JPH11343998A - Axial flow compressor - Google Patents

Abstract

PROBLEM TO BE SOLVED: To reduce loss of impulse waver at a tip part by constituting a camber line to be a center line of a blade surface by two or more of circular arcs and providing a circular arc connection location of a frontmost edge side on the front edge side of the half of a blade cord length. SOLUTION: When a camber line 12 to be a center line of a blade is constituted of two circular arcs, a circular arc connection location 13 is made a front edge 16 side of the half of a blade cord length 11 and a center angle of the circular arc located on a frontmost edge side is constituted to be larger than a center angle of the circular arc located on a rear edge side, a radius of curvature of a back side blade surface 14 up to the camber connection location 13 is small and a radius of curvature of a back side blade surface 14 from the camber connection location 13 up to a rear edge 17 is large. Acceleration of flow in the blade surface area of the small radius of curvature is large, and conversely, when the radius of curvature is large, acceleration is small. Therefore, when the camber connection location 13 is made a front edge side of the half location of the blade cord length 11, an acceleration area becomes short and the number of mach can be reduced.

Description

DETAILED DESCRIPTION OF THE INVENTION

[0001]

BACKGROUND OF THE INVENTION 1. Field of the Invention The present invention relates to a gas turbine or industrial axial flow compressor, and more particularly to an axial flow compressor having a row of moving blades operating at a transonic speed.

[0002]

2. Description of the Related Art A multi-stage axial compressor includes a rotating rotor 44 having a plurality of moving blade rows 42 attached thereto and a plurality of stationary blade rows 41 as shown in the schematic diagram of the axial flow compressor shown in FIG. An annular flow path is formed by the casing 43 attached, and the rotor 44, the casing 43, and each of the blade rows 41 and 42. While the inflow airflow 1 passes through this annular flow path,
It is compressed by each cascade to become a high-temperature and high-pressure outflow airflow 2. In the design of the axial compressor, the flow path shape and the cascade are designed so that a predetermined flow rate and pressure ratio are achieved with high efficiency.

In the cascade design of the axial flow compressor, a sufficient blade root thickness is provided in order to cope with the stress acting on the blade root portion due to the centrifugal force due to rotation and the fluid force received from the airflow. At the tip of the low-pressure stage rotor blade of the axial compressor, the peripheral speed is high, so that the inflow Mach number to the cascade is high, and a region where the inflow velocity exceeds the sonic speed occurs. In the section of the tip of such a cascade, see the document "Turbo Blower and Compressor"
(Takefumi Ikui, Masahiro Inoue, 1988, Corona Co., Ltd.), the shock wave by adopting a double arc wing or a multiple arc wing with the maximum thickness position on the trailing edge side and the wing leading edge side thinned The loss has been reduced.

[0004]

An increase in the flow rate and an increase in the peripheral speed of the axial flow compressor lead to an increase in the blade length, which increases the fluid force received by the blade from the airflow and the stress caused by the centrifugal force in the rotor blade. . Therefore, if steel is used as the material and a sufficient base thickness is secured to secure the reliability in strength, it may be difficult to secure the planned flow rate. In other words, when the blade root is made thicker, the throat width of the section near the root becomes narrower,
From the limit value of (throat width / inlet passage width) for the inflow Mach number described in the document "Turbo Blower and Compressor" (Takefumi Ikui, Masahiro Inoue, 1988, Corona Co.) (Throat width / inlet flow path width) becomes small, resulting in a choke flow state, and a predetermined flow rate cannot be secured. If high strength and lightweight materials such as titanium alloy with large specific strength are used, even if the root of the blade is thinned, the reliability of strength can be secured and the flow rate can be easily secured, but the cost increases and it is not economical . In addition, even if the material is steel and the root remains thick, if the tip shape is a double arc blade with a large choke margin, it is easy to secure the flow rate, but the shock wave loss in the supersonic inflow region in the tip region increases and the shaft loss increases. This leads to reduced efficiency of the flow compressor.

Accordingly, an object of the present invention is to provide a thick blade which ensures reliability in strength at low cost, avoids choke flow, secures a predetermined flow rate, and reduces shock wave loss at the tip. It is an object of the present invention to provide an axial compressor having a transonic cascade.

[0006]

According to the present invention, there is provided a supersonic inflow section of at least one moving blade row of a transonic moving blade row, wherein two or more warpage lines which are center lines of the blade surface are formed. The arc connection position on the forefront edge side is provided on the front edge side from half of the chord length, and the center angle of the arc located on the front edge side is the center angle of the arc other than the arc located on the front edge side. By setting the arc connection position on the leading edge side toward the leading edge side toward the blade tip, the predetermined flow rate is ensured, and the shock wave loss at the tip part is reduced.

In the supersonic inflow cross section of at least one of the transonic rotor cascades, a warp line, which is a center line of the blade surface, is formed such that the curvature radius of the warp line is equal to the chord length from the leading edge. Start increasing from less than half of the chord length at the position less than half the chord length from the leading edge, or decrease or remain constant from the trailing edge to the chord length. By setting the position where the radius of curvature starts to increase toward the leading edge toward the tip of the blade, a predetermined flow rate is ensured, and the shock wave loss at the tip is reduced.

[0008]

DESCRIPTION OF THE PREFERRED EMBODIMENTS An embodiment of the present invention will be described with reference to FIGS. FIG. 4 shows an axial compressor to which the present invention is applied. The axial flow compressor shown in FIG. 4 has a casing 43 to which a stationary blade row 41 is attached, and a rotating rotor 44 to which a moving blade row 42 is attached, and the inflow airflow 1 has a plurality of stationary blade rows 41 and a plurality of rotating blades. After passing through the row 42, a high-temperature, high-pressure outflow air stream 2 is formed. The supersonic inflow region 45 is located on the tip side of the transonic rotor cascade as shown in FIG. 4, and the inflow Mach number is generally higher at the tip of the blade.

[0009] The supersonic inflow section of the rotor row (for example, section A-
A) has a shape as shown in FIG. 5 and defines a warp line 12 which is the center line of the wing surface. FIG. 5 shows an example in which the warp line 12 is composed of two arc warp lines connected at the warp connection position 13 and each having a center angle φ1, φ2 of a circular warp line and a radius of curvature R1, R2.

In FIG. 1A, a warp line 12 which is a center line of a blade is formed by two arcs, and an arc connection position (a frontmost arc connection position) 13 is shifted from a half of a chord length 11 to a front edge 16. And the center angle φ of the arc located at the leading edge (most leading edge)
This is an example in which 1 is larger than the central angle φ2 (sum of central angles) of an arc located on the trailing edge side (an arc smaller than the arc on the leading edge side). When such a warp line is formed, the radius of curvature of the back side wing surface 14 up to the warp connection position (leading edge side warp connection position) 13 is small, and from the warp connection position (leading edge side warp connection position) 13 to the trailing edge 17. The radius of curvature of the back side wing surface 14 is large. Here, the radius of curvature of the wing surface does not take into account sudden changes in curvature near the leading edge and the trailing edge, which are common in wings.

[0011] The acceleration of the flow in the wing surface region having a small radius of curvature is large, while the acceleration in the wing surface region having a large radius of curvature is small. Therefore, if the warp connection position (the connection position of the warp line on the forefront edge side) 13 is on the front edge side from a half position of the chord length 11, the acceleration region becomes shorter, and the Mach number can be reduced. In other words, the acceleration near the leading edge increases, but the region is short, so that the Mach number at the throat position 18 where the interblade shock wave is generated can be kept low, and the shock wave loss can be reduced.

Further, the warp connection position (arc connection position on the forefront edge side) 13 is set to the front edge side from half of the chord length 11, and the center angle φ1 of the front edge side (the arc located on the forefront edge side) is set to the rear edge side. Warp line 12 which is the center line of the wings from two arcs (two or more arcs) so as to be larger than or equal to the central angle φ2 (the sum of the central angles of the arcs other than the arc located on the leading edge side).
Therefore, the degree of convexity of the front-side abdominal wing surface 15 is reduced or the surface of the rear-side wing surface 14 of the rear-edge side is reduced in curvature, and the degree of convexity is reduced. Therefore, the width of the throat 18 formed by the points on the trailing edge-side dorsal wing surface and the points on the leading edge-side ventral wing surface can be increased, and the value of (throat width / inlet flow path width) can be increased.
Therefore, it is possible to provide a large margin with respect to the choke limit near the tip, thereby canceling the choke state near the root and securing the planned flow rate.

Further, as shown in FIG. 1 (b), the warp connection position toward the blade tip (connection position of the warp line on the forefront edge side)
Decreasing the ratio of the to the chord length would shorten the acceleration range to offset the increase in the inflow Mach number,
The maximum Mach number can be suppressed at the entire wing height, and the shock wave loss can be reduced. In FIG. 1 (b), the ratio of the warp connection position to the chord length is set to 0.5 in the subsonic inflow region (near the blade root), but may be of any blade cross section without being directly related to the present invention.

FIG. 2A shows the relationship between the pressure ratio and the flow rate in comparison with the conventional double-arc blade in order to show the effect of this embodiment. According to the present invention, it can be seen that a predetermined flow rate can be ensured by the above-described operation. FIG.
(B) shows a comparison between the shock wave loss of the conventional example and the shock wave loss of the embodiment of the present invention. The Mach number is reduced by the above-described operation at all blade heights according to the present invention, so that the shock wave loss is reduced. You can see that it is done.

In the above description, the case where the warp line is constituted by two arcs has been described. However, the warp line, which is the center line of the blade, is constituted by two or more arcs, and the arc connection position on the leading edge side is determined by the chord. Provided on the leading edge side of half of the length, if the center angle of the arc located on the leading edge side is equal to or greater than the sum of the center angles of the arcs other than the arc located on the leading edge side, a predetermined flow rate is secured by the same action Thus, shock wave loss can be reduced.

FIG. 3 shows a second embodiment of the present invention applied to a supersonic inflow section (AA section) of an axial compressor as shown in FIG.
This is an embodiment of the invention. As shown in FIG. 3A, the radius of curvature of the warp line, which is the center line of the wing, starts increasing from a position less than half the chord length from the leading edge, and a position less than half the chord length from the leading edge. In this case, a warp line, which is the center line of the wing surface, is configured so that it increases to a maximum value and is constant until the trailing edge.

When such a warp line is formed, a position 3 where the radius of curvature of the back side wing surface is small and the curvature radius of the warp line is constant up to a position 31 where the radius of curvature of the warp line starts increasing.
The radius of curvature of the back side wing surface after 2 is large. On the dorsal wing surface, acceleration continues from the leading edge to the point where the radius of curvature reaches a maximum, but the degree of acceleration is lower downstream than that because the radius of curvature of the wing surface is large. Therefore, if the position where the curvature reaches a constant value (the position where the rate of increase of the radius of curvature decreases) is closer to the leading edge than the position where the chord length is half, the acceleration region becomes shorter, and the Mach number can be reduced.

That is, the acceleration near the leading edge increases, but the region is short, so that the Mach number at the throat position where the interblade shock wave is generated can be suppressed low, and the shock wave loss can be reduced. Further, since the radius of curvature of the warp line is small near the leading edge, the degree of convexity on the leading edge side ventral wing surface is reduced, or the leading edge side ventral wing surface becomes concave. In addition, the radius of curvature is large on the rear side wing surface on the trailing edge side, and the degree of convexity is reduced.

Therefore, the width of the throat formed by the points on the trailing edge-side dorsal wing surface and the points on the leading edge-side ventral wing surface can be increased, and the value of (throat width / inlet flow path width) can be increased. In the vicinity, it is possible to have a large margin with respect to the choke limit, thereby canceling the choke state near the root and securing the planned flow rate.

Further, as shown in FIG. 3B, the position 31 where the curvature radius of the warp line starts to increase is set to the leading edge side toward the blade tip (the ratio to the chord length is reduced).
Thus, the acceleration region is shortened so as to cancel the increase in the inflow Mach number, so that the maximum Mach number can be suppressed at all blade heights, and the shock wave loss can be reduced.
In FIG. 3 (b), in the subsonic inflow region (near the blade root), the ratio of the position where the radius of curvature of the warp starts to increase to the chord length is 0.5, which is directly related to the present invention. Instead, any wing section may be used.

In the above description, the radius of curvature is constant at a position less than half the chord length from the leading edge, but the rate of increase with respect to the chord length decreases at a position less than half the chord length from the leading edge. Even if it does, by the same action, ensure a predetermined flow rate,
Shock wave loss at the tip can be reduced.

The embodiment described above is applied to an axial flow compressor having a transonic rotor cascade in which the inflow velocity ranges from subsonic to supersonic, but the tip Mach number is 1.1 to 1.1. More preferably, the present invention is applied to the transonic bucket row No. 3.

[0023]

As described above, according to the present invention, a high-performance axial flow having a transonic cascade, which is low in cost, secures strength reliability, avoids choke flow, and reduces shock wave loss. A compressor can be provided.

[Brief description of the drawings]

FIGS. 1 (a) and 1 (b) are a cross-sectional view of a transonic rotor cascade of an axial compressor according to a first embodiment of the present invention, and a characteristic diagram showing a relationship between an inflow Mach number and a row connection position. .

FIGS. 2A and 2B are characteristic diagrams showing the effect of the present invention.

FIGS. 3 (a) and 3 (b) are characteristic diagrams showing a relationship between the axial flow compressor and a transonic rotor cascade according to the present invention.

FIG. 4 is a schematic view of an axial compressor to which the present invention is applied.

FIG. 5 is a diagram illustrating a warp connection position of a bucket row according to the present invention.

[Explanation of symbols]

DESCRIPTION OF SYMBOLS 1 ... Inflow airflow, 2 ... Outflow airflow, 3 ... Rotation, 11 ... Chord length, 12 ... Warp line, 13 ... Warp connection position, 14 ... Dorsal wing surface, 15 ... Ventral wing surface, 16 ... Leading edge, 17 ... Trailing edge, 18 ...
Throat, 31 ... Position where the radius of curvature of the warp line starts to increase, 32 ... Position where the radius of curvature of the warp line is maximum and constant or the rate of increase decreases, 41 ... Stator blade row, 42 ... Rotating blade row, 43 ... Casing, 44: rotor, 45: supersonic inflow area, φ1, φ2: arc center angle.

Claims (2)

[Claims] An axial flow compressor having a transonic moving blade cascade having an inflow velocity ranging from subsonic to supersonic speeds, wherein a supersonic inflow section of at least one moving blade cascade of the transonic moving cascade has a blade surface. The center line of the arc is located at two or more arcs, and the arc connection position on the forefront edge side is provided closer to the front edge side than half of the chord length, and the center angle of the arc located on the forefront edge side is the front edge side. An axial flow compressor characterized in that the center angle of the arc other than the arc located at the center is equal to or greater than the sum of the center angles of the arcs, and the arc connection position on the frontmost edge side is on the front edge side toward the blade tip. 2. An axial flow compressor having a transonic rotor cascade with an inflow velocity ranging from subsonic to supersonic, wherein at least one of the transonic rotor cascades has a supersonic inflow cross section, and a blade surface The radius of curvature of the warp line starts increasing from a position less than half the chord length from the leading edge, and the curvature line relative to the chord length at a position less than half the chord length from the leading edge. An axial flow compressor, wherein an increasing rate is reduced or constant up to a trailing edge, and a position where a curvature radius of a warp line starts increasing is set to a leading edge side toward a blade tip.