonation turbine engine according to claim 6, characterized in that: The compressor includes multiple stages of moving blades and multiple stages of stationary blades, and the moving blades and the stationary blades are arranged alternately in sequence; wherein, the number of stages of the moving blades is one more than the number of stages of the stationary blades, so that the primary blades and the final blades of the compressor are both the moving blades.

8. The floating detonation turbine engine according to claim 6, characterized in that: The turbine is an axial flow turbine without guide vanes.

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