GB2151310A

GB2151310A - Gas turbine engine blade

Info

Publication number: GB2151310A
Application number: GB08430785A
Authority: GB
Inventors: John George Nourse; John Joseph Bourneuf; David Robert Abbott; Jack Reid Martin
Original assignee: General Electric Co
Current assignee: General Electric Co
Priority date: 1983-12-12
Filing date: 1984-12-06
Publication date: 1985-07-17
Also published as: FR2556409B1; IT1178658B; GB2151310B; GB8430785D0; SE8406320D0; IT8423829A0; SE8406320L; DE3444810C2; IT8423829A1; FR2556409A1; DE3444810A1

Abstract

A blade 10 for a gas turbine engine has a non-linear stacking axis 30 with portions 40, 42 of differing slopes. The slopes of the portions 40, 42 are of opposite signs. The arrangement provides a compressive component of bending stress on the trailing and leading edges due to centrifugal forces when in use. <IMAGE>

Description

SPECIFICATION Gas turbine engine blade This invention relates generally to blades for a gas turbine engine and, more particularly, to an improved blade effective for reducing stresses due to centrifugal force to improve the useful life of the blade.

An axial flow gas turbine engine conventionally includes a plurality of rows of alternating stationary vanes and rotating blades. The rotating blades are typically found in fan, compressor, and turbine sections of the engine, and inasmuch as these blades rotate for performing work in the engine, they are subject to stress due to centrifugal forces.

The centrifugal stress in a blade is relatively substantial and includes a substantially uniform centrifugal tensile stress and centrifugal bending stress including a tensile component and a compressive component which are added to the uniform tensile stress.

In a turbine section of the gas turbine engine, turbine blades are also subject to relatively hot, pressurized combustion gases.

These gases induce bending stresses due to the pressure of the combustion gases acting across the turbine blades, which stresses are often relatively small when compared to the centrifugal stresses. The relatively hot gases also induce thermal stress due to any temperature gradient created in the turbine blade.

A turbine blade, in particular, has a useful life, i.e., total time in service after which time it is removed from service, conventionally determined based on the above-described stresses and high-cycle fatigue, low-cycle fatigue, and creep-rupture considerations. A typical turbine blade has an analytically determined life-limiting section wherein failure of the blade is most likely to occur. However, blades are typically designed to have a useful life that is well in advance of the statistically determined time'of failure for providing a safety margin.

A significant factor in determining the use ful life of a turbine blade is the conventionally known creep-rupture strength, which is primarily proportional to material properties, tensile stress, temperature, and time. Notwithstanding that the relatively high temperatures of the combustion gases can induce thermal stess due to gradients thereof, these temperatures when acting on a blade under centrifugal tensile stress are a significant factor in the creep consideration of the useful life. In an effort to improve the useful life of turbine blades, these blades typically include internal cooling for reducing the temperatures experienced by the blade.However, the internal cooling is primarily most effective in cooling center portions of the blade while allowing leading and trailing edges of the blade to remain at relatively high temperatures with respect to the center portions thereof. Unfortunately, the leading and trailing edges of the blade are also, typically, portions of the blade subject to the highest stresses and therefore, the life-limiting section of a blade typically occurs at either the leading or trailing edges thereof.

Furthermore, a primary factor in designing turbine blades is the aerodynamic surface contour of the blade which is typically determined substantially independently of the mechanical strength and useful life of the blade. The aerodynamic performance of a blade is a primary factor in obtaining acceptable performance of the gas turbine engine. Accordingly, the aerodynamic surface contour that defines a turbine blade may be a significant limitation in the design of the blade from a mechanical strength and useful life consideration. With this aerodynamic performance restriction, the useful life of a blade may not be an optimum, which, therefore, results in the undesirable replacement of blades at less than optimal intervals.

Figure 1 is a perspective view of an axial entry blade for a gas turbine engine.

Figure 2 is a sectional view of the blade of Fig. 1 taken along line 2-2.

Figure 3 is a graphical representation of the stacking axis of the blade of Fig. 1 in a Y-Z plane.

Figure 4 is a perspective end view of the blade of Fig. 1 taken along line 4-4.

Figure 5 is a graphical representation of the stacking axis of the blade of Fig. 1 in an X-Y plane.

Figure 6 is a side view of the blade of Fig.

1 in the X-Z plane.

Figure 7 is a perspective view of an other axial entry blade for a gas turbine engine.

Figure 8 is a sectional view of the blade of Fig. 7 taken along line 8-8.

Figure 9 is a graphical representation of the stacking axis of the blade of Fig. 7 in a Y-Z plane.

Figure 10 is a graphical representation of the stacking axis of the blade of Fig. 7 in an X-Y plane.

Illustrated in Fig. 1 is a generally perspective view of an exemplary axial entry turbine blade 10 mounted in a turbine disk 11 of a gas turbine engine (not shown). The blade 10 includes an airfoil portion 12, a dovetail portion 14 and an optional platform 16. The airfoil portion 1 2 of the blade 10 comprises a plurality of transverse sections including a tip section 18, an intermediate section 20 and a root section 22, each of which has a center of gravity (C.g.) 24, 26 and 28, respectively.

The locus of the centers of gravity of the airfoil portion 1 2 define a stacking axis 30, which in accordance with the present invention is non-linear, e.g. bowed, and is described in further detail below.

The blade 10 further includes a conven tional reference XYZ coordinate system having an origin at the C.g. 28 of the root section 22. This coordinate system includes: an X, axial axis, which is aligned substantially parallel to a longitudinal centerline axis of the gas turbine engine; a Y, tangential axis, which is normal to the X axis and has a positive sense in the direction of rotation of the turbine disk 11; and a Z, radial axis, which represents a longitudinal axis of the blade 10 which is aligned coaxially with a radial axis of the gas turbine engine.

As illustrated in Figs. 1 and 2, the airfoil portion 1 2 of the blade 10 has an aerodynamic surface contour defined by and including a leading edge 32 and a trailing edge 34, between which extend a generally convex suction side 36 and a generally concave pressure side 38. The pressure side 38 faces generally in a negative direction with respect to the reference tangential axis Y; the suction side 36 faces generally in a positive direction with respect thereto.

Each of the plurality of transverse sections of the airfoil portion 1 2 of the blade 10 has its own conventionally known principal coordinate system. Illustrated in Fig. 2 is an exemplary principal coordinate system for the intermediate section 20 including an I axis and an Imjn axis. The principal coordinate system has an origin at the C.g. 26 of the intermediate section 20. 1ma represents an axis of maximum moment of inertia about which the intermediate section 20 has a maximum stiffness or resistance to bending and I,,,rn represents an axis of minimum moment of inertia about which the intermediate section 20 has a minimum stiffness or resistance to bending.

A conventional method of designing the blade 10 includes designing the airfoil portion 1 2 for obtaining a preferred aerodynamic surface contour as represented by the suction side 36 and the pressure side 38. The stacking axis 30 of the airfoil portion 1 2 would be conventionally made linear and coaxial with the reference radial axis Z.A suitable dovetail 14 and an optional platform 1 6 would be added and the entire blade 10 would then be analyzed for defining a life-limiting section, which, for example, may be the intermediate section 20, which is typically located between about 40 percent to about 70 percent of the distance from the root 22 to the tip 1 8 of the airfoil portion 1 2. Of course, analyzing the blade 10 for defining a life-limiting section is relatively complex and may include centrifugal, gas and thermal loading of the blade 10, which is accomplished by conventional methods.

However, in accordance with the present invention, the method of designing the blade 10 further includes redesigning the blade having the linear stacking axis, i.e., the reference blade, for obtaining a non-linear, tilted stacking axis 30 which is effective for introducing a compressive component of bending stress in the predetermined, life-limiting section.

More specifically, it will be appreciated from an examination of Figs. 1 and 2 that if the stacking axis 30 is spaced from the reference radial axis Z, that upon centrifugal loading of the airfoil portion 12, centrifugal force acting on the centers of gravity, C.g. 26 for example, will tend to rotate or bend the stacking axis 30 toward the reference radial axis Z thus introducing or inducing bending stress.

It will be appreciated from the teachings of this invention, that by properly tilting and spacing the stacking axis 30 with respect to the reference radial axis Z a compressive component of bending stress can be induced at both the leading edge 32 and the trailing edge 34 of the intermediate section 20 due to bending about the I mn axis as illustrated in Fig. 2. Of course, due to equilibrium of forces, an off-setting tensile component of bending stress is simultaneously introduced in the suction side 36 of the intermediate section 20 and generally at positive values of the 1max axis.

Illustrated in more particularity in Fig. 3 is an exemplary embodiment of the stacking axis 30 in accordance with the present invention and as viewed in the Y-Z plane. The stacking axis 30 is described as being non-linear from the C.g. 28 of the root section 22 to the C.g.

24 of the tip section 1 8 and may include either linear or curvilinear portions therebetween. As long as the stacking axis 30 has portions which extend away from and are spaced from the reference radial axis Z in a positive direction with respect to the reference tangential axis Y compressive components of bending stress will be introduced at the leading edge 32 and the trailing edge 34 of the airfoil portion 1 2.

The stacking axis 30 includes a first portion 40 extending from the C.g. 28 of the root section 22 to the C.g. 26 of the intermediate section 20, and a second portion 42 extending from C.g. 26 of the intermediate section 20 to the C.g. 24 of the tip section 18. Also illustrated is a reference, linearly tilted stacking axis 44 extending from the C.g. 28 of the root section 22 to the C.g. 24 of the tip section 18. The stacking axis 30 has an average slope represented by dashed line 46 which, as illustrated, is larger in magnitude than the slope of the reference axis 44 and is disposed between the reference radial axis Z and the reference stacking axis 44.

Assuming, for example, that the life-limiting section of the airfoil portion 1 2 is located at the intermediate section 20 it will be apparent from the teachings herein that a compressive component of bending stress can be introduced in the intermediate section 20 by using either the linear stacking axis 44 or the nonlinear stacking axis 30. To introduce the de sired bending stress at the intermediate section 20, the stacking axis 30 must be tilted with respect to the reference radial axis Z at those sections radially outwardly from the intermediate section 20, i.e., the second portion 42 of the stacking axis 30.

The slope of the stacking axis 30 is generally inversely proportional to the amount of bending stress realizable at the intermediate section 20. Accordingly, in the first embodiment of the invention illustrated in Figs. 1-6, relatively low values of the slope of the second portion 42 are preferred and result in relatively large values of induced bending stress in the intermediate section 20. However, a relatively large value of the average slope 46 is also preferred so that relatively low bending stress is simultaneously induced in the root section 22. Additionally, the second portion 42 of the stacking axis 30 has less of a slope than that of a comparable portion 44a of the reference linear stacking axis 44, which indicates that relatively more bending stress can be introduced thereby at the intermediate section 20.

However, not only is the reference linear stacking axis 44 less effective in introducing the desired bending stress to the intermediate section 20, but inasmuch as the reference stacking axis 44 is linear fronm C.g. 28 to the C.g. 24, substantial, undesirable bending stresses are also introduced at the root section 22. These increased bending stresses at the root section 22 are a limit to the amount of bending stress introducible by the reference linear stacking axis 44 in the life-limiting section of the airfoil portion 1 2 in that the lifelimiting section may thereby be relocated from the intermediate section 20 to the root section 22.

In contrast, inasmuch as the average slope line 46 of the non-linear stacking axis 30 has a magnitude greater than that of the reference stacking axis 44, it will be appreciated that not only does the non-linear stacking axis 30 provide for increased bending stress at the intermediate section 20 but less of a bending stress at the root 22 as compared to that provided by the reference linear stacking axis 44. Accordingly, a non-linear stacking axis 30 is more effective for introducing the desired compressive components of bending stress at the life-limiting section without adversely increasing the bending stresses at the root section 22.

More specifically, the stacking axis 30 according to the exemplary embodiment illustrated in Fig. 3 includes portions thereof disposed on two sides of the reference radial axis Z which are effective for obtaining increased bending stress at the intermediate section 20 without adversely increasing bending stress at the root section 22. The first portion 40 has a first average slope between C.g. 28 and C.g.

26, and the second portion 42 has a second average slope between the C.g. 26 and the C.g. 24, wherein the second slope has a negative sense with respect to the first slope.

Furthermore, the first portion 40 extends from the C.g. 28 and is tilted away from the reference radial axis Z in a generally negative Y axis direction, thusly, resulting in the first slope having a negative value. The second portion 42 extends from the C.g. 26 in a positive Y direction and with a positive slope which allows the second portion 42 to intersect the reference radial axis Z at one point and extend into the positive side of the Y axis.

Inasmuch as the stacking axis 30 has portions on both sides of the reference radial axis Z, it will be appreciated that the average slope line 46 of the stacking axis 30 will have a relatively larger value than would otherwise occur if the stacking axis 30 were disposed solely on one side of the reference radial axis Z. This arrangement is effective for allowing the second portion 42 to have a relatively small second slope for introducing substantially more compressive component of bending stress at the leading edge 32 and the trailing edge 34, for example, at the intermediate section 20.

The embodiment of the invention illustrated in Fig. 3, therefore, not only allows for an increase in the desired compressie stress at the intermediate section 20 but also results in reduced stresses at the root section 22 inasmuch as the average slope line 46 can be made substantially close to, if not coaxial with, the reference radial axis Z.

Fig. 4 illustrates an end view of the airfoil portion 1 2 from the trailing edge 34. The airfoil portion 1 2 further includes a substantially flat, relatively thin and flexible plate-like trailing edge portion 48 which extends radially inwardly from the tip portion 18 and may extend to the root portion 22 as illustrated.

The trailing edge portion 48 defines a trailing edge plane and is disposed at an angle B from the X axis toward the Y axis. In accordance with another feature of the present invention, the trailing edge portion 48 is not tilted in a transverse direction and is oriented in a substantially radial direction as additionally illustrated in Fig. 2. This is preferred for minimizing centrifugal bending stresses in the trailing edge portion 48 which would otherwise be generated if the trailing edge portion 48 was disposed at an angle with respect to the radial axis Z. This is effective for preventing distortion of the trailing edge portion 48, which would otherwise occur, for, thereby, preventing substantial changes in the aerodynamic contour thereof as well as for preventing localized creep distortion.

Accordingly, in order to maintain the preferred radial orientation of the trailing edge portion 48, and in order to introduce the desired compressive components of bending stress in the leading edge 32 and the trailing edge 34, the stacking axis 30 is tilted or disposed in a direction primarily parallel to the orientation of the trailing edge portion 48 and, therefore, lies substantially in a plane aligned substantially parallel to the trailing edge plane.

More specifically, the stacking axis 30 as illustrated in Fig. 5 is disposed at an angle B with respect to the X axis toward the Y axis.

The angle B represents the orientation of the trailing edge portion 48 in the X-Y plane as illutrated in Figs. 2 and 4. Although the stacking axis 30 is not disposed in a direction substantially parallel to the Y axis, it includes components in the positive Y axis direction which will thus introduce the preferred compressive component of bending stress in the leading edge 32 and the trailing edge 34.

Another advantage in accordance with the present invention from tilting the stacking axis 30 primarily in a direction parallel to the orientation of the trailing edge portion 48 is illustrated in Fig. 6. More specifically, by tilting the stacking axis 30 as above described, it will be appreciated that for a given aerodynamic surface contour, the leading edge 32 will be tilted away from the reference radial axis Z and the trailing edge 34 will be tilted toward the reference radial axis Z. As a result, the tilted airfoil portion 1 2 in accordance with the present invention when compared with an untilted airfoil portion represented partly in dashed line as 50 will no longer have a trailing edge tip region 52 disposed directly radially outwardly of a trailing edge intermediate region 54.

More specifically, the airfoil portion 1 2 includes the leading edge tip region 56 disposed radially outwardly of the leading edge intermediate region 58 and in a positive X direction therefrom. Similarly, the trailing edge tip region 52 extends in a positive X direction from the trailing edge intermediate 54 but, however, is not disposed directly radially outwardly therefrom, thusly, leaving a space 52' which would otherwise be a trailing edge tip region of the airfoil portion 1 2. The significance of this feature is that the trailing edge intermediate region 54 will be therefore subject to less centrifugal loading, and stresses therefrom, inasmuch as centrifugal loading from the trailing edge tip region 52 is primarily dispersed through a center region 60 of the airfoil portion 1 2. Although the leading edge intermediate region 58 must now absorb the centrifugal loading due to the leading edge tip region 56 disposed thereover, the increase in stress at the leading edge intermediate region 58 is relatively small inasmuch as the leading edge intermediate region 58 is substantially larger in cross-sectional area than the trailing edge intermediate region 54.

Illustrated in Fig. 7 is a generally perspective view of an other exemplary axial-entry turbine blade 110 mounted in a turbine disk 111 of a gas turbine engine (not shown). The blade 110 includes an airfoil portion 112, a dovetail portion 114 and an optional platform 116. The airfoil portion 112 of the blade 110 comprisea a plurality of transverse sections including a tip section 118, an intermediate section 1 20 and a root section 122, each of which has a center of gravity (C.g.) 124, 1 26 and 128, respectively. The locus of the centers of gravity of the airfoil portion 11 2 define a stacking axis 130, which in accordance with the present invention is non-linear, e.g.

bowed, and is described in further detail below.

The blade 110 further includes a conventional reference XYZ coordinate system having an origin at the C.g. 1 28 of the root section 1 22. This coordinate system includes: an X, axial axis, which is aligned substantially parallel to a longitudinal centerline axis of the gas turbine engine; a Y, tangential axis, which is normal to the X axis and has a positive sense in the direction of rotation of the turbine disk 111; and a Z, radial axis, which represents a longitudinal axis of the blade 110 which is aligned coaxially with a radial axis of the gas turbine engine.

As illustrated in Figs. 7 and 8, the airfoil portion 11 2 of the blade 110 has an aerodynamic surface contour defined by and including a leading edge 1 32 and a trailing edge 134, between which extend a generally convex suction side 1 36 and a generally concave pressure side 1 38. The pressure side 1 38 faces generally in a negative direction with respect to the reference tangential axis Y; the suction side 1 36 faces generally in a positive direction with respect thereto.

Each of the plurality of transverse sections of the airfoil portion 11 2 of the blade 110 has its own conventionally known principal coordinate system. Illustrated in Fig. 8 is an exemplary principal coordinate system for the intermediate section 1 20 including an 1max axis and an Imm axis.The principal coordinate system has an origin at the C.g. 1 26 of the intermediate section 1 20. 1max represents an axis of maximum moment of inertia about which the intermediate section 1 20 has a maximum stiffness or resistance to bending and lm,n represents an axis of minimum moment of inertia about which the intermediate section 1 20 has a minimum stiffness or resistance to bending.

A conventional method of designing the blade 110 includes designing the airfoil portion 11 2 for obtaining a preferred aerodynamic surface contour as represented by the suction side 1 36 and the pressure side 1 38.

The stacking axis 1 30 of the airfoil portion 11 2 would be conventionally made linear and coaxial with the reference radial axis Z. A suitable dovetail 11 4 and an optional platform 11 6 would be added and the entire blade 110 would then be analyzed for defining a life-limiting section, which, for example, may be the intermediate section 120, which is typically located between about 40 percent to about 70 percent of the distance from the root 1 22 to the tip 11 8 of the airfoil portion 11 2. Of course, analyzing the blade 110 for defining a life-limiting section is relatively complex and may include centrifugal, gas and thermal loading of the blade 110, which is accomplished by conventional methods.

However, in accordance with the present invention, the method of designing the blade 110 further includes redesigning the blade having the linear stacking axis, i.e., the reference blade, for obtaining a non-linear, tilted stacking axis 1 30 which is effective for introducing a compressive component of bending stress in the predetermined, life-limiting section.

More specifically, it will be appreciated from an examination of Figs. 7 and 8 that if the stacking axis 1 30 is spaced from the reference radial axis Z, that upon centrifugal loading of the airfoil portion 11 2, centrifugal force acting on the centers of gravity, C.g. 26 for example, will tend to rotate or bend the stacking axis 1 30 toward the reference radial axis Z thus introducing or inducing bending stress.

It will be appreciated from the teachings of this invention, that by properly tilting and spacing the stacking axis 1 30 with respect to the reference radial axis Z a compressive component of bending stress can be induced at both the leading edge 1 32 and the trailing edge 1 34 of the intermediate section 1 20 due to bending about the Im, axis as illustrated in Fig. 8. Of course, due to equilibrium of forces, an off-setting tensile component of bending stress is simultaneously introduced in the suction side 1 36 of the intermediate section 1 20 and generally at positive values of the 1max axis.

Illustrated in more particularity in Fig. 9 is an exemplary embodiment of the stacking axis 1 30 in accordance with the present invention and as viewed in the Y-Z plane. The stacking axis 30 extends away from and is spaced from the reference radial axis Z in a positive direction with respect to the reference tangential axis Y from, but not including, the root section 1 22 to the tip section 11 8. The stacking axis 1 30 is generally defined as including a first portion 1 40 extending from the C.g. 128 of the root section 122 to the C.g. 1 26 of the intermediate section 20, and a second portion 142 extending from C.g.

1 26 of the intermediate section 1 20 to the C.g. 124 of the tip section 118. Also illustrated is a reference, linear, tilted stacking axis 144 extending from C.g. 1 28 of the root section 1 22 to the C.g. 1 24 of the tip section 118. The stacking axis 130 has an average slope represented by dashed line 146 which, as illustrated, is larger in magnitude than the slope of the reference axis 144 and is disposed between the reference radial axis Z and the reference stacking axis 144.

Assuming, for example, that the life-limiting section of the airfoil portion 11 2 is located at the intermediate section 1 20 it will be apparent from the teachings herein that a compressive component of bending stress can be introduced in the intermediate section 1 20 by using either the linear stacking axis 144 or the non-linear stacking axis 1 30. To introduce the desired bending stress at the intermediate section 120, the stacking axis 1 30 must be tilted with respect tothe reference radial axis Z at those sections radially outwardly from the intermediate section 120, i.e., the second portion 42 of the stacking axis 1 30. The slope of the stacking axis 1 30 is generally inversely proportional to the amount of bending stress realizable at the intermediate section 120.

As illustrated in Fig. 9, the first portion 140 has a first, average slope, and the second portion 142 has a second, average slope, the first slope being greater than the second slope. This is effective for obtaining increased bending stress at the intermediate section 1 20 without adversely increasing bending stress at the root section 1 22. Additionally, the second portion 142 of the stacking axis 1 30 has less of a slope than a comparable portion 1 44a of the reference linear stacking axis 144, which indicates that more bending stress can be introduced thereby at the intermediate sectin 1 20.

However, not only is the reference linear stacking axis 1 44 less effective in introducing the desired bending stress to the intermediate section 120, but inasmuch as the reference stacking axis 144 is linear from C.g. 1 28 to the C.g. 124, substantial, undesirable bending stresses are also introduced at the root section 1 22. These increased bending stresses at the root 1 22 are a limit to the amount of bending stress introducible by the reference linear stacking axis 1 44 in the lifelimiting section of the airfoil portion 11 2 in that the life-limiting section may thereby be relocated from the intermediate section 1 20 to the root section 1 22.

In contrast, inasmuch as the average slope line 1 46 for the non-linear stacking axis 1 30 has a greater magnitude than that of the reference stacking axis 144, it will be appreciated that not only does the non-linear stacking axis 1 30 provide for increased bending stress at the intermediate secion 1 20 but less of a bending stress at the root 1 22 as compared to that provided by the reference linear stacking axis 144. Accordingly, a non-linear stacking axis 1 30 is more effective for introducing the desired compressive components of bending stress at the life-limiting section without adversely increasing the bending stresses at the root sectin 1 22.

Fig. 9 illustrates in more particularity the non-linear stacking axis 1 30 according to the present invention. The stacking axis 1 30 is described as being non-linear from the C.g.

128 of the root section 122 to the C.g. 1 24 of the tip section 11 8 and may include either linear or curvilinear portions therebetween.

As long as the stacking axis 1 30 has portions which extend away from and are spaced from the reference radial axis Z in a positive direction with respect to the reference tangential axis Y compressive components of bending stress will be introduced at the leading edge 1 32 and the trailing edge 1 34 of the airfoil portion 11 2.

Optimally, the magnitude of compressive stress induced is preferably made equal to approximately the compressive yield strength of the blade material. This will provide maximum compressive stress in the leading edge 1 32 and the trailing edge 1 34 during operation which will provide improved fatigue life.

Furthermore, the stacking axis 1 30 can be tilted also to induce stresses initially greater than the compressive yield strength, which stresses will yield thereto after the first few initial cycles of operation, so that manufacturing inaccuracies do not prevent the induced stress from reaching the optimal value.

More specifically, and as additionally illustrated in Fig. 8, the tangential reference axis Y is generally aligned with the I man axes of the transverse sections of the airfoil portion 11 2, the Iman axis of the intermediate section 120, for example. Accordingly, in operation, centrifugal forces act at each of the centers of gravity of the airfoil portion 1 2 and will thus tend to straighten the airfoil portion 11 2 to bring the stacking axis 1 30 closer to the reference radial axis Z.For example, when the average slope line 146 of the stacking axis 1 30 is spaced, from reference radial axis Z in a generally positive direction with respect to the tangential axis Y and the Ima, axis, compressive components of bending stress will be introduced at the leading edge 1 32 and trailing 1 34.

Illustrated in Fig. 10 is a view of the stacking axis 1 30 with respect to the X-Y plane. The stacking axis 1 30 preferably lies substantially in a plane defined by the reference radial axis Z and tangential axis Y, is preferably linear in the X-Y plane and is preferably aligned along the positive Y axis.

This is preferred so that the aerodynamic surface contour and orientation of the airfoil portion 1 2 does not significantly change as the stacking axis 1 30 is tilted.

Alternatively, the spacing of the stacking axis 1 30 from the reference radial axis Z could also be positive in magnitude and be substantially oriented along the I ma, direction for each of the transverse sections and might look like a stacking axis 1 30a illustrated in Fig. 1 0. However, the relative twist of the airfoil portion 112, i.e., its orientation with respect to the reference axial axis X, would change from that of an untilted blade, thusly changing the aerodynamic surface contour of the airfoil portion 11 2.

Claims

1. A blade for a gas turbine engine comprising an airfoil portion including a non-linear stacking axis having a first portion having a first slope and a second portion having a second slope, said second slope having a negative sense with respect to said first slope.

2. A blade according to Claim 1 wherein said airfoil portion further comprises a leading edge and a trailing edge and said stacking axis is effective for generating a compressive component of bending stress in said trailing edge and said leading edge due to centrifugal force acting on said blade.

3. A blade according to Claim 2 wherein said airfoil portion further comprises: a plurality of transverse sections including a root section, an intermediate section, and a tip section, each having a center of gravity; reference radial and tangential axes extending outwardly from said center of gravity of said root section; and wherein said stacking axis extends from said center of gravity of said root section and is spaced from said reference radial axis at said tip section.

4. A blade according to Claim 3 wherein said first portion of said stacking axis extends from said root section to said intermediate section, said second portion of said stacking axis extends from said intermediate section to said tip section and said second portion of said stacking axis intersects said reference radial axis.

5. A blade according to Claim 2 wherein said airfoil portion further comprises: a pressure side facing generally in a negative direction with respect to said reference tangential axis; a suction side facing generally in a position direction with respect to said reference tangential axis; wherein said first portion of stacking axis extends away from said reference radial axis in a negative direction with respect to said reference tangential axis and said second portion thereof extends in a positive direction thereto.

6. A blade according to Claim 5 wherein said airfoil portion further comprises a substantially flat trailing edge portion defining a trailing edge plane aligned generally in a radial direction and said stacking axis lies substantially in a plane aligned substantially parallel to said trailing edge plane.

7. A blade according to Claim 6 wherein said trailing edge portion is aligned substantially in a radial direction.

8. A blade for a gas turbine engine comprising an airfoil portion including a leading edge, a trailing edge, a pressure side, a suction side, and a plurality of transverse sections including a root section, an intermediate section, and a tip section, each of said plurality of sections having a center of gravity, the locus of which define a stacking axis, said blade further including reference radial and tangential axes extending outwardly from said center of gravity of said root section toward said tip section and said suction side, respectively, said stacking axis being non-linear and being effective for introducing a compressive component of bending stress in said trailing edge and said leading edge due to centrifugal force acting on said blade, said stacking axis including a first portion having a first slope and a second portion having a second slope, said second slope having a negative sense with respect to said first slope.

9. A blade according to Claim 8 wherein said airfoil portion further comprises a substantially flat trailing edge portion defining a trailing edge plane aligned substantially parallel in a radial direction and said stacking axis lies substantially in a plane aligned substantially parallel to said trailing edge plane.

1 0. A method of designing blade for a gas turbine engine comprising the steps of: designing a reference blade including an airfoil portion having a linear stacking axis and an aerodynamic surface contour; analyzing said reference blade for defining a life-limiting section of said airfoil portion; redesigning said reference blade for obtaining a non-linear stacking axis effective for introducing a compressive component of bending stress in said life-limiting section of said airfoil portion, said stacking axis including a first portion having a first slope and a second portion having a second slope, said second slope having a negative sense with respect to said first slope.

11. A blade for a gas turbine engine comprising an airfoil portion including a non-linear stacking axis.

1 2. A blade according to Claim 11 wherein said airfoil portion further comprises a leading edge and a trailing edge and said stacking axis is effective for introducing a compressive component of bending stress in said trailing edge and said leading edge due to centrifugal force acting on said blade.

1 3. A blade according to Claim 1 2 wherein said stacking axis is effective for generating a compressive stress in said trailing edge and said leading edge exceeding a compressive yield strength of said blade.

1 4. A blade according to Claim 1 2 wherein said airfoil portion further comprises: a plurality of transverse sections including a root section, an intermediate section, and a tip section, each having a center of gravity; reference radial and tangential axes extending outwardly from said center of gravity of said root section; and wherein said stacking axis extends from said center of gravity of said root section and is spaced from said reference radial axis at said tip section.

15. A blade according to Claim 14 wherein said stacking axis is spaced from said reference radial axis from said root section to said tip section.

1 6. A blade according to Claim 14 wherein said airfoil portion further comprises: a pressure side facing generally in a negative direction with respect to said reference tangential axis; a suction side facing generally in a positive direction with respect to said reference tangential axis; wherein said stacking axis extends away from said reference radial axis in a positive direction with respect to said reference tangential axis.

1 7. A blade according to Claim 1 6 wherein said stacking axis lies substantially in a plane defined by said reference radial and tangential axes.

18. A blade according to Claim 14 wherein said stacking axis further includes a first portion extending from said root section to said intermediate section and a second portion extending from said intermediate section to said tip section, said first portion having a first slope and said second portion having a second slope, said first slope being greater than said second slope.

19. A blade according to Claim 14 wherein said airfoil portion further includes a predetermined life-limiting section having an Im, axis and an 1ma, axis and wherein said stacking axis is spaced from said reference radial axis in a positive direction with respect to said I ma, axis.

20. A blade for a gas turbine engine comprising an airfoil portion including a leading edge, a trailing edge, a pressure side, and a suction side, and a plurality of transverse sections including a root section, an intermediate section, and a tip section, each of said plurality of sections having a center of gravity, the locus of which define a stacking axis, said blade further including reference radial and tangential axes extending outwardly from said center of gravity of said root section toward said tip section and said suction side, respectively, said stacking axis being non-linear and spaced from said reference radial axis from said root section to said tip section and being effective for introducing a compressive component of bending stress in said trailing edge and said leading edge due to centrifugal force acting on said blade.

21. A blade according to Claim 21 wherein said stacking axis includes a first portion extending from said root section to said intermediate section and a second portion extending from said intermediate section to said tip section, said first portion having a first slope and said second portion having a second slope, said first slope being greater than said second slope.

22. A blade according to Claim 21 wherein said stacking axis lies substantially in a plane defined by said reference radial and tangential axes.

23. A method of designing a blade for a gas turbine engine comprising the steps of: designing a reference blade including an airfoil portion having a linear stacking axis and an aerodynamic surface contour; analyzing said reference blade for defining a life-limiting section of said airfoil portion; and redesigning said reference blade for obtaining a non-linear stacking axis effective for introducing a compressive component of bending stress in said life-limiting section of said airfoil portion.

24. A method of designing a blade as claimed in claim 10 or claim 23 and substantially as hereinbefore described.

25. A blade substantially as hereinbefore described with reference to and as illustrated in Figs. 1 to 6 or Figs. 7 to 10 of the drawings.