Combustion chamber and turbine guider integrated composite flow channel with multi-stage axial staged combustion device

Abstract

The invention relates to the field of integrated design of combustion chambers and turbine guiders of gas turbines, and particularly discloses a composite flow channel structure with a multi-stage axial staged combustion device. A low-speed combustion area is constructed on the front section of the combined type flow channel through a gradually-shrunk molded surface, micro-mixing combustion units distributed in a multi-stage rotational flow gradient mode are arranged in the axial direction, and fuel staged injection and step-by-step mixing combustion are achieved through an embedded type inserting connection structure. The rear section retains the pneumatic profile of a loading blade profile behind the guider, and gas expansion acceleration and turning are controlled through a curvature continuous flow channel. The mechanical interface of the transition section is eliminated through structural fusion, flow separation and thermal stress concentration are effectively avoided, the non-uniformity of a temperature field is reduced and the generation of nitrogen oxides is inhibited by adopting an axial staged combustion structure, and the independence of a combustion area and a flow guide turning area is ensured by the pneumatic decoupling characteristic of the post-loading blade profile. The combined type flow channel with the multi-stage axial staged combustion device improves the combustion efficiency and the pneumatic performance, and enhances the structural reliability under the high-temperature working condition.

Description

Technical Field

[0001] The present invention belongs to the technical field of heavy-duty ground gas turbines, and in particular is a composite flow passage integrating a combustion chamber with a multi-stage axially graded combustion device and a turbine guide. Background Art

[0002] Conventional gas turbine designs connecting the combustor and turbine guide vane via a transition section suffer from significant combustion flow field discontinuity. As an independent functional component, the transition section simultaneously performs high-temperature gas flow guidance, pressure maintenance, and structural support. When the high-temperature gas at the combustor outlet enters the transition section, the discontinuous flow path and abrupt geometric changes can lead to reduced flame front stability, uneven distribution of combustion products, and even localized backflow. Most notably, the diffuser structure at the front of the transition section, while reducing gas velocity, exacerbates turbulent mixing uncertainty, distorting the combustor outlet temperature field. This distortion is directly transmitted to the turbine guide vane inlet, resulting in uneven heat load distribution along the guide vane leading edge and the potential for high-temperature hotspots in certain areas, impacting combustion efficiency and emissions performance. The mixing process between the thermal boundary layer and the mainstream gas on the inner wall of the transition section is difficult to precisely control, which can easily lead to secondary combustion of unburned fuel or an extended NOx generation zone, further hindering the coordinated optimization of the combustor and turbine components. In the prior art, complex film cooling or multi-stage cyclone designs are usually used to solve such problems, but these methods do not fundamentally eliminate the flow field fragmentation problem caused by the transition section.

[0003] From an aerodynamic design perspective, the traditional transition section, as a separate component, introduces additional flow separation and total pressure losses. As the combustion chamber exit gas flows through the transition section, the discontinuous change in cross-section of the fan-shaped inlet flow path from the combustion chamber outlet to the turbine guide vane causes the gas to form localized low-speed regions and separated vortex structures under the influence of viscosity and pressure gradients. While the convergent section at the rear end of the transition section can accelerate the flow to meet the requirements of the turbine guide vane throat, the excessive contraction rate can create a localized supersonic region, leading to wave interference within the guide vane and exacerbating momentum exchange losses between the mainstream gas and the endwall boundary layer. Furthermore, the assembly gap between the transition section and the guide vane induces gap leakage flow, which mixes with the mainstream gas to form an unsteady wake structure. This not only reduces the aerodynamic efficiency of the turbine stage but also induces periodic excitation forces on the guide vane blade surface, threatening structural reliability. Existing technologies primarily attempt to alleviate these issues by optimizing the transition section profile or adopting non-axisymmetric endwall designs. However, these approaches fail to overcome the inherent constraints of the combustion chamber and turbine guide vane as independent modules, and thus fail to achieve aerodynamic integration of the entire flow path.

[0004] In the traditional split structure, the combustion chamber and turbine guide vane as independent functional modules bear the combustion process and energy conversion function in the thermodynamic cycle respectively. This design paradigm leads to the transition section becoming the main source of structural redundancy. The transition section needs to meet the thermal protection requirements of the combustion chamber outlet and the aerodynamic load bearing requirements of the turbine guide vane. The wall thickness design and support structure of the transition section often adopts a compromise scheme. In order to withstand the high-temperature gas impact of the combustion chamber outlet, the transition section is usually designed as a double-wall structure and is equipped with multiple rows of impingement cooling holes, but these strengthening structures significantly increase the weight of the component, and form a stiffness mutation between the thin-walled aerodynamic profile of the turbine guide vane leading edge, which is prone to thermal stress concentration under high temperature gradient. The split structure needs to set flange connection or bolt fastening device, and these mechanical connection interfaces are prone to creep deformation in high temperature environment, which leads to the increase of gas leakage rate and vibration. In the prior art, the reliability is improved by increasing the temperature resistance of the transition section material or optimizing the connection structure, but the split design cannot fundamentally eliminate the contradiction between the thermal and mechanical loads between the combustion chamber and the turbine guide vane, and it is also difficult to realize the connection of the combustion endothermic process and the turbine expansion work process. From the perspective of emission control, the separate architecture of the traditional combustion chamber and turbine guide vane seriously limits the implementation effect of the staged combustion technology. In order to realize low nitrogen oxide emission, modern gas turbines generally adopt lean premixed staged combustion technology, but the flow field distortion caused by the transition section can destroy the mixing uniformity of the premixed fuel and air. The residence time of high-temperature gas at the outlet of the combustion chamber in the transition section is prolonged, which leads to the increase of local thermal nitrogen oxide generation, thereby weakening the emission reduction effect of the staged combustion.

[0005] In summary, the traditional gas turbine adopts the split architecture of combustion chamber-transition section-turbine guide vane, which has technical contradictions in combustion organization, aerodynamic matching and structural bearing in three dimensions, which results in the physical separation of the combustion reaction region and the turbine energy conversion region, and causes the forced interruption of the continuity of the thermodynamic process. The prior art can only achieve marginal improvement through transition section profile modification, cooling structure strengthening, etc., and cannot break through the fundamental limitation of the split architecture on the coordinated optimization of the combustion-expansion process. Therefore, it is necessary to construct an integrated flow channel integrating the combustion chamber and the turbine guide vane, which realizes the continuity of combustion organization and energy conversion by eliminating the transition section; at the same time, the flow channel should integrate the axial staged combustion device to realize seamless transition from the combustion chamber outlet to the guide vane inlet in the flow channel form. This integrated architecture will become a key technical path to break through the bottleneck of thermal efficiency and emission performance of gas turbines. SUMMARY

[0006] In view of the above problems, the present application provides a combustion chamber and turbine guide vane integrated composite flow channel, which integrates a multi-stage axial staged combustion device. Through structural fusion and aerodynamic decoupling design, the coordination control of combustion and expansion process is realized. The core features of the composite flow channel include:

[0007] 1. Continuous flow passage formed by leading edge structure modification and profile reconstruction;

[0008] 2. Blade profile parameter inheritance and local correction method based on aerodynamic equivalence criterion;

[0009] 3. Embedded multi-stage micro-mixing combustion unit layout with gradient distribution along the flow direction.

[0010] The purpose of the present application is achieved as follows:

[0011] The blunt nose region of the wing profile leading edge of the front section with equal thickness design is cut off by using the method of overall forming and local fine-tuning. Then, the profile curvature continuous reconstruction technology is applied to make the remaining profile smooth and closed to form a continuous flow passage containing the combustion chamber functional section and the aerodynamic section of the guide vane. Among them, the front section constitutes a low-speed combustion zone to improve the internal enthalpy energy of the gas, and the rear section forms an expansion acceleration zone to convert into mechanical energy, thereby realizing the integration of the turbine guide vane and the combustion chamber.

[0012] It should be noted that the pressure gradient of the integrated flow passage is jointly dominated by the pressure difference between the inlet and outlet, the pressure difference between the side wall and the end face, and the pressure difference between the upper and lower end faces, so that the overall flow passage presents a composite feature of axial contraction, circumferential torsion and radial oblique flow.

[0013] The rear-loaded aerodynamic shape of the guide vane profile is retained in the middle and rear sections of the flow passage, and the continuity of the flow passage curvature is controlled to realize the expansion and acceleration of the gas and the direction turning. The throat area ratio of the blade profile and the outlet flow angle inherit the original guide vane design parameters to ensure the equivalence of the aerodynamic function. The equivalence of the aerodynamic function is verified by the computational fluid dynamics (CFD) method. First, the baseline simulation of the complete turbine guide vane model is performed to obtain the outlet Mach number, flow coefficient and other key parameters; then the initial model of the integrated composite flow passage is generated, and the deviation of the outlet parameters is compared with the baseline model. If the deviation of the aerodynamic parameters exceeds the allowed range, profile correction is applied to the rear-loaded area of the middle section of the blade profile, and the flow disturbance caused by the cutting of the leading edge is compensated by adjusting the camber of the blade back and the distribution of the pressure gradient. Through iterative correction, the parameters such as the flow angle and the Mach number at the outlet of the integrated flow passage are equivalent to the original guide vane design indicators.

[0014] In terms of combustion organization, according to the CFD velocity field analysis results, multi-stage micro-mixing combustion units are arranged along the axial direction in the low-speed zone of the two side walls of the front section of the composite flow passage, and fuel is injected for combustion. The first-stage combustion unit is designed with low swirl number, and the fuel injection direction is at a small angle with the main flow direction to promote premixing. The swirl intensity and injection angle of each subsequent stage of combustion unit are gradually increased to utilize the shear layer to promote complete combustion. The combustion unit adopts an embedded plug-in structure, and the fuel injection angle of each stage of unit relative to the main flow direction shows an increasing trend, thereby realizing the step-by-step mixing and combustion intensification of unburned fuel.

[0015] The flow path adopts a rear-loaded airfoil configuration. The front section has a gentle curvature to maintain a low-speed uniform flow field, providing a stable airflow environment for the combustion section. The middle and rear sections of the airfoil form an acceleration channel to achieve the expansion and deflection of the combustion gas. The aerodynamic characteristics of the rear-loaded airfoil make it relatively insensitive to changes in the leading edge structure, ensuring that the aerodynamic performance of the combustion and guide vane sections can be independently controlled.

[0016] In summary, the integrated composite flow channel of the combustion chamber and turbine guide vane with a multi-stage axially staged combustion device provided by the present invention solves the inherent contradictions of the traditional split design: the physical fusion of the combustion section and the guide vane section eliminates the flow field distortion caused by the transition section; the axially staged combustion device realizes the coordinated organization of combustion heat release and combustion gas expansion in the continuous flow channel; the aerodynamic decoupling characteristics of the post-loaded blade shape ensure that the optimization of the combustion section structure will not significantly interfere with the guide vane performance.

[0017] Beneficial effects of the present invention:

[0018] 1. Improved aerodynamic efficiency and stability: The integrated composite flow path of the combustion chamber and turbine guide vane effectively reduces the flow separation and total pressure loss caused by the traditional transition section. The aerodynamic decoupling characteristics of the afterloaded blade ensure the stability of the guide vane outlet airflow angle and flow rate, thereby improving the aerodynamic efficiency and turbine efficiency of the integrated component.

[0019] 2. Optimized combustion and thermal management: The axially staged combustion unit achieves spatiotemporal matching of combustion heat release and gas expansion within a continuous flow channel. The multi-stage mixing mechanism reduces the non-uniformity of the outlet temperature field while suppressing the formation of nitrogen oxides. The continuity of the flow channel eliminates the secondary combustion phenomenon that may exist in the traditional transition section.

[0020] 3. Enhanced structural reliability: The mechanical connection interface of the transition section is eliminated through integral molding, avoiding the risk of thermal stress concentration and high-temperature creep; the overall load-bearing structure enhances structural reliability through variable thickness wall design and coordinated control of thermal expansion. BRIEF DESCRIPTION OF THE DRAWINGS

[0021] Figure 1 and Figure 2 A schematic diagram of a composite flow passage integrating a combustion chamber and a turbine guide vane with a multi-stage axially staged combustion device according to the present invention;

[0022] Figure 3 This is a partially enlarged view of the multi-stage axially graded combustion device. DETAILED DESCRIPTION

[0023] The present invention will be described in more detail below with reference to the accompanying drawings:

[0024] The integrated flow channel is composed of the hub profile, casing profile, and the guide vane pressure and suction profiles, exhibiting a composite characteristic of axial contraction, circumferential twist, and radial oblique flow. The hub and casing profiles adopt an axially tapered design, the pressure surface achieves aerodynamic optimization through a three-stage curvature variation, and the suction surface adopts an S-shaped compound curvature design. An even number of combustion units (14, 16, 18, 20, 22, 24, 26, 28, and 30) are arranged around the entire circumference, dynamically adapting their number to the length of the front combustion zone (150-280 mm). The length of the rear expansion acceleration zone is reversely adjusted to stabilize the vane density of the front combustion zone within the range of 1.2-1.3, meeting the requirements of the gas turbine's split-half structure.

[0025] The combustion zone utilizes a three-stage organization scheme. The main micro-mixing burner is embedded in the fan-shaped inlet surface, with a depth of 100±0.5mm. It contains 30-50 honeycomb-shaped precision fuel holes with a diameter of 1.0-1.5mm and a tolerance of ±0.02mm. A hyperbolic staggered injection layout distributes 65%-70% of the total fuel flow. A two-stage axially staged burner is added to the low-speed section of the casing: the first stage, located at an isentropic Mach number of 0.2, utilizes a double-row staggered fuel hole layout with a hole spacing gradient of 1.2–1.5mm, distributing 20%–25% of the fuel flow. The second stage, located at a Mach number of 0.3, features a single row of linearly arranged fuel holes, distributing 5%–10% of the fuel flow. Fuel is supplied from external casing piping. The combustion zone profile utilizes a constant thickness design, with the hub and casing profiles maintaining a slightly expanded feature, with an expansion angle of less than 3°, to ensure low-speed flow stability.

[0026] The guide section achieves efficient energy conversion through refined surface control. The pressure surface curvature radius is controlled segmentally along the chord length, decreasing gradually from 3.8 to 4.2 mm in the forward section (0–30% chord length); maintaining a constant 85 ± 2 mm in the midsection (30%–70% chord length); and increasing by 25%–30% in the trailing edge section (70%–100%). The suction surface adopts an S-shaped compound curvature, with the point of maximum curvature shifted aft to 60%–65% chord length, resulting in a curvature reduction of 15%–20% compared to conventional blade profiles. The throat section features an asymmetric convergence design, with convergence angles of 20°–25° on the pressure side and 30°–35° on the suction side. By matching the ratio of the expansion section axial length to the throat height (2.8–3.2), the exit Mach number is precisely controlled within the range of 0.87–0.89. The outlet airflow angle is controlled by the camber line of the rear-loaded blade profile, with a bend angle gradient of 0.25°–0.35° / mm, constrained within the range of 69°–71°. Simultaneously, the cascade density in the aft guide zone follows a gradient of 1.20→1.25→1.30 from root to tip, reducing secondary flow losses in the tip region by 40%–50% compared to conventional designs. The entire flow channel profile is generated using Bezier curves and NURBS surface technology, with curvature strictly meeting the second-order derivative continuity criterion. The inlet and outlet curvature radii are 120±5 mm and 85±3 mm, respectively.

[0027] Three-dimensional CFD simulations have verified that the design has a cold-state total pressure recovery coefficient greater than 0.97 and an outlet flow field nonuniformity less than 8%. The fluid-driven enthalpy difference established by the main micro-mixed burner, combined with axial staged combustion, synergizes to enhance the gas's ability to work. The throat design and afterloaded blade profile control jointly ensure flow stability and outlet airflow angle accuracy. Compared to a split-type structure, this integrated flow path significantly improves energy conversion efficiency and suppresses secondary flow losses through deep aerodynamic-combustion synergy, meeting the high-load operation requirements of advanced heavy-duty gas turbines.

Claims

1. The combustion chamber and turbine guide vane are integrated into a composite flow channel, characterized by: The integrated composite flow channel of the combustion chamber and the turbine guide vane is formed by the hub profile, the casing profile, the guide vane pressure profile and the suction profile. The flow channel is divided into a front combustion zone and a rear guide expansion acceleration zone to form a fuel-injection combination; the hub profile and the casing profile adopt an axially tapered shape, the curvature of the pressure profile changes in sections along the chord length, and the suction profile adopts an S-shaped composite curvature; the number of fuel-injection combination units arranged around the entire circumference is adapted to the length of the front combustion zone, the number of units is an even number, and the length of the front combustion zone and the length of the rear guide zone are adjusted in opposite directions to maintain the stability of the blade density; the outlet airflow angle and the outlet Mach number of the fuel-injection combination are collaboratively controlled by the flow channel profile.

2. Multi-stage axial combustion system and fuel distribution, characterized by: The combustion zone includes an embedded main-fuel micro-mixing burner and a two-stage axially staged burner; the main-fuel micro-mixing burner has a honeycomb fuel hole array, the fuel holes are arranged in a staggered manner and at a predetermined angle to the wall, distributing the main part of the total fuel amount; the first-stage staged burner is arranged in the low-speed area of ​​the casing surface and adopts a double-row staggered fuel hole structure, and the second stage is arranged in the flow field transition area and adopts a single-row fuel hole structure.

3. The aerodynamic design of the diversion expansion acceleration zone is characterized by: The radius of curvature of the pressure surface is controlled in sections along the chord length, including a decreasing zone, a constant zone and a sudden increase zone; the suction surface adopts an S-shaped compound curvature, and the maximum curvature point is located at the rear section of the chord length; the throat section is asymmetric and convergent, and the convergence angles of the pressure side and the suction side are set differently; the arc line of the rear-loaded blade has a gradually increasing bending angle feature, and the blade density increases gradually from the blade root to the blade tip.

4. The surface generation and parameter control method is characterized by: All surfaces are generated using Bezier curves and NURBS surfaces, with curvature continuous and second-order differentiable; the inlet and outlet curvature radii of the hub surface are configured according to a predetermined ratio, and the three-dimensional torsion angle is distributed along a radial gradient; the circumferential torsion angle of the casing surface varies according to a progressive function; the outlet airflow angle is precisely constrained by controlling the gradient of the blade bending angle.

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