EP1927724A2 - Turbomachine blade - Google Patents

Abstract

The method involves defining a skeleton curve by a skeleton line angle over a chord length, leading edge (2), blade height (h) and a blade spike (6). The curve is run in a blade profile section, which is present in an area running out of the blade spike up to 30 percentages of the blade height, in a sectional line angle distribution area lying between upper and lower limit curves. A pressure load is produced along a blade surface in the distribution area. Dimensionless skeleton line angles at a position of the chord length are provided for the limit curves from a specific formula.

Description

The invention relates to the airfoil design of the blades and vanes of a turbomachine, in particular a gas turbine engine, defined by the progression of the skeletal line defined by the skeletal line angle over the chord length and blade height, and the leading edge profile and the blade tip terminating at an air gap.

The blade of engine blades is composed under the aspect of a fluidically optimal shape design of a plurality of over the blade blade height threaded individual profiles into a three-dimensional blade shape, the individual profile sections are characterized by a particular skeleton line and a certain material thickness on both sides of the skeleton line. The course of the skeleton line representing a center line in the respective profile section is designed for a minimum profile pressure loss and a maximum working range in the respective blade area. These requirements meet the currently used CDA (Controlled Diffusion Airfoil) blade profiles and their derivatives in the blade tip, that is, in the near-gap blade area, but not, since the aerodynamically adverse influence of the gap between the blade tip and the machine housing or hub at The blade shapes used today is not sufficiently considered. By flow around and overflow of the blade tip occurs in this blade area to form vortices that limit the stable operation of the machine and thus flow and power losses caused by a - in terms of weight and cost disadvantageous - increase the number of blades must be compensated.

The object of the invention is to design the airfoil profiles of rotor blades and guide vanes of a turbomachine in such a way that the flow disturbances which occur due to the flow disturbances occurring close to the gap, which lead to power losses, are minimized.

According to the invention, the object is achieved with an airfoil design according to the features of patent claim 1. Advantageous developments of the invention are the subject of the dependent claims.

The gist of the invention is that in a gap-proximate region of up to 30% of the blade height starting from the blade tip, the blade profile cuts are characterized by a specific skeleton line profile defined by the skeleton line angle with respect to the chord length of the blade profile, at which in the vicinity of the gap a uniform pressure distribution along the blade section and thus a stable gap vortex is achieved. The uniform load distribution in the near-gap blade area has lower gap losses, that is, an increase in the power and the stability limit or a reduction in the number of blades and thus the weight at a constant power and ultimately the cost.

The dimensionless skeleton line angles for the inventively optimal skeleton line profile for blade profile cuts that fall within the above-mentioned 30% range are in a certain skeletal line angle distribution range that is in one of the chord length (x-axis, in percent) and the dimensionless skeleton line angle ( y-axis) formed coordinate system is arranged, wherein the upper and the lower limit curve of the skeleton line angle distribution are determined by the equations given in claim 1.

The dimensionless skeleton line angle results from the relationship given in claim 2.

If the skeleton lines or the corresponding skeleton line angles in the blade profile sections close to the gap lie within the limits defined by the limit curves, the disturbances and losses caused by the gap are greatly reduced. The formation of the skeleton lines according to the invention is not limited to certain leading edge profiles of the blades.

Fig. 1

eine Seitenansicht einer Laufschaufel mit gepfeilter Vorderkante und durch waagerechte Linien angedeuteten Profilschnittebenen;

Fig. 2

eine Darstellung eines Schaufelprofils mit Skelettlinie in einem durch die dimensionslose Sehnenlänge (x-Achse) und den dimensionslosen Skelettlinienwinkel (y-Achse) definierten Koordinatensystem;

Fig. 3

den von einer oberen und einer unteren Grenzkurve begrenzten Bereich der Skelettlinienwinkelverteilung für einen von der Schaufelspitze ausgehenden begrenzten Schaufelteil; und

Fig. 4

eine Gegenüberstellung eines erfindungsgemäß und eines nach dem Stand der Technik ausgebildeten

Schaufelprofils im spaltnahen Bereich mit der jeweiligen Belastungsverteilung.An embodiment of the invention will be explained in more detail with reference to the drawing. Show it:

Fig. 1

a side view of a blade with swept leading edge and indicated by horizontal lines profile section planes;

Fig. 2

a representation of a blade profile with skeleton line in a coordinate system defined by the dimensionless chord length (x-axis) and the dimensionless skeleton line angle (y-axis);

Fig. 3

the area of the skeleton line angle distribution bounded by upper and lower limit curves for a limited blade part emanating from the blade tip; and

Fig. 4

a comparison of an inventively and of the prior art trained

Blade profiles in the gap-near area with the respective load distribution.

Fig. 1 shows a side view of an airfoil 1 with swept course of the leading edge 2 of a blade of the compressor of a gas turbine engine. A plurality of sectional planes distributed over the blade height "h" can be seen. According to the skeleton line 4 (FIG. Fig. 2 ) with in each reference point on both sides of the same material thickness "d" is defined by threading the corresponding blade profile sections 5 in the cutting planes 3, the shape of the airfoil 1.

worin

α_i (1)

der jeweilige lokale Winkel bei einem bestimmten Wert 1_x der Sehnenlänge,

BIA

der Eintrittswinkel und

BOA

der Austrittswinkel

sind.The skeleton line 4 in Fig. 2 is defined in the form of the dimensionless skeleton line angle α (1) along the chord length "1", which is likewise dimensionless as a percentage, and results from $$\mathrm{\alpha }\left(\mathrm{l}\right)\mathrm{=}{\mathrm{\alpha }}_{\mathrm{i}}\left(\mathrm{l}\right)\mathrm{-}\mathrm{BIA}\mathrm{/}\mathrm{BOA}\mathrm{-}\mathrm{BIA}\mathrm{.}$$

wherein

α _i (1)

the respective local angle at a certain value 1 _x the chord length,

BIA

the entrance angle and

BOA

the exit angle

are.

In an area extending from the blade tip 6, which comprises approximately 30% of the blade height "h" ( Fig. 1 ) and in which the cutting planes 3 are arranged closer, the skeleton lines 4 of the respective blade profile section 5 are designed such that their dimensionless specified skeleton line angles α (1) = 0.0 to 1.0 at all points above the dimensionless chord length 1 = 0 to 100% of the respective blade profile section 5 in one predetermined limit range between an upper limit curve 7 (oG) and a lower limit curve 8 (uG) are. If the skeleton lines of the airfoil 1 in the upper gap close range of up to 30% of the blade height "h" in this limited skeletal line angle distribution range, despite the gap and in three-dimensional blade shape and independent of leading edge profile of the blade, a flow-optimal blade profile is achieved in which Pressure load on the blade is uniform and thus the gap vortex stabilized and the gap losses are minimized.

und für die untere Grenzkurve 8 aus $${\mathrm{\alpha }}_{\mathrm{uG}}\mathrm{=}\mathrm{3}\mathrm{,}\mathrm{97581923552676}\mathrm{\times }{\mathrm{10}}^{\mathrm{11}}\mathrm{\times }{{\mathrm{l}}_{\mathrm{x}}}^{\mathrm{6}}\mathrm{-}\mathrm{1}\mathrm{,}\mathrm{02257586096638}\mathrm{\times }{\mathrm{10}}^{\mathrm{-}\mathrm{8}}\mathrm{\times }{{\mathrm{l}}_{\mathrm{x}}}^{\mathrm{5}}\mathrm{+}\mathrm{9}\mathrm{,}\mathrm{81093271630595}\mathrm{\times }{\mathrm{10}}^{\mathrm{-}\mathrm{7}}\mathrm{\times }{{\mathrm{l}}_{\mathrm{x}}}^{\mathrm{4}}\mathrm{-}\mathrm{0}\mathrm{,}\mathrm{000042865320363461}\mathrm{\times }{{\mathrm{l}}_{\mathrm{x}}}^{\mathrm{3}}\mathrm{\times }\mathrm{0}\mathrm{,}\mathrm{00082697833059342}\mathrm{\times }{{\mathrm{l}}_{\mathrm{x}}}^{\mathrm{2}}\mathrm{-}\mathrm{0}\mathrm{,}\mathrm{000113440630116202}\mathrm{\times }{\mathrm{l}}_{\mathrm{x}}\mathrm{.}$$

The skeleton line angle α _oG for a plurality of values lying between 0 and 100% l _x , that is, l _x1 , l _x2 , etc., of the chord length "1" results for the upper limit curve 7 $${\mathrm{\alpha }}_{\mathrm{oG}}\mathrm{=}\mathrm{1}\mathrm{.}\mathrm{2893686702647}\mathrm{\times }{\mathrm{10}}^{\mathrm{-}\mathrm{9}}\mathrm{\times }{{\mathrm{l}}_{\mathrm{x}}}^{\mathrm{5}}\mathrm{-}\mathrm{3}\mathrm{.}\mathrm{17452341597451}\mathrm{\times }{\mathrm{10}}^{\mathrm{-}\mathrm{7}}\mathrm{\times }{{\mathrm{l}}_{\mathrm{x}}}^{\mathrm{4}}\mathrm{+}\mathrm{0}\mathrm{.}\mathrm{0000293283473623007}\mathrm{\times }{{\mathrm{l}}_{\mathrm{x}}}^{\mathrm{3}}\mathrm{-}\mathrm{0}\mathrm{.}\mathrm{00129356647808443}\mathrm{\times }{{\mathrm{l}}_{\mathrm{x}}}^{\mathrm{2}}\mathrm{+}\mathrm{0}\mathrm{.}\mathrm{0345950133223312}\mathrm{\times }{\mathrm{l}}_{\mathrm{x}}$$

and for the lower limit curve 8 off $${\mathrm{\alpha }}_{\mathrm{uG}}\mathrm{=}\mathrm{3}\mathrm{.}\mathrm{97581923552676}\mathrm{\times }{\mathrm{10}}^{\mathrm{11}}\mathrm{\times }{{\mathrm{l}}_{\mathrm{x}}}^{\mathrm{6}}\mathrm{-}\mathrm{1}\mathrm{.}\mathrm{02257586096638}\mathrm{\times }{\mathrm{10}}^{\mathrm{-}\mathrm{8th}}\mathrm{\times }{{\mathrm{l}}_{\mathrm{x}}}^{\mathrm{5}}\mathrm{+}\mathrm{9}\mathrm{.}\mathrm{81093271630595}\mathrm{\times }{\mathrm{10}}^{\mathrm{-}\mathrm{7}}\mathrm{\times }{{\mathrm{l}}_{\mathrm{x}}}^{\mathrm{4}}\mathrm{-}\mathrm{0}\mathrm{.}\mathrm{000042865320363461}\mathrm{\times }{{\mathrm{l}}_{\mathrm{x}}}^{\mathrm{3}}\mathrm{\times }\mathrm{0}\mathrm{.}\mathrm{00082697833059342}\mathrm{\times }{{\mathrm{l}}_{\mathrm{x}}}^{\mathrm{2}}\mathrm{-}\mathrm{0}\mathrm{.}\mathrm{000113440630116202}\mathrm{\times }{\mathrm{l}}_{\mathrm{x}}\mathrm{,}$$

In Fig. 4 are - each with the corresponding schematic pressure load - two blade profile cuts 5 in the near-gap region facing each other, namely of a blade according to the prior art (zigzag line hatching) and of a blade according to the invention (slash-hatching). The indicated pressure load is substantially uniform in the blade according to the invention and has the shape of a triangle in the blade according to the prior art, which leads to flow disturbances and losses.

LIST OF REFERENCE NUMBERS

1

airfoil

2

leading edge

3

cutting planes

4

skeleton line

5

Aerofoil section

6

blade tip

7

Upper limit curve

8th

Lower limit curve

H

blade height

d

material thickness

α (l)

Skeleton line angles

α _i

local skeleton line angle

l

chord length

l _x

certain value of chord length

Claims (3)

An airfoil design for the blades and vanes of a turbomachine, in particular a gas turbine engine, characterized by the progression of the skeleton line (4) defined by the skeletal line angle (α) over the chord length (1) and the leading edge profile and the blade height (h) and at an air gap characterized in that the skeleton line (4) in the blade profile cuts (5), which are in a range of the blade tip (6) outgoing range of up to 30% of the blade height (h) in one extending between upper and lower limit curves (7, 8), wherein the dimensionless skeleton line angle (α) at the respective location (lx) of the chord length (1) for the upper limit curve (7) $${\mathrm{\alpha }}_{\mathrm{oG}}\mathrm{=}\mathrm{1}\mathrm{.}\mathrm{2893686702647}\mathrm{\times }{\mathrm{10}}^{\mathrm{-}\mathrm{9}}\mathrm{\times }{{\mathrm{l}}_{\mathrm{x}}}^{\mathrm{5}}\mathrm{-}\mathrm{3}\mathrm{.}\mathrm{17452341597451}\mathrm{\times }{\mathrm{10}}^{\mathrm{-}\mathrm{7}}\mathrm{\times }{{\mathrm{l}}_{\mathrm{x}}}^{\mathrm{4}}\mathrm{+}\mathrm{0}\mathrm{.}\mathrm{0000293283473623007}\mathrm{\times }{{\mathrm{l}}_{\mathrm{x}}}^{\mathrm{3}}\mathrm{-}\mathrm{0}\mathrm{.}\mathrm{00129356647808443}\mathrm{\times }{{\mathrm{l}}_{\mathrm{x}}}^{\mathrm{2}}\mathrm{+}\mathrm{0}\mathrm{.}\mathrm{0345950133223312}\mathrm{\times }{\mathrm{l}}_{\mathrm{x}}$$

is and for the lower limit curve (8) $${\mathrm{\alpha }}_{\mathrm{uG}}\mathrm{=}\mathrm{3}\mathrm{.}\mathrm{97581923552676}\mathrm{\times }{\mathrm{10}}^{\mathrm{11}}\mathrm{\times }{{\mathrm{l}}_{\mathrm{x}}}^{\mathrm{6}}\mathrm{-}\mathrm{1}\mathrm{.}\mathrm{02257586096638}\mathrm{\times }{\mathrm{10}}^{\mathrm{-}\mathrm{8th}}\mathrm{\times }{{\mathrm{l}}_{\mathrm{x}}}^{\mathrm{5}}\mathrm{+}\mathrm{9}\mathrm{.}\mathrm{81093271630595}\mathrm{\times }{\mathrm{10}}^{\mathrm{-}\mathrm{7}}\mathrm{\times }{{\mathrm{l}}_{\mathrm{x}}}^{\mathrm{4}}\mathrm{-}\mathrm{0}\mathrm{.}\mathrm{000042865320363461}\mathrm{\times }{{\mathrm{l}}_{\mathrm{x}}}^{\mathrm{3}}\mathrm{\times }\mathrm{0}\mathrm{.}\mathrm{00082697833059342}\mathrm{\times }{{\mathrm{l}}_{\mathrm{x}}}^{\mathrm{2}}\mathrm{-}\mathrm{0}\mathrm{.}\mathrm{000113440630116202}\mathrm{\times }{\mathrm{l}}_{\mathrm{x}}\mathrm{,}$$

is. An airfoil design according to claim 1, characterized in that the dimensionless skeleton line angle (α) is defined by the equation α _i BIA / BOA-BIA, where (α _i ) is the local angle at a location (l _x ) of the chord length (1) and BIA and BOA is the entry and exit angles, respectively, of the skeleton line (4) at the beginning and end of the tendon. Blade design according to claim 1, characterized in that the course of the skeleton lines (4) within the area defined by the upper and lower limit curves (7, 8) is independent of the course of the leading edge (2) of the airfoil (1).

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