A combustion chamber device for ground-based heavy-duty gas turbines

Abstract

This invention belongs to the technical field of ground-based heavy-duty gas turbines, and provides a combustion chamber device for ground-based heavy-duty gas turbines. It includes several circumferentially distributed combustion chamber stages, each comprising a first channel and a second channel interconnected from front to back. A micro-mixing combustion chamber is disposed within the first channel, used to mix and burn the mainstream gas from the compressor with fuel oil. An integrated baffle blade is disposed within both the first and second channels, arranged along the flow direction of the mainstream gas flow. A cooling assembly includes a cold source and several microchannels disposed within the integrated baffle blade. The cold source cools the integrated baffle blade by supplying cooling gas into the microchannels. This invention can significantly reduce the amount of cooling gas used during gas turbine operation while maintaining considerable combustion efficiency.

Description

Technical Field

[0001] This invention belongs to the field of ground-based heavy-duty gas turbine technology, and particularly relates to a combustion chamber device for ground-based heavy-duty gas turbines. Background Technology

[0002] Gas turbines operate on similar principles to aircraft engines. However, because gas turbines are always ground-based, they don't focus on maneuvering parameters such as thrust and thrust-to-weight ratio. To achieve industrial-grade power generation, ground-based gas turbines often prioritize high efficiency to reduce power generation costs and improve operational profits. Gas turbines utilize the Brayton cycle. Based on the principles of this cycle, to improve efficiency, ground-based gas turbines must continuously increase the turbine inlet temperature. As the turbine inlet temperature increases, the optimal pressure ratio should also increase. To reduce combustion pollutant emissions, lean combustion technology is employed, primarily to reduce nitrogen oxide emissions during combustion.

[0003] However, the enhanced performance and lower-emission combustion conditions also present a series of challenges to gas turbine design. Higher performance leads to higher turbine inlet temperatures, exceeding the material's temperature resistance limit. This necessitates employing cooling design methods to reduce blade temperature. This temperature increase requires either more efficient cooling methods or a larger volume of cooling gas. While lean-mixed combustion, used to test lower-emission combustion, can reduce emissions, it also introduces new challenges, particularly regarding the cooling structure design of the first-stage guide vanes, due to the varying temperatures at the combustion chamber outlet. Uneven gas distribution will lead to hot spots at the inlet. The localized high temperature of these hot spots subject the guide vanes to higher temperature loads, increasing the risk of localized burn-out of the blades. Simultaneously, hot and cold gas separation occurs within the moving blades. Furthermore, as the axial length of the engine combustion chamber gradually shortens, the swirling flow generated by the combustion chamber vortex generator does not completely dissipate at the outlet. Residual swirling flow will enter the turbine components with the mainstream flow, forming a significant secondary flow at the blade's mid-span. During its downstream propagation, the powerful swirling flow entrains nearby film cooling gas, and the secondary vortex carries the cooling gas downwards, resulting in insufficient cooling gas coverage on the blade surface and significantly reducing cooling efficiency. Current research indicates that the combined effect of swirling flow and hot spots greatly impacts turbine blade cooling efficiency. Hot spots themselves have higher temperatures, and the swirling flow weakens the turbine blade's thermal protection by entraining the cooling film. The interaction between the two is further manifested in the following ways: the hot spot, under the entrainment of the swirling flow, continues to propagate downstream; the high turbulence of the swirling flow enhances heat transfer between the high-temperature fluid and the wall plate; and this interaction further deteriorates the thermal environment of the turbine blades. How to resolve the conflict between the high performance and high operating conditions of heavy-duty gas turbines and the design difficulties is an urgent problem that needs to be solved in the field.

[0004] The current mainstream research focuses on exploring the flow field characteristics of the combustion chamber turbine under coupled relationships, thereby obtaining the migration characteristics of swirl and hot spots in the flow field, and finally optimizing the cooling design of the turbine blades. The focus is on: designing cooling structures or changing upstream structures to serve the cooling design of downstream guide vanes. Other studies mostly treat the combustion chamber, turbine, and compressor as separate entities, emphasizing the adjustment and adaptation of multiple components to other components to improve performance.

[0005] In summary, how to design and propose a scheme that integrates the guide vanes and combustion chamber upstream and downstream into a single design, significantly reducing the amount of cooling air required while ensuring considerable combustion chamber efficiency, and taking into account structural issues such as processing and assembly, is an urgent problem to be solved in this field. Summary of the Invention

[0006] The purpose of this invention is to provide a combustion chamber device for ground-based heavy-duty gas turbines to solve the above-mentioned problems and achieve the goal of significantly reducing the amount of cooling air used while ensuring considerable combustion efficiency during gas turbine operation.

[0007] To achieve the above objectives, the present invention provides the following solution: a combustion chamber device for a ground-based heavy-duty gas turbine, comprising:

[0008] The combustion chamber stage is circumferentially distributed, and the combustion chamber stage includes a first channel and a second channel that are interconnected from front to back. A micro-mixing combustion chamber is provided in the first channel. The micro-mixing combustion chamber is used to mix and burn the mainstream gas from the compressor with fuel. An integrated baffle blade is provided in the first channel and the second channel. The integrated baffle blade is arranged along the flow direction of the mainstream gas.

[0009] A cooling assembly includes a cold source and several microchannels disposed within the integrated baffle blade body. The cold source cools the integrated baffle blade body by supplying cooling gas into the several microchannels.

[0010] Preferably, an angle is provided between the first channel and the second channel.

[0011] Preferably, the integrated baffle blade includes two sets of integrated microchannel cooling baffles. The near ends of the two sets of integrated microchannel cooling baffles are connected to form a cavity. The upper and lower ends of the cavity penetrate the walls of the first channel and the second channel, respectively. An internal front and rear cavity baffle of the integrated baffle is provided in the cavity. The internal front and rear cavity baffle of the integrated baffle is fixedly connected to the two sets of integrated microchannel cooling baffles.

[0012] Preferably, a plurality of cooling grooves are formed at equal intervals on the wall of the microchannel along the extension direction of the microchannel.

[0013] Preferably, the cold source includes a central cavity formed downwardly at the top of the front and rear cavity partitions inside the integrated partition, and a plurality of the microchannels communicate with the central cavity.

[0014] Preferably, the inner diameter of the microchannel is set between 150μm and 300μm, the length of the cooling groove along the extension direction of the microchannel is set between 250μm and 450μm, the distance between two adjacent cooling grooves is set between 300μm and 500μm, and the wall thickness of the integrated microchannel cooling partition is set between 3mm and 8mm.

[0015] Preferably, the distance between the two opposite outer walls of the front and rear cavity partitions inside the integrated partition is set between 5mm and 10mm, and the side wall thickness of the front and rear cavity partitions inside the integrated partition is set between 1mm and 2mm.

[0016] Preferably, the integrated microchannel cooling baffle has a plurality of axially graded combustion holes in the middle of its sidewall, and the plurality of axially graded combustion holes are arranged sequentially from top to bottom and are respectively connected to the interior of the second channel.

[0017] Preferably, one end of each of the microchannels near the micro-mixing combustion chamber passes through the integrated microchannel cooling baffle and communicates with the interior of the micro-mixing combustion chamber.

[0018] Compared with existing technologies, this invention has the following advantages and technical effects: The main function of the micro-mixing combustion chamber is to mix fuel and air separately from different microtube heads, improving the mixing effect; the main function of the integrated baffle blade body is to replace the turbine guide vanes, serving to streamline the airflow. Simultaneously, by placing the integrated baffle blade body within the combustion chamber stage, a compact integrated design is achieved, which can shorten the axial dimension of heavy-duty gas turbines; the main function of the microchannel is to allow cooling gas to flow within the integrated baffle blade body, controlling the temperature within a reasonable range. Overall, this invention reduces the axial distance by combining the combustion chamber and turbine guide vanes into an integrated combustion chamber stage. The integrated baffle blade body achieves the effects of airflow guidance, cooling, and axial combustion gradation, effectively reducing the amount of cooling gas used while ensuring aerodynamic efficiency. Attached Figure Description

[0019] To more clearly illustrate the technical solutions in the embodiments of the present invention or the prior art, the drawings used in the embodiments will be briefly introduced below. Obviously, the drawings described below are only some embodiments of the present invention. For those skilled in the art, other drawings can be obtained based on these drawings without creative effort.

[0020] Figure 1 This is a schematic diagram of the combustion chamber stage of the present invention;

[0021] Figure 2 This is a left view of the combustion chamber stage of the present invention;

[0022] Figure 3 for Figure 2 AA section view in the middle;

[0023] Figure 4 This is a schematic diagram of the microchannel of the present invention;

[0024] Figure 5 This is a schematic diagram of the gas flow path in this invention;

[0025] The components include: 1. Micro-mixing combustion chamber; 2. Integrated baffle blade; 3. Lower wall of combustion chamber stage; 4. Upper wall of combustion chamber stage; 5. Integrated microchannel cooling baffle; 6. Axial staged combustion holes; 7. Internal front and rear cavity baffles of integrated baffle; 8. Microchannel; 9. Cooling groove. Detailed Implementation

[0026] The technical solutions of the embodiments of the present invention will be clearly and completely described below with reference to the accompanying drawings. Obviously, the described embodiments are only some embodiments of the present invention, and not all embodiments. Based on the embodiments of the present invention, all other embodiments obtained by those skilled in the art without creative effort are within the scope of protection of the present invention.

[0027] To make the above-mentioned objects, features and advantages of the present invention more apparent and understandable, the present invention will be further described in detail below with reference to the accompanying drawings and specific embodiments.

[0028] Reference Figures 1-5 The present invention provides a combustion chamber device for a ground-based heavy-duty gas turbine, comprising:

[0029] The combustion chamber stage is circumferentially distributed. The combustion chamber stage includes a first channel and a second channel that are interconnected from front to back. A micro-mixing combustion chamber 1 is provided in the first channel. The micro-mixing combustion chamber 1 is used to mix and burn the mainstream gas from the compressor with fuel. An integrated baffle blade 2 is provided in the first channel and the second channel. The integrated baffle blade 2 is arranged along the flow direction of the mainstream gas.

[0030] The cooling assembly includes a cold source and several microchannels 8. The several microchannels 8 are disposed inside the integrated baffle blade body 2. The cold source cools the integrated baffle blade body 2 by delivering cooling gas into the several microchannels 8.

[0031] The main function of the micro-mixing combustion chamber 1 is to mix fuel and air separately from different microtube heads, improving the mixing effect. The main function of the integrated baffle blade 2 is to replace the turbine guide vanes, serving to streamline the airflow. Furthermore, by placing the integrated baffle blade 2 within the combustion chamber stage, a compact, integrated design is achieved, shortening the axial dimension of the heavy-duty gas turbine. The main function of the microchannel 8 is to allow cooling gas to flow within the integrated baffle blade 2, controlling the temperature within a reasonable range. Overall, this invention reduces the axial distance by combining the combustion chamber and turbine guide vanes into an integrated combustion chamber stage. The integrated baffle blade achieves the effects of airflow guidance, cooling, and axial combustion gradation, effectively reducing cooling gas consumption while maintaining aerodynamic efficiency.

[0032] Further optimization involves a micro-mixing combustion chamber 1 composed of multiple micropores. These micropores facilitate more dense mixing of fuel and air. The mainstream gas from the compressor is mixed with fuel in the micro-mixing combustion chamber 1 before combustion. Here, the fuel has fewer components, resulting in lean combustion and relatively fewer pollutant emissions.

[0033] The scheme is further optimized. The second channel includes the lower wall surface 3 of the combustion chamber stage and the upper wall surface 4 of the combustion chamber stage. The micro-mixing combustion chamber 1 and the integrated baffle blade 2 are fixedly connected between the lower wall surface 3 and the upper wall surface 4 of the combustion chamber stage.

[0034] Further optimization of the scheme resulted in a combustion chamber stage with structural dimensions that conform to the actual structural design of heavy-duty gas turbines. The main gas flow passage formed by the lower wall 3 and the upper wall 4 of the combustion chamber stage flows from the high radius to the low radius. The integrated baffle blade 2 is also an integrated baffle similar to the original guide vanes, obtained by halving the number of guide vanes and enlarging the volume based on the aerodynamic parameters of a certain aircraft engine turbine guide vanes. The aerodynamic parameter derivative design involved will not be elaborated here.

[0035] The design was further optimized by setting an angle between the first and second channels.

[0036] By setting an angle between the first and second channels, the front half of the integrated baffle blade 2 can be straight, which will not accelerate the gas. The rear half of the integrated baffle blade 2 is curved, and the gas will be accelerated when it flows through the rear half, resulting in a flow change similar to that inside a turbine guide vane.

[0037] Further optimization of the scheme: the integrated baffle blade 2 includes two sets of integrated microchannel cooling baffles 5. The near ends of the two sets of integrated microchannel cooling baffles 5 are connected to form a cavity. The upper and lower ends of the cavity pass through the walls of the first channel and the second channel, respectively. The cavity is provided with an integrated baffle internal front and rear cavity baffle 7. The integrated baffle internal front and rear cavity baffle 7 is fixedly connected to the two sets of integrated microchannel cooling baffles 5.

[0038] The primary function of the integrated microchannel cooling baffle 7 is to separate the cavity into two chambers, front and rear. Cooling air is introduced from top to bottom into both chambers to cool the walls of the integrated microchannel cooling baffle 5 via convection cooling. The cooling air for the front and rear chambers is supplied separately by corresponding structures within the gas turbine, and the cooling air from the front chamber flows into the micro-mixing combustion chamber 1 to participate in combustion. Secondly, the integrated baffle 7 also provides support for the integrated baffle blade 2, thereby improving structural strength.

[0039] To further optimize the design, several cooling grooves 9 are equally spaced on the wall of the microchannel 8 along the extension direction of the microchannel 8.

[0040] The scheme is further optimized so that the cold source includes a central cavity opened downward from the top of the front and rear cavity partitions 7 inside the integrated partition, and several microchannels 8 are connected to the central cavity.

[0041] The cooling airflow enters the microchannel 8 through the central cavity of the front and rear cavity partition 7 inside the integrated partition and flows towards both ends of the integrated partition blade 2 within the microchannel 8. As the cooling airflow flows within the microchannel 8, it forms vortices within the cooling tank 9. The resulting air resistance prevents the airflow located in the middle of the microchannel 8 from flowing along the wall of the microchannel 8, thereby changing the original gas-solid contact into a gas-gas contact. This greatly reduces flow resistance, giving the microchannel of the present invention low flow resistance characteristics.

[0042] Further optimization of the scheme: the inner diameter of the microchannel 8 is defined as d, which is set between 150μm and 300μm; the length of the cooling tank 9 along the direction of the microchannel 8 is defined as x, which is set between 250μm and 450μm; the distance between two adjacent cooling tanks 9 is defined as l, which is set between 300μm and 500μm; and the wall thickness of the integrated microchannel cooling partition 5 is defined as k, which is set between 3mm and 8mm.

[0043] Further optimization of the scheme: the distance between the two opposite outer walls of the front and rear cavity partition 7 inside the integrated partition is defined as p, and p is set between 5mm and 10mm. The side wall thickness of the front and rear cavity partition 7 inside the integrated partition is defined as m, and m is set between 1mm and 2mm.

[0044] The range of values ​​for the wall thickness k of the integrated microchannel cooling baffle 5 and the side wall thickness m of the front and rear cavity baffles 7 inside the integrated baffle is determined based on the current cooling structure and the minimum allowable thickness considering structural strength. The selection ranges for the inner diameter d of the microchannel 8, the length x of the cooling groove 9, the distance l between two adjacent cooling grooves 9, and the distance p between the two opposite outer walls of the front and rear cavity baffles 7 inside the integrated baffle are obtained through simulation. Under the premise of not compromising structural strength, the smoothness of airflow can be ensured to the greatest extent, guaranteeing the supply of cooling airflow.

[0045] To further optimize the design, the integrated microchannel cooling baffle 5 has several axially graded combustion holes 6 in the middle of its side wall. These axially graded combustion holes 6 are arranged sequentially from top to bottom and are connected to the interior of the second channel.

[0046] The gas initially combusted in the micro-mixing combustion chamber 1 moves towards the rear end along the integrated baffle blade 2 and passes through the middle of the integrated baffle blade 2. Fuel is then injected outward through the axial staged combustion holes 6. After further mixing with the previously lean combustion gas, further combustion occurs, thereby achieving axial staged combustion. This minimizes combustion pollutants while maximizing the combustion potential of the fuel.

[0047] In a further optimized design, several microchannels 8 are inserted through an integrated microchannel cooling baffle 5 at one end near the micro-mixing combustion chamber 1 and are connected to the interior of the micro-mixing combustion chamber 1.

[0048] The cold airflow delivered to the front end in the microchannel 8 flows into the micro-mixing combustion chamber 1 after passing through the integrated baffle blade 2, and mixes with the mainstream gas to aid combustion.

[0049] The working process of this embodiment is as follows:

[0050] In this embodiment, the distance l between two adjacent cooling tanks 9 is 350. The length x of the cooling tank 9 along the extension direction of the microchannel 8 is 300. The inner diameter d of microchannel 8 is set to 250. The distance p between the two opposite outer walls of the integrated partition 7 in the front and rear cavities is 7mm, the side wall thickness m of the integrated partition 7 in the front and rear cavities is 1.5mm, and the wall thickness k of the integrated microchannel cooling partition 5 is set to 5mm.

[0051] like Figure 5 As shown, the mainstream gas and fuel oil are mixed and burned in the micro-mixing combustion chamber 1 before entering the mainstream gas passage. When the mainstream gas passes through the middle of the integrated baffle blade 2, staged fuel oil is injected outward through the axial staged combustion holes 6, where it mixes with the mainstream gas and undergoes secondary combustion. The gas turbine's cooling gas supply component vertically injects cooling gas into the middle cavity of the front and rear chamber baffles 7 inside the integrated baffle, and flows into the microchannel 8. The cooling gas is transferred towards the front and rear ends in the microchannel 8, cooling the interior of the integrated microchannel cooling baffle 5. The cooling gas transferred to the rear end fills the entire rear end of the integrated baffle blade 2, including the trailing edge, while the cooling gas transferred to the front end flows through the microchannel 8 and into the micro-mixing combustion chamber 1, where it mixes with the mainstream gas. Simultaneously, the cooling gas also passes through the front and rear chambers to cool the outer wall of the integrated microchannel cooling baffle 5. The cooling gas in the front chamber flows into the micro-mixing combustion chamber 1 after passing through and participates in combustion.

[0052] In the description of this invention, it should be understood that the terms "longitudinal", "lateral", "up", "down", "front", "rear", "left", "right", "vertical", "horizontal", "top", "bottom", "inner", "outer", etc., indicate the orientation or positional relationship based on the orientation or positional relationship shown in the accompanying drawings, and are only for the convenience of describing this invention, and are not intended to indicate or imply that the device or element referred to must have a specific orientation, or be constructed and operated in a specific orientation, and therefore should not be construed as a limitation of this invention.

[0053] The embodiments described above are merely preferred embodiments of the present invention and are not intended to limit the scope of the present invention. Various modifications and improvements made by those skilled in the art to the technical solutions of the present invention without departing from the spirit of the present invention should fall within the protection scope defined by the claims of the present invention.

Claims

1. A combustion chamber device for a ground-based heavy-duty gas turbine, characterized in that, include: A plurality of combustion chamber stages are distributed circumferentially. Each combustion chamber stage includes a first channel and a second channel that are interconnected from front to back. A micro-mixing combustion chamber (1) is provided in the first channel. The micro-mixing combustion chamber (1) is used to mix and burn the mainstream gas from the compressor with fuel oil. An integrated baffle blade (2) is provided in the first channel and the second channel. The integrated baffle blade (2) is arranged along the flow direction of the mainstream gas. The cooling assembly includes a cold source and several microchannels (8), the several microchannels (8) are disposed in the integrated baffle blade body (2), and the cold source cools the integrated baffle blade body (2) by delivering cooling gas into the several microchannels (8); An angle is provided between the first channel and the second channel; The integrated baffle blade (2) includes two sets of integrated microchannel cooling baffles (5). The near ends of the two sets of integrated microchannel cooling baffles (5) are connected to each other to form a cavity. The upper and lower ends of the cavity pass through the walls of the first channel and the second channel, respectively. An integrated baffle internal front and rear cavity baffle (7) is provided in the cavity. The integrated baffle internal front and rear cavity baffle (7) is fixedly connected to the two sets of integrated microchannel cooling baffles (5). Several cooling grooves (9) are equally spaced on the wall of the microchannel (8) along the extension direction of the microchannel (8). The cold source includes a central cavity that is opened downward from the top of the front and rear cavity partitions (7) inside the integrated partition, and a plurality of the microchannels (8) are connected to the central cavity; The integrated microchannel cooling baffle (5) has several axially graded combustion holes (6) in the middle of its side wall. The several axially graded combustion holes (6) are arranged from top to bottom and are respectively connected to the interior of the second channel. One end of each of the microchannels (8) near the micro-mixing combustion chamber (1) passes through the integrated microchannel cooling baffle (5) and is connected to the interior of the micro-mixing combustion chamber (1).

2. The combustion chamber device for a ground-based heavy-duty gas turbine according to claim 1, characterized in that: The inner diameter of the microchannel (8) is set at 150. Up to 300 Between, the cooling tank (9) is set at a length of 250 along the extension direction of the microchannel (8). Up to 450 The spacing between two adjacent cooling tanks (9) is set at 300 mm. Up to 500 The wall thickness of the integrated microchannel cooling baffle (5) is set between 3 mm and 8 mm.

3. The combustion chamber device for a ground-based heavy-duty gas turbine according to claim 1, characterized in that: The distance between the two opposite outer walls of the front and rear cavity partitions (7) inside the integrated partition is set between 5mm and 10mm, and the side wall thickness of the front and rear cavity partitions (7) inside the integrated partition is set between 1mm and 2mm.

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