Gas turbine combustion and spraying combined body low-consistency blade flow guide combustion chamber device

Abstract

The invention belongs to the technical field of combustion and spraying combinations, and provides a gas turbine combustion and spraying combination low-consistency blade flow guide combustion chamber device which comprises a plurality of combustion chamber stages distributed in the circumferential direction, each combustion chamber stage comprises a circulation channel, and a micro mixing combustion chamber is arranged at the front end of each circulation channel. The micro mixed combustion chamber is used for mixing and combusting mainstream gas from a gas compressor and fuel oil, integrated partition plates are arranged in the circulation channel and are arranged in the circulation direction of the mainstream gas, and the micro mixed combustion chamber is arranged between the two integrated partition plates. According to the invention, the flame tube wall of the traditional combustion chamber and the first-stage stationary blade of the turbine are integrally designed, so that the cold air consumption is reduced, and the circulation efficiency of the gas turbine is improved.

Description

Technical Field

[0001] The invention belongs to the technical field of fuel-injection combination, and in particular relates to a low-consistency blade guide combustion chamber device of a fuel-injection combination of a gas turbine. Background Art

[0002] Gas turbine engine is a typical thermal power device. It has been widely used in modern aircraft propulsion, ship power and industrial production due to its powerful output power and high thermal efficiency. Although the requirements and working conditions of gas turbine engines in different application fields are different, their development direction is to obtain higher cycle thermal efficiency and specific power. The working principle of the Brayton cycle points out that the improvement of the cycle efficiency and specific power of gas turbine engines depends on the increase of turbine inlet temperature. Since the beginning of the new century, gas turbine technology, especially its thermal efficiency, has been in a process of continuous improvement. At present, the turbine inlet temperature of commercialized international advanced heavy-duty gas turbine products (such as Mitsubishi's J-class, GE and Siemens' H-class) has reached 1600℃, which is far higher than the high temperature limit of blade materials (the long-term use temperature of turbine blade metal materials is 900-950℃). In order to ensure the safe and reliable operation of gas turbine turbines at higher operating temperatures, for conventional configuration gas turbines, it is currently mainly achieved by improving the upper temperature limit of hot end component materials and advanced cooling technology. Assuming that the turbine inlet temperature reaches 1650℃, the temperature resistance limit of the high-temperature alloy considering economy and availability is about 950℃. The thermal barrier coating (TBC) can provide about 100℃ of temperature resistance improvement, and the remaining 600℃ of temperature resistance mainly depends on efficient cooling technology. Without considering the limitation of cold air flow, the existing advanced cooling technology can fully achieve this indicator. However, according to the design experience of gas turbines, for every 1% increase in cold air consumption, the simple cycle efficiency of the gas turbine will decrease by about 0.12%. Therefore, using more cold air to effectively cool the turbine blades also restricts the improvement of the thermal efficiency of the gas turbine to a certain extent.

[0003] In the design process of traditional gas engines, the design and development of the combustion chamber and the turbine are carried out relatively independently, ignoring the gas-heat coupling characteristics of the combustion chamber and the turbine. With the application of advanced combustion organization methods, the outlet characteristics of the advanced gas engine combustion chamber are quite different from those of the traditional lean premixed combustion chamber. The higher temperature and complex aerodynamic-heat transfer field at the combustion chamber outlet will bring greater challenges to the cooling design of the turbine blades. The adoption of advanced combustion technology leads to the flow characteristics of the combustion chamber outlet showing non-uniform gas-heat parameter distribution with non-uniform temperature (hot spots), strong swirl and high turbulence. The circumferential airflow angle and radial pitch angle of the combustion chamber outlet swirl are up to about ±50°; the swirl in the combustion chamber causes its outlet turbulence to reach a maximum of about 40%. In order to improve the efficiency of aircraft engines, the turbine inlet temperature continues to rise, resulting in an increase in the heat load of the high-pressure turbine stage; in addition, due to the mutual interference of the swirls downstream of adjacent nozzles and the interaction between the swirl and the hot spots, the combustion chamber outlet presents a distorted hot spot temperature distribution. Therefore, in order to further improve the aero-thermal performance of high-temperature key components, it is necessary to take into account and comprehensively consider the interaction between the combustion chamber and the turbine, carry out an integrated aero-thermal design of the combustion chamber and the turbine based on the component coupling structure, and use the component coupling design to improve the performance of the entire machine. Summary of the invention

[0004] The object of the present invention is to provide a low-consistency blade guide combustion chamber device for a gas turbine fuel injection combination to solve the above-mentioned problems, thereby achieving the purpose of integrating the traditional combustion chamber flame tube wall and the first-stage turbine stator blades, reducing the amount of cold air used, and improving the gas turbine cycle efficiency.

[0005] To achieve the above object, the present invention provides the following solution: A low-consistency blade guide combustion chamber device of a gas turbine fuel injection combination, comprising:

[0006] A plurality of combustion chamber stages are circumferentially distributed, each of the combustion chamber stages comprising a circulation channel, a micro-mixing combustion chamber being arranged at the front end of the circulation channel, the micro-mixing combustion chamber being used for mixing and burning the mainstream gas from the compressor with the fuel, an integrated baffle being arranged in the circulation channel, the integrated baffle being arranged along the circulation direction of the mainstream fuel gas, the integrated baffle being used for guiding the airflow to flow downstream and playing the role of a guide vane, thereby guiding the airflow to the moving blades downstream of the gas turbine, and the micro-mixing combustion chamber being arranged between two of the integrated baffles.

[0007] Preferably, the number of the circumferential arrays of combustion chamber stages is between 8 and 24.

[0008] Preferably, the circulation channel includes an upper wall surface of the fuel-injection combination and a lower wall surface of the fuel-injection combination, and the micro-mixing combustion chamber and the integrated partition are fixedly connected between the upper wall surface of the fuel-injection combination and the lower wall surface of the fuel-injection combination.

[0009] Preferably, the integrated partition includes a partition, which is arranged along the flow direction of the mainstream fuel gas and is arranged close to the micro-mixing combustion chamber, and one end of the partition away from the micro-mixing combustion chamber is fixedly connected with a guide vane.

[0010] Preferably, the baffle is provided with a straight section at the front end and a contraction section at the rear end along the flow direction of the mainstream gas. The front section between the two baffles is straight to make the airflow flow smoothly, and the rear section between the two baffles is contraction to accelerate the airflow.

[0011] Preferably, the micro hybrid combustion chamber is composed of a plurality of circular tubular channels in a rectangular array.

[0012] Preferably, the diameter of the circular tube is 7 mm.

[0013] Compared with the prior art, the present invention has the following advantages and technical effects: the present invention integrates the combustion chamber flame tube and the first-stage stator blades to ensure the diversion of hot gas while preventing the blades from being directly exposed to the hot gas channel, thereby protecting the blades from direct impact of hot spots and solving the problem of reduced film cooling efficiency coefficient caused by interference of the flame tube wake on the blades. The back heat transfer coefficient of the guide vane on the integrated partition is significantly lower than that of the split first-stage guide vane. The integrated design simplifies the structure of the engine combustion chamber and turbine section, reduces the cooling gas volume, and is of great help in improving the engine's cycle efficiency. BRIEF DESCRIPTION OF THE DRAWINGS

[0014] In order to more clearly illustrate the embodiments of the present invention or the technical solutions in the prior art, the drawings required for use in the embodiments will be briefly introduced below. Obviously, the drawings described below are only some embodiments of the present invention. For ordinary technicians in this field, other drawings can be obtained based on these drawings without paying creative work.

[0015] Figure 1 It is a front view of the combustion chamber stage of the present invention;

[0016] Figure 2 is a schematic diagram of an integrated partition of the present invention;

[0017] Figure 3 is a schematic diagram of a micro hybrid combustion chamber of the present invention;

[0018] Among them, 1. micro-mixing combustion chamber; 2. upper wall of the fuel-injection combination; 3. lower wall of the fuel-injection combination; 4. integrated partition; 401. guide vane; 402. partition. DETAILED DESCRIPTION

[0019] The following will be combined with the drawings in the embodiments of the present invention to clearly and completely describe the technical solutions in the embodiments of the present invention. Obviously, the described embodiments are only part of the embodiments of the present invention, not all of the embodiments. Based on the embodiments of the present invention, all other embodiments obtained by ordinary technicians in this field without creative work are within the scope of protection of the present invention.

[0020] In order to make the above-mentioned objects, features and advantages of the present invention more obvious and easy to understand, the present invention is further described in detail below with reference to the accompanying drawings and specific embodiments.

[0021] Reference Figure 1-Figure 3 The present invention provides a low-consistency blade guide combustion chamber device of a gas turbine fuel injection combination, comprising:

[0022] A plurality of combustion chamber stages are circumferentially distributed, each of which includes a circulation channel, a micro-mixing combustion chamber 1 being arranged at the front end of the circulation channel, the micro-mixing combustion chamber 1 being used for mixing and burning the mainstream gas from the compressor with the fuel, an integrated baffle 4 being arranged in the circulation channel, the integrated baffle 4 being arranged along the circulation direction of the mainstream fuel gas, the integrated baffle 4 being used for guiding the airflow to flow downstream and playing the role of a guide vane, thereby guiding the airflow to the moving blades downstream of the gas turbine, and the micro-mixing combustion chamber 1 being arranged between two integrated baffles 4.

[0023] The main function of the integrated baffle 4 is to combine the combustion chamber baffle and guide vane structure of the traditional gas turbine. By halving the number of the original turbine guide vanes and enlarging the blade shape accordingly, the front part of the guide vane is connected to the baffle to form an integrated baffle 4, so as to achieve the effect of integrating the combustion chamber and the turbine guide vane structure, reduce the heat load of the high-temperature gas from the combustion chamber on the guide vane, and reduce the amount of cold air used in the combustion chamber as a whole. On the whole, the present invention integrates the combustion chamber flame tube and the first-stage stator blades, ensures the flow guidance effect on the hot gas, and prevents the blades from being directly exposed to the hot gas channel, protects the blades from being directly impacted by the hot spot, solves the problem of the blade being disturbed by the flame tube wake, resulting in a reduction in the air film cooling efficiency coefficient, and the back heat transfer coefficient of the guide vane on the integrated baffle is significantly lower than that of the split first-stage guide vane; the integrated design simplifies the structure of the engine combustion chamber and turbine section, reduces the amount of cooling gas, and is of great help in improving the cycle efficiency of the engine.

[0024] To further optimize the solution, the end of the integrated partition 4 is arranged to correspond to the downstream moving blade.

[0025] The combustion chamber stage in the present invention adopts a compact design, which further shortens the axial size of the heavy-duty gas turbine. The combustion chamber adopts an advanced micro-mixing combustion chamber, which reduces the emission of nitrogen and oxygen pollutants.

[0026] According to a further optimization scheme, the number of circumferential arrays of several combustion chamber stages is between 18 and 24.

[0027] According to a further optimized solution, the flow channel includes an upper wall surface 2 of the fuel-injection combination and a lower wall surface 3 of the fuel-injection combination, and the micro-mixing combustion chamber 1 and an integrated partition plate 4 are fixedly connected between the upper wall surface 2 of the fuel-injection combination and the lower wall surface 3 of the fuel-injection combination.

[0028] According to a further optimization scheme, the integrated partition 4 includes a partition 402 , which is arranged along the flow direction of the mainstream gas, and is arranged close to the micro-mixing combustion chamber 1 , and a guide vane 401 is fixedly connected to one end of the partition 402 away from the micro-mixing combustion chamber 1 .

[0029] According to a further optimization scheme, the partition 402 is provided with a straight section at the front end and a contraction section at the rear end along the flow direction of the mainstream gas. The front section between the two partitions 402 is straight to ensure smooth flow of the airflow, and the rear section between the two partitions 402 is contraction to accelerate the flow of the airflow.

[0030] According to a further optimized solution, the micro hybrid combustion chamber 1 is composed of a plurality of rectangular arrays of circular tubular channels.

[0031] In this embodiment, the micro-mixing combustion chamber 1 is composed of 8×10 circular tubular channels.

[0032] The solution was further optimized and the diameter of the round tube was 7 mm.

[0033] The working process of this embodiment is as follows: the mainstream gas from the compressor passes through the micro-mixing combustion chamber 1 to mix with the fuel, and at the same time, a part of the cooling gas in the front chamber also flows through the micro-mixing combustion chamber 1, and is burned after mixing. At this time, due to the small number of components in the fuel, it belongs to lean combustion, and relatively few pollutant emissions are generated. The low-pollutant mainstream gas flow passes through the flow channel composed of two adjacent integrated partitions 4, the upper wall surface 2 of the fuel-injection joint body, and the lower wall surface 3 of the fuel-injection joint body. The flow channel is straight in the straight section, and the airflow will not be accelerated here. The flow speed is slow, which is conducive to full combustion. After that, the airflow enters the contraction section of the flow channel. The contraction section presents a contraction feature, which is conducive to wrapping the flame. At the same time, graded combustion can be designed on the integrated partition 4 of this section, which can efficiently reuse the blade cooling gas, optimize the combustion strategy, and increase the combustion temperature; when the airflow flows through the rear section of the flow channel, that is, when it flows through the guide vane 401, it undergoes flow changes similar to those in the turbine guide vane, the flow channel profile changes gently, and the flow channel sections are connected smoothly, reducing the flow loss of the airflow. The integrated baffle integrates the combustion chamber and turbine structure, reducing the impact of the high-temperature combustion gas in the combustion chamber on the guide vanes. Combined with efficient cooling technology, it reduces the overall cooling air consumption at the combustion chamber level. At the same time, it can compact the structural design of the entire machine and shorten the axial size of the gas turbine.

[0034] In the description of the present invention, it should be understood that the terms "longitudinal", "lateral", "up", "down", "front", "back", "left", "right", "vertical", "horizontal", "top", "bottom", "inside" and "outside" etc., indicating orientations or positional relationships, are based on the orientations or positional relationships shown in the accompanying drawings, and are only for the convenience of describing the present invention, and do not indicate or imply that the device or element referred to must have a specific orientation, be constructed and operated in a specific orientation, and therefore should not be understood as a limitation on the present invention.

[0035] The embodiments described above are only descriptions of the preferred modes of the present invention, and are not intended to limit the scope of the present invention. Without departing from the design spirit of the present invention, various modifications and improvements made to the technical solutions of the present invention by ordinary technicians in this field should all fall within the protection scope determined by the claims of the present invention.

Claims

1. A low-consistency blade guide combustion chamber device for a gas turbine fuel injection combination, characterized in that: include: A plurality of combustion chamber stages are circumferentially distributed, wherein the combustion chamber stage comprises a circulation channel, wherein a micro-mixing combustion chamber (1) is arranged at the front end of the circulation channel, wherein the micro-mixing combustion chamber (1) is used to mix and burn the mainstream gas from the compressor with the fuel, wherein an integrated baffle (4) is arranged in the circulation channel, wherein the integrated baffle (4) is arranged along the circulation direction of the mainstream fuel gas, wherein the integrated baffle (4) is used to guide the airflow to flow downstream and play the role of a guide vane, thereby guiding the airflow to the moving blades downstream of the gas turbine, wherein the micro-mixing combustion chamber (1) is arranged between two of the integrated baffles (4).

2. A low-consistency blade guide combustion chamber device for a gas turbine fuel injection combination according to claim 1, characterized in that: The number of the circumferential arrays of the combustion chamber stages is between 18 and 24.

3. A low-consistency blade guide combustion chamber device for a gas turbine fuel injection combination according to claim 1, characterized in that: The circulation channel comprises an upper wall surface (2) of the fuel-injection combination and a lower wall surface (3) of the fuel-injection combination, and the micro-mixing combustion chamber (1) and the integrated partition plate (4) are fixedly connected between the upper wall surface (2) of the fuel-injection combination and the lower wall surface (3) of the fuel-injection combination.

4. A low-consistency blade guide combustion chamber device for a gas turbine fuel injection combination according to claim 1, characterized in that: The integrated partition (4) comprises a partition (402), the partition (402) being arranged along the flow direction of the mainstream fuel gas, and the partition (402) being arranged close to the micro-mixing combustion chamber (1), and a guide vane (401) being fixedly connected to one end of the partition (402) away from the micro-mixing combustion chamber (1).

5. A low-consistency blade guide combustion chamber device for a gas turbine fuel injection combination according to claim 1, characterized in that: The partition (402) is provided with a straight section at the front end and a contraction section at the rear end along the flow direction of the mainstream fuel gas; the front section between the two partitions (402) is straight for making the air flow flow smoothly, and the rear section between the two partitions (402) is contraction for making the air flow flow faster.

6. A low-consistency blade guide combustion chamber device for a gas turbine fuel injection combination according to claim 1, characterized in that: The micro-mixing combustion chamber (1) is composed of a plurality of circular tubular channels in a rectangular array.

7. A low-consistency blade guide combustion chamber device for a gas turbine fuel injection combination according to claim 6, characterized in that: The diameter of the round tube is 7 mm.

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