Turbine blade

A front engagement face 27f of a front engagement member 27 and a rear engagement face 29f of a rear engagement member 29 are respectively configured to be located a little back from a virtual plane VF coplanar with one side of a platform 15 and also to be substantially parallel to the axial direction of a male dovetail 25, and an end face of a front wall Wf integrally molded in the vicinity of a base portion of a front seal fin 21 and an end face of a rear wall Wr integrally molded in the vicinity of a base portion of a rear seal fin 23 are respectively configured to be coplanar with the virtual plane Vf.

CROSS REFERENCE TO RELATED APPLICATION

This application claims benefit of priority under 35USC § 119 to Japanese Patent Application No. 2003-159175, filed on Jun. 4, 2003, the entire contents of which are incorporated by reference herein.

BACKGROUND OF THE INVENTION

1. Field of the Invention

The present invention relates to a turbine blade to be installed into a female dovetail of a turbine disk of an aircraft engine.

2. Description of the Related Art

The constitution of a typical turbine blade to be installed into a female dovetail of a turbine disk of an aircraft will be described below.

A typical turbine blade includes a blade airfoil as a blade base, one side of the blade airfoil being a convex suction surface and the other side of the blade being a concave pressure surface. A platform is integrally molded on the hub side (at the base end portion) of the blade and recesses are formed respectively on both sides of the platform. A front seal fin protruding forward is formed at the front end of the platform and a rear seal fin protruding backward is formed at the back end of the platform.

A male dovetail is integrally disposed on the hub side (at the base end portion) of the platform, the dovetail has an engagement portion able to engage with a female dovetail of a turbine disk, and the engagement portion is usually formed by grinding, whereby, a jig is used for grinding, and one side of the platform can be engaged against a platform-locating portion of the jig.

The manufacturing process of the typical turbine blade will be described below. The greater part of the turbine blade with the engagement portion remaining unfinished (an unfinished turbine blade) is molded by casting. Next, the unfinished turbine blade is located in the jig so as to make the dovetail axial direction perpendicular to the repulsive force due to work-resistance during grinding, by letting the pressure surface of the blade airfoil be supported with a support portion of the jig. Further, setting of the unfinished turbine blade onto the jig is completed by pressing the pressure surface of the blade airfoil against the location portion by means of a clamp of the jig. Then, the turbine blade is finished by forming the engagement portion along the dovetail axial direction by grinding.

SUMMARY OF THE INVENTION

In order for the blade airfoil to be disposed obliquely against the engine axial direction of the aircraft engine, both sides of the platform are angled to the dovetail axial direction of the engagement portion. Consequently, during formation of the engagement portion along the dovetail axial direction by grinding, a component of a force that may cause a displacement of the blade airfoil from the jig is generated. Therefore, there is a problem in that a machining tolerance of the engagement portion is degraded, lowering the quality of the turbine blade, because the unfinished turbine blade is displaced from the jig owing to an increase in the magnitude of the component of a force during grinding.

According to the present invention the unfinished turbine blade is never displaced from the jig, forming an engagement portion with tight machining tolerance, thus enhancing the quality of the turbine blade.

According to a first technical aspect of the present invention, a turbine blade to be installed into an engaged member of a turbine disk of an aircraft engine is characterized in that it comprises a blade airfoil, one side of which having a convex suction surface and the other side having a concave pressure surface; a platform integrally molded on the hub side of the blade wherein a recess is formed on one side of the platform; a front seal fin formed protruding forward at the front end of the platform; and a rear seal fin formed protruding backward at the back end of the platform, an engagement member integrally molded on the hub side of the platform wherein the engagement member has a engagement face which is able to be engaged with the engaged member of turbine disk and is formed by grinding, a front engagement member integrally molded in the vicinity of a base portion of the front seal fin wherein the front engagement member has a front engagement face able to engage with a front locating portion of a jig to be used for the grinding, and the front engagement face located back from a virtual plane including one side of the platform, a front wall integrally molded in the vicinity of the base portion of the front seal fin wherein the front wall surrounds a front side-edge portion of the front engagement member, a rear engagement member integrally molded in the vicinity of a base portion of the rear seal fin wherein the rear engagement member has a rear engagement face able to engage with a rear locating portion of the jig, and the rear engagement face located back from the virtual plane, and a rear wall integrally molded in the vicinity of the base portion of the rear seal fin, where the rear wall surrounds a rear side-edge portion of the rear engagement member, wherein an end face of the front wall and an end face of the rear wall are respectively configured to be coplanar with the virtual plane.

According to a second technical aspect of the present invention, the turbine blade is characterized in that the front engagement face and the rear engagement face are respectively configured to be substantially parallel to the longitudinal direction of the engagement member.

According to a third technical aspect of the present invention, the turbine blade is further characterized in that the spacing between the front edge of the front engagement face and the rear edge of the rear engagement face are configured to be longer than the longitudinal length of the engagement member.

According to a fourth technical aspect of the present invention, the turbine blade is further characterized in that the depth that the front engagement face is located back from the virtual plane and the depth that the rear engagement face is located back from the virtual plane are respectively configured to be in a range of less than or equal to 0.7 mm.

DESCRIPTION OF THE PREFERRED EMBODIMENT

An embodiment of the present invention will be described below referring toFIG. 1toFIG. 5.FIG. 1shows a turbine blade according to an embodiment of the present invention;FIG. 2is an enlarged view of the arrowed portion II inFIG. 1;FIG. 3shows a state where the turbine blade according to the embodiment of the present invention is set on a jig;FIG. 4is a schematic view of the arrowed portion IV inFIG. 3; andFIG. 5shows a state where the turbine blade according to the embodiment of the present invention is installed into a female dovetail of a turbine disk. Herein, “front and rear (or back)” refers to the right hand side and left hand side inFIG. 1andFIG. 2, and refers to the left hand side and right hand side inFIG. 4.

As shown inFIGS. 1,2and5, the turbine blade1relating to the embodiment of the present invention is one to be installed into a female dovetail5of a turbine disk3of a low-pressure turbine for an aircraft engine and comprises a blade airfoil7as a main body of the turbine blade1. One side (the front side inFIG. 1) of the blade airfoil7is a convex suction surface7faand the other side (the back side inFIG. 1) of the blade airfoil7is a concave pressure surface7fb.

A shroud9is integrally molded on the tip side (the outer end portion, the upside inFIG. 1) of the blade airfoil7and the shroud9has a couple of seal fins11,13.

A platform15is integrally molded on the hub side (the inner end portion, the downside inFIG. 1) of the blade airfoil7, and recesses17,19each having a face are formed respectively on both sides (the one side and the other side) the platform15. Moreover, a front seal fin21protruding forward is formed at the front end of the platform, and a rear seal fin23protruding backward is formed at the back end of the platform15. A so-called shank portion is also included in the platform.

Further, a male dovetail25as an engagement member is integrally molded on the hub side of the platform15and the male dovetail25has an engagement groove (an engagement face)25swhich is able to be engaged with an engaged protrusion (an engaged portion)5bof the female dovetail5as an engaged member, and the engagement groove25sis formed by grinding.

A front engagement member27is integrally molded within the recess17in the vicinity of a base portion of the front seal fin21and the front engagement member27has a planar front engagement face27f. Further, as shown with diagonal lines inFIG. 2, a thin front semi-circular shaped wall Wf surrounding a front side-edge portion of the front engagement member27is integrally molded in the vicinity of the base portion of the front seal fin21.

Moreover, a rear engagement member29is integrally molded within the recess17in the vicinity of a base portion of the rear seal fin23and the rear engagement member29has a planar rear engagement face29f. Further, as shown with diagonal lines inFIG. 2, a thin rear wall semi-circular shaped Wr, which surrounds a rear side-edge portion of the rear engagement member29, is integrally molded in the vicinity of the base portion of the rear seal fin23.

As shown inFIG. 3andFIG. 4, the front engagement face27fof the front engagement member27is able to be engaged by a front locator pin33of a jig31to be used for the grinding, and the rear engagement face29fof the rear engagement member29is able to be engaged against a rear locator pin35of the jig31. Moreover, the front engagement face27fof the front engagement member27and the rear engagement face29fof the rear engagement member29are respectively configured to be located slightly back from a virtual plane VF including one side of the platform15and also to be substantially parallel to the dovetail axial direction (longitudinal direction) of the male dovetail25.

The front engagement face27fis offset forward from the face of the recess toward the virtual plane VF. The front engagement face27fis located in a first plane positioned back from the virtual plane VF. The rear engagement face29fis located in a second plane, different from the first plane, back from the virtual plane VF and offset forward from the face of the recess toward the virtual plane VF. Particularly, the distance (depth) of which the front engagement face27fis located back from the virtual plane VF and the distance (depth) of which the rear engagement face29fis located back from the virtual plane VF are respectively configured to be in a range of less than or equal to 0.7 mm. In other words, each of the front engagement face27fof the front engagement member27and the rear engagement face29fof the rear engagement member29has a recess in a range of less than or equal to 0.7 mm. Further, the spacing between the front edge of the front engagement face27fand the rear edge of the rear engagement face29fare configured to be longer than the length of the male dovetail25in the dovetail axial direction.

An end face of the front wall Wf and an end face of the rear wall Wr are respectively configured to be coplanar with the virtual plane VF.

The jig31includes the front locator pin33and a rear locator pin35as well as a locating roller37for locating the suction surface7fanear the tip of the blade airfoil7, a locating pin41for clipping the male dovetail25at the rear side thereof and a clip39for clipping the male dovetail25at the front side thereof, and a clamp45for pressing the pressure surface7fbnear the hub of the blade airfoil7downward via a rubber pad43, an engagement roller47able to be engaged against the back end of the shroud9, a contact bolt49able to be contacted with the front end of the shroud9.

The operation (mainly manufacturing of the turbine blade1) of an embodiment of the present invention will be described below. The greater part of the turbine blade1, with machining portions of the engagement groove25sand the shroud9remaining unfinished (an unfinished turbine blade1′), is molded by casting. Since the vicinity of base portion of the front seal fin21is configured to be thicker in consideration of the strength of the fin, the front engagement member27and the front wall Wf are molded utilizing the thicker portion. Also, since the vicinity of base portion of the rear seal fin23is configured to be thicker in consideration of the strength of the fin, the rear engagement member29and the rear wall Wr are molded utilizing the thicker portion.

Since each of the front engagement face27fand the rear engagement face29fhas a recess in a range of less than or equal to 0.7 mm, casting defects do not easily occur in the vicinity of the front engagement face27fand the rear engagement face29f. After molding the greater part of the turbine blade1, the machining portion of the shroud9is operated by appropriate machining.

Next, the front engagement face27fof the front engagement member27and the rear engagement face29fof the rear engagement member29are engaged by the front locator pin33of the jig31and the rear locator pin35of the jig31respectively, and the suction surface7faof the blade airfoil7is made to be located with the locating roller37of the jig31. Thereby, the unfinished turbine blade1′ can be located in the jig31so as to make the dovetail axial direction perpendicular to the repulsive force due to work-resistance during grinding. Besides, location of the unfinished turbine blade1′ relative to the jig31in the machining direction is also performed by engaging the back end of the shroud9with the engagement roller47of the jig31to make the contact bolt49contact with the front end of the shroud9.

Further, the male dovetail25is located at the rear end of the dovetail25with a locating pin41of the jig31, and clipped with a clip39at the front end of the dovetail25, and the pressure surface7fbnear the hub of the blade airfoil7is pressed downward with a clamp45of the jig31via a rubber pad43. Thereby, the setting of the unfinished turbine blade1′ onto the jig31is completed. Since the spacing between the front edge of the front engagement face27fand the rear edge of the rear engagement face29fhas been configured to be longer than the longitudinal length of the engagement member, the loaded state of the unfinished turbine blade1′ on the jig31is further stabilized.

Then, the manufacturing of the turbine blade1is finished by forming the engagement groove25salong the dovetail axial direction by the grinding. Since the front engagement face27fof the front engagement member27and the rear engagement face29fof the rear engagement member29are respectively configured to be substantially parallel to the dovetail axial direction, only the repulsive force, which is due to work resistance and is perpendicular to the dovetail axial direction, will occur on the front engagement face27fand the rear engagement face29fin a case where the engagement groove25sis formed along the dovetail axial direction by grinding. Therefore, substantially no repulsive force, which displaces the longitudinal direction of the dovetail, will occur.

In addition to the operation described above, the front engagement face27fof the front engagement member27and the rear engagement face29fof the rear engagement member29are respectively configured to be located slightly back from the virtual plane VF including one side of the platform15. And also, the end face of the front wall Wf and the end face of the rear wall Wr are respectively configured to be coplanar with the virtual plane VF. Therefore, the spacing between adjacent turbine blades1will not be widened locally when a number of turbine blades1are installed into the turbine disk3.

According to the embodiment of the present invention, since only the repulsive force being perpendicular to the dovetail axial direction will occur on the front engagement face27fand the rear engagement face29fin a case where the engagement groove25sis formed along the dovetail axial direction by grinding, enhancement in quality of the turbine blade1can be achieved because the unfinished turbine blade1′ will not be displaced from the jig31during grinding, thus preventing degradation of the machining tolerance of the engagement groove25s. Particularly, since the set state of the unfinished turbine blade1′ is further stabilized, a further improvement in quality of the turbine blade1can be achieved by enhancing the machining tolerance of the engagement groove25s.

According to the embodiment of the present invention, the front engagement member27and the front wall Wf are molded utilizing the thicker portion which is configured to be thicker in consideration to the strength of the front seal fin21, and the rear engagement member29and the rear wall Wr are molded utilizing the thicker portion which is configured to be thicker in consideration to the strength of the rear seal fin23. Therefore, the addition of the front engagement member27, the front wall Wf, the rear engagement member29and the rear wall Wr to the components of the turbine blade1does not cause any increase in a weight of the turbine blade1.

Further, since casting defects can do not easily occur in the vicinity of the front engagement face27fand in the vicinity of the rear engagement face29f, the occurrence of rejects of the turbine blade1can be reduced.

Moreover, since the spacing between adjacent turbine blades1is not be extended locally when a number of turbine blades1are installed into the turbine disk3, a flow of high-temperature gas from the main stream toward the center of the engine can be prevented during the operation of the aircraft engine to increase the engine efficiency of the aircraft engine and to avoid temperature increase of the turbine disk3.

Further, the present invention should not be limited to the description of the above embodiment of the invention, but it can be applicable in various modes through causing the appropriate conversion thereof, for example, an application of the turbine blade1to a turbine blade for a high-pressure turbine of the aircraft engine, etc.

Modifications and variations of the embodiments described above will occur to those skilled in the art, in light of the teachings. The scope of the invention is defined with reference to the following claims.