US20190376392A1

US20190376392A1 - Gas turbine

Info

Publication number: US20190376392A1
Application number: US16/462,466
Authority: US
Inventors: Karl Schreiber
Original assignee: Rolls Royce Deutschland Ltd and Co KG
Current assignee: Rolls Royce Deutschland Ltd and Co KG
Priority date: 2016-12-01
Filing date: 2017-12-01
Publication date: 2019-12-12
Also published as: WO2018100160A1; EP3548704A1; DE102016123248A1

Abstract

A gas turbine having at least one disk, wherein turbine blade elements are connected via connection means to the at least one disk, wherein the connection means are arranged in the interior of the turbine blade elements in a region radially above the disk in the radial direction of the turbine blade elements, in particular in a region which is in the driving airflow during the operation of the gas turbine, and the turbine blade elements have at least two zones composed of different materials, wherein the at least two zones adjoin one another in particular in the radial direction, and in that a zone with a material suited to compressive stress, in particular a ceramic, in particular an yttrium-stabilized zirconium oxide, is arranged radially below the connection means, and a zone with a material suited to tensile stress, in particular CMSX 4, is arranged radially above the connection means.

Description

The invention relates to a gas turbine having the features of claim 1.
Gas turbines, such as aircraft engines or stationary gas turbines, are units subject to high thermal and mechanical loads.
The efficiency of a gas turbine is greatly affected by the thermal and mechanical loading capacity of the gas turbine. There is a known practice in the prior art, e.g. WO 2012 160819 A1, of connecting the turbine blades positively to a turbine disk by means of a “firtree root”.
This positive joint requires a considerable volume of material and has a considerable effect on the weight and loading capacity of the gas turbine. In a comparison with compressor designs, it is possible to estimate that - as compared with an integral blisk design—about 30% of the total weight of a stage is required for the positive joint. Apart from use in small turboshaft engines, a blisk design is not worthwhile and also not common, owing to very large differences in the requirements as regards materials for the turbine blade and the turbine disk.
The object is therefore to adapt gas turbines to the special conditions of use.
This object is achieved by a gas turbine having the features of claim 1.
The gas turbine has at least one disk, wherein turbine blade elements are connected via connection means to the at least one disk.
The connection means are arranged in the interior of the turbine blade elements, wherein the turbine blade elements are arranged in a region radially above the disk in the radial direction, in particular are arranged in a region which is in the driving airflow during the operation of the gas turbine.
The turbine blade elements have at least two zones composed of different materials, wherein the at least two zones adjoin one another in particular in the radial direction, and a zone with a material suited to compressive stress, in particular a ceramic, in particular an yttrium-stabilized zirconium oxide, is arranged radially below the connection means, and a zone with a material suited to tensile stress, in particular CMSX 4, is arranged radially above the connection means.
This enables the materials to be selected to match the loads. In this case, at least one parting line can be arranged between the at least two zones of different material in the turbine blade elements radially below the connection means.
Thus, the connection means are situated, in particular in a region of the turbine blade elements which is exposed to the hot gas flow, i.e. in the aerodynamically effective region (airfoil region) of the turbine blade elements. Thus, the connection means are arranged in a region in which comparatively small masses have to be transferred.
In this case, the cores for the turbine blade elements can be manufactured in a materially integral manner from the disk, for example. Only in a radial region which is in the hot gas flow during operation is the positive connection made to the surrounding airfoil region. This leads to an approximately 70% reduction in the mass which has to be transferred via the positive connection and to a considerable reduction in stress in the airfoil region. Tensile stresses occur only in the blade region situated radially on the outside of the positive connection. The region situated radially below is subject to compressive stresses.
Another aspect of such a construction is improved clearance control (i.e. the clearance being the distance between the blade tip and the surrounding housing) by virtue of lower thermal and elastic expansion inherent in the design.
In one embodiment, the connection means are arranged radially on the inside, radially in the center or radially on the outside in the region radially above the disk, in particular in the region of the turbine blade elements which is in the driving airflow during the operation of the gas turbine.
Fundamentally, it is possible for the connection means to be designed for positive joining, nonpositive joining and/or material joining.
To form the positive connection means, it is possible in some embodiments to use positive joining means, in particular shoulders, projections and/or undercuts for the axial and/or radial fixing of the turbine blade elements.
Nonpositive connection means can have a wedge connection, in particular in the disk, and/or shrink-fit joints to achieve frictional joining.
Material connection means can have a laser-welded joint between the turbine blade elements and the disk.
It is also possible to have wedge elements for clamping the disk to the turbine blade elements and/or for producing positive joining between the disk and the turbine blade elements. The material of the core is then more elastic, for example, than the material of the wedging means. The wedging means presses the core against the inner side of the turbine blade elements, for example.
To connect the turbine blade elements to the cores, a welded joint can be arranged radially on the outside.
The turbine blade elements can consist of two parts, wherein the positive connections are established only when the parts are assembled.
The gas turbines can be designed as an aircraft engine, as a vehicle propulsion system, as a ship's propulsion system or as a stationary gas turbine.
Here, the turbine blade elements or the disk can be especially configured and designed for use as claimed in at least one of claims 1 to 15.

The invention will be discussed in connection with the exemplary embodiments illustrated in the figures. In the figures:

FIG. 1 shows a schematic illustration of a gas turbine, in this case an aircraft turbine;

FIG. 2A shows a horizontal section through a turbine blade element;

FIG. 2B shows a section through the turbine blade element according to FIG. 2A along the line A-A;

FIG. 3A shows an alternative embodiment of the turbine blade element with a wedging means for the installation of a turbine blade element;

FIG. 3B shows the embodiment according to FIG. 3A with a driven-in wedging means for fixing the turbine blade element by means of positive joining;

FIG. 4A shows an alternative embodiment of the turbine blade element with a disk without a wedging means;

FIG. 4B shows an alternative embodiment of the turbine blade element with a disk having a wedging means for clamping the turbine blade element to the disk.

As illustrated in FIG. 1, the individual components of the gas turbine 100 are arranged in series along a rotational axis or central axis M, wherein the gas turbine 100 is designed as a turbofan engine. At an inlet or intake E of the gas turbine 100, air is drawn in along an inlet direction R by means of a fan F. This fan F, which is arranged in a fan casing FC, is driven by means of a rotor shaft S, to which rotation is imparted by a turbine TT of the gas turbine 100. In this arrangement, the turbine TT is adjacent to a compressor V, which comprises a low pressure compressor 11 and a high pressure compressor 12, for example. The fan F feeds air to the compressor V and to the bypass duct B. In this arrangement, the bypass duct B runs around a core engine, which comprises the compressor V and the turbine TT and comprises a primary flow duct for the air fed to the core engine by the fan F.
The air fed into the primary flow duct via the compressor V enters a combustor section BK of the core engine, in which the driving energy for driving the turbine TT is generated. For this purpose, the turbine TT has a high pressure turbine 13, a medium pressure turbine 14 and a low pressure turbine 15. Here, the energy released during combustion is used by the low pressure turbine 15 to drive the rotor shaft S and hence the fan F in order to produce the required thrust by means of the air fed into the bypass duct B. Both the air from the bypass duct B and the exhaust gases from the primary flow duct of the core engine flow out via an outlet A at the end of the engine T. In this arrangement, the outlet A generally has a thrust nozzle with a centrally arranged outlet cone C.
As is known, rotor blade assemblies rotating around the central axis M, each of which has a row of rotor blades and in which the rotor blades are provided on an annular or disk-shaped blade carrier, are used both in the region of the (axial-) compressor with its low pressure compressor 11 and its high pressure compressor 12 and in the region of the turbine TT. In this arrangement, it is possible in principle for the annular or disk-shaped blade carrier to have integral blades and thus to be produced as a bling or blisk. As an alternative, individual rotor blades can be fixed on an annular or disk-shaped blade carrier by means of their respective blade roots.
The embodiments of the invention which are illustrated below relate to the connection of the rotor blades in the region of the turbine TT of the gas turbine 100.
FIG. 2A shows a horizontal section through a rotor blade, in this case a turbine blade element 1. In this arrangement, the turbine blade element 1 surrounds, in the interior, a core 4 connected integrally to a disk 5 of the turbine TT. This can be seen more precisely in the sectional view in FIG. 2B.
The disk 5 has the core 4, which projects into the interior of the turbine blade element 1. Here, the connection between the disk 5 and the turbine blade element 1 is made by means of a positive connection means 2 in the interior of the turbine blade element 1. In this case, the connection means 2 in the embodiment illustrated here has a positive joining means 3, by means of which the turbine blade element 1 is fixed axially and/or radially. A positive connection means 2 can be combined with nonpositive and/or material connection means 2.
Here, the positive joining means 3 is formed on the radially outer edge of the core 4 as a mushroom-shaped feature forming an offset. A corresponding projection, which enters into engagement with the offset of the positive joining means 3 on the core 4, is formed in the interior of the turbine blade element 1. The positive joining means 3 can also have an undercut, for example.
During operation, considerable radial forces act on the turbine blade elements 1 owing to the centrifugal force. The positive connection 2 is therefore designed to withstand these radial forces. A typical weight for a turbine blade element 1 in an aircraft engine is between 50 and 150 g. In the case of stationary gas turbines, the weight may be significantly higher.
Here, the tip of the turbine blade elements 1 extends radially away from the disk 4 over a height H.
In the embodiment illustrated here, the positive connection means 2 is about half way up the blade height or half the region H1 of the turbine blade element 1 which is exposed to the hot driving airflow L during the operation of the gas turbine. Thus, the turbine blade element 1 can be divided into two zones Z1, Z2 in the radial direction.
The first zone Z1 extends from the base of the turbine blade element 1 to the positive connection means 2. The second zone Z2 extends from the positive connection means 2 to the blade tip.
In the embodiment illustrated, a parting line T extends between the zones Z1, Z2, between different materials.
In the first zone Z1 below the positive connection means 2 it is primarily compressive stresses which act, and therefore materials that are particularly resistant to compressive stresses can be employed here. One example of these are ceramics, for instance, especially an yttrium-stabilized zirconium oxide or CMC (ceramic matrix composites).
A monocrystalline material CMSX-4, for example, is used in the second zone Z2 above the positive connection means 2. A typical composition for this nickel base alloy is:

6.5% by weight of Cr,
5.6% by weight of Al,
1.0% by weight of Ti,
6.5% by weight of Ta,
6.4% by weight of W,
0.6% by weight of Mo,
9.6% by weight of Co,
3.0% by weight of Re,
0.07% by weight of Hf.

This material is particularly temperature-stable. In principle, however, other superalloys resistant to high temperatures can also be used in the second zone Z2.
In the embodiment illustrated, the positive connection means 2 is arranged substantially half way up the turbine blade element 1. As an alternative, however, it is also possible for the positive connection means 2 to be arranged closer to the base, i.e. closer to the disk 5, or closer to the tip of the turbine blade element 1.
FIG. 2B furthermore illustrates that cooling air from the region of the disk 5 can enter radially outward into the interior of the turbine blade element 1.
FIG. 3A illustrates a detail of an embodiment of a connection between a turbine blade element 1 and a disk 5. Here, the core 4 of the disk 5 has a wedging means 7, which can be driven into a corresponding gap in the core 4.
FIG. 3A illustrates the position of the wedging means 7 when it has not yet been driven in (the core 4 then having a small cross section). In this state, the turbine blade element 1 can be mounted.
FIG. 3B illustrates a driven-in position of the wedging means 7, i.e. the cross section of the core 4 increases, thus establishing a positive connection means 2 between the core 4 and the turbine blade element 1.
FIGS. 4A and 4B illustrate another alternative embodiment. Here, the wedging means 7 is used to produce a nonpositive connection means 2 between the core 4 and the disk 5.
FIG. 4A illustrates a sectional view in which the core 4 is illustrated without the clamping wedging means 7. The core 4 has only a pre-produced gap (depicted here in dashed lines), into which the wedging means 7 can be inserted. In the position illustrated here, there is no joint between the turbine blade elements 1 and the core 4 at the side walls.
If, on the other hand, there is a wedging means 7 in the core, as illustrated in FIG. 4B, the core 4 expands, thus ensuring that there is frictional engagement at the side walls of the core 4 with the insides of the turbine blade element 1.
Such a joint can be combined, for example, with a positive joining means 3 (as in the embodiments illustrated in FIGS. 3A, 3B). It is also possible, as an alternative or in addition, to produce a material joint.
The configuration shown in FIGS. 3 and 4 can be produced in such a way that the sleeve-shaped turbine blade elements 1 are placed over the core 4, i.e. the turbine blade elements 1 are open at the radially outer end. After placement, the positive joining, the nonpositive joining and/or the material joining can then be produced.
In FIGS. 3 and 4, the turbine blade elements 1 are each of closed design at the radially outer edge. This can be accomplished, for example, by welding on a cover after the production of the joint—as described above. However, it is also possible for the radially outer end to remain open, thus allowing cooling air K which enters the turbine blade elements 1 at the bottom to escape at the top.

LIST OF REFERENCE SIGNS

1 Turbine blade element
2 Connection means
3 Positive joining means
4 Core
5 Disk
7 Wedging means, wedge connection
11 Low pressure compressor
12 High pressure compressor
13 High pressure turbine
14 Medium pressure turbine
15 Low pressure turbine
100 Gas turbine
A Outlet
B Bypass duct
BK Combustor section
C Outlet cone
E Inlet/Intake
F Fan
FC Fan casing
H Height of the turbine blade tip, measured radially from the disk
H1 Region of the turbine blade in the airflow
K Cooling air
L Driving airflow
M Rotational axis
R Inlet direction
T Parting line
TT Turbine
V Compressor
Z1 First zone
Z2 Second zone

Claims

1. A gas turbine having at least one disk, wherein turbine blade elements are connected via connection means to the at least one disk, wherein

the connection means are arranged in the interior of the turbine blade elements in a region radially above the disk in the radial direction of the turbine blade elements, in particular in a region which is in the driving airflow during the operation of the gas turbine, and

the turbine blade elements have at least two zones composed of different materials, wherein the at least two zones adjoin one another in particular in the radial direction, and

in that a zone with a material suited to compressive stress, in particular a ceramic, in particular an yttrium-stabilized zirconium oxide, is arranged radially below the connection means, and a zone with a material suited to tensile stress, in particular CMSX 4, is arranged radially above the connection means.

2. The gas turbine as claimed in claim 1, wherein the connection means are arranged radially on the inside, radially in the center or radially on the outside in the region radially above the disk, in particular in the region of the turbine blade elements which is in the driving airflow during the operation of the gas turbine.

3. The gas turbine as claimed in claim 1, wherein the connection means are designed for positive joining, nonpositive joining and/or material joining.

4. The gas turbine as claimed in claim 3, wherein the positive connection means have positive joining means, in particular shoulders, projections and/or undercuts for the axial and/or radial fixing of the turbine blade elements.

5. The gas turbine as claimed in claim 3, wherein, in the case of nonpositive connection means, frictional joining can be achieved by means of a wedge connection, in particular in the disk, and/or by means of shrink-fit connections.

6. The gas turbine as claimed in claim 3, wherein, in the case of material connection means, there is a laser-welded joint between the turbine blade elements and the disk.

7. The gas turbine as claimed in claim 1, wherein at least one parting line is arranged between two zones of different material in the turbine blade elements radially below the connection means.

8. The gas turbine as claimed in claim 1, characterized by wedge elements for clamping the disk to the turbine blade elements and/or for producing positive joining between the disk and the turbine blade elements.

9. The gas turbine as claimed in claim 1, wherein the turbine blade elements are connected radially on the outside via a welded joint to the at least one disk.

10. The gas turbine as claimed in claim 1, wherein the turbine blade elements can be assembled from at least two parts.

11. The gas turbine as claimed in claim 1, wherein it is designed as an aircraft engine, as a vehicle propulsion system, as a ship's propulsion system or as a stationary gas turbine.

12. A turbine blade element or disk, in particular configured and designed for use in a gas turbine as claimed in claim 1.