BLADE LEADING EDGE TIP RIB

Abstract

A rotor blade for a gas turbine engine includes a leading edge tip rib projecting outwardly from an airfoil of the blade at a tip region thereof. The tip rib continuously surrounds a leading edge of the airfoil and extends rearwardly from the leading edge along respective pressure and suction side surfaces to thereby alter the blade tip leakage vortex structure and strength, resulting in a stage efficiency benefit.

Description

TECHNICAL FIELD

The application relates generally to gas turbine engines and more particularly, to a rotor blade for such engines.

BACKGROUND OF THE ART

A rotor blade for a gas turbine engine typically includes an attachment and an airfoil. The airfoil extends between the attachment and a tip and has a concave pressure side surface, a convex suction side surface, a leading edge and a trailing edge. The airfoil is sized such that when it is configured within the engine, a clearance gap is defined between the blade tip and the surrounding static structure. Blade tip clearance for an unshrouded turbine blade, and the resultant air flow that flows from the pressure side surface to the suction side surface through this space, is a significant contributor to energy loss in an axial flow turbine stage. This “leakage” airflow mixes with the suction side airflow to form a vortex. The vortex disburses, causing relatively significant flow disturbances along most of the suction side surface. The cumulative result of these flow disturbances is a reduction in engine efficiency. Therefore, it is a challenging task to disrupt, block and/or reduce the leakage flow in order to mitigate the resultant damaging effects.

Accordingly, there is a need for an improved rotor blade configuration to reduce tip gap flow leakage.

SUMMARY

In one aspect, there is provided a rotor blade for a gas turbine engine, comprising: an attachment; an airfoil defining a pressure side surface and a suction side surface, extending between the attachment and a tip, a base region disposed adjacent to the attachment, a tip region, and a transition region located between the base region and the tip region, the tip defining a chord line, and the pressure side surface adjacent to the tip being substantially planar; and wherein the tip region includes a tip rib projecting outwardly from the airfoil, the tip rib continuously surrounding a leading edge of the airfoil, extending from the leading edge rearwardly along the respective pressure side surface and suction side surface, to define a first axial extension on the pressure side surface and a second axial extension on the suction side surface, each of the first and second axial extensions being in a range between 5% and 35% of the length of the chord line immediately downstream of the leading edge.

In another aspect, there is provided a gas turbine engine, comprising: a compressor section; a combustor section; and a turbine section; wherein the turbine section includes a plurality of rotors having a plurality of radially disposed rotor blades, each of the rotor blades including an attachment and an airfoil, the airfoil defining a pressure side surface and a suction side surface, extending between the attachment and a tip, a base region disposed adjacent to the attachment, a tip region, and a transition region located between the base region and the tip region, the tip region defining a chord line, and the pressure side surface in the tip region being substantially planar; and wherein the tip region includes a tip rib projecting outwardly from the airfoil, continuously surrounding a leading edge of the airfoil, extending from the leading edge rearwardly along the respective pressure side surface and suction side surface to define a first axial extension on the pressure side surface and a second axial extension on the suction side surface, each of the first and second axial extensions being in a range between 5% and 35% of the length of the chord line immediately downstream of the leading edge.

DESCRIPTION OF THE DRAWINGS

Reference is now made to the accompanying figures in which:

FIG. 1 is a partial schematic side cross-sectional view of a gas turbine engine, illustrating an example of the application of the described subject matter;

FIG. 2 is a schematic illustration of a rotor blade for the gas turbine engine in FIG. 1;

FIG. 3 is a schematic illustration of a cross-sectional slice of the airfoil of the rotor blade of FIG. 2;

FIG. 4 is a partial schematic front elevational view of the rotor blade of FIG. 2; and

FIG. 5 is a top plan view of a tip of the rotor blade of FIG. 2.

It will be noted that throughout the appended drawings, like features are identified by like reference numerals.

DETAILED DESCRIPTION

FIG. 1 illustrates a turbofan gas turbine engine according to one embodiment. The engine includes a housing or nacelle 10, a core casing 13, a low pressure spool assembly (not numbered) which includes a fan rotor 14, a low pressure compressor assembly 16 and a low pressure turbine assembly 18 connected by a shaft 12, and a high pressure spool assembly (not numbered) which includes a high pressure compressor assembly 22 and a high pressure turbine assembly 24 connected by a turbine shaft 20. The housing or nacelle 10 surrounds the core casing 13 and in combination with the housing 10 and the core casing 13, defines an annular bypass duct 28 for directing a bypass flow. The core casing 13 surrounds the low and high pressure spool assemblies to define a core fluid path 30 therethrough. In the core fluid path 30 there is provided a combustor 26 to form a combustion gas generator assembly which generates combustion gases in order to power the high pressure turbine assembly 24 and the low pressure turbine assembly 18. The engine therefore defines in series a compressor section (not numbered), a combustor section (not numbered) and a turbine section (not numbered). Low and high pressure compressor assemblies 16, 22 in the compressor section and the high and low pressure turbine assemblies 18, 24 in the turbine section, each include a plurality of radially disposed rotor blades 32 attached to a rotor disc. The radially disposed rotor blades 32 are rotatable along a longitudinally extending axis 33 of the engine.

FIG. 2 is a schematic illustration of one embodiment of the rotor blade 32 for use, for example in the high pressure turbine assembly 24 of the gas turbine engine. The rotor blade 32 includes an attachment 34, a platform 35, and an airfoil 36. Some embodiments of the rotor blade 32 may not include the platform 35. To simplify the description herein, the attachment 34 may be considered to include the platform 35 for purpose of defining the beginning of the airfoil 36. The rotor blade attachment 34 is adapted to be received within a slot disposed within the rotor disc. Rotor blade attachments are well known in the art and will not be described herein. The described subject matter is not limited to any particular attachment configuration.

The airfoil 36 has a leading edge 38, a trailing edge 40, a pressure side or pressure side surface 42, and a suction side or suction side surface 44. The pressure side surface 42 and the suction side surface 44 extend radially between the attachment 34 (including the platform 35) and a tip 46. The airfoil 36 includes a base region 50, disposed adjacent to the attachment 34, a tip region 54 disposed adjacent to the tip 46 and a transition region 52 disposed between the base region 50 and the tip region 54. For convenience of description, the base region 50 may be defined between a first end 60 and a second end 62 thereof. The first end 60 of the base region 52 is located at a cross-sectional “slice” of the airfoil 36 where the base region 50 abuts the attachment 34. The second end 62 of the base region 52 is located at a cross-sectional “slice” of the airfoil 36 where the base region 50 abuts the transition region 52. The transition region 52 is defined between a first end 62a and a second end 64 thereof. The first end 62a of the transition region 52 is located at the same cross-sectional “slice” of the airfoil 36 as the second end 62 of the base region 50. The second end 64 of the transition region 52 is located at a cross-sectional “slice” of the airfoil 36 where the transition region 52 abuts the tip region 54. The tip region 54 is defined between a first end 64a and a second end 66 thereof. The first end 64a of the tip region 54 is located at the same cross-sectional “slice” of the airfoil 36 as the second end 64 of the transition region 52. The second end 66 of the tip region 54 is located at the tip 46 of the airfoil 36.

FIG. 3 schematically illustrates the cross-sectional “slices” of the airfoil 36 of FIG. 2, forming the respective ends defining the base, transition and tip regions 50, 52 and 54. Stagger angles 56, 56a, 56b and 56c in the respective cross-sectional “slices” are defined as the angle between a respective chord line 48, 48a, 48b or 48c of the airfoil 36 and an axis (e.g. the longitudinally extending axis 33 of the gas turbine engine, etc.). In accordance with one embodiment the stagger angle may increase in a direction defined by a line (not shown) beginning at the attachment 34 and extending along the span of the airfoil 36 toward the tip 46. In one embodiment the chord lines of the airfoil 36 such as indicated by 48, 48a, 48b or 48c as well as their related stagger angles 56, 56a, 56b and 56c, may increase as the airfoil extends from the attachment 34 to the tip 46. A maximum camber line thickness of the airfoil 36 is measured by the maximum distance between a “mean camber line” 69, 69a, 69b or 69c and the corresponding chord line 48, 48a, 48b or 48c, and is indicated by respective arrows 68, 68a, 68b, 68c. FIG. 3 illustrates that the maximum camber line airfoil thickness may vary along the airfoil span from the cross-sectional “slice” representing the first end 60 of the base region 50, to the cross-sectional slice representing the second end 66 of the tip region 54, such as by decreasing. The pressure side surface 42 adjacent to the tip 46 may be substantially planar or may be substantially parallel to the chord line 48c defined in the tip 46.

The embodiment of the airfoil 32 as above-described is similar to the rotor blade described in US Patent Publication 2011/0135482 A1, published on Jun. 9, 2011 which is incorporated by reference herein. The rotor blade as above-described is a rotor blade with very low or no camber at the tip or a rotor blade having a substantially un-cambered symmetrical airfoil tip. The difference between the embodiment illustrated in FIGS. 2 and 3 and the rotor blade described in US 2011/0135482 A1 lies in that the rotor blade 32 as shown in FIGS. 2 and 3 includes a novel feature of a blade leading edge tip rib 70.

Shrouded blade tips with fins are used to counter the blade tip clearance leakage however, this is not feasible for high pressure turbines. For an un-shrouded blade tip in high pressure turbines, various winglet configurations on the pressure and suction sides have been studied. The pull-stress created by the actual weight of the winglet configurations has been shown to be a drawback to its use in actual aero engines. It is even more challenging and has not been attempted, to provide a suitable winglet design for a blade with a very low or no camber at the tip. However, it has been found by computational fluid dynamics (CFD) analysis on a specific blade design, such as the airfoil 36 described above, that a mini-winglet at the leading edge, referred to herein as a leading edge tip rib 70, smoothly blended on the pressure and suction side surfaces 42, 44, has shown to provide up to 80% of the improvement in stage efficiency with respect to a full tip rib (not shown) which encompasses the leading edge, pressure and suction side surfaces and trailing edge. Thus, most of the benefit is obtained for the leading edge tip rib 70 with minimum increase in pull-stress at the blade tip 46 and low impact on modal frequencies.

In accordance with one embodiment illustrated in FIGS. 3, 4 and 5, the rotor blade 32 may have a specific highly loaded and front-loaded airfoil tip loading distribution. The front-loaded airfoil tip loading footprint is that the minimum pressure may occur between 5% and 20% axial chord on the suction side surface 44 when the leading edge tip rib 70 is not incorporated with the airfoil 36. The tip region 54 may include the leading edge tip rib 70 projecting outwardly from the airfoil 36, continuously surrounding the leading edge 38 of the airfoil 36 and extending from the leading edge 38 rearwardly along the respective pressure side surface 42 and suction side surface 44 to define a first axial extension 72 on the pressure side surface 42 and a second axial extension 74 on the suction side surface 44. The leading edge tip rib 70 in the tip region 54, for example may be flush with the tip 46 of the airfoil 36 and smoothly blended along the pressure and suction side surfaces 42 and 44, for example by a suitable casting requirement fillet 76. In some embodiments the leading edge tip rib 70 may have a lateral extension indicated for example by arrows 78, which may or may not be equal one to another on the leading edge 38, pressure and suction side surfaces 42, 44. The axial extension 72 and 74 of the leading edge tip rib 70 on the respective pressure and suction side surfaces 42, 44 may or may not be equal to each other. However, equal lateral and equal axial extents may be required from a structural point of view. The lateral extension 78 and the axial extensions 72, 74 of the leading edge tip rib 70 may be selected depending on the blade pressure loading distribution and thus tend to be an optimization for a particular blade design.

In one embodiment of the rotor blade 32 the dimension for the lateral extension 78 may be selected between a range of 5% and 30% as thick as the maximum airfoil thickness 75 along the mean camber line 69c at the tip 46 (the preferred embodiment was 25%). The dimensions for the axial extension 72 or 74 of the leading edge tip rib 70 may be selected in a range of 5% and 35% of the length of the chord line immediately downstream of the leading edge 38. Diminishing returns in stage efficiency gains have been shown outside these ranges and thus, optimum results are obtained with a minimum leading edge tip rib in size and weight that can lend itself applicable to high pressure turbines within structural limitations. The previous assertion assumed smooth blending of the axial extension 72 or 74 of the leading edge tip rib 70. Without smooth axial blending, the first axial extension 72 on the pressure side surface 42 would be up to a range between 5% and 10% and the axial extension 74 on the suction side surface 44 would be up to a range between 5% to 20% of the length of the chord line, both normal to the airfoil walls, but this would be difficult to cast as a blade shape and maintain structural integrity when being ground for final blade tip radius at assembly. With the optimized leading edge tip rib 70, the pass resistance for the leakage of fluid as it migrates from the pressure side surface to the suction side surface and from the leading edge to the suction side surface, is increased, thus altering the blade tip leakage vortex structure and strength. The latter leads to a stage efficiency benefit.

The above description is meant to be exemplary only, and one skilled in the art will recognize that changes may be made to the embodiments described without departing from the scope of the described subject matter. Still other modifications will be apparent to those skilled in the art, in light of a review of this disclosure and such modifications are intended to fall within the appended claims.

Claims

1. A rotor blade for a gas turbine engine, comprising:

an attachment; and

an airfoil defining a pressure side surface and a suction side surface, extending between the attachment and a tip, a base region disposed adjacent to the attachment, a tip region, and a transition region located between the base region and the tip region, the tip defining a chord line, and the pressure side surface adjacent to the tip being substantially planar; and

wherein the tip region includes a tip rib projecting outwardly from the airfoil, the tip rib continuously surrounding a leading edge of the airfoil, extending from the leading edge rearwardly along the respective pressure side surface and suction side surface, to define a first axial extension on the pressure side surface and a second axial extension on the suction side surface, each of the first and second axial extensions being in a range between 5% and 35% of the length of the chord line immediately downstream of the leading edge.

2. The rotor blade as defined in claim 1 wherein the tip rib has a lateral extension in a range between 5% and 30% as thick as the maximum airfoil thickness along the mean camber line.

3. The rotor blade as defined in claim 1 wherein the tip rib has a lateral extension which is 25% as thick as the maximum airfoil thickness along the mean camber line.

4. The rotor blade as defined in claim 1 wherein each of the first and second axial extensions is 35% of the length of the chord line immediately downstream of the leading edge.

5. The rotor blade as defined in claim 1 wherein tip rib is flush with the tip of the airfoil.

6. The rotor blade as defined in claim 1 wherein the chord line is substantially parallel to the pressure side surface.

7. The rotor blade as defined in claim 1 wherein a chord line increases as the airfoil extends from the attachment to the tip.

8. The rotor blade as defined in claim 1 wherein the first and second axial extensions of the tip rib are equal.

9. The rotor blade as defined in claim 1 wherein the first and second axial extensions of the tip rib are different.

10. The rotor blade as defined in claim 3 wherein the lateral extension on the respective leading edge, pressure side surface and suction side surface are equal.

11. The rotor blade as defined in claim 3 wherein the lateral extension varies on the leading edge, pressure side surface and suction side surface.

12. A gas turbine engine, comprising:

a compressor section;

a combustor section; and

a turbine section;

wherein the turbine section includes a plurality of rotors having a plurality of radially disposed rotor blades, each of the rotor blades including an attachment and an airfoil, the airfoil defining a pressure side surface and a suction side surface, extending between the attachment and a tip, a base region disposed adjacent to the attachment, a tip region, and a transition region located between the base region and the tip region, the tip region defining a chord line and the pressure side surface in the tip region being substantially planar; and

wherein the tip region includes a tip rib projecting outwardly from the airfoil, continuously surrounding a leading edge of the airfoil, extending from the leading edge rearwardly along the respective pressure side surface and suction side surface to define a first axial extension on the pressure side surface and a second axial extension on the suction side surface, each of the first and second axial extensions being in a range between 5% and 35% of the length of the chord line immediately downstream of the leading edge.

13. The gas turbine engine as defined in claim 12 wherein the tip rib has a lateral extension in a range between 5% and 30% as thick as the maximum airfoil thickness along the mean camber line.