KR100713252B1 - Rotor blade for axial-flow turbine - Google Patents

Abstract

The present invention is designed for the axial flow turbine that can improve the efficiency and thrust of the turbine by designing the shape of the blade leading edge in order to reduce the influence of the gradient shock wave generated in the rotor blade leading edge and the flow separation in the flow path It is an object to provide a rotor blade.

In order to achieve the above object, the rotor blade for an axial turbine according to the present invention, the inner peripheral surface for generating a static pressure between the leading edge (trailing edge); And an outer circumferential surface generating a negative pressure between the leading and trailing edges, wherein the inner circumferential surface and the outer circumferential surface at the leading edge are chamfered to be connected to each other in a plane, and the chamfered plane The angle between the extended surface and the outer circumferential surface is smaller than the angle between the outer circumferential surface and the vertical surface including the leading edge.

Turbine, Rotor Blade, Leading Edge, Chamfering, Cutting Angle

Description

Rotor blade for axial turbine {ROTOR BLADE FOR AXIAL-FLOW TURBINE}

1A shows a propellant supply system, and FIG. 1B shows a rotor of a supersonic turbine of the propellant supply system of FIG. 1A.

2 shows a cross-sectional shape of an axial turbine;

Figures 3a and 3b is a view of the rotor blade shape of the existing supersonic turbine.

Figure 4 is a numerical value of the blade formed so that the angle between the inner peripheral surface and the outer peripheral surface at the leading edge in the rotor blade of the conventional supersonic turbine coincides with the angle between the chamfered inner peripheral surface and the vertical plane including the leading edge. Graph showing the analysis results in isometric lines.

Fig. 5 is a partially enlarged view showing the chamfering cutting angle of the rotor blades according to the prior art (left view) and the present invention (right view), respectively.

6a to 6d are graphs showing the numerical results of the rotor blade chamfering angle according to the cutting angle.

Figure 7 is a graph of the thrust distribution according to the rotor blade chamfering cutting angle change according to the present invention.

8 is a graph showing the distribution of the total pressure loss coefficient according to the change of the cutting angle of the rotor blade chamfering according to the present invention.

Figure 9 is a schematic diagram for the flow visualization experiment of the turbine according to the present invention.

10A and 10B are visualized images according to changes in cutting edge chamfering angles of rotor blades, respectively.

<Brief description of the major symbols in the drawings>

1: outer circumferential surface (extrados) 2: inner circumferential surface (intrados)

3: chamfered plane

The present invention relates to a rotor blade for an axial turbine, and more particularly to a rotor blade for a supersonic impulse axial turbine for improving turbine efficiency and thrust.

In general, in a liquid propulsion rocket for a space launch vehicle, a propellant supply system for supplying a propellant to a combustion chamber becomes an essential component. These propellant supply systems usually employ high pressure turbopumps, which are also a key component of guided weapons. Therefore, development of turbopump needs to be preceded for the development of space and defense industry.

In this turbopump system, most of the turbines driving the pumps are partially suction type and impulse type axial turbines. Impulse axial turbines are widely used when weight is important because they are small in size, light in weight, and have a large output. To produce high output at such a small size, the impulse axial turbine operates in supersonic flow conditions and adopts a partial suction type, unlike conventional axial turbines used for general power generation or industrial gas turbines. Therefore, its flow characteristics lead to severe back pressure gradients and large secondary flows in both axial, radial and rotational directions, and flow separation due to back pressure gradients, resulting in passage vortex, horseshoe vortex Various vortices, such as horseshoe vortex and corner vortex, that occur in front of or inside the collar passage, mix with wakes in the latter half of the blade row to form complex three-dimensional, viscous, turbulent flows. It is very different from the characteristics. In general, axial turbines have been widely studied experimentally and numerically, and their performance characteristics are well known, and they are widely used in the prediction and design of turbines. And prediction of performance characteristics is very difficult.

FIG. 1A shows the propellant supply system, and FIG. 1B shows the rotor of the supersonic turbine of the propellant supply system of FIG. 1A. In particular, the efficiency of the supersonic turbine is greatly influenced by the gradient shock wave generated at the leading edge of the rotor blades and the separation of flow generated in the flow path.

2 shows a cross-sectional shape of an axial turbine. This axial turbine consists of a stage followed by a rotor followed by a stationary blade row called a nozzle or stator as shown in FIG. 2 or a conical nozzle as shown in FIG. 1. On the other hand, the gas at high temperature and high pressure expands and accelerates while passing through the nozzle. At this time, the speed of the circumferential direction is greatly increased, and the rotor rotates to generate power.

The turbine output per unit, which can be obtained per unit mass flow rate, which is called the specific power output in the rotor under constant axial speed, is given by Equation 1 below.

: 비출력 (specific power output)

Specific power output

U : Rotational Speed of Rotor

W _w1 : Relative speed at rotor inlet in rotor rotation direction

W _w2 : Relative speed at rotor exit in rotor rotation direction

C _a : Axial speed at the rotor

tan β ₁ : relative flow angle at rotor inlet

tan β ₂ : relative outlet angle at rotor exit

In Equation 1, it can be seen that a large specific power can be obtained by increasing the change of the blade rotation speed, the axial flow speed, and the rotor angle.

However, the increase in the rotational speed causes an increase in the tensile stress due to the centrifugal force, so there is a limit in terms of structure. In particular, since turbines operate at high temperatures, the rotational speed must be determined by accurately understanding the properties of the material with temperature. Often this speed is also limited by the performance characteristics of a rotating compressor, such as a turbine.

In addition, increasing the rotational speed causes an increase in the flow velocity, and an increase in the axial velocity also means an increase in the flow velocity. In case of supersonic flow due to increased flow velocity, shock wave is generated and flow loss is largely generated.

Therefore, in order to maintain excellent performance, the rotational speed of the rotor and the axial speed of the fluid should be limited to a range in which the loss due to the shock wave is not large even if no supersonic flow occurs or a localized supersonic region occurs in the feather channel.

In the rotor blades of supersonic turbines currently commonly used, the shape of leading edge (LE) and trailing edge (TE) is treated by arc-shaped curve processing (or 0 °). 3A and 3B are views of the rotor blade shape of a conventional supersonic turbine, and FIG. 3A shows a case in which the shapes of the leading edge and the trailing edge are treated with arc-shaped curves, and FIG. 3B shows the shapes of the leading edge and the trailing edge in the cord direction. The case of 90 degree cutting process is shown. That is, the rotor blade for an axial turbine shown in FIG. 3B includes an inner circumferential surface for generating a positive pressure between the leading edge and the trailing edge, and an outer circumferential surface for generating a negative pressure between the leading edge and the trailing edge, The inner circumferential surface and the outer circumferential surface are connected to each other in a plane, and the angle between the inner circumferential surface and the outer circumferential surface is formed to coincide with the angle between the chamfered inner circumferential surface and the vertical surface including the leading edge.

In the rotor leading edge of such a supersonic turbine, oblique shock waves or arcuate shock waves are generated at the leading edge as shown in FIG. 4 due to the influence of the accelerated supersonic flow at the nozzle outlet, and the separation at about 40% of the blade cord length is also separated. Phenomenon occurs. On the other hand, Figure 4 is formed so that the angle between the inner circumferential surface and the outer circumferential surface coincides with the angle between the chamfered inner circumferential surface and the vertical plane including the leading edge, i.e., the back of the blade cut in the cord direction 90 °. As a graph on the diagram, these shock waves or flow separations greatly influence the efficiency and specific power output of the turbine.

On the other hand, as a prior art regarding a turbine blade shape, the invention disclosed in Unexamined-Japanese-Patent No. 57-113906, Unexamined-Japanese-Patent No. 7-125904, Unexamined-Japanese-Patent No. 2002-138801, US Patent Publication US 6,666,654, etc. are mentioned.

The present invention is designed to change the shape of the blade leading edge in order to reduce the influence of the gradient shock wave generated in the rotor blade leading edge (LE) and flow separation in the flow path to improve the efficiency and thrust of the turbine. It is an object of the present invention to provide a rotor blade for an axial turbine.

In order to achieve the above object, the rotor blade for an axial turbine according to the present invention, the inner circumferential surface (intrados) for generating a positive pressure between the leading edge (trailing edge); And an outer circumferential surface generating a negative pressure between the leading edge and the trailing edge, wherein the inner circumferential surface and the outer circumferential surface at the leading edge are chamfered to be connected to each other in a plane. And an angle between the extension surface of the chamfered plane and the outer peripheral surface is smaller than the angle between the outer peripheral surface and the vertical surface including the leading edge.

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Further, the difference between the angle between the extended surface and the outer peripheral surface of the chamfered plane and the angle between the outer peripheral surface and the vertical surface including the leading edge is more preferably 1 ° to 4 °.

Hereinafter, embodiments of the present invention will be described in detail with reference to the accompanying drawings.

In the present invention, in order to reduce the influence of the gradient shock wave generated in the rotor blade leading edge and the separation in the flow path (chamfer), the shape of the blade leading edge chamfering the shape of the blade leading edge as shown in FIG. By changing the cutting angle, the efficiency and thrust of the turbine were improved.

5 is a partially enlarged view showing a chamfering cutting angle of the rotor blade according to the prior art and a right side of the present invention, wherein the rotor blade for the axial turbine has a positive pressure between the leading and trailing edges. An inner circumferential surface (2) for generating a and an outer circumferential surface (1) for generating a negative pressure between the leading and trailing edge.

Here, the inner circumferential surface 2 and the outer circumferential surface 1 at the leading edge are chamfered so as to be connected to each other in a plane 3, and an extension surface and an outer surface of the chamfered plane 3. angle (γ) between the peripheral surface (1) it can be seen that is formed to be smaller than the angle (γ _vertical) between the outer peripheral surface (1) and a virtual vertical plane, including a leading edge (vertical dotted line in Fig. 5) . That is, the leading edge of the rotor blade according to the present invention has a sharp shape sharply by the chamfering through the chamfering process.

Further, the chamfered cutting angle (chamfering angle), that is, the angle γ between the extended surface and the outer peripheral surface 1 of the chamfered plane 3 and the outer peripheral surface 1 and the leading edge It is preferable that the difference ( γ _vertical -γ) with the angle ( γ _vertical ) between the imaginary vertical surfaces to be included is substantially 1 ° to 12 °. As will be described later, as the chamfered cutting angle ( γ _vertical -γ) increases, the strength of the gradient shock wave generated at the blade leading edge is weakened and pushed backward in the flow direction, thereby acting as an important factor for increasing the thrust of the supersonic turbine. However, in terms of processing and mechanical aspects, the chamfering at angles substantially greater than 12 ° approaches the manufacturing limits of the rotor blades, so that the chamfered cutting angle is preferably 12 ° or less.

Further, the chamfered cutting angle (chamfering angle; γ _vertical -γ ), that is, the angle γ between the extension surface of the chamfered plane 3 and the outer peripheral surface 1 and the outer peripheral surface ( 1) and the angle between the vertical plane containing the leading edge (the difference between the _vertical γ) (γ _vertical -γ), it is substantially more preferred in the 1 ° to 4 °. This is because the chamfered cutting angle for the rotor blades of the supersonic turbine exhibits the largest thrust increase in the above range, and the intensity of the shock wave is no longer weakened in the above range.

On the other hand, in order to investigate the shock wave and the fluid peeling phenomenon occurring in the leading edge of the rotor blade according to the present invention was used a numerical analysis and supersonic wind tunnel experiment using Fine ^TM / Turbo of Numeca, an expert program for turbomachinery analysis.

Rotor  blade Chamfering  Numerical Analysis of Cutting Angle

First, numerical analysis using Fine ^TM / Turbo was performed by changing the rotor blade chamfering cutting angle from 0 ° to 12 °. As described above, the chamfering cutting angle of 0 ° is for the prior art, and in terms of processing and mechanics, chamfering at angles substantially greater than 12 ° may reach the manufacturing limits of the rotor blades. ° was set at the critical angle.

6a to 6d show the numerical results of the rotor blade chamfering cutting angles in the equilateral line diagram, (a) shows the chamfering cutting angle of 0 °, and (b) shows the chamfering cutting angle of 4 °, ( c) shows the result of the chamfering cutting angle of 8 ° and (d) of the chamfering cutting angle of 12 °. As shown in FIGS. 6A to 6D, as the chamfering cutting angle increases, the strength of the sloped shock wave generated at the blade leading edge becomes weaker, and thus shows a phenomenon of being pushed backward in the flow direction. This phenomenon is an important factor to increase the thrust of the supersonic turbine.

On the other hand, Figure 7 is a graph of the thrust distribution according to the change of the cutting angle of the rotor blade chamfering, Figure 8 shows a graph of the distribution of voltage loss coefficient according to the change of the cutting angle of the rotor blade chamfering.

In other words, when the chamfering cutting angle is changed from 0 ° to 4 °, the voltage loss coefficients representing the thrust value and the shock wave intensity of the turbine rotor blade according to the chamfering angle are shown in FIGS. 7 and 8. The change in thrust is the largest and when it is greater than 4 °, the thrust is somewhat constant. The resultant voltage loss factor is also higher than 4 °, indicating a small change, indicating that the intensity of the shock wave generated at the leading edge of the blade is no longer weakened. Therefore, it can be seen that the rotor blade chamfering cutting angle of the supersonic turbine shows the largest thrust increase in the range of 1 ° to 4 °.

Rotor  blade Chamfering  Supersonic speed for cutting angle wind tunnel  Experiment

Similar to the above numerical results, the experimental results using the supersonic wind tunnel test apparatus under the same conditions will be described below.

The Schlieren system was used for the supersonic turbine flow visualization according to the present invention. In addition, a single mirror Schlieren system was used to acquire shadowgraph images. A 150W tungsten continuous light was used as a light source and a high speed camera system (Kodak, SR Ultra-C) was used to acquire the visualized image. The camera is operated at the same time as the start of the flow visualization experiment, and the image obtained by the flow visualization experiment is corrected with an optical filter and stored as a digital image.

FIG. 9 is a schematic diagram of a flow visualization experiment. Specifically, the experiment is a combination of two types of blades having different leading edge chamfering angles by combining a nozzle and a blade row as shown in FIG. 10a has a cutting angle of 0 ° and FIG. 10b has a cutting angle of 8 °], and the nozzle exit flow angle is changed according to the nozzle pressure ratio. Rate: PR) was performed in 5.1.

10A and 10B are shadowgraph images obtained through experiments. The black area at the top right is the nozzle part, the bent part at the top left is the nozzle top surface, and the pointed protruding part at the right middle part represents the end of the nozzle.

As described above, FIG. 10A illustrates a visualized image when the leading chamfering cutting angle value is 0 °. The jet interface can be observed at the lower end of the nozzle outlet and the interface meets the leading edge of blade 6. And it can be observed that the arc (Bow) shock wave is generated in each blade leading edge, and the extension wave is generated in the portion bent by the corner angle of the blade leading edge, just behind each arch shock wave. The arcuate shock angles of the blades 3 to 5 except the blade 6 are similar to each other because the flow angles are similar while passing the arcuate shock wave and the extension wave. The arcuate shock wave generated at the leading edge of the cascade is reflected back to the nozzle top surface and can be seen entering the cascade flow path. In addition, it can be seen that the shock wave generated at the leading edge of the blade row is concentrated at the outlet of the upper surface of the nozzle, and a complicated flow appears.

On the other hand, it can be observed that flow separation occurs at about 40% of the axial cord of each blade channel suction surface. And it can be seen that the difference in the flow separation area is not a big difference except the flow separation generated in the flow path between the blades 2 and 3. In the cascade flow path, the shock waves are reflected and exit to the cascade wake. And the tail-tail (Fish-tail) shock wave can be seen that occurs.

10B illustrates a visualized image when the leading edge chamfering cutting angle value is 8 °. It can be seen that the arcuate shock wave generated at the blade leading edge is slightly more toward the flow path than when the leading chamfering cutting angle value is 0 ° as the blade leading edge becomes sharper. In addition, it can be seen that the flow separation in the casing flow channel occurs at about 40% of the axial chord of the suction surface with almost no change from the previous case, and the peeling region is also similar.

10A and 10B, when the chamfering cutting angle of the leading edge of the rotor blade is gradually increased and changed from 0 ° to 8 °, the strength of the inclined shock wave generated at the blade leading edge is weakened and is observed to be pushed in the flow direction. can do. Therefore, as the chamfered cutting angle increases, the strength of the gradient shock wave generated at the blade leading edge is weakened and pushes backward in the flow direction. This phenomenon acts as an important factor for increasing the thrust of the supersonic turbine.

The present invention is not limited to the above-described embodiments, and various modifications and changes can be made by those skilled in the art, which should be regarded as included in the spirit and scope of the present invention as defined in the appended claims. will be.

As described above, by forming the rotor blade chamfering cutting angle in the supersonic turbine according to the present invention, the efficiency of the turbine by reducing the influence of the gradient shock wave generated in the rotor blade leading edge (LE) and the flow separation in the flow path And thrust can be improved.

In addition, since the cutting angle of the rotor blade chamfering of the supersonic turbine according to the present invention is formed within a range of 1 ° to 4 °, the largest thrust increase effect can be created.

Claims (3)

In the rotor blade for an axial turbine, An inner circumferential surface for generating a static pressure between the leading edge and the trailing edge; And An outer circumferential surface generating a negative pressure between the leading and trailing edges; Including, The inner circumferential surface and the outer circumferential surface at the leading edge are chamfered to be connected to each other in a plane. The angle between the extension surface and the outer peripheral surface of the chamfered plane is smaller than the angle between the outer peripheral surface and the vertical surface including the leading edge, The difference between the angle between the extended surface of the chamfered plane and the outer circumferential surface and the angle between the outer circumferential surface and the vertical surface comprising the leading edge is preferably 1 ° to 4 °. Rotor blades. delete delete

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