JPH0712094A - Blade shape for compressor blade cascade - Google Patents

Abstract

PURPOSE:To restrain effectively separation of a boundary layer on the back surface side by forming the blade center line of mutually continued three circular arcs having an optional radius of curvature, and setting the maximum blade thickness position in an optional position. CONSTITUTION:A blade body 10 has a back surface 2 continuing to the rear end 5 from the front end 4 on one side surface, and has a belly surface 3 continuing to the rear end 5 from the front end 4 on the other side surface. The blade center line 7 of the blade body 10 is formed of mutually continued three circular arcs 11, 12 and 13 having respectively different radiuses of curvature. The maximum blade thickness position Xmax of the blade body 10 is set in an optional position. The flow speed distribution on the back surface side B of air 9 flowing to the rear end side T from the front end side L can be changed more optionally than a conventional blade body, and a speed reduction rate of the air 9 on the back surface side B can be adjusted. In this way, separation of a boundary layer on the back surface side B can be restrained effectively, so that compressor efficiency is not deteriorated.

Description

Detailed Description of the Invention

[0001]

BACKGROUND OF THE INVENTION 1. Field of the Invention The present invention relates to an airfoil of a compressor cascade.

[0002]

2. Description of the Related Art Conventionally, a wide range of systematic experimental studies have been conducted by a blade cascade on an airfoil applied to a compressor blade.
The ACA65 series blade type or the double arc blade type is applied.

FIG. 4 shows an example of a wing shape. 1 is a wing body, 2 is a back surface which forms one side surface of the wing body 1, 3 is an abdominal surface which forms the other side surface of the wing body 1, and 4 is The leading edge 5 of the wing body 1 represents the trailing edge of the wing body 1.

Reference numeral 6 represents a chord line, which is a straight line connecting the leading edge 4 and the trailing edge 5.

The length of the chord line 6 is the chord length c of the wing body 1.

Reference numeral 7 denotes a blade center line, and the blade center line 7 is a curve connecting the centers of circles 8 inscribed to the back surface 2 and the abdominal surface 3.

The length from the back intersection to the abdominal intersection of the perpendicular line orthogonal to the tangent to the blade center line 7 is the blade thickness t, and when the chord line 6 is the x axis, the x axis direction. Of the blade thickness t of the blade body 1 becomes the blade thickness distribution of the blade body 1.

Further, Tmax represents the maximum blade thickness.

In such a wing body 1, since the back surface 2 has a large curvature with respect to the ventral surface 3, the velocity of the air 9 flowing from the leading edge side L to the trailing edge side T of the wing body 1 is such that the back surface side B is the ventral surface side F. The pressure on the back side B becomes lower than that on the abdominal side F.

On the other hand, when the rate of deceleration of the air 9 flowing from the front edge side L to the rear edge side T on the back side B increases, the boundary layer is separated and the compressor efficiency is likely to be lowered, so that the blade body 1 It is necessary to reduce the rate of deceleration of the air 9 flowing on the back surface side B.

[0011]

However, since the NACA65 series blade type or the double arc blade type applied to the compressor blade row has a small number of parameters that specify the blade thickness distribution, the blade thickness distribution is limited. Can be set, but an arbitrary blade thickness distribution cannot be set.

Therefore, in the compressor blade row to which the NACA65 series blade type or the double arc blade type is applied, the deceleration rate of the air 9 on the back side B of the blade body 1 cannot be set arbitrarily, and the boundary layer It is difficult to effectively suppress the peeling of.

An object of the present invention is to provide an airfoil that effectively suppresses separation of the boundary layer on the back side in a compressor blade row.

[0014]

In order to achieve the above object, in the present invention, a wing having a back surface extending from the leading edge to the trailing edge on one side surface and an abdominal surface extending from the leading edge to the trailing edge on the other side surface. In the airfoil of a compressor blade cascade including a body, a blade center line is formed by three arcs that are continuous with each other and have an arbitrary radius of curvature, and a maximum blade thickness position is set at an arbitrary position.

[0015]

The maximum blade thickness position of the blade body is set to an arbitrary position, and the blade center line of an arbitrary shape is formed by three arcs which are continuous with each other and have an arbitrary radius of curvature. The rate of deceleration is reduced and the separation of the boundary layer is suppressed.

[0016]

Embodiments of the present invention will be described below with reference to the drawings.

FIGS. 1 to 3 show an embodiment of the airfoil of a compressor blade row according to the present invention, and the portions denoted by the same reference numerals as those in FIG. 4 represent the same components.

Reference numeral 10 represents a blade body, and blade center lines 7 of the blade body 10 are continuous with each other and have different radii of curvature r ₁ ,
Three arcs 1 with r ₂ and r ₃ and central angles φ ₁ , φ ₂ , and φ ₃
It is formed by 1, 12, and 13 (see FIG. 2).

Further, the maximum blade thickness position Xmax of the blade body 10 is set to an arbitrary position.

The blade thickness distribution of the blade body 10 is from the leading edge 4 to the maximum blade.
The blade thickness t toward the trailing edge side T is up to the thickness position Xmax
From the increasing part that becomes larger and the maximum blade thickness position Xmax, the maximum blade
Boundary of the decreasing part arbitrarily set on the trailing edge side T from the thickness position Xmax
Position X_LTo the trailing edge side T up to
The first reduced portion where the blade thickness t decreases and the reduced portion boundary position X
_LFrom the trailing edge 5 to the trailing edge 5
So that it can be divided into a second decreasing part that decreases in a substantially linear fashion
Configured (see Figure 3).

The inflection point position Xh of the quartic curve of the first decreasing portion
It is optionally set between the reduced portion boundary position X _L and the maximum blade thickness position Xmax.

In FIG. 3, R _LE represents the leading edge radius of the blade body 10, and R _TE represents the trailing edge radius of the blade body 10.

In this embodiment having the above construction, the maximum blade thickness position Xmax is set to an arbitrary position, and the blade center lines 7 are connected to each other and are formed by three arbitrary arcs 11, 12 and 13. The flow velocity distribution of the air 9 flowing from the leading edge side L to the trailing edge side T on the back side B is the same as that of the conventional blade body 1.
(See FIG. 4), and the rate of deceleration of the air 9 on the back side B can be adjusted. Therefore, the separation of the boundary layer on the back side B can be effectively suppressed, and the compressor efficiency does not decrease.

The airfoil of the compressor blade cascade of the present invention is not limited to the above-described embodiment, but the maximum blade thickness position Xmax and the linear shape of the blade center line may be changed as appropriate. Needless to say, various changes can be made without departing from the scope of the invention.

[0025]

As described above, according to the airfoil of the compressor blade cascade of the present invention, various excellent effects as described below can be obtained.

(1) Since the maximum blade thickness position of the blade body is set to an arbitrary position and the blade center line is formed by three arcs having an arbitrary radius of curvature, air deceleration on the back surface side of the blade body is reduced. The proportion can be adjusted.

(2) Since the rate of air deceleration on the back side of the blade body can be adjusted, the separation of the boundary layer on the back side is effectively suppressed, and the compressor efficiency does not decrease.

[Brief description of drawings]

FIG. 1 is a wing body shape diagram showing an example of a wing shape of a compressor blade row according to the present invention.

FIG. 2 is a blade centerline diagram of an example of the airfoil of the compressor blade cascade of the present invention.

FIG. 3 is a blade thickness distribution diagram in an example of the airfoil of the compressor blade cascade of the present invention.

FIG. 4 is a wing body shape diagram showing an example of a wing shape.

[Explanation of symbols]

2 Rear surface 3 Ventral surface 4 Leading edge 5 Trailing edge 7 Blade center line 10 Blade body 11, 12, 13 Arc L Leading edge side T Trailing edge side Xmax Maximum blade thickness position r ₁ , r ₂ , r ₃ Radius of curvature

Claims (1)

[Claims] 1. An airfoil of a compressor cascade comprising: a blade body having a back surface extending from a leading edge to a trailing edge on one side surface and an abdomen surface extending from a leading edge to a trailing edge on the other side surface, wherein An airfoil for a compressor cascade, wherein a blade centerline is formed by three arcs having an arbitrary radius of curvature, and a maximum blade thickness position is set at an arbitrary position.