U.S. patent application number 14/559668 was filed with the patent office on 2015-06-11 for turbine rotor blade of a gas turbine and method for cooling a blade tip of a turbine rotor blade of a gas turbine.
The applicant listed for this patent is Rolls-Royce Deutschland Ltd & Co KG, Rolls-Royce plc. Invention is credited to Knut LEHMANN, Anthony RAWLINSON, Jens TAEGE.
Application Number | 20150159488 14/559668 |
Document ID | / |
Family ID | 51868051 |
Filed Date | 2015-06-11 |
United States Patent
Application |
20150159488 |
Kind Code |
A1 |
LEHMANN; Knut ; et
al. |
June 11, 2015 |
TURBINE ROTOR BLADE OF A GAS TURBINE AND METHOD FOR COOLING A BLADE
TIP OF A TURBINE ROTOR BLADE OF A GAS TURBINE
Abstract
The present invention relates to a turbine rotor blade of a gas
turbine with a blade tip, on which means for duct-type guidance of
cooling air extending from a front suction-side area of the blade
tip to a rear area of the blade tip are provided, and to a method
for cooling a blade tip of a turbine rotor blade of a gas turbine,
where air from a hot gas flow is guided from a front suction-side
area of a blade tip to a rear area of the blade tip through a
duct-type guidance.
Inventors: |
LEHMANN; Knut; (Berlin,
DE) ; RAWLINSON; Anthony; (Derby, GB) ; TAEGE;
Jens; (Schoeneiche, DE) |
|
Applicant: |
Name |
City |
State |
Country |
Type |
Rolls-Royce Deutschland Ltd & Co KG
Rolls-Royce plc |
Blankenfelde-Mahlow
London |
|
DE
GB |
|
|
Family ID: |
51868051 |
Appl. No.: |
14/559668 |
Filed: |
December 3, 2014 |
Current U.S.
Class: |
416/1 ;
416/97R |
Current CPC
Class: |
F01D 5/18 20130101; F01D
5/187 20130101; F05D 2240/80 20130101; F01D 5/20 20130101; F01D
11/10 20130101 |
International
Class: |
F01D 5/18 20060101
F01D005/18 |
Foreign Application Data
Date |
Code |
Application Number |
Dec 5, 2013 |
DE |
10 2013 224 998.5 |
Claims
1. Turbine rotor blade of a gas turbine with a blade tip, on which
means for duct-type guidance of cooling air extending from a front
suction-side area of the blade tip to a rear area of the blade tip
are provided.
2. Turbine rotor blade in accordance with claim 1, wherein the
means are provided at least over part of the length of the blade
tip.
3. Turbine rotor blade in accordance with claim 1, wherein the
means for the inlet of cooling air from the hot gas flow are
provided in the front suction-side area of the blade tip.
4. Turbine rotor blade in accordance with claim 1, wherein the
means for the outlet of cooling air are provided in the
pressure-side area of the rear blade tip or in the suction-side
area of the rear blade tip or in the area of the blade trailing
edge.
5. Turbine rotor blade in accordance with claim 1, wherein a
sealing edge and/or an overhang are/is provided at the blade
tip.
6. Turbine rotor blade in accordance with claim 1, wherein the
means are designed in the form of a cover arranged at the blade tip
and forming a flow duct.
7. Turbine rotor blade in accordance with claim 1, wherein the
means are designed for the introduction of air exiting at least one
air duct extending inside the turbine rotor blade.
8. Turbine rotor blade in accordance with claim 1, wherein at least
one web or support is provided at the blade tip for guiding the
cooling air in the flow duct.
9. Turbine rotor blade in accordance with claim 6, wherein the
cover is arranged on the web.
10. Method for cooling a blade tip of a turbine rotor blade of a
gas turbine, where air from a hot gas flow is guided from a front
suction-side area of a blade tip to a rear area of the blade tip
through a duct-type guidance.
Description
[0001] This application claims priority to German Patent
Application DE102013224998.5 filed Dec. 5. 2013, the entirety of
which is incorporated by reference herein.
[0002] This invention relates to a turbine rotor blade of a gas
turbine with a blade tip. Furthermore, this invention relates to a
method for cooling such a blade tip of a turbine rotor blade.
[0003] It is known from the state of the art that a leakage mass
flow caused by the pressure difference from a blade pressure side
to a blade suction side arises at a radial gap between a turbine
rotor and a casing. Attempts are therefore being made to design the
blade tip of the turbine rotor such that the leakage mass flow is
reduced. Another objective is to reduce the negative effect of the
blade tip leakage vortex caused by the leakage mass flow on the
turbine aerodynamics.
[0004] To improve the flow over the blade tips of the turbine
rotor, circumferential sealing edges (squealers) are used. Designs
are also known, where overhangs at the blade tip (winglets) are
provided. The circumferential sealing edges can contribute to an
improvement in the aerodynamics. The overhangs on the suction side
and/or on the pressure side can reduce the leakage mass flow and
also improve the aerodynamics around the blade tip.
[0005] It is furthermore necessary in turbine rotor blades,
particularly under operating conditions with high thermal loads due
to the hot gas flow, to achieve high durability and a long service
life for the blade tips. To that end, it is known to cool the
turbine rotor blades internally by convection using compressor air
and to discharge this air for the purpose of film cooling on the
blade outer surfaces or blade tips. Besides the blade leading edge,
the thermally most highly loaded area of a blade tip is often the
rear pressure-side area. It is known to provide on the blade tip
pressure-side holes or recesses through which blade-internal film
cooling air is passed out near the blade tip. In addition, cooling
air can be passed out through openings on the blade tip (dust
holes), in order to prevent dirt accumulations in the internal
cooling air passages and to achieve additional cooling.
[0006] The measures known from the state of the art entail a number
of disadvantages:
[0007] When film cooling air is used, it is necessary to consume a
large quantity of "expensive" film cooling air in order to protect
the blade tip from the high temperatures of the hot gas. The film
cooling air must be taken from the compressor of the gas turbine,
thereby reducing the efficiency of the thermodynamic cyclic process
inside the gas-turbine engine.
[0008] Convective internal cooling of the blade tip by internal
cooling air passages is relatively ineffective in the area of the
blade tip. This is partly due to the fact that only insufficiently
high internal heat transfer coefficients can be generated directly
underneath the blade tip. In addition, conventional casting
techniques, using which the turbine rotor blades are manufactured,
require relatively high minimum wall thicknesses, so that the
temperature in the outer wall area is relatively close to the hot
gas temperature.
[0009] Cooling air that must be forcibly discharged at the blade
tip in order to prevent dirt accumulations in the internal blade
ducts contributes only to a limited extent to the cooling of the
blade tip, since it cannot reach the thermally most highly loaded
points of the blade tip.
[0010] The object underlying the present invention is to provide a
method for cooling a blade tip of a turbine rotor blade of a gas
turbine as well as a turbine rotor blade suitable for carrying out
the method, which, while being simply designed guarantees effective
cooling.
[0011] It is a particular object of the present invention to
provide solution to the above problematics by a combination of the
features of the independent Claims. Further advantageous
embodiments of the invention become apparent from the
sub-claims.
[0012] In accordance with the invention, it is thus provided in
respect of the turbine rotor blade that means for duct-type
guidance of cooling air extend from its front suction-side area to
a rear area. The invention thus provides for the thermally highly
loaded rear area of the turbine rotor blade tip to be supplied at
least partly with passive air, i.e. with relatively cold air from
the hot gas flow. It is self-evident that the hot gas temperature
must be below the maximum temperature sustainable by the blade
material. There is an area with relatively low hot gas temperature
on the front suction-side blade tip. In particular in the case of
additional discharge of cooling air at the turbine casing upstream
of the rotor blades, the hot gas temperature in this area can be
considerably below the highest permissible metal temperature.
[0013] In accordance with the invention, the relatively "cold" hot
gas flow is thus used for cooling the blade tip by being routed to
the thermally most highly loaded pressure-side rear area of the
blade tip. This is achieved by the means provided in accordance
with the invention for duct-type guidance of this cooling air. In a
particularly favourable embodiment of the invention, it is provided
that a cover forming a flow duct is attached at the blade tip. This
creates a duct or a cavity on the blade tip, through which the
cooling air can be passed. This duct or cavity, forming the means
in accordance with the invention for duct-type guidance of cooling
air, preferably has an inlet opening on the suction-side blade
leading edge. It is however also possible to introduce cooling air
from the blade interior. In the area of the blade trailing edge
too, an opening is provided through which the cooling air flows
out. The pressure difference over the blade row ensures here a flow
of relatively cold hot gas through the duct or cavity.
[0014] The means provided in accordance with the invention can
extend over the entire length or only over part of the length of
the blade tip.
[0015] The cooling air can exit the duct or cavity in accordance
with the invention at the blade tip, depending on the required
counter-pressure level, in the area of the pressure-side or the
suction-side rear blade tip. The discharge at the pressure side has
the further advantage, besides lower aerodynamic losses, that the
cooling air discharged there is in turn sucked into the blade tip
gap, so that the external blade tip surface too is supplied with
relatively cold air.
[0016] In a preferred embodiment of the invention, a protective
cover as mentioned is fitted onto a circumferential sealing edge of
the blade tip (winglet, squealer) and fastened there. The result is
an effective, flat and contourless blade tip geometry with a duct
or cavity underneath. In an even more advantageous embodiment of
the invention, it is provided that the cover is suitably shaped or
contoured such that the blade tip can retain the contour of the
circumferential sealing edge.
[0017] It is particularly advantageous when the wall thicknesses of
the cover are designed relatively thin, so that its maximum metal
temperature can be kept as low as possible.
[0018] The protective cover and/or the cavity on the blade tip can
be but do/does not necessarily have to be, designed up to the blade
leading edge. Since the pressure level in the front blade tip area
is relatively constant and only drops steeply towards the blade
trailing edge, it can be advantageous to design the protective
cover and hence the duct or cavity only starting from a middle
position of the blade tip. With this embodiment, the dust holes can
then be near the cover and hence close to the inflow area of the
duct or cavity, in order to promote the aspiration of the cold dust
hole air into the duct or cavity. It is thus possible to use the
dust hole air, otherwise not readily usable for cooling, to cool
the pressure-side blade tip close to the trailing edge.
[0019] In a further embodiment of the invention, it is possible to
design the protective cover at the blade leading edge closed
(without opening). Hence only one opening of the cavity or of the
duct is provided close to the blade trailing edge. With this
embodiment of the invention, the cavity or duct can be flooded
completely with cold blade-internal air. This embodiment is
suitable in particular for very high hot gas temperatures, when
blade-external air is no longer usable for cooling the blade
tip.
[0020] For guiding the flow inside the blade tip cavity or in the
flow duct provided at the blade tip, additional webs or supports
can be used. By means of these webs, the air can be guided to the
thermally most highly loaded areas. The webs or supports are
furthermore used as fastening surfaces for the protective cover and
thereby contribute to the mechanical stability of the blade tip.
Furthermore, the webs or supports can increase the internal heat
transfer in the cavity or duct, so that the cooling effect can be
further improved.
[0021] The following advantages result in accordance with the
invention, as already partially explained above:
[0022] With the invention, the quantity of film cooling air
required for cooling the rotor blade tip can be considerably
reduced.
[0023] Due to the more effective cooling of the blade tip and the
reduction of the blade tip temperatures that this entails, the wear
on the blade tip can be reduced and hence the service life of the
turbine blade extended.
[0024] Due to the reduced wear on the blade tip during operation,
the decrease in turbine efficiency over the period of operation can
be reduced.
[0025] In accordance with the invention the operating costs of the
gas turbine are reduced.
[0026] The present invention is described in the following in light
of the accompanying drawing showing exemplary embodiments. In the
drawing,
[0027] FIG. 1 shows a schematic representation of a gas-turbine
engine in accordance with the present invention,
[0028] FIG. 2 shows a simplified sectional view of a blade tip
designed in accordance with the present invention,
[0029] FIG. 3 shows a view, by analogy with FIG. 2, of a further
exemplary embodiment of the present invention,
[0030] FIGS. 4 shows a further exemplary embodiment, by analogy
with FIGS. 2 and 3 without lateral blade tip overhang, and
[0031] FIGS. 5 to 12 show simplified perspective representations of
exemplary embodiments in accordance with the present invention.
[0032] The gas-turbine engine 10 in accordance with FIG. 1 is a
generally represented example of a turbomachine where the invention
can be used. The engine 10 is of conventional design and includes
in the flow direction, one behind the other, an air inlet 11, a fan
12 rotating inside a casing, an intermediate-pressure compressor
13, a high-pressure compressor 14, a combustion chamber 15, a
high-pressure turbine 16, an intermediate-pressure turbine 17 and a
low-pressure turbine 18 as well as an exhaust nozzle 19, all of
which being arranged about a central engine axis 1.
[0033] The intermediate-pressure compressor 13 and the
high-pressure compressor 14 each include several stages, of which
each has an arrangement extending in the circumferential direction
of fixed and stationary guide vanes 20, generally referred to as
stator vanes and projecting radially inwards from the engine casing
21 in an annular flow duct through the compressors 13, 14. The
compressors furthermore have an arrangement of compressor rotor
blades 22 which project radially outwards from a rotatable drum or
disk 26 linked to hubs 27 of the high-pressure turbine 16 or the
intermediate-pressure turbine 17, respectively.
[0034] The turbine sections 16, 17, 18 have similar stages,
including an arrangement of fixed stator vanes 23 projecting
radially inwards from the casing 21 into the annular flow duct
through the turbines 16, 17, 18, and a subsequent arrangement of
turbine rotor blades 24 projecting outwards from a rotatable hub
27. The compressor drum or compressor disk 26 and the blades 22
arranged thereon, as well as the turbine rotor hub 27 and the
turbine rotor blades 24 arranged thereon rotate about the engine
axis 1 during operation.
[0035] FIG. 2 shows a simplified sectional view of a blade tip 29
of a turbine rotor blade 24. The reference numeral 35 shows the
suction side, while the reference numeral 36 indicates the pressure
side. Front-side sealing edges 33 are provided on the blade tip 29.
Furthermore, the blade tip 29 can have lateral overhangs (winglets)
34. A flow duct 37 is provided on the front side of the blade tip
29 and closed by a protective cover 38. The protective cover 38 is
designed profiled, so that any contouring of the blade tip 29 can
be retained.
[0036] FIG. 3 shows a view, by analogy with FIG. 2, where
additional webs or supports 39 are provided which support the
protective cover 38. The webs or supports 39 furthermore enable
several flow ducts 37 to be formed, or the airflow through the flow
duct 37 to be optimized.
[0037] FIG. 4 shows a view, by analogy with FIGS. 2 and 3, where in
the exemplary embodiment of FIG. 4 the blade tip has no lateral
overhang (winglet overhang).
[0038] FIGS. 5 to 12 show differing exemplary embodiments of the
invention, where the perspective view is schematic and where the
protective cover 38 is only shown in simplified form in order to
make clear the flow through the flow duct 37.
[0039] In all exemplary embodiments of FIGS. 5 to 12, the front
suction-side area of the blade tip 29 has the reference numeral 30,
while the rear area has the reference numeral 31. A blade trailing
edge 32 is designed in the usual way.
[0040] FIG. 5 shows an exemplary embodiment in which the protective
cover 38 is provided on the entire front-side area of the blade tip
29. The flow necessary for cooling the blade tip is introduced
centrally at the blade leading edge, while the outflow through a
suction-side opening takes place in the rear area of the blade
tip.
[0041] FIG. 6 shows a design variant of the exemplary embodiment of
FIG. 5, where the protective cover 38 extends over only part of the
total length of the blade tip 29.
[0042] Complementing the design variant of FIG. 5, a centric
support 39 is provided in the exemplary embodiment of FIG. 7, which
divides the flow duct 37.
[0043] In the exemplary embodiment of FIG. 8, the outflow is
provided, in a variation of the exemplary embodiment of FIG. 7 on
the pressure side of the rear area 31 of the blade tip 29.
[0044] The exemplary embodiment of FIG. 9 represents a variant of
the exemplary embodiments of FIGS. 7 and 8 and has a centric outlet
opening in the area of the blade trailing edge 32.
[0045] In the exemplary embodiment of FIG. 10, the protective cover
extends completely over the front area of the blade tip 29 and has
an outlet opening only on the suction side of the rear area 31. The
cooling air is supplied from the blade interior via ducts.
[0046] The exemplary embodiment of FIG. 11 also shows a protective
cover closed in the front area, by analogy with FIG. 10, where a
central web 39 divides the flow duct 37.
[0047] In the exemplary embodiment of FIG. 12, it is provided, in a
variation from the exemplary embodiments of FIGS. 10 and 11, that
individual supports acting as turbulators are arranged inside the
flow duct 37.
[0048] The exemplary embodiments of FIGS. 10 to 12 each show the
supply of the flow through cooling air holes (dust holes) 41.
List of Reference Numerals
[0049] 1 Engine axis [0050] 10 Gas-turbine engine/core engine
[0051] 11 Air inlet [0052] 12 Fan [0053] 13 Intermediate-pressure
compressor (compressor) [0054] 14 High-pressure compressor [0055]
15 Combustion chamber [0056] 16 High-pressure turbine [0057] 17
Intermediate-pressure turbine [0058] 18 Low-pressure turbine [0059]
19 Exhaust nozzle [0060] 20 Guide vanes [0061] 21 Engine casing
[0062] 22 Compressor rotor blades [0063] 23 Stator vanes [0064] 24
Turbine rotor blades [0065] 25 -- [0066] 26 Compressor drum or disk
[0067] 27 Turbine rotor hub [0068] 28 Exhaust cone [0069] 29 Blade
tip [0070] 30 Front suction-side area of blade tip 29 [0071] 31
Rear area of blade tip 29 [0072] 32 Blade trailing edge [0073] 33
Sealing edge [0074] 34 Overhang [0075] 35 Suction side [0076] 36
Pressure side [0077] 37 Flow duct [0078] 38 Cover/protective cover
[0079] 39 Web/support [0080] 40 Casing [0081] 41 Cooling air
hole
* * * * *