US20150037161A1

US20150037161A1 - Method for mounting a gas turbine blade in an associated receiving recess of a rotor base body

Info

Publication number: US20150037161A1
Application number: US14/445,847
Authority: US
Inventors: Sebastian Kaltenbach; Armin Kammerer; Alexander Boeck
Original assignee: MTU Aero Engines AG
Current assignee: MTU Aero Engines AG
Priority date: 2013-07-30
Filing date: 2014-07-29
Publication date: 2015-02-05
Also published as: DE102013214933A1; EP2907976A1

Abstract

A gas turbine blade in an associated receiving recess of a rotor base body includes a slot in the blade root and has the shape of a straight cylinder. A blade retaining plate has a first securing region and a second securing region interconnected by a shaft region; the shaft region having the shape of a straight cylinder configured to conform to the slot of the blade root. The shaft region comes to rest on a bottom of the receiving recess of the rotor base body; the first securing region being reshaped to form a first limit stop for a first end wall of the blade root; and the second securing region being reshaped to allow the blade root to be inserted along the insertion direction into the receiving recess. The second securing region of the blade retaining plate is subsequently reshaped to allow the second securing region to form a second limit stop for a second end wall of the blade root facing opposite the first end wall, and the blade root to be secured in position in the receiving recess.

Description

This claims the benefit of German Patent Application DE 10 2013 214 933.6, filed Jul. 30, 2013 and hereby incorporated by reference herein.
The present invention relates to a method for mounting a gas turbine blade in an associated receiving recess of a rotor base body. The present invention also relates to a rotor for a gas turbine, as well as to a gas turbine having such a rotor.

BACKGROUND

It is generally known from the related art to configure a blade retaining plate underneath a blade root of a gas turbine blade between the bottom of a blade root and the bottom of a receiving recess of a rotor base body and to bend the same in a way that secures the gas turbine blade against displacement relative to the rotor base body. A plurality of gas turbine blades are thereby attached in corresponding receiving recesses on the rotor base body until the resulting rotor is completely bladed. The gas turbine blade—rotor base body connection is also referred to as “live rim.” The rotor base body may be configured as a rotor disk or rotor ring, for example. Alternatively, it is known from the German Patent Application DE 1 426 797 A1 to use a resilient securing element having terminal transverse ribs that is configured in a slot in the rotor base body which deepens obliquely toward an end face of the rotor base body to secure the gas turbine blade in position.

SUMMARY OF THE INVENTION

The inherent condition of the known methods and rotors that is to be considered as disadvantageous is that critical stress concentrations can occur in what is generally referred to as the live rim of the rotor during operation of an associated gas turbine in response to a tilting or displacement of the blade retaining plate.
It is an object of the present invention to provide a method for mounting a gas turbine blade in an associated receiving recess of a rotor base body that will make it possible to reliably prevent critical stress concentrations from occurring during later operation of the rotor. It is a further object of the present invention to provide a rotor of this kind, as well as a gas turbine that is at least equipped with such a rotor.
A method for mounting a gas turbine blade in an associated receiving recess of a rotor base body includes at least the steps of providing the gas turbine blade; this including a slot in the blade root thereof that extends along the same and has the shape of a straight cylinder whose longitudinal axis extends in parallel to a direction of insertion of the blade root into the receiving recess; providing a blade retaining plate that has a first securing region and a second securing region that are interconnected by a shaft region; the shaft region having the shape of a straight cylinder that is configured to conform to the slot of the blade root; configuring the blade retaining plate to allow the shaft region to come to rest on a bottom of the receiving recess of the rotor base body; the first securing region being reshaped to form a first limit stop for a first end wall of the blade root; and the second securing region being reshaped to allow the blade root to be inserted along the insertion direction into the receiving recess; the blade root of the gas turbine blade being inserted along the insertion direction into the receiving recess of the rotor base body until the first end wall of the blade root rests against the limit stop formed by the reshaped first securing region; the shaft region of the blade retaining plate being configured within the slot of the blade root and filling the same when the first end wall comes to rest against the limit stop; and the second securing region of the blade retaining plate being reshaped to enable the second securing region to form a second limit stop for the second end wall of the blade root facing opposite the first end wall, and the blade root to be fixed in position in the receiving recess. In other words, in contrast to the related art, it is provided that the blade root of the gas turbine blade features a slot for fully accommodating the shaft region of the blade retaining plate. Both the slot and the shaft region have the shape of a straight cylinder and are configured to conform to one another. In the context of the present invention, a straight cylinder is understood to be a body that has two parallel, plane, congruent base surfaces and one lateral, respectively cylindrical surface that is formed by parallel straight lines; both the parallel straight lines, as well as the longitudinal axis of the cylinder extending orthogonally to the base surfaces. The length of the cylinder, and thus the length of the longitudinal axis thereof is defined by the distance between the two base surfaces. In general, the base surfaces may have any desired plane shape, and have a square, rectangular, circular, elliptical, etc., form, for example. The base surfaces preferably have a square or rectangular shape, so that the shaft region has a parallelepiped, respectively cuboid shape in this case. Thus, both the slot, as well as the shaft region are free of local raised portions, indentations, oblique surfaces or the like. At least during mounting, respectively in the mounted state, both the longitudinal axis of the shaft region, as well as the longitudinal axis of the slot extend in parallel to an insertion direction, respectively an insertion vector of the blade root, into the associated receiving recess, thereby making it readily possible for the blade root to be linearly inserted into the receiving recess that has the blade retaining plate mounted therein. It is understood that it is generally not necessary to reshape the second securing region if it is also possible for the blade root to be inserted along the insertion direction into the receiving recess without deforming the second securing region. Since the shaft region of the blade retaining plate fills the slot of the blade root when the gas turbine blade is mounted, and since the securing regions of the blade retaining plate rest against the mutually opposing end walls of the blade root, the blade retaining plate is located in a defined position within the slot of the blade root, as well as within the receiving recess, making a tilting or displacement of the blade retaining plate relative to the receiving recess or relative to the gas turbine blade impossible during later operation of an associated gas turbine. Thus, this also precludes any critical stress concentrations in the live rim area of the finished rotor. Moreover, the gas turbine blade may be attached to the rotor base body with the aid of one single further component, namely the blade retaining plate. There is no need for any further components. Finally, the blade retaining plate that is accommodated in certain regions of the slot also reduces the leakage cross section in the area of the gas turbine blade—rotor base body connection. The gas turbine blade may generally be disassembled in a reverse sequence.
One advantageous embodiment of the present invention provides that the slot in the blade root of the gas turbine blade be produced by the deposition and/or ablation of blade material. In terms of structural design, this provides a simple method for producing the slot in the blade root. Moreover, this makes it possible to furnish existing gas turbine blades, which do not yet have a slot in the blade root thereof, with a slot, thereby allowing the use thereof in the context of the present invention. When the slot is produced by the deposition of blade material, the blade root is enlarged radially inwardly. Here the advantage is derived that the geometric cross section of passage is advantageously reduced in the slot bottom of the receiving recess, thereby achieving a reduction in leakage flows and a corresponding increased efficiency of the gas turbine during operation of the associated gas turbine. The blade retaining plate is hereby advantageously used both for securing the gas turbine blade in position, as well as for sealing the receiving recess.
Other advantages are derived by rounding at least one edge in the area of the blade root slot prior to placement of the shaft region of the blade retaining plate within the slot. In other words, a gas turbine blade is used that does not have any sharp edges, at least in the slot region. Due to the absence of sharp edges in the area of the blade root and, thus, in the area of the blade retaining plate configured in the slot, material of the blade root, of the blade retaining plate and/or of the receiving recess is reliably prevented from breaking away or tearing.
In a further embodiment of the present invention, a mechanically especially stable connection of the gas turbine blade to the rotor base body may be achieved by providing a gas turbine blade whose blade root has at least one dovetail portion and/or at least one fir tree portion. It may be provided that the blade root of the gas turbine blade feature a plurality of dovetail and/or fir tree portions in the radial direction.
A second aspect of the present invention relates to a rotor for a gas turbine that includes a rotor base body having at least one receiving recess; a blade root of an associated gas turbine blade being configured in the receiving recess and secured in position by a blade retaining plate. It is at least provided in accordance with the present invention that the gas turbine blade include a slot in the blade root thereof that extends along the same and has the shape of a straight cylinder, whose longitudinal axis is disposed in parallel to a direction of insertion of the blade root into the receiving recess. In addition, it is provided that the blade retaining plate have a first securing region and a second securing region that are interconnected by a shaft region; the shaft region having the shape of a straight cylinder that is configured to conform to the slot of the blade root. Finally, the blade retaining plate is configured in the receiving recess in a way that allows the shaft region thereof to be configured within the slot of the blade root and fill the same; the first securing region being reshaped in a way that allows it to form a first limit stop for a first end wall of the blade root; and the second securing region being reshaped in a way that allows the second securing region to form a second limit stop for a second end wall of the blade root facing opposite the first end wall. Critical stress concentrations in the live rim of the rotor blade body are hereby prevented in the case of the rotor according to the present invention since the blade retaining plate is located in a defined position and is not able to tilt within the slot in the blade root. The rotor base body may be configured as a rotor disk or rotor ring, for example. Further advantages resulting herefrom and features thereof will become apparent from the descriptions of the first inventive aspect; advantageous embodiments of the first inventive aspect being considered to be advantageous embodiments of the second inventive aspect and vice versa.
One advantageous embodiment of the present invention provides that the blade root and the shaft region of the blade retaining plate fill the receiving recess of the rotor base body at least virtually completely. A geometrically especially small leakage cross section is hereby provided in the slot bottom of the gas turbine blade - rotor base body connection, thereby advantageously enhancing the efficiency of a gas turbine equipped with the rotor according to the present invention. It is alternatively or additionally provided that the shaft region of the blade retaining plate be designed to conform to the slot of the blade root. In other words, it is provided that the geometry of the shaft region correspond exactly to that of the slot, so that, on the one hand, the shaft region is completely configured within the slot and, on the other hand, also completely fills the same.
Further advantages are derived when the blade root of the gas turbine blade is designed to be edge-free at least in the area of the slot thereof. In other words, it is provided that the blade root not have a sharp-edged form at least in the area of the slot thereof, whereby material is reliably prevented from breaking away or tearing during operation of the rotor.
A third aspect of the present invention relates to a gas turbine, in particular an aircraft engine, having at least one rotor that includes a rotor base body having at least one receiving recess in which a blade root of an associated gas turbine blade is configured and secured in position by a blade retaining plate. The present invention provides that the gas turbine blade be mounted in the receiving recess of the rotor base body using a method in accordance with the first inventive aspect and/or that the rotor be designed in accordance with the second inventive aspect. The features derived herefrom and the advantages thereof are to be inferred from the descriptions of the first, respectively the second inventive aspect.

BRIEF DESCRIPTION OF THE DRAWINGS

Other features of the present invention are derived from the claims, the exemplary embodiment, as well as in light of the drawings. The aforementioned features and combinations of features, as well as the features and combinations of features mentioned in the exemplary embodiment may be used not only in the particular stated combination, but also in other combinations without departing from the scope of the present invention. Specifically,

FIG. 1 shows a schematic section of a gas turbine; one portion of a rotor being discernible;

FIG. 2 shows a schematic cross section of a blade root of a conventional gas turbine blade that is configured in a receiving recess of a rotor base body;

FIG. 3 shows a schematic cross section of a blade root of a gas turbine blade that is mounted in a receiving recess of a rotor base body with the aid of a blade retaining plate that is configured in a slot of the blade root;

FIG. 4 shows a schematic longitudinal section of the blade root and blade retaining plate shown in FIG. 3;

FIG. 5 shows a schematic frontal view of detail V-V shown in FIG. 4;

FIG. 6 shows a schematic frontal view of detail VI shown in FIG. 5; and

FIG. 7 shows a schematic frontal view of detail VII shown in FIG. 6.

DETAILED DESCRIPTION

FIG. 1 shows a schematic section of a gas turbine 10 primarily in the form of an aircraft engine whose basic design is known as such. Discernible is one section of a disk-shaped rotor 12 that includes a rotor base body 16 which is provided with a plurality of conventional gas turbine blades 14. In a manner known per se, each gas turbine blade 14 features a blade leaf 18, a radially subjacent blade platform 20, as well as a radially subjacent blade root 22. An enlarged representation of the area of blade root 22 is illustrated as a schematic cross section in FIG. 2. Discernible is that blade root 22 of conventional gas turbine blade 14 is configured as what is generally referred to as a dovetail root and is located in an associated receiving recess 24 of rotor base body 16. The contour of receiving recess 24 is shaped to substantially conform to the outer contour of blade root 22. The illustrated blade—disk connection is also referred to as “live rim.” Located underneath blade root 22 between blade root 22 and the bottom of receiving recess 24 is a blade retaining plate 26 that is also referred to as “blade retainer.” Discernible underneath and laterally of blade retaining plate are hollow spaces that are clearly indicated by arrows II and that render possible significant leakage flows through receiving recess 24 during operation of gas turbine 10. Due to the hollow spaces, the relative position of blade retaining plate 26 is not uniquely defined, neither with respect to blade root 22, nor with respect to receiving recess 24, and is subject to relatively minor constraints. Therefore, critical stress concentrations may occur in the live rim area of rotor 12 during operation of gas turbine 10 in response to slippage or tilting of blade retaining plate 26. As is also discernible in FIG. 1, guide vane segments 28, which are likewise known per se, are located upstream and downstream of rotor 12 and are sealed by sealing elements 30 against drums 32 that are connected to rotor 12. Sealing elements 30 may, for example, be in the form of honeycomb seals.
FIG. 3 shows a schematic cross section of a blade root 34 of a gas turbine blade 14 where blade root 34 is designed in accordance with the present invention. It is clarified in the following in connection with FIG. 4 which shows a schematic longitudinal section of blade root 34 and blade retaining plate 26 shown in FIG. 3. It is discernible that blade root 34 includes a slot 36 in which a shaft region 38 of blade retaining plate 26 is configured. To form slot 36, blade root 34 is enlarged radially inwardly by material deposition in the region of the slot bottom of the blade—disk connection (live rim). This is symbolized by arrows III. On the one hand, slot 36 is formed by material deposition. On the other hand, lateral hollow spaces II shown in FIG. 2 are sealed, so that merely radially lower hollow space II at least essentially remains as a potential leakage site in the exemplary embodiment shown in FIG. 3.
To assemble gas turbine blade 14 provided with blade root 34 according to the present invention, blade retaining plate 26 is first placed on the bottom of receiving recess 24, allowing shaft region 38 to extend therethrough. Beforehand or subsequently, a first securing region 40 a is bent radially upwards to form a limit stop for a first end wall 42 a of blade root 34. As is discernible in FIG. 4, if the need arises, second securing region 40 b, which is joined via shaft region 38 to first securing region 40 a, is bent slightly radially downwards, to enable blade root 34 to be inserted in the insertion direction symbolized by arrow IV. Blade root 34 of gas turbine blade 14 is thereby inserted along insertion direction IV into receiving recess 24 of rotor base body 16 until first end wall 42 a thereof rests against the limit stop formed by reshaped first securing region 40 a. In this end position of blade root 34, it is discernible that shaft region 38 of blade retaining plate 26 is completely located within slot 36 of blade root 34 and fills the same. As is also discernible in FIG. 4, both slot 36, as well as shaft region 38 are formed to be free of raised portions, indentations, oblique surfaces and the like.
In particular in FIG. 5, which shows a schematic frontal view of detail V-V shown in FIG. 4, one discerns that both slot 36, as well as shaft region 38 each have the shape of a straight cylinder whose longitudinal axis L extends in parallel to insertion direction IV during insertion and in the end position of blade root 34 shown in FIG. 4. In the present case, the straight cylinders are in the form of cuboids, shaft region 38 being designed to conform to slot 36. Once the end or mounting position shown in FIG. 4 is reached, second securing region 40 b of blade retaining plate 26 is also bent radially upwards, analogously to first securing region 40 b, for the final securing in position of blade root 34 in receiving recess 24, so that second securing region 40 b forms a second limit stop for second end wall 42 b of blade root 34, and blade root 34 is consequently secured in position in receiving recess 24. In the illustrated exemplary embodiment, securing regions 40 a, 40 b are configured as T-shaped brackets, respectively retention brackets. Generally, however, other geometries may also be provided. Securing regions 40 a, 40 b may also have different designs.
To prevent damage to blade retaining plate 26 or blade root 34, both during assembly, respectively disassembly, as well as during operation of associated gas turbine 10, blade root 34 is designed to be free of sharp edges and the like, in particular in the area of slot 36 thereof, to prevent material from breaking away or tearing. Sharp edges may form, for example, in the manufacturing of slot 36 at blade root 34. Such sharp edges are, therefore, preferably rounded. To this end, FIG. 6 shows a schematic frontal view of detail VI shown in FIG. 5; while FIG. 7 shows a schematic frontal view of detail VII shown in FIG. 6. It is discernible that blade root 34 is designed to be free of sharp edges, and, to the greatest degree possible, to have round, respectively rounded surfaces and surface transitions.

Claims

What it claimed is:

1. A method for mounting a gas turbine blade in an associated receiving recess of a rotor base body, comprising the following steps:

providing the gas turbine blade having a blade root, the turbine blade including a slot in the blade root, the slot extending along the blade root and having the shape of a straight cylinder whose longitudinal axis extends in parallel to a direction of insertion of the blade root into the receiving recess;

providing a blade retaining plate having a first securing region and a second securing region interconnected by a shaft region, the shaft region having the shape of a further straight cylinder configured to conform to the slot of the blade root;

configuring the blade retaining plate to allow the shaft region to come to rest on a bottom of the receiving recess of the rotor base body, the first securing region being reshaped to form a first limit stop for a first end wall of the blade root and the second securing region being reshaped to allow the blade root to be inserted along the insertion direction into the receiving recess;

inserting the blade root of the gas turbine blade along the insertion direction into the receiving recess of the rotor base body until the first end wall of the blade root rests against the limit stop formed by the reshaped first securing region, the shaft region of the blade retaining plate being configured within the slot of the blade root and filling the same when the first end wall comes to rest against the limit stop; and

reshaping the second securing region of the blade retaining plate to allow the second securing region to form a second limit stop for a second end wall of the blade root facing opposite the first end wall and the blade root to be secured in position in the receiving recess.

2. The method as recited in claim 1 wherein the slot in the blade root of the gas turbine blade is produced by the deposition and/or ablation of blade material.

3. The method as recited in claim 1 wherein at least one edge in the area of the slot of the blade root is rounded prior to placement of the shaft region of the blade retaining plate within the slot.

4. The method as recited in claim 1 wherein the blade root has at least one dovetail portion or at least one fir tree portion.

5. A rotor for a gas turbine comprising:

a rotor base body including at least one receiving recess;

a gas turbine blade having a blade root configured and secured in position in the at least one receiving recess by a blade retaining plate, the gas turbine blade including a slot in the blade root extending along the blade root and having the shape of a straight cylinder whose longitudinal axis extends in parallel to a direction of insertion of the blade root into the receiving recess;

the blade retaining plate having a first securing region and a second securing region interconnected by a shaft region, the shaft region having the shape of a further straight cylinder configured to conform to the slot of the blade root, the blade retaining plate being configured in the receiving recess in such a way that the shaft region thereof is configured within the slot of the blade root and fills the slot, the first securing region being reshaped to form a first limit stop for a first end wall of the blade root, and the second securing region being reshaped to allow the second securing region to form a second limit stop for a second end wall of the blade root facing opposite the first end wall.

6. The rotor as recited in claim 5 wherein the blade root and the shaft region of the blade retaining plate fill the receiving recess of the rotor base body completely, or the shaft region of the blade retaining plate is designed to conform to the slot of the blade root.

7. The rotor as recited in claim 5 wherein the blade root of the gas turbine blade is designed to be edge-free at least in the area of the slot thereof.

8. A gas turbine comprising the rotor as recited in claim 5.

9. An aircraft engine comprising the gas turbine as recited in claim 8.

10. A gas turbine comprising at least one rotor including a rotor base body having at least one receiving recess in which a blade root of an associated gas turbine blade (14) is configured and secured in position by a blade retaining plate, wherein the gas turbine blade is mounted in the receiving recess of the rotor base body according to the method recited in claim 1.

11. An aircraft engine comprising the gas turbine as recited in claim 10.