GB2106192A - Turbomachine blade - Google Patents

Abstract

A highly twisted fan blade has a camber over at least a part of the radial length of the aerofoil adjacent the root portion which is, when viewed from the high pressure side of the blade, concave adjacent leading edge 23 and convex adjacent trailing edge 25. <IMAGE>

Description

SPECIFICATION Rotor blade for a turbomachine This invention relates to rotor blades for turbomachines and is particularly concerned with fan blades for by-pass type gas turbine engines.

The recent trend in the design of fan blades for by-pass gas turbine aeroengines having high bypass ratios is towards thinner more highly twisted and cambered rotor blades. The centrifugal loads on such blades tends to cause the blades to untwist. Much of the stiffness of such blades derives from their resistance to the longitudinal deformation which accompanies the twisting. As the blade untwists the blade tip attempts to line up over the root, and the leading and trailing edges, in trying to line up radially, are compressed whilst to compensate, the centre of the section goes into tension. There is thus a chordwise stress distribution which is compressive at leading and trailing edges and tensile in the middie.The aerofoil section warps freely away from discontinuities but at discontinuities such as where the aerofoil section merges with the blade root platform the warping is restrained and very high longitudinal stresses are incurred.

The invention as claimed reduces the warping stresses by aligning the leading and trailing edges of the aerofoil, at least over the region adjacent the root portion, more closely to radial planes.

In one embodiment the camber of the blade over its whole length is more closely aligned to radial planes.

The present invention will now be described, by way of an example only, with reference to the accompanying drawings, in which: Figure 1 illustrates a known highly twisted add cambered blade not constructed in accordance with the present invention, Figure 2 illustrates a blade incorporating the present invention, and, Figure 3 illustrates the cross-sectional shape of the aerofoil of the blade of Figure 2 at various radii shown as AA, BB, and CC in Figure 2.

Referring to Figure 1 the blade illustrated is a front fan blade of a by-pass type of gas turbine aeroengine. The blade 10 comprises a root portion 11 of conventional fir-tree configuration which, in use, is inserted into a complementary shaped slot in a fan hub of the engine.

The blade 10 incorporates a blade root platform 1 2 from which protrudes the aerofoil 1 3. The aerofoil is highly twisted and cambered with the result that the leading and trailing edges 14, 15, respectively, are highly curved. At the region where the aerofoil 13 merges into the platform 12 very high stresses are produced due to the centrifugal loads trying to untwist the blade. As the leading and trailing edges 14, 1 5 of the blade try to straighten out in radial direction very high stresses are generated at the blade platform, because of the drastic change in cross-sectional shapes between the aerofoil and the root portion.

To lessen these stresses the blade 1 6 of the present invention shown in Figure 2 comprises a root portion 1 7 which may be of conventional firtree configuration, and an aerofoil 18. The camber of the aerofoil 18 (see Figure 3) over at least the radially inner extent 19, adjacent the blade root platform 20 is when viewed from the high pressure side 21 of the blade 1 6 is concave over a region 22 adjacent the leading edge 23 and is convex over a region 24 adjacent the trailing edge 25 of the aerofoil. This has the effect of straightening out the leading and trailing edges of the blade at least over the radially inner extent 19 of the blade and the modified cross-sectional area of the aerofoil provides a better restraint against warping. This in turn lessens the maximum warping restraint stresses generated at the leading and trailing edges of the blade.

It will be appreciated that by altering the camber at the trailing edge of the blade to bring the trailing edge into a radial plane, the whole aerofoil can be repositioned relative to the blade platform to position the leading edge in a radial plane.

1. A blade for a turbomachine comprising a root portion and a twisted, cambered, aerofoil shaped portion, wherein the camber of the aerofoil, over at least a part of the length of the aerofoil adjacent the root portion, is, when viewed from the high pressure side of the blade, concave at a leading edge region of the aerofoil and convex at a trailing edge region of the aerofoil.

2. A blade according to Claim 1 wherein the leading and trailing edges of the blade extend radially from the root portion.

3. A blade substantially as hereindescribed with reference to Figures 2 and 3 of the accompanying drawings.

**WARNING** end of DESC field may overlap start of CLMS **.

Claims (3)

**WARNING** start of CLMS field may overlap end of DESC **. SPECIFICATION Rotor blade for a turbomachine This invention relates to rotor blades for turbomachines and is particularly concerned with fan blades for by-pass type gas turbine engines. The recent trend in the design of fan blades for by-pass gas turbine aeroengines having high bypass ratios is towards thinner more highly twisted and cambered rotor blades. The centrifugal loads on such blades tends to cause the blades to untwist. Much of the stiffness of such blades derives from their resistance to the longitudinal deformation which accompanies the twisting. As the blade untwists the blade tip attempts to line up over the root, and the leading and trailing edges, in trying to line up radially, are compressed whilst to compensate, the centre of the section goes into tension. There is thus a chordwise stress distribution which is compressive at leading and trailing edges and tensile in the middie.The aerofoil section warps freely away from discontinuities but at discontinuities such as where the aerofoil section merges with the blade root platform the warping is restrained and very high longitudinal stresses are incurred. The invention as claimed reduces the warping stresses by aligning the leading and trailing edges of the aerofoil, at least over the region adjacent the root portion, more closely to radial planes. In one embodiment the camber of the blade over its whole length is more closely aligned to radial planes. The present invention will now be described, by way of an example only, with reference to the accompanying drawings, in which: Figure 1 illustrates a known highly twisted add cambered blade not constructed in accordance with the present invention, Figure 2 illustrates a blade incorporating the present invention, and, Figure 3 illustrates the cross-sectional shape of the aerofoil of the blade of Figure 2 at various radii shown as AA, BB, and CC in Figure 2. Referring to Figure 1 the blade illustrated is a front fan blade of a by-pass type of gas turbine aeroengine. The blade 10 comprises a root portion 11 of conventional fir-tree configuration which, in use, is inserted into a complementary shaped slot in a fan hub of the engine. The blade 10 incorporates a blade root platform 1 2 from which protrudes the aerofoil 1 3. The aerofoil is highly twisted and cambered with the result that the leading and trailing edges 14, 15, respectively, are highly curved. At the region where the aerofoil 13 merges into the platform 12 very high stresses are produced due to the centrifugal loads trying to untwist the blade. As the leading and trailing edges 14, 1 5 of the blade try to straighten out in radial direction very high stresses are generated at the blade platform, because of the drastic change in cross-sectional shapes between the aerofoil and the root portion. To lessen these stresses the blade 1 6 of the present invention shown in Figure 2 comprises a root portion 1 7 which may be of conventional firtree configuration, and an aerofoil 18. The camber of the aerofoil 18 (see Figure 3) over at least the radially inner extent 19, adjacent the blade root platform 20 is when viewed from the high pressure side 21 of the blade 1 6 is concave over a region 22 adjacent the leading edge 23 and is convex over a region 24 adjacent the trailing edge 25 of the aerofoil. This has the effect of straightening out the leading and trailing edges of the blade at least over the radially inner extent 19 of the blade and the modified cross-sectional area of the aerofoil provides a better restraint against warping. This in turn lessens the maximum warping restraint stresses generated at the leading and trailing edges of the blade. It will be appreciated that by altering the camber at the trailing edge of the blade to bring the trailing edge into a radial plane, the whole aerofoil can be repositioned relative to the blade platform to position the leading edge in a radial plane. CLAIMS

1. A blade for a turbomachine comprising a root portion and a twisted, cambered, aerofoil shaped portion, wherein the camber of the aerofoil, over at least a part of the length of the aerofoil adjacent the root portion, is, when viewed from the high pressure side of the blade, concave at a leading edge region of the aerofoil and convex at a trailing edge region of the aerofoil.

2. A blade according to Claim 1 wherein the leading and trailing edges of the blade extend radially from the root portion.

3. A blade substantially as hereindescribed with reference to Figures 2 and 3 of the accompanying drawings.