U.S. patent application number 16/462466 was filed with the patent office on 2019-12-12 for gas turbine.
The applicant listed for this patent is Rolls-Royce Deutschland Ltd & Co KG. Invention is credited to Karl SCHREIBER.
Application Number | 20190376392 16/462466 |
Document ID | / |
Family ID | 60888356 |
Filed Date | 2019-12-12 |
United States Patent
Application |
20190376392 |
Kind Code |
A1 |
SCHREIBER; Karl |
December 12, 2019 |
GAS TURBINE
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.
Inventors: |
SCHREIBER; Karl; (Am
Mellensee, DE) |
|
Applicant: |
Name |
City |
State |
Country |
Type |
Rolls-Royce Deutschland Ltd & Co KG |
Blankenfelde-Mahlow |
|
DE |
|
|
Family ID: |
60888356 |
Appl. No.: |
16/462466 |
Filed: |
December 1, 2017 |
PCT Filed: |
December 1, 2017 |
PCT NO: |
PCT/EP2017/081193 |
371 Date: |
May 20, 2019 |
Current U.S.
Class: |
1/1 |
Current CPC
Class: |
Y02T 50/60 20130101;
F01D 5/3007 20130101; F05D 2300/6033 20130101; F01D 5/323 20130101;
F01D 5/147 20130101; F01D 5/3084 20130101; F01D 5/284 20130101 |
International
Class: |
F01D 5/14 20060101
F01D005/14; F01D 5/28 20060101 F01D005/28 |
Foreign Application Data
Date |
Code |
Application Number |
Dec 1, 2016 |
DE |
10 2016 123 248.3 |
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.
Description
[0001] The invention relates to a gas turbine having the features
of claim 1.
[0002] Gas turbines, such as aircraft engines or stationary gas
turbines, are units subject to high thermal and mechanical
loads.
[0003] 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".
[0004] 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.
[0005] The object is therefore to adapt gas turbines to the special
conditions of use.
[0006] This object is achieved by a gas turbine having the features
of claim 1.
[0007] The gas turbine has at least one disk, wherein turbine blade
elements are connected via connection means to the at least one
disk.
[0008] 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.
[0009] 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.
[0010] 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.
[0011] 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.
[0012] 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.
[0013] 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.
[0014] 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.
[0015] Fundamentally, it is possible for the connection means to be
designed for positive joining, nonpositive joining and/or material
joining.
[0016] 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.
[0017] Nonpositive connection means can have a wedge connection, in
particular in the disk, and/or shrink-fit joints to achieve
frictional joining.
[0018] Material connection means can have a laser-welded joint
between the turbine blade elements and the disk.
[0019] 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.
[0020] To connect the turbine blade elements to the cores, a welded
joint can be arranged radially on the outside.
[0021] The turbine blade elements can consist of two parts, wherein
the positive connections are established only when the parts are
assembled.
[0022] 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.
[0023] 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.
[0024] The invention will be discussed in connection with the
exemplary embodiments illustrated in the figures. In the
figures:
[0025] FIG. 1 shows a schematic illustration of a gas turbine, in
this case an aircraft turbine;
[0026] FIG. 2A shows a horizontal section through a turbine blade
element;
[0027] FIG. 2B shows a section through the turbine blade element
according to FIG. 2A along the line A-A;
[0028] FIG. 3A shows an alternative embodiment of the turbine blade
element with a wedging means for the installation of a turbine
blade element;
[0029] 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;
[0030] FIG. 4A shows an alternative embodiment of the turbine blade
element with a disk without a wedging means;
[0031] 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.
[0032] 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.
[0033] 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.
[0034] 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.
[0035] 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.
[0036] 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.
[0037] 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.
[0038] 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.
[0039] 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.
[0040] Here, the tip of the turbine blade elements 1 extends
radially away from the disk 4 over a height H.
[0041] 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.
[0042] 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.
[0043] In the embodiment illustrated, a parting line T extends
between the zones Z1, Z2, between different materials.
[0044] 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).
[0045] 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: [0046] 6.5% by weight of
Cr, [0047] 5.6% by weight of Al, [0048] 1.0% by weight of Ti,
[0049] 6.5% by weight of Ta, [0050] 6.4% by weight of W, [0051]
0.6% by weight of Mo, [0052] 9.6% by weight of Co, [0053] 3.0% by
weight of Re, [0054] 0.07% by weight of Hf.
[0055] This material is particularly temperature-stable. In
principle, however, other superalloys resistant to high
temperatures can also be used in the second zone Z2.
[0056] 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.
[0057] 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.
[0058] 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.
[0059] 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.
[0060] 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.
[0061] 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.
[0062] 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.
[0063] 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.
[0064] 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.
[0065] 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.
[0066] 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
[0067] 1 Turbine blade element [0068] 2 Connection means [0069] 3
Positive joining means [0070] 4 Core [0071] 5 Disk [0072] 7 Wedging
means, wedge connection [0073] 11 Low pressure compressor [0074] 12
High pressure compressor [0075] 13 High pressure turbine [0076] 14
Medium pressure turbine [0077] 15 Low pressure turbine [0078] 100
Gas turbine [0079] A Outlet [0080] B Bypass duct [0081] BK
Combustor section [0082] C Outlet cone [0083] E Inlet/Intake [0084]
F Fan [0085] FC Fan casing [0086] H Height of the turbine blade
tip, measured radially from the disk [0087] H1 Region of the
turbine blade in the airflow [0088] K Cooling air [0089] L Driving
airflow [0090] M Rotational axis [0091] R Inlet direction [0092] T
Parting line [0093] TT Turbine [0094] V Compressor [0095] Z1 First
zone [0096] Z2 Second zone
* * * * *