A fan stator and transition section integrated design layout structure

Abstract

The present invention provides a layout structure of a fan stator and a transition section fusion design, the layout structure includes an inner end wall, an outer end wall and a fan stator blade, the inner end wall, the outer end wall and the fan stator blade form a transition section; the inner end wall and the outer end wall form an annular channel, the annular channel inlet is connected to the fan rotor inner channel outlet, and the annular channel outlet is connected to the inner channel compressor inlet; the fan stator blades are uniformly distributed in the annular channel, and each blade tip is connected to the outer end wall, and the blade root is connected to the inner end wall; the inner end wall adopts a non-axisymmetric end wall design in the fan stator blade channel area, and the non-axisymmetric end wall uses trigonometric functions for curved surface modeling, which has abundant degrees of freedom. The present invention can effectively shorten the axial length of the engine and simplify the engine structure under the premise of ensuring that the performance of the compression system remains basically unchanged.

Description

Technical Field

[0001] The present invention relates to the technical field of micro-turbofan engines, and in particular to a layout structure of a fan stator and a transition section fusion design. Background Art

[0002] In the field of small intelligent cruise ammunition, specific usage scenarios determine that the power system should have the characteristics of simple structure, low cost, and high reliability. Therefore, the power form of the above-mentioned flight platform is mainly a micro-turbojet engine. One of the development trends of small intelligent cruise ammunition in the future is longer flight time and higher survivability, but the inherent technical characteristics of micro-turbojet engines such as high fuel consumption and high exhaust temperature cannot meet the demand for power systems in the above development trends.

[0003] With the development of technology, the space for reducing fuel consumption and exhaust temperature through conventional optimization design methods is very limited. Another feasible technical means is to use turbofan engines to replace turbojet engines. Compared with turbojet engines, due to the different engine configuration, turbofan engines have lower fuel consumption, exhaust temperature and noise. Therefore, in the fields of manned aircraft and cruise missiles, turbofan engines have gradually replaced turbojet engines and have been more and more widely used. However, in the field of small intelligent cruise ammunition, due to the miniaturization of turbofan engines, there are still problems such as complex structure, low thrust-to-weight ratio and high cost, so micro turbofan engines have not yet been used in the above fields.

[0004] In order to solve the problems existing in the miniaturization of turbofan engines, domestic and foreign scholars have proposed many solutions. For example, the patent "Single-rotor micro-turbofan engine using axial-flow oblique-flow tandem composite compression system" (CN201310048204.8) describes a composite compression system for a micro-turbofan engine. The axial-flow fan rotor and the oblique-flow compressor rotor are arranged in series, and the intermediate casing is arranged behind the axial-flow rotor to divide the axial-flow rotor outlet airflow into two. Compared with the conventional design, the axial-flow fan stator and transition section are omitted in the inner channel, the axial length of the system is shortened, and the system structure is simplified. However, there is no transition section between the axial-flow fan rotor and the inner channel oblique-flow rotor, and the inlet and outlet height difference of the oblique-flow rotor is limited under the same outer diameter; there is no stator between the axial-flow fan rotor and the oblique-flow rotor to twist the airflow into the axial direction, and the incoming flow of the oblique-flow rotor has a large positive pre-swirl. The above two factors make it difficult to improve the pressure ratio of the oblique-flow rotor. The inner channel pressure ratio is limited to the pressure ratio range that can be achieved by a single-stage oblique flow. Therefore, the fuel consumption rate of the engine has no obvious advantage over the turbojet engine.

[0005] According to the above analysis, in order to make the internal pressure ratio of the engine reach the conventional design level, the internal pressure transition section and the fan stator are indispensable. An effective way to reduce the complexity of the structure is to combine the internal pressure transition section and the fan stator blades into one, and arrange the fan stator blades in the transition section.

[0006] In the transition section, due to the change of wall curvature and channel area, there are radial and flow pressure gradients. The low-energy fluid in the wall boundary layer tends to gather at the hub in the second half, inducing the separation of the hub boundary layer. In addition, there is a circumferential pressure gradient in the stator blade channel, and the low-energy fluid in the blade surface and end wall boundary layer accumulates in the corner area formed by the blade suction surface and the hub, which is very likely to cause three-dimensional corner separation. After the transition section and the stator blade are combined into one, the two flow phenomena are superimposed, resulting in a significant deterioration of the flow field in the channel, and an increase in the airflow separation range in the corner area of ​​the blade suction surface at the hub, which seriously affects the performance of the components.

[0007] In response to this phenomenon, the patent "A high-low pressure turbine transition section layout structure and design method" (CN201410004116.2) proposed a transition section layout structure and design method that integrates the support plate and the low-pressure turbine guide vane. The key lies in the integrated design of the support plate blades and the low-pressure turbine guide vane. At the same time, non-axisymmetric modeling is used on the transition section hub wall to regulate the pressure gradient in the channel and reduce flow losses. The application object of this patent is the transition section between the high and low pressure turbines. Since there is a clear difference between the flow in the turbine guide vane and the flow in the fan stator (the airflow is accelerated and expanded in the turbine guide vane, and decelerated and expanded in the fan stator), the design method of this patent is not applicable to the integrated design of the fan stator and the transition section; in addition, the patent does not provide a specific design method for the non-axisymmetric end wall modeling. Summary of the invention

[0008] Purpose of the invention: In order to solve the problem that a large airflow separation is likely to occur in the transition section when the fan stator and the transition section are integrated, and to take into account both compactness and high aerodynamic performance, the present invention provides a layout structure of a fan stator and a transition section integration design, the layout structure comprising an inner end wall, an outer end wall and a fan stator blade, the inner end wall, the outer end wall and the fan stator blade forming a transition section;

[0009] The inner end wall and the outer end wall form an annular channel, the annular channel inlet is connected to the inner channel outlet of the fan rotor, and the annular channel outlet is connected to the inner channel compressor inlet;

[0010] The fan stator blades are evenly distributed in the annular channel in the circumferential direction, and the tip of each blade is connected to the outer end wall, and the root of each blade is connected to the inner end wall;

[0011] The inner end wall of the blade channel formed by adjacent fan stator blades adopts a non-axisymmetric end wall design, and the non-axisymmetric end wall shape of the inner end wall of each blade channel is the same.

[0012] The transition section is divided into a first turning section and a second turning section according to the curvature distribution; in the first turning section, the outer wall surface is a concave surface and the inner wall surface is a convex surface; in the second turning section, the outer wall surface is a convex surface and the inner wall surface is a concave surface.

[0013] At the inner end wall, the wall surface changes from a convex wall surface to a concave wall surface.

[0014] The non-axisymmetric end wall surface of the inner end wall of each fan stator blade channel is obtained by superimposing a radial disturbance amplitude on the original axisymmetric end wall surface. In the cylindrical coordinate system, the radial coordinate r, axial coordinate z, and circumferential coordinate θ are defined. The radial disturbance amplitude Δr at each position will be represented by the following radial disturbance amplitude control function:

[0015] Δr(z,θ)=A(z)C(θ)

[0016] Among them, A(z) and C(θ) are two independent amplitude control functions, among which A(z) is the axial amplitude control function; C(θ) is the circumferential amplitude control function.

[0017] The calculation formula of the axial amplitude control function A(z) is:

[0018]

[0019] Among them, h is the height of the stator blade; β is the end wall amplitude control coefficient; z0 and z1 are the axial starting position and ending position of the non-axisymmetric end wall shaping area respectively; n is the coefficient for controlling the axial relative position of the maximum disturbance amplitude point, n>0. The shaping area refers to the area where the non-axisymmetric design is performed, and also refers to the inner end wall area of ​​the fan stator blade channel.

[0020] The calculation formula of the circumferential amplitude control function C(θ) is:

[0021]

[0022] Among them, θ0 and θ1 are the circumferential starting position and ending position of the non-axisymmetric end wall shaping area at an axial position respectively; the coefficient is the phase, which is used to control the shape of the disturbance amplitude curve in the circumferential direction.

[0023] The value of β is 0.04~0.08, the value of n is 1.5~4, The value ranges from 0.25π to 0.75π.

[0024] The ratio k of the average radius difference between the inlet and outlet of the transition section and the axial distance is defined as:

[0025]

[0026] Among them, r in is the average radius of the transition section inlet, r ex is the average radius of the transition section outlet, L is the axial length of the transition section, and the value of k is 0.4~0.5.

[0027] The half-fall ratio e of the transition section is defined as:

[0028]

[0029] Among them, r mi It is the average radius at the 1 / 2 axial position of the transition section, and the value of e is 0.55~0.75.

[0030] The area along the transition section adopts a distribution pattern of first increasing and then decreasing.

[0031] Beneficial effects: The layout structure designed by the present invention combines the fan stator and the transition section into one. Compared with the conventional layout structure, the present invention can effectively shorten the axial length of the engine compression system and simplify the engine structure. The fusion design makes the flow field of the stator blade channel complex and the flow separation intensifies. The present invention further limits the half-length drop ratio and area distribution law of the transition section, and adopts a non-axisymmetric end wall design for the inner end wall to reduce the accumulation of low-energy fluid on the suction surface of the stator blade and the inner end wall corner area, thereby reducing flow losses. According to simulation calculations, compared with the conventional layout structure, the present invention can keep the aerodynamic efficiency of the compression system at the same level while reducing the length of the engine shaft system by 10%, thereby improving the engine thrust-to-weight ratio and greatly improving the rotor dynamics performance. In addition, compared with the fusion layout structure in which the inner end wall of the transition section does not adopt a non-axisymmetric design, the present invention can increase the compressor efficiency by 0.5% and the stability margin by 5%; at the same time, the airflow at the transition section outlet is more uniform, which is conducive to improving the inter-stage matching. BRIEF DESCRIPTION OF THE DRAWINGS

[0032] The present invention will be further described in detail below in conjunction with the accompanying drawings and specific embodiments, and the above and / or other advantages of the present invention will become more clear.

[0033] Figure 1 A meridian view showing the location of the present invention in a turbofan engine compression system;

[0034] Figure 2 is the parameter r in the transition segment meridian view in 、r mi 、r ex And a schematic diagram of the definition of L;

[0035] Figure 3 Define the schematic diagram for the first turning section and the second turning section of the transition section meridian view;

[0036] Figure 4 It is a schematic diagram of the non-axisymmetric end wall curved surface modeling of the inner end wall of the transition section;

[0037] Figure 5It is the schematic diagram of the definition of z0 and z1 in the non-axisymmetric end wall modeling function A(z) for the transition section B2B;

[0038] Figure 6 It is a partial enlarged view of a cross section at a certain axial position of the transition section, and a schematic diagram of the definition of θ0 and θ1 in the non-axisymmetric inner end wall shaping function C(θ);

[0039] Figure 7 The changing curve of function A(z) under different values ​​of n;

[0040] Figure 8 For different The changing curve of function C(θ) under different values. DETAILED DESCRIPTION

[0041] The present invention proposes a layout structure of a fan stator and a transition section fusion design, wherein the layout structure comprises an inner end wall, an outer end wall and a fan stator blade, wherein the inner end wall, the outer end wall and the fan stator blade constitute a transition section;

[0042] The inner end wall and the outer end wall form an annular channel, the annular channel inlet is connected to the inner channel outlet of the fan rotor, and the annular channel outlet is connected to the inner channel compressor inlet. Figure 1 As shown (1 is the fan rotor, 2 is the splitter ring, 3 is the inner duct transition section, 4 is the outer duct stator, and 5 is the inner duct compressor). The fan stator blades are evenly distributed circumferentially in the annular channel, and each blade tip is connected to the outer end wall, and the blade root is connected to the inner end wall.

[0043] In order to highlight the compact structure of the present invention, the transition section should complete the change of the radius drop within a relatively short axial length. The ratio k of the average radius difference between the inlet and outlet of the transition section and the axial distance is defined as follows:

[0044]

[0045] Among them, r in is the average radius of the transition section inlet, r ex is the average radius of the transition section outlet, and L is the axial length of the transition section. Figure 2 As shown (31 is the outer end wall, 32 is the inner end wall, and 33 is the fan stator blade), the value of k is 0.4 to 0.5.

[0046] In the transition section, the larger average radius difference (r in -r ex ) and the axial distance (L) make the end wall profile of the transition section have a larger curvature. The transition section is divided into the first turning section (concave surface on the outer wall and convex surface on the inner wall) and the second turning section (convex surface on the outer wall and concave surface on the inner wall) according to the curvature distribution. Figure 3At the inner end wall, the wall surface changes from a convex wall surface to a concave wall surface, and the fluid near the inner end wall has a large flow-reverse pressure gradient, and the boundary layer thickens rapidly; at the same time, at the second turn, there is a radial pressure gradient from the inner end wall to the outer end wall. The fluid near the inner end wall tends to flow from the inner end wall to the outer end wall under the influence of the radial pressure gradient, resulting in a further increase in the thickness of the boundary layer of the inner end wall.

[0047] In terms of the flow in the fan stator blade channel, near the trailing edge of the stator blade (the second turning section), the channel area suddenly increases due to the sudden decrease in blade thickness, and the adverse pressure gradient further increases; in addition, there is a lateral pressure gradient between the blade pressure surface and the adjacent blade suction surface. Under the action of this pressure gradient, a lateral secondary flow will be formed in the boundary layer of the inner end wall, pointing from the pressure surface to the adjacent blade suction surface.

[0048] The superposition of the above two flow phenomena causes low-energy flow to accumulate in the corner area formed by the suction surface of the blade and the inner end wall, which is prone to large corner separation. Driven by the inertia of the mainstream, the low-energy flow will roll up in the corner area and form a channel vortex across the entire blade channel, seriously affecting the pressure expansion capacity of the stator blade and causing large flow losses.

[0049] To solve the above problems, in the design of the wall profile of the transition section, the present invention adopts a design in which the average radius drop in the first half of the transition section is larger and the average radius drop in the second half is smaller. Specifically, the half-length drop ratio e is defined as follows:

[0050]

[0051] where r mi is the average radius at the 50% axial position of the transition section, such as Figure 2 As shown, the value of e is 0.55~0.75.

[0052] Since the boundary layer in the first half of the transition section has a large momentum, it can resist a higher adverse pressure gradient and is less likely to separate; as the flow develops, the boundary layer in the second half thickens, and there is a radial pressure gradient from the inner end wall to the outer end wall, and the airflow on the inner end wall is very likely to separate. Therefore, appropriately increasing the half-length height ratio to increase the curvature of the first half and reduce the curvature of the second half can reduce the adverse pressure gradient of the fluid in the second half, thereby inhibiting the flow separation in the second half and reducing the flow loss.

[0053] Similarly, the area along the transition section increases first and then decreases, which also plays a role in reducing the adverse pressure gradient in the second half, thereby further reducing the loss in the transition section.

[0054] In the design of the fan stator blades, the present invention adopts a non-axisymmetric end wall design on the inner end wall of the blade channel composed of adjacent fan stator blades, and the non-axisymmetric end wall shape of the inner end wall of each blade channel is the same, such as Figure 4 shown.

[0055] To facilitate modeling, non-axisymmetric end wall surfaces need to adopt a parametric design method, that is, using modeling functions to generate surfaces. The patent "A non-axisymmetric end wall modeling method for compressor / turbine annular blades" (CN201110459987.X) proposes a non-axisymmetric end wall modeling method for compressor / turbine annular blades, which uses axial and circumferential control curve functions in the blade channel to construct a non-axisymmetric end wall surface. The axial control curve function in this modeling method can only adjust the maximum amplitude of the curve, but cannot adjust the axial relative position of the maximum amplitude (the function has determined that the axial relative position of the maximum disturbance is 50% of the axial position), with few degrees of freedom and certain limitations.

[0056] In view of the above situation, the present invention proposes a curved surface modeling method of a non-axisymmetric end wall with a higher degree of freedom, which is as follows:

[0057] In the cylindrical coordinate system, the radial coordinate r, axial coordinate z, and circumferential coordinate θ are defined. The surface modeling of the non-axisymmetric end wall is obtained by superimposing a radial disturbance amplitude on the original axisymmetric end wall surface. The radial disturbance amplitude control function is:

[0058] Δr(z,θ)=A(z)C(θ)

[0059] In the above formula, A(z) and C(θ) are two independent amplitude control functions, among which A(z) is the axial amplitude control function, which is a function of the axial position z of each point on the end wall and is defined as follows:

[0060]

[0061] In the above formula, h is the blade height at the leading edge of the stator blade; β is the end wall amplitude control coefficient, which indicates the ratio of the maximum disturbance amplitude of the end wall to the blade height; z0 and z1 are the axial starting position (the axial position of the blade leading edge point) and the ending position (the axial position of the blade trailing edge point) of the non-axisymmetric end wall shaping area, respectively, as shown in Figure 5 As shown (331 is the blade pressure surface, 332 is the adjacent blade suction surface); the coefficient n is the coefficient for controlling the relative axial position of the maximum disturbance amplitude point, and the value range is n>0.

[0062] C(θ) is the circumferential amplitude control function, which is a function of the circumferential position θ of each point on the end wall and is defined as follows:

[0063]

[0064] In the above formula, θ0 and θ1 are the circumferential starting position (blade pressure surface) and the ending position (blade suction surface) at a certain axial position in the modeling area, respectively, as Figure 6 shown; coefficient is the phase, which is used to control the shape of the circumferential disturbance amplitude curve. Its value range is

[0065] This end wall modeling function has rich degrees of freedom. Changing β can control the amplitude of the maximum disturbance, changing n can control the axial relative position of the maximum disturbance, and changing The circumferential relative position of the maximum disturbance can be controlled.

[0066] Specifically, when β = 0.08, h = 50, and n is 1.5, 2, and 3 respectively, the axial amplitude A(z) decreases with The change curve of Figure 7 As shown. When the phase When the circumferential disturbance amplitude C(θ) is 0.25π, 0.375π, 0.5π, 0.625π, and 0.75π respectively, The change curve of Figure 8 shown.

[0067] According to the optimization design results, the variable value range of the surface modeling function of the non-axisymmetric end wall of the present invention is: β = 0.04 ~ 0.08, n = 1.5 ~ 3, At this time, the radial disturbance amplitude of the curved surface of the non-axisymmetric end wall in the axial direction increases first and then decreases from the leading edge to the trailing edge of the stationary blade, and the maximum amplitude is located in the middle and rear section of the stationary blade; in the circumferential direction, the radial disturbance amplitude shows a trend of concave on the inner end wall close to the suction side of the blade and convex on the inner end wall close to the pressure side of the blade, as shown in Figure 4 shown.

[0068] Under the curved surface shape of the non-axisymmetric end wall, the inner end wall on the suction side of the blade is concave, which slows down the fluid and increases the static pressure; at the same time, the inner end wall on the pressure side of the blade is convex, which accelerates the fluid and reduces the static pressure. Therefore, the non-axisymmetric end wall reduces the lateral pressure gradient near the inner end wall of the cascade channel, reduces the accumulation of low-energy fluid in the boundary layer of the inner wall surface at the corner area of ​​the blade suction surface and the inner end wall, thereby weakening the airflow separation on the inner end wall and reducing flow losses.

[0069] The present invention provides a fan stator and transition section fusion design layout structure. There are many methods and ways to implement the technical solution. The above is only a preferred embodiment of the present invention. It should be pointed out that for ordinary technicians in this technical field, several improvements and modifications can be made without departing from the principle of the present invention. These improvements and modifications should also be regarded as the protection scope of the present invention. All components not specified in this embodiment can be implemented by existing technologies.

Claims

1. A fan stator and transition section integrated design layout structure, characterized in that: The layout structure includes an inner end wall, an outer end wall and fan stator blades, and the inner end wall, the outer end wall and the fan stator blades form a transition section; The inner end wall and the outer end wall form an annular channel, the annular channel inlet is connected to the inner channel outlet of the fan rotor, and the annular channel outlet is connected to the inner channel compressor inlet; The fan stator blades are evenly distributed in the annular channel in the circumferential direction, and the tip of each blade is connected to the outer end wall, and the root of each blade is connected to the inner end wall; The inner end wall of the blade channel formed by adjacent fan stator blades adopts a non-axisymmetric end wall design, and the non-axisymmetric end wall shape of the inner end wall of each blade channel is the same; The non-axisymmetric end wall surface of the inner end wall of each fan stator blade channel is obtained by superimposing a radial disturbance amplitude on the original axisymmetric end wall surface. In the cylindrical coordinate system, the radial coordinate r, axial coordinate z, and circumferential coordinate θ are defined. The radial disturbance amplitude Δr at each position will be represented by the following radial disturbance amplitude control function: Δr(z,θ)=A(z)C(θ) Among them, A(z) and C(θ) are two independent amplitude control functions, where A(z) is the axial amplitude control function; C(θ) is the circumferential amplitude control function; The calculation formula of the axial amplitude control function A(z) is: Among them, h is the height of the stator blade; β is the end wall amplitude control coefficient; z0 and z1 are the axial positions of the leading edge point of the blade and the trailing edge point of the blade respectively; n is the coefficient for controlling the relative axial position of the maximum disturbance amplitude point, β is 0.04~0.08, and n is 1.5~4; The calculation formula of the circumferential amplitude control function C(θ) is: Among them, θ0 and θ1 are the circumferential starting position and ending position at an axial position of the shaping area, respectively. The circumferential starting position is the blade pressure surface, and the ending position is the adjacent blade suction surface; the coefficient is the phase, which is used to control the shape of the disturbance amplitude curve in the circumferential direction. The value ranges from 0.25π to 0.75π.

2. A fan stator and transition section integrated design layout structure according to claim 1, characterized in that: The ratio k of the average radius difference between the inlet and outlet of the transition section and the axial distance is defined as: Among them, r in is the average radius of the transition section inlet, r ex is the average radius of the transition section outlet, L is the axial length of the transition section, and the value of k is 0.4~0.

5.

3. A fan stator and transition section integrated design layout structure according to claim 2, characterized in that: The half-fall ratio e of the transition section is defined as: Among them, r mi It is the average radius at the 1 / 2 axial position of the transition section, and the value of e is 0.55~0.75.

Images