Angled cooling divider wall in blade attachment

A rotor blade of a gas turbine engine includes a blade root defining a cooling airflow entry cavity therein, in fluid communication with internal cooling air passages through the blade. The cavity includes opposed side walls and at least one divider wall extending therebetween. At least one end portion of the divider walls adjoins one of the side walls in an angled direction relative to a perpendicular direction of the side walls.

TECHNICAL FIELD

The invention relates generally to gas turbine engines, and more particularly to a cooled turbine rotor assembly.

BACKGROUND OF THE ART

A conventional gas turbine engine includes various rotor blades in the fan, compressor, and turbine sectors thereof, which are removably mounted to respective rotor discs. Each of the rotor blades includes a blade root at the radially inner end thereof. Each of the blade roots conventionally includes one or more pairs of lobes which can axially slide into and be retained in one of a plurality of axially extending attachment slots in the periphery of the rotor disc, thereby forming the attachment of the rotor blade. In a cooled turbine rotor assembly, the attachment or blade root of each rotor blade defines a cooling air entry cavity therein for receiving cooling air and bringing cooling air into the airfoil of the rotor blade for cooling same. In order to maintain the structural stiffness of the attachment, a given number of divider walls or ribs extending within the cavity is usually required because a centrifugal load which is born by the blade attachment, is generated as the blade rotates around the main engine axis. Nevertheless, conventional divider walls or ribs have limited effect. The centrifugal load generated by the high rotational speed of the rotor assembly results in not only large compressive stresses on the ribs, but also buckling and shear effects which can initiate cracks in the blade attachment structure.

Accordingly, there is a need to provide an improved blade root structure for cooled turbine rotor assemblies of gas turbine engines in order to meet the demanding requirements of various aspects of high efficiency gas turbine engines.

SUMMARY OF THE INVENTION

It is therefore an object of the present invention to provide an improved blade attachment structure for a rotor assembly of a gas turbine engine.

In one aspect, the present invention provides a rotor blade having internal cooling air passages for a gas turbine engine, which comprises an airfoil section defining the cooling air passages therethrough, a blade root having at least one side projection on each of opposed sides thereof extending between leading and trailing ends of the blade root, and platform segments extending laterally from opposed sides of the airfoil section. The blade root defines a cavity therein with an opening thereof in a bottom of the blade root. The cavity is in fluid communication with the cooling air passages through the airfoil section. The cavity includes opposed side walls substantially parallel to a main longitudinal axis of the blade root and at least one divider wall extending from the opening inwardly into the cavity and extending between the side walls. At least one end portion of the divider wall adjoins one of the side walls in an angled direction relative to a perpendicular direction of the side walls.

In another aspect, the present invention provides a turbine rotor assembly for a gas turbine engine, which comprises a rotor disc defining a plurality of attachment slots circumferentially spaced apart one from another and extending axially through a periphery thereof, and an array of rotor blades extending outwardly from the periphery of the rotor disc. Each of the rotor blades includes an airfoil section defining internal cooling air passages therethrough, a blade root affixed within the attachment slots of the rotor disc, and platform segments extending laterally from sides of the airfoil section into opposing relationship with corresponding platform segments of adjacent rotor blades. Each of the blade roots defines a cavity therein with an opening thereof in a bottom of the blade root and in fluid communication with the cooling air passages, and includes means defined within the cavity for reducing a torsion effect on the blade root resulting from a rotational speed of the turbine rotor assembly during engine operation, thereby stiffening the blade root.

In another aspect, the present invention provides a turbine rotor assembly for a gas turbine engine, which comprises a rotor disc and an array of rotor blades extending outwardly from a periphery of the rotor disc. The rotor disc defines a plurality of attachment slots circumferentially spaced apart one from another. Each of the attachment slots together with at least one pair of side recesses in respective side walls of the attachment slot, extends axially and circumferentially through the periphery thereof. Each of the rotor blades includes an airfoil section defining internal cooling air passages therethrough, a blade root having at least one side projection on each of opposed sides thereof affixed in one attachment slot of the rotor disc, and platform segments extending laterally from sides of the airfoil section into opposing relationship with corresponding platform segments of adjacent rotor blades. Each of the blade roots further defines a cavity therein with an opening in a bottom of the blade root and in fluid communication with the cooling air passages. The cavity includes opposed side walls substantially parallel to the attachment slot receiving the blade root. At least one divider wall extends from the opening inwardly into the cavity and extends between the side walls in an angled direction relative to a perpendicular direction of the side walls.

Further details of these and other aspects of the present invention will be apparent from the detailed description and figures included below.

DETAILED DESCRIPTION OF THE PREFERRED EMBODIMENTS

A turbofan engine illustrated schematically inFIG. 1, presented as an example of the application of the present invention, includes a housing or nacelle10, a low pressure spool assembly seen generally at12which includes a fan14, a low pressure compressor16and a low pressure turbine18, a high pressure spool assembly seen generally at20which includes a high pressure compressor22and a high pressure turbine24. A core casing28surrounds the low and high pressure spool assemblies12and20to define a main fluid path (not indicated) therethrough. In the main fluid path there is provide a combustor seen generally at26with fuel injecting means (not indicated) to constitute a gas generator section. The compressors16and22drive a main airflow (not indicated) along the main fluid path and provide bleed airflow as a cooling air source for cooling the combustor26as well as the turbines18and24.

It should be noted that similar components of the different embodiments shown in the accompanying Figures are indicated by similar numerals for convenience of description of the present invention. Only those components different in one embodiment from the other will be separately described with reference to additional numerals.

Referring toFIGS. 1 and 2, a rotor assembly, for example, a turbine rotor assembly24aof the high pressure turbine24is described herein according to one embodiment of the present invention. The turbine rotor assembly24aincludes a turbine rotor disc40mounted to a rotating shaft (not indicated) of the high pressure spool assembly20and is rotatable about a main longitudinal axis30of the engine. An array of rotor blades33(only one shown inFIG. 2) extend outwardly from a periphery of the turbine rotor disc40. Each of the rotor blades33includes an airfoil section42, a root section44and platform segments46extending laterally from opposed sides of the airfoil section42into opposing relationship with corresponding platform segments of adjacent rotor blades (not shown).

The rotor assembly24ais now described in greater detail with reference toFIGS. 1-6. The turbine rotor disc40includes a web section48extending radially outwardly from a hub (not shown) mounted to the rotating shaft (not indicated) of the high pressure spool assembly20ofFIG. 1, and a rim section50extending radially outwardly from the web section48. Rim section50has an axial thickness defined between a front face52and a rear face54.

The root section44of each turbine rotor blade33includes at least one projection on each of opposed sides thereof, which in this embodiment are, for example, formed by a series of lobes56,58and60, having decreasing circumferential widths from the radially outermost lobe56to the radially innermost lobe60, with the radially central lobe58disposed therebetween and having an intermediate lobe width (SeeFIG. 4). The root section44of such a multi-lobed. type is often referred to as a firtree, because of this characteristic shape. The root section44is adapted for attachment to the rotor disc40and therefore is generally referred to as a blade attachment.

For aerodynamic benefits, each of the blades33is preferably positioned in an angled direction relative to the main axis30of the engine. The angle between the angled direction of the blade33and the main axis30of the engine is referred to as a broach angle B hereinafter throughout the description and appended claims of this application. Therefore, a main longitudinal axis32of the root section44extends in a direction of a broach angle B relative to the main axis30of the engine. The at least one projection on each of opposed sides of the root section44or the lobes56,58and60, extend between leading and trailing ends34,36of the root section44and are substantially parallel to the main longitudinal axis32of the root section44.

The turbine rotor disc40further includes a plurality of attachment slots62(only one shown) circumferentially spaced apart one from another and extending axially and circumferentially in an angled direction of the broach angle B relative to the main axis30of the engine, through the periphery of the turbine rotor disc40which is the entire axial thickness of the rim50in this embodiment. Each axial attachment slot62includes a series of axial recesses or fillets56a,58aand60adefined in opposed side walls (not indicated) of attachment slot62, which substantially conform in both shape and direction to the firtree of root section44, so as to form abutting returning surfaces of the respective root section44and attachment slot62for retaining rotor blade33in the turbine rotor assembly24aunder the high temperature, high stress environment of the rotating turbine. The abutting retaining forces will be further described in detail hereinafter.

The turbine rotor blade33preferably further includes internal cooling airflow passages which are not shown but are indicated by broken line arrows76, for directing pressurized cooling airflow through the airfoil section42of the turbine rotor blades33, and discharging same through a plurality of openings78on the trailing edge80of the airfoil section42, into the gas path, and/or through a plurality of openings called film holes/slots (seeFIG. 3, not indicated) on the airfoil side walls (not indicated) at the airfoil section42into the gas path. In particular, the root section44defines a cooling air entry cavity64therein with an opening (not indicated) thereof in a bottom68of the root section44and in fluid communication with the cooling airflow passages76through the airfoil section42. The cavity64includes opposed side walls63substantially parallel to the main longitudinal axis32of the root section44. The cavity64further includes at least one divider wall, but preferably a plurality of divider walls66extending from the opening inwardly into the cavity64and extending between the side walls63. At least one end portion of the divider wall66adjoins one of the side walls63in an angled direction relative to a perpendicular direction of the side walls63. The angled direction is indicated by the angle A inFIG. 6. In this embodiment, the entire divider wall66extends between the opposed side walls63in the angled direction indicated by the angle A. Therefore, the cooling air entry cavity64receives a pressurized cooling airflow (indicated by arrow84) which is delivered from a pressurized cooling air source (not shown) such as bleed air from compressor assembly16or22. The cooling airflow84is guided between a front cover plate74and turbine rotor disc40, and is then directed into a space72defined between the bottom68of the root section44and a bottom70of the attachment slot62. The cooling air entry cavity64further directs the received cooling airflow through the passages defined by the divider walls66therein into the inner cooling airflow passages76within the airfoil section42to cool the rotor blade33.

Particularly referring toFIGS. 4-6, the centrifugal load indicated by arrow81, on the root section44, caused by the rotational speed of the rotor assembly24aofFIG. 2, produces retaining forces on the projections or the lobes56,58and60of the root section44, as indicated by arrows82. The retaining forces82include radial components which counter the centrifugal load81, and circumferential components resulting in high compressive stresses on the root section44. Therefore, divider walls66are used within the cavity64against the compressive stresses, in order to stiffen the hollow structure of the root section44.

Nevertheless, the centrifugal load81caused by the rotational speed of the rotor assembly24ofFIG. 2not only causes compressive stresses by traction effect but also further results in a torsion effect on the root section44, as indicated by a pair of arrows86inFIG. 6, which can create shear effects on a conventional divider wall (as shown by broken lines inFIG. 6) which extends between the opposed side walls63of the cavity64in a perpendicular direction. A wall or rib reinforcing element can effectively bear a compressive load but is much less effective for bearing a torsional load. The broaching angle of the root section44will further increase the torsion effect caused by the rotational speed of the rotor assembly24a(seeFIG. 2). Therefore, in accordance with one embodiment of the present invention, the divider wall66extends in an angled direction relative to the perpendicular direction of the side walls63of the cavity64. Thus, the torsion effect caused by the rotational speed of the rotor assembly24aofFIG. 2creates more compressive load on the divider wall66and less shear loads on same, in contrast to the torsion effect on the conventional divider wall perpendicularly adjoining the side walls63of the cavity64, thereby more effectively stiffening the hollow structure of the root secton44.

FIG. 7illustrates a further embodiment of the present invention in which the divider wall66ais itself angled, in contrast to the flat configuration of the divider wall66ofFIG. 6. The divider wall66ahas only one end portion adjoining one of the side surfaces63in the angled direction of angle A relative to the perpendicular direction of the side walls63. The other end of the divider wall66aadjoins the other of the side walls63perpendicularly.

FIG. 8illustrates a still further embodiment of the present invention, in which the divider wall66bhas a first end portion thereof adjoining one of the side walls63in a first angled direction of angle A1and a second end portion thereof adjoining the other of the side walls63in a second angled direction of angle A2, relative to the perpendicular direction of the side walls63. As an example of the present invention, the illustrated divider wall66bincludes a middle portion (not indicated) interconnecting the first and second end portions of the divider wall66b, the angles A1and A2being different. Alternatively, angles A1and A2of respective first and second end portions of divider wall66b, could be equal. Also alternatively, angles A1and A2could be different, with the first second end portions connected to each other directly, without a middle portion therebetween.

Referring now toFIGS. 2 and 3, as a secondary effect, the angled orientation of the divider wall66is beneficial to the cooling flow84into the internal cooling passages76of the airfoil, by reducing pressure loss when the cooling air flow84enters the cooling air entry cavity64. With a more appropriate feed pressure, blade durability can be improved or the required quantity of airfoil cooling airflow can be reduced, thereby improving engine performance.

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 department from the scope of the invention disclosed. For example, the divider walls can be configured differently from those described, provided that at least one end portion thereof adjoins one of the side walls of the cavity in angled direction relative to the perpendicular direction of the side walls of the cavity. Still other modifications which fall within the scope of the present invention 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.