US01079422432

( 12) United States Patent ( 10) Patent No .: US 10,794,224 B2

Schiessl (45 ) Date of Patent : Oct. 6, 2020
(54) GAS TURBINE AND METHOD OF ( 56 ) References Cited
ATTACHING A TURBINE NOZZLE GUIDE
VANE SEGMENT OF A GAS TURBINE U.S. PATENT DOCUMENTS
( 71 ) Applicant : Rolls -Royce Deutschland Ltd & Co 2,488,867 A 11/1949 Judson
3,620,641 A 11/1971 Keen et al .
KG , Blankenfelde -Mahlow (DE ) 3,623,736 A * 11/1971 Petrie FO1D 11/025
277/422
( 72 ) Inventor: Thomas Schiessl , Rangsdorf (DE ) (Continued )
( 73 ) Assignee : ROLLS - ROYCE DEUTSCHLAND FOREIGN PATENT DOCUMENTS
LTD & CO KG , Blankenfelde -Mahlow
( DE ) DE 3541606 A1 6/1986
DE 102005045459 Al 4/2006
( * ) Notice: Subject to any disclaimer, the term of this ( Continued )
patent is extended or adjusted under 35
U.S.C. 154 (b ) by 565 days . OTHER PUBLICATIONS
( 21 ) Appl. No .: 15 /682,034 German Search Report dated May 2 , 2017 for counterpart German
Application No. 10 2016 115 610.8 .
( 22 ) Filed : Aug. 21 , 2017 (Continued )
( 65 ) Prior Publication Data
US 2018/0058262 A1 Mar. 1 , 2018
Primary Examiner Moshe Wilenksy
Assistant Examiner Jason Mikus
( 30 ) Foreign Application Priority Data (74 ) Attorney, Agent, or Firm Shuttleworth &
Ingersoll, PLC ; Timothy J. Klima
Aug. 23 , 2016 ( DE ) 10 2016 115 610
( 57 ) ABSTRACT
(51 ) Int . Cl .
FOID 25/24 ( 2006.01 ) A gas turbine, including: a combustion chamber ; a high
FOID 9/04 ( 2006.01 ) pressure turbine with a first turbine guide vane ring that is
F02C 3/04 ( 2006.01 ) arranged downstream of the combustion chamber, wherein
( 52 ) U.S. CI . the first turbine guide vane ring has a plurality of turbine
CPC FO1D 25/246 (2013.01 ) ; F01D 9/041 nozzle guide vane segments that respectively include at least
( 2013.01 ) ; FO2C 3/04 (2013.01 ) ; F05D one guide vane, an outer platform , and an inner platform ;
2220/32 (2013.01 ) ; F05D 2220/3212 and an outer housing . Provision is made that the turbine
( 2013.01 ) ; F05D 2240/11 (2013.01 ) ; F05D nozzle guide vane segments are fixed in the radial direction
2240/12 (2013.01 ) at the outer housing, wherein occurring radial loads are
( 58 ) Field of Classification Search transferred into the outer housing . The invention further
CPC F01D 25/246 ; F01D 25/243 ; FO1D 25/24 ; relates to a method for attaching a turbine nozzle ide le
FO1D 25/26 ; FO1D 25/28 segment of a gas turbine.
USPC 415 /209.2
See application file for complete search history . 15 Claims , 9 Drawing Sheets
52

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 US 10,794,224 B2
Page 2

( 56 ) References Cited
U.S. PATENT DOCUMENTS
4,300,868 A 11/1981 Wilkinson et al.
4,391,565 A 7/1983 Speak
4,522,557 A 6/1985 Bouiller et al .
4,720,236 A * 1/1988 Stevens FO1D 9/042
415/116
5,129,783 A 7/1992 Hayton
6,425,738 B1 7/2002 Shaw
7,160,078 B2 * 1/2007 Coign FO1D 9/041
29/ 889.22
8,038,389 B2 10/2011 Arness et al.
9,188,062 B2 11/2015 Tsutsumi
2008/0080970 A1 4/2008 Cooke et al .
2009/0129917 A1 5/2009 Hazevis et al .
FOREIGN PATENT DOCUMENTS
DE 102007001459 Al 7/2007
DE 112012006864 T5 5/2015
EP 2278125 A2 1/2011

OTHER PUBLICATIONS
European Search Report dated Jan. 30 , 2018 from counterpart
European App No. 17187224.5 .
* cited by examiner
U.S. Patent Oct. 6 , 2020 Sheet 1 of 9 US 10,794,224 B2

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U.S. Patent Oct. 6 , 2020 Sheet 2 of 9 US 10,794,224 B2
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U.S. Patent Oct. 6 , 2020 Sheet 3 of 9 US 10,794,224 B2

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U.S. Patent Oct. 6 , 2020 Sheet 4 of 9 US 10,794,224 B2

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U.S. Patent Oct. 6 , 2020 Sheet 5 of 9 US 10,794,224 B2

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U.S. Patent Oct. 6 , 2020 Sheet 6 of 9 US 10,794,224 B2

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U.S. Patent Oct. 6 , 2020 Sheet 7 of 9 US 10,794,224 B2

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U.S. Patent Oct. 6.2020 Sheet 8 of 9 US 10,794,224 B2

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U.S. Patent Oct. 6 , 2020 Sheet 9 of 9 US 10,794,224 B2

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 US 10,794,224 B2
1 2
GAS TURBINE AND METHOD OF direction at the outer housing , with any occurring radial
ATTACHING A TURBINE NOZZLE GUIDE loads being transferred into the outer housing .
VANE SEGMENT OF A GAS TURBINE Aspects of the invention are thus based on the idea to fix
the turbine nozzle guide vane segments not at the combus
REFERENCE TO RELATED APPLICATION 5 tion chamber inner casing , but instead at an outer housing of
the gas turbine in the radial direction . Radial forces are
This application claims priority to German Patent Appli correspondingly guided directly into the outer housing. In
cation No. 10 2016 115 610.8 filed on Aug. 23 , 2016 , the this manner, thermal relative movements and tolerances at
entirety of which is incorporated by reference herein . the radially outer boundary of the main flow path are
10 reduced to a minimum through the gas turbine . For, accord
BACKGROUND ing to the invention , the turbine nozzle guide vane segments
and the adjoining components and structures of the high
The invention relates to a gas turbine and to a method for pressure turbine in the main flow path are fixed in the radial
attaching a turbine nozzle guide vane segment of a gas direction at the same structure , namely at the outer housing .
turbine. 15 A further advantage associated with the invention is that
It is known to fix the turbine nozzle guide vane segments an installation space in the rim area of a turbine disc can be
of the stage 1 , which form the first guide vane ring of the obtained , which supports the rotor blades of stage 1 of the
high - pressure turbine of a gas turbine that is arranged high - pressure turbine and is arranged at a small axial dis
downstream of the combustion chamber, at the combustion tance to the inner platform of the turbine nozzle guide vane
chamber inner casing in the radial direction and in the 20 segments. For, the radial fixation of the turbine nozzle guide
circumferential direction, thus at a structure that is located vane segments at the combustion chamber inner casing , as
radially inside with respect to the main flow path through the it is known from the state of the art, has been realized
gas turbine . As a re the turbine nozzle guide vane through structures in the area of the inner platform of the
segments have to be suspended with a relatively large radial turbine nozzle guide vane segments. Such an additional
play with respect to the casing of the high -pressure turbine, 25 installation space can for example be used to provide
at which the adjoining structures of the high -pressure turbine pre -swirl nozzles for cooling the turbine disc .
are formed or arranged, delimiting the main flow path in this The outer housing of the gas turbine at which the turbine
area radially outside together with the turbine nozzle guide nozzle guide vane segments are fixed can be formed by a
vane segments. Thus, a long tolerance chain results due to turbine casing of the high -pressure turbine and a combustion
the fixation of the turbine nozzle guide vane segments at the 30 chamber outer casing of the combustion chamber that are
combustion chamber inner casing , leading via the combus connected to one another, wherein the turbine nozzle guide
tion chamber inner casing and a high -pressure compressor vane segments are fixed at the turbine casing or at the
exit nozzle ring with a diffusor to the combustion chamber combustion chamber outer casing .
outer casing and from there to the casing of the high According to one embodiment of the invention , provision
pressure turbine. This tolerance chain and differing thermal 35 is made that the turbine nozzle guide vane segments are
expansions lead to radial relative movements at the radially supported in the axial direction at the outer casing of the
outer boundary of the main flow path , for which a corre high -pressure turbine as well as at a combustion chamber
sponding installation space has to be provided . inner casing of the combustion chamber, wherein occurring
In the case that the rotor blade ring arranged downstream axial loads are guided into the outer housing as well as into
of the turbine guide vane ring has rotor blades that are 40 the combustion chamber inner casing . Correspondingly,
formed without a tip shroud , there is the associated problem only the fixation of the turbine nozzle guide vane segments
that at the radially outer boundary of the main flow path, in the radial direction ( and possibly in the circumferential
between the turbine nozzle guide vane segments and a direction , as will be explained in the following ) is realized
radially outer boundary of the main flow path that is fixed at only at the outer housing . In contrast, the fixation of the
the casing of the high -pressure turbine, a step is present that 45 turbine nozzle guide vane segments in the axial direction ,
can be optimized only for one operating point and in which is realized separately of the radial fixation , is realized
addition is subject to the mentioned relatively long tolerance at the structures of the outer housing (or at structures that are
chain . connected to the outer housing) and also at structures of the
The invention is based on the objective of providing a gas combustion chamber inner casing ( or at structures that are
turbine and a method for attaching a turbine nozzle guide 50 connected to the combustion chamber inner casing ).
vane segment of a gas turbine, by means of which the high Here, the fixation in the axial direction is realized for
tolerances as well as the differing radial thermal expansions example by a support, i.e. an axially adjoining element that
or resulting displacements between the guide vane ring of prevents any axial movement. Alternatively, an axial fixation
stage 1 and adjoining structures, as they have been existing can for example be realized by means of a groove - finger
so far, can be reduced . 55 connection .
Thus, according to this exemplary embodiment, any
SUMMARY occurring axial loads are diverted into the outer housing as
well as into the combustion chamber inner casing . This is
An embodiment of the invention relates to a gas turbine associated with the advantage that axial loads are distributed
that has a combustion chamber as well as a high - pressure 60 better.
turbine with a first turbine guide vane ring that is arranged In a further embodiment of the invention , provision is
downstream of the combustion chamber. The turbine guide made that, at its inner platform , a turbine nozzle guide vane
vane ring has a plurality of turbine nozzle guide vane segment forms a structure that extends in the radial direction
segments which respectively comprise at least one guide inwards and that comprises a sealing edge that extends in the
vane, an inner platform , and an outer platform . Further, the 65 circumferential direction , wherein the structure extending
gas turbine has an outer housing. Provision is made that the inwards in the radial direction is arranged inside a recess of
turbine nozzle guide vane segments are fixed in the radial a structure of the combustion chamber inner casing that
 US 10,794,224 B2
3 4
extends the radial direction and that is arranged so as to means of at least one hook . For this purpose, the hooks are
be displaceable inside the same in the radial direction , arranged respectively inside an axial groove of the turbine
wherein the recess that extends in the radial direction forms casing , for example.
an axial stop collar of the combustion chamber inner casing Further, it can be provided that the mentioned groove of
for the sealing edge of the structure, and thus for the turbine 5 the turbine casing further serves for the radial fixation of a
nozzle guide vane segment. In this embodiment, it is ensured guide structure that is arranged downstream of the first turbine
that, for one thing, the turbine nozzle guide vane segments vane ring and provides the outer boundary of the main
that are fixed in the radial direction at the outer housing, can flow path for the rotor blades of a first turbine blade ring
be radially displaced with respect to the combustion cham 10 arranged downstream of the first turbine guide vane ring.
This structure forms a shroud segment of the outer main flow
ber inner casing , so that radial tolerances that are still present path boundary, for example.
can be accommodated . On the other hand , through this In the case that a turbine nozzle guide vane segment is
embodiment, an axial stop collar for the axial support of the fixed in the radial direction respectively by means of at least
turbine nozzle guide vane segments with respect to the one hook , a fixation of the turbine nozzle guide vane
combustion
manner .
chamber inner casing is provided in a simple 15 segment in the circumferential direction can be realized my
means of one or multiple pins . The mentioned structure,
In a further embodiment, provision is made that, at its which is arranged downstream of the first turbine guide vane
outer platform , a turbine nozzle guide vane segment is ring and provides the outer boundary of the main flow path
axially supported at a structure which provides the outer for the rotor blades of a first turbine blade ring that is
boundary of the main flow path for the rotor blades of a first 20 arranged downstream of the first turbine guide vane ring, can
turbine guide vane ring that is arranged downstream of the be supported in the circumferential direction at the lateral
first turbine guide vane ring . This structure can for example surfaces of the hooks of the turbine nozzle guide vane
be a shroud segment that provides towards the main flow segment.
path an abradable coating for the rotor blades of the first In general , the means ( pins , hooks ) for fixing the turbine
turbine blade ring which is arranged downstream of the first 25 nozzle guide vane segments in the radial direction at the
turbine guide vane ring . outer housing can be formed at a structure that extends
Here, it can be provided that, at its outer platform , a outwards at the outer platform in the radial direction .
turbine nozzle guide vane segment forms a structure that According to one embodiment of the invention , provision
extends outwards in the radial direction and that has a is made that the turbine nozzle guide vane segments are
sealing edge that extends in the circumferential direction and 30 fixed in the radial direction directly at the outer housing .
is axially supported . The sealing edge can for example be However, this is not necessarily the case . Thus, it can be
supported at the just mentioned structure, or alternatively at provided in alternative embodiments that the turbine nozzle
a structure of the outer housing, or alternatively at a support guide vane segments are radially fixed the outer housing
ring that is connected to the outer housing . by means of an intermediate structure . This may forexample
For fixing the turbine nozzle guide vane segments in the 35 be advantageous with respect to constructional specifica
radial direction , it can be provided that the guide vane tions or assembly requirements.
segments are respectively fixed at the outer housing by In a first embodiment variant to that, provision is made
means of at least one pin , in particular two pins or bolts . that the intermediate structure is formed by a guide vane
Here, it can be provided that the one pin is arranged inside support ring at which the turbine nozzle guide vane seg
a circular hole and the other pin is arranged inside an 40 ments are radially fixed . Here, the guide vane support ring
elongated hole . Here, the connection via the elongated hole is connected to the turbine casing .
provides a fixation only in the radial direction . The connec Here, it can be provided according to one embodiment
tion via the circular hole additionally provides a fixation of variant that the guide vane support ring has a wall area that
the turbine nozzle guide vane segment also in the circum is provided with axially extending slits . In this manner , it can
ferential direction . 45 be ensured that the guide vane support ring maintains a
Thus, provision is made according to one embodiment of connection to the outer housing under all operational con
the invention that the turbine nozzle guide vane segments ditions , and that the radial relative movements between the
are fixed in the radial direction and in the circumferential guide vane and neighboring structures are minimized . Thus,
direction at the outer housing , wherein occurring radial and through such a guide vane support ring provided with slits ,
tangential loads are guided into the outer housing. Here, it 50 it can for example be achieved that the thermal movements
can be provided that the same fixation means are at least of the outer housing are not obstructed by a support ring
partially used for the radial fixation as well as for the fixation which is formed as a full ring without any such slits .
in the circumferential direction . As just mentioned, this In one embodiment variant to this , provision is made that
fixation means may for example be a circular hole that is a radial area of the guide vane support ring is fixed radially
formed in a wall of the outer platform of the turbine nozzle 55 between the outer housing and a structure that adjoins the
guide vane segment and inside of which a pin is arranged turbine nozzle guide vane segment in the axial direction. For
that is mounted in the outer housing. example, the outer housing and the mentioned structure
For fixing the turbine nozzle guide vane segment in the form a groove into which a radial area of the guide vane
radial direction at the outer housing, it can alternatively be support ring meshes , thus forming a groove - finger connec
provided that the turbine nozzle guide vane segments are 60 tion . In this manner, a radial fixation of the guide vane
respectively fixed in the radial direction inside an axial support ring at the outer housing can be provided, whereby
groove in the outer housing by means of at least one hook . radial relative movements between the turbine nozzle guide
A fixation in the circumferential direction can also be vane segment and the structure adjoining it in the axial
realized through one or multiple pins . direction can be minimized .
At that, provision is made in one embodiment that the 65 Alternatively, it can be provided that the guide vane
turbine nozzle guide vane segments are fixed in the radial support ring is formed without slits . At that, radial relative
direction directly at the turbine housing respectively by movements between the guide vane and neighboring struc
 US 10,794,224 B2
5 6
tures are still more controlled and have smaller deflections FIG . 2b shows a rendering of the radial fixation of the
as compared to a radial fixation of the turbine nozzle guide turbine nozzle guide vane segment of FIG . 2a , in the area of
vane segments at the combustion chamber inner casing its inner platform , at a combustion chamber inner casing by
according to the state of the art. means of two pins ;
In a second embodiment variant for providing an inter- 5 FIG . 3a shows a first exemplary embodiment of a turbine
mediate structure, provision is made that the intermediate nozzle guide vane segment of a nozzle guide vane ring of
structure has a combustion chamber support ring by means stage 1 of a high -pressure turbine, wherein the turbine
of which the intermediate structure is additionally connected nozzle guide vane segment is radially fixed directly at a
to the combustion chamber, in particular to the outer com combustion chamber outer casing ;
bustion chamber wall . In this way, a fixation and attachment 10 FIG . 3b shows a second exemplary embodiment of a
of the combustion chamber can be provided in a simple turbine nozzle guide vane segment of a nozzle guide vane
manner. Here, it can be provided that the combustion ring of stage 1 of a high -pressure turbine, wherein the
chamber sur port ring also has a wall area that is provided turbine nozzle guide vane ent is also radially fixed
with axially extending slits . directly at a combustion chamber outer casing ;
In all mentioned embodiment variants, it can be provided 15 FIG . 3c shows a third exemplary embodiment of a turbine
that the intermediate structure forms a heat shield that nozzle guide vane segment of a nozzle guide vane ring of
provides shielding against the high temperatures which the stage 1 of a high -pressure turbine, wherein the turbine
gases have in the main flow path directly behind the com nozzle guide vane segment is radially fixed directly at a
bustion chamber. turbine casing:
The present invention provides an attachment of turbine 20 FIG . 4 shows a fourth exemplary embodiment of a turbine
nozzle guide vane segments independently of the manner in nozzle guide vane segment of a nozzle guide vane ring of
which the rotor blade ring of stage 1 following the turbine stage 1 of a high -pressure turbine, wherein the turbine
guide vane ring in the flow direction is formed . In principle, nozzle guide vane segment is radially fixed at a turbine
this rotor blade ring can be realized in an embodiment with casing by means of a slit guide vane support ring ;
a tip shroud as well as in an embodiment without a tip 25 FIG . 5 shows a fifth exemplary embodiment of a turbine
shroud . In one embodiment of the invention , provision is nozzle guide vane segment of a nozzle guide vane ring of
made that the rotor blades of the rotor blade ring arranged stage 1 of a high -pressure turbine, wherein the turbine
downstream of the first turbine guide vane ring is formed nozzle guide vane segment is radially fixed at a turbine
without a tip shroud . Particularly in such an embodiment, casing by means of a non -slit guide vane support ring ;
there is the danger that a radial step is formed between the 30 FIG . 6 shows a sixth exemplary embodiment of a turbine
turbine guide vane ring and the structure adjoining it in the nozzle guide vane segment of a nozzle guide vane ring of
axial direction , which provides abradable coating towards stage 1 of a high -pressure turbine, wherein the turbine
the main flow path. Through such an abradable coating, the nozzle guide vane segment is radially fixed at a turbine
gap of the blade tips to the outer boundary of the main flow casing by means of a slit guide vane support ring that
path can be minimized . 35 additionally forms a support arm which is connected to the
According to a further aspect of the invention, the inven combustion chamber; and
tion relates to a method for attaching a turbine nozzle guide FIG . 7 shows a seventh exemplary embodiment of a
vane segment of a gas turbine, wherein the gas turbine has turbine nozzle guide vane segment of a nozzle guide vane
a combustion chamber, a high -pressure turbine with a first ring of stage 1 of a high -pressure turbine, wherein the
turbine guide vane ring arranged downstream of the com- 40 turbine nozzle guide vane segment is radially fixed at a
bustion chamber, and an outer housing . The method com turbine casing by means of a non - slit guide vane support ring
prises: that additionally forms a support arm which is connected to
fixing a turbine nozzle guide vane segment in the radial the combustion chamber.
direction at the outer housing ,
fixing the turbine nozzle guide vane segment in the 45 DETAILED DESCRIPTION
circumferential direction also at the outer housing, and
fixing the turbine nozzle guide vane segment in the axial FIG . 1 shows , in a schematic manner , a turbofan engine
direction at the outer housing as well as at a combustion 100 that has a fan stage with a fan 10 as the low - pressure
chamber inner casing of the gas turbine. compressor, an intermediate -pressure compressor 420 , a

Here, provision is made in one embodiment that the 50 high

fixation in the radial direction and the fixation in the cir
high --pressure
pressure compressor 30 , a combustion chamber 40 ,
turbine 50 , an intermediate -pressure turbine
a

cumferential direction is performed at least partially by 60 , and a low -pressure turbine 70 .

means of the same fixation means . The intermediate -pressure compressor 420 and the high
pressure compressor 30 respectively have a plurality of
BRIEF DESCRIPTION OF THE DRAWINGS 55 compressor stages that respectively comprise a rotor stage
and a stator stage . The turbofan engine 100 of FIG . 1 further
The invention will be explained in more detail on the basis has three separate shafts, a low -pressure shaft 81 that
of exemplary embodiments with reference to the accompa connects the low -pressure turbine 70 to the fan 10 , a
nying drawings in which: intermediate -pressure shaft 82 that connects the intermedi
FIG . 1 shows a simplified schematic sectional view of a 60 ate - pressure turbine 60 to the intermediate -pressure com
turbofan engine in which the present invention can be pressor 420 , and a high -pressure shaft 83 that connects the
realized ; high -pressure turbine 50 to the high -pressure compressor 30 .
FIG . 2a shows a rendering of a turbine nozzle guide vane However, this is to be understood to be merely an example.
segment of a guide vane ring of stage 1 of a high -pressure If, for example, the turbofan engine has no intermediate
turbine according to the state of the art, which is realized so 65 pressure compressor and no intermediate -pressure turbine ,
as to be adjoining the combustion chamber in the main flow only a low -pressure shaft and a high -pressure shaft would be
path ; present.
 US 10,794,224 B2
7 8
The turbofan engine 100 has an engine nacelle 1 that flow path 5 radially outside , and an inner platform 23 that
comprises an inlet lip 14 and forms an engine inlet 11 at the delimits the main flow path 5 radially inside . Here , a
inner side , supplying inflowing air to the fan 10. The fan 10 segment 20 can have one or multiple guide vanes 21 .
has a plurality of fan blades 101 that are connected to a fan Together, the segments 20 that are arranged next to each
disc 102. Here, the annulus of the fan disc 102 forms the 5 other in the circumferential direction form the turbine guide
radially inner boundary of the flow path through the fan 10 . vane ring of the first stage of the high -pressure turbine.
Radially outside , the flow path is delimited by the fan casing The rotor blades 71 are arranged at a turbine disc 72 at a
2. Upstream of the fan - disc 102 , a nose cone 103 is arranged. distance to each other in the circumferential direction . At
Behind the fan 10 , the turbofan engine 100 forms a their radially outer ends, they are provided with a tip shroud
secondary flow channel 4 and a primary flow channel 5. The 10 73. The rotor blades 71 form a rotor blade ring of the first
primary flow channel 5 leads through the core engine ( gas stage of the high -pressure turbine .
turbine ) that comprises the intermediate -pressure compres The turbine nozzle guide vane segments 20 are fixed at the
sor 20 , the high -pressure compressor 30 , the combustion combustion chamber inner casing 32 with respect to the
chamber 40 , the high -pressure turbine 50 , the intermediate radial direction . For this purpose , the inner platform 23 of
pressure turbine 60 , and the low -pressure turbine 70. At that, 15 the segment 20 forms a substantially radially extending wall
the intermediate -pressure compressor 20 and the high -pres 230 that [ is ] fixed at a structure 32 ' , which is a part of the
sure compressor 30 are surrounded by a circumferential combustion chamber inner casing 32 , by means of two pins
housing 29 which forms an annulus surface at the internal 61 , 62. The type of this radial fixation can be seen in FIG .
side , delimitating the primary flow channel 5 radially out 2b , in which a view from the front onto the turbine nozzle
side . Radially inside, the primary flow channel 5 is delimi- 20 guide vane segment 20 is shown . Accordingly, in the shown
tated by corresponding rim surfaces of the rotors and stators exemplary embodiment, the segment 20 has two guide vanes
of the respective compressor stages , or by the hub or by 21 that extend radially between the outer platform 22 and the
elements of the corresponding drive shaft connected to the inner platform 23. The circumferential direction is indicated
hub . by u , the radial direction by r , and the axial direction by x .
During operation of the turbofan engine 100 , a primary 25 According to FIG . 2b , the inner platform 23 forms a
flow flows through the primary flow channel 5 (also referred circular hole 231 , on the one hand , and an extended hole
to as the main flow channel in the following ). The secondary 232 , on the other, in the area of the wall 230. Respectively
flow channel 4 , which is also referred to as the partial - flow one pin 61 , 62 , which is fixed in the structure 32 ' , is inserted
channel, sheath flow channel, or bypass channel, guides air into the circular hole 231 and the extended hole 232. By
sucked in by the fan 10 during operation of the turbofan 30 means of the pin 62 and the extended hole 232 , the segment
engine 100 past the core engine. 20 is fixed only in the radial direction r. By means of the pin
The described components have a common symmetry 61 and the circular hole 231 , the segment 20 is fixed in the
axis 90. The symmetry axis 90 defines an axial direction of radial direction r as well as in the circumferential direction
the turbofan engine. A radial direction of the turbofan engine u . A fixation in the axial direction is realized neither by
extends perpendicularly to the axial direction . 35 means of the pin 61 nor the pin 62 .
In the context of the present invention , the configuration In contrast, a fixation in the axial direction is realized
of the high -pressure turbine 50 , in particular of the first stage through contact surfaces at axially neighboring structures .
of the high - pressure turbine 50 , is of importance. For example, at the outer platform 22 , the segment 20 forms
At first, the suspension of turbine nozzle guide vane a wall 225 that extends in the radial direction and that is
segments according to the state of the art is described based 40 axially supported in a groove - like structure 53 of the turbine
on FIGS . 2a and 25 to provide a better understanding of the casing 51 of the high -pressure turbine and is arranged so as
present invention . to be movable in the radial direction . The movability in the
FIG . 2a shows a partial section of a main flow path 5 radial direction makes it possible to prevent constraint forces
through a gas turbine that is part of an aircraft engine. The that are generated by tolerances or different thermal expan
shown partial section illustrates the with respect to the 45 sions . Here , the occurring tolerances can be relatively
flow direction rear section of a combustion chamber 3 , a strong, since a long tolerance chain results due to the fixation
turbine nozzle guide vane segment 20 of a turbine guide of the turbine nozzle guide vane segments 20 at the com
vane ring that is arranged directly downstream of the com bustion chamber inner casing 32 , leading via the combustion
bustion chamber 3 , and a rotor blade 71 of a turbine blade chamber inner casing 32 and the radial supports, e.g. a
ring . Here, the guide vane ring and the rotor blade ring form 50 compressor exit nozzle ring ( not shown) , to the combustion
the first stage of the high -pressure turbine . chamber outer casing 34 , and from there to the turbine
The combustion chamber 3 has a radially inner combus casing 51 of the high -pressure turbine .
tion chamber wall 31 and a radially outer combustion A further disadvantage of the embodiment described in
chamber wall 33. Structurally, the combustion chamber 3 is FIGS . 2a , 2b is that, when it comes to concepts of rotor
supported by a combustion chamber inner casing 32 and a 55 blades 71 that, in contrast to the ones shown in FIG . 2a , are
combustion chamber outer casing 34. Here, the radially formed without a tip shroud , a radial step is present at the
outer combustion chamber wall 33 is connected via a wall 35 radially outer boundary of the main flow path 5 , which can
to the combustion chamber outer casing 34. In a correspond be optimized for only one operating point and in addition is
ing manner, the radially inner combustion chamber wall 31 subject to a relatively long tolerance chain . In constructions
is connected to the combustion chamber inner casing 32 via 60 where a precise circumferential position of the turbine
a wall ( not shown ). nozzle guide vane segments 20 relative to the combustion
The combustion chamber outer casing 34 is connected to chamber 3 or other structures of the high -pressure turbine is
the casing 51 of the high -pressure turbine in a connection of crucial importance , the relatively long tolerance chain is
area 52 , for example by means of bolts . The casing 51 of the also disadvantageous. The aerodynamic forces in the cir
high -pressure turbine is an outer housing . 65 cumferential direction have to be absorbed via the combus
The turbine nozzle guide vane segment 20 comprises a tion chamber inner casing 32 and e.g. a compressor exit
guide vane 21 , an outer platform 22 that delimits the main nozzle ring into the outer housing 34 , 51 , which has to be
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taken into account when designing these components. After separating them from the same . This gap can be minimized
all , especially with smaller engines, the installation space in by providing an abradable material 75 , into which the blade
the rim area of the turbine disc 72 is restricted through the tips can work themselves in . The adjoining structure 76 is
internal attachment with the components 230 , 231 , 232 , 61 , also referred to as a shroud segment of the outer main flow
62 , and renders any additional cooling of the turbine disc 72 , 5 path boundary.
as it may possibly be necessary , impossible . The individual rotor blades 71 are arranged at the outer
FIG . 3a shows a first exemplary embodiment of the circumference of a turbine disc 72 .
arrangement of the turbine nozzle guide vane segments 20 in FIG . 3b shows an exemplary embodiment that corre
the core engine. Like FIG . 2 , FIG . 3 shows a partial section sponds to the exemplary embodiment of FIG . 3a , except for
that comprises the rear area of a combustion chamber 3 , a 10 the type of radial fixation of the turbine nozzle guide vane
turbine nozzle guide vane segment 20 of a guide vane ring, segments 20 at the outer housing . Thus, provision is made in
and a rotor blade 71 of a rotor blade ring . The combustion the exemplary embodiment of FIG . 3b that the turbine
chamber 3 comprises a combustion chamber inner wall 31 nozzle guide vane segments 20 are fixed in the radial
and a combustion chamber outer wall 33 . direction inside an axial groove 340 in the combustion
The combustion chamber inner wall 31 is clad with heat 15 chamber outer casing 34 that extends about the circumfer
shingles 310 , and the combustion chamber outer wall 33 is ential direction respectively by means of the at least one
clad with heat shingles 330. Structurally, the combustion hook 63 that is formed at the wall section 220 of the outer
chamber 3 is supported by a combustion chamber inner platform 22 extending in the radial direction . Here , two
casing 32 and a combustion chamber outer casing 34 , in a hooks 63 per turbine nozzle guide vane segment 20 are
comparable manner as the one described with regard to FIG . 20 provided, for example . Alternatively, a hook 63 that is
2a . continuous in the circumferential direction can be provided .
Each turbine nozzle guide vane segment 20 comprises one In this embodiment, a fixation of a turbine nozzle guide
or multiple guide vanes 21 , an outer platform 22 , and an vane segment 20 in the circumferential direction can be
inner platform 23. The inner platform 23 forms a wall realized by means of one or multiple axial pins . For
section 230 that extends in the radial direction and meshes 25 example, one pin is inserted into an axial slot that is formed
with the recess 320 (e.g. a groove) of a wall of the com by one of the hooks 63 ( not shown ).
bustion chamber inner casing 32 , which also extends in the In an alternative embodiment, the outer platform 22 can
radial direction . The wall section 230 forms a sealing edge be radially fixed inside an axial groove of the turbine casing
233 that extends in the circumferential direction . Here, the 51 of the high -pressure turbine by means of at least one
recess 320 that extends in the radial direction forms an axial 30 hook . In this context, FIG . 3c shows an exemplary embodi
stop collar for the sealing edge 233 , and thus for the turbine ment that corresponds to the exemplary embodiment of FIG .
nozzle guide vane segment 20. The wall section 230 is not 3b , except for the type of radial fixation of the turbine nozzle
fixed in the radial direction with respect to the wall 32 . guide vane segments 20. According the exemplary
The outer platform 22 comprises a wall section 220 that embodiment of FIG . 3c , the wall section 220 of the outer
extends in the radial direction . It is directly connected to the 35 platform 22 that extends in the radial direction forms at least
combustion chamber outer casing 34 by means of two pins one axially extending hook 64 , which is fixed at the turbine
61 , 62. Here , the connection is realized via a circular hole casing 51 inside an axial groove 511 of the turbine casing 51
and an extended hole corresponding to the rendering of FIG . extending about the circumferential direction . By arranging
2b . Thus, the embodiments of FIG . 2b may be referred to the hook 64 in the groove 511 , a fixation of the turbine
here . 40 nozzle guide vane segments 20 in the radial direction is
The combustion chamber outer casing 34 forms a part of realized directly at the turbine casing 51. Here, two hooks 64
the outer housing of the gas turbine, which is formed by the per turbine nozzle guide vane segment 20 are provided, for
combustion chamber outer casing 34 and the turbine casing example . Alternatively, a hook 64 that is continuous in the
51 of the high -pressure turbine, wherein the combustion circumferential direction can be provided.
chamber outer casing 34 and the turbine casing 51 are 45 It is to be understood that the groove 511 of the turbine
connected to each other in the connection area 52 , for casing 51 is also provided for the radial fixation of the
example by means of bolts . In alternative embodiments, the shroud segment 76. For this purpose, the hooks 64 of the
outer platform 22 is directly connected to the turbine casing turbine nozzle guide vane segment 20 and the hooks 761 of
51 of the high -pressure turbine. the shroud segment 76 are seated alternating inside the
It is to be understood that the turbine nozzle guide vane 50 groove 511 of the turbine casing. Due to the turbine nozzle
segment 20 is also fixed in the circumferential direction guide vane segment 20 as well as the shroud segment 76
through the connection of the circular hole and the pin . being attached in a common groove 511 of the turbine casing
For an axial fixation of the turbine nozzle guide vane 51 , the tolerances between these two components are mini
segment 20 in the area of the upper platform 22 and thus mized .
radially outside with respect to the main flow path 5 , the wall 55 In the embodiment of FIG . 3c , a fixation of a turbine
section 220 forms a contact surface in the form of sealing nozzle guide vane segment 20 in the circumferential direc
edge 221 that extends in the circumferential direction and is tion can be realized by means of one or multiple pins . A
axially supported at a wall 76 of an adjoining structure . This shroud segment 76 can be supported in the circumferential
structure may for example be a structure that is connected to direction at the lateral surfaces of the hooks 64 of the turbine
the turbine casing 51 of the high - pressure turbine and that 60 nozzle guide vane segment 20 , whereby a separate fixation
serves for receiving an abradable material 75 that is formed of the shroud segment 76 by pins can be waived , and a
by the structure towards the main flow path 5. Thus, provi particularly effective embodiment is achieved .
sion is made in the embodiment of Thus, in the exemplary embodiments of FIGS . 3a , 3b and
FIG . 3 that the rotor blades 71 of the rotor blade ring of 3c , a fixation of the turbine nozzle guide vane segments 20
the first stage of the high -pressure turbine are formed 65 in the radial direction at the outer housing 34 , 51 is realized .
without a tip shroud, so that the blade tips are arranged Correspondingly, any occurring radial loads are guided into
across from the outer flow path boundary with a gap the outer housing 34 , 51. This leads to the radial tolerances
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of the arrangement of the turbine nozzle guide vane segment combustion chamber support arm 94 is provided, which is
20 relative to the adjoining structures 75 , 76 at different connected to the combustion chamber outer wall 33 of the
thermal expansions being considerably reduced , since the combustion chamber 3. Here , the combustion chamber sup
turbine nozzle guide vane segment 20 as well as the adjoin port arm 94 starts at the first radial area 92 of the support ring
ing structures are connected to the outer housing 34 , 51 . 5 9 .
FIG . 4 shows an exemplary embodiment in which , unlike By providing a combustion chamber support arm 94 , a
in FIG . 3 , the turbine nozzle guide vane segments are fixed fixation of the combustion chamber outer wall 33 is pro
at the outer housing not directly, but rather via an interme vided in a simple manner, without having to provide addi
diate structure . Here , the fixation is realized at the turbine tional components for that purpose . Here, provision can be
casing 51. In the shown exemplary embodiment, this inter- 10 made that the combustion chamber support arm 94 as well
mediate structure is a guide vane support ring 9. It comprises as the wall area 91 and the radial area 92 are provided with
a first radial wall area 92 in which the guide vane support slits to achieve an improved accommodation of radial rela
ring 9 is connected in the described manner to the wall tive movements .
section 220 of the outer platform 22 by means of two pins FIG . 7 shows an exemplary embodiment that corresponds
62. Further, the support ring 9 comprises a second radial 15 to the exemplary embodiment of FIG . 6 , except for the fact
wall area 93 by which the support ring 9 is radially con that the wall area 91 and the radial area 92 are formed not
nected in the attachment area 52 to the turbine casing 51 and in a slit - like , but in a continuous manner, just as is the case
also to the combustion chamber outer casing 34. A third wall with FIG . 5 .
area 91 , which connects the two radial areas 92 , 93 , extends The present invention is not limited in its embodiment to
obliquely and thus with an axial directional component 20 the previously described exemplary embodiments . For
between the two radial wall areas 92 , 93 . example, it is pointed out that the concrete shape and
Here, provision is made according to the exemplary embodiment of the guide vanes 21 , the outer platform 22 and
embodiment of FIG . 4 that the wall areas 91 , 92 are slit , i.e. the inner platform 23 are to be understood merely as
that they are provided with individual slits and segmented in examples .
this way. In this manner , the radial movability is increased 25 Further, it is to be understood that the features of the
if radial displacements of the turbine casing 51 occur . individual described exemplary embodiments of the inven
In this manner, it is facilitated that a connection to the tion can be combined with each other into different combi
turbine casing 51 is maintained under all operational con nations. As far as ranges are defined , they comprise all
ditions , for example also in the event that strong thermal values within these ranges , as well as all partial ranges
movements of the turbine casing 51 occur, and thus the 30 falling within a range.
radial relative movements between the turbine nozzle guide The invention claimed is :
vane segment 20 , or its outer platform 22 , and the adjoining 1. A gas turbine, comprising:
structures 75 , 76 are minimized . a combustion chamber including a combustion chamber
Here, according to one advantageous embodiment, the casing;
embodiment of the wall areas 91 , 92 with slits is realized in 35 a high - pressure turbine including;
combination with a groove - finger connection 95 , wherein a turbine casing including an axial groove with respect
the outer wall of the groove is formed by the inner wall 510 to a rotational axis of the gas turbine; the axial
of the turbine casing 51 and the inner wall of the groove is groove being defined by two opposing side walls and
formed by an outer wall 760 of the structure 76 adjoining the a connecting base surface and
turbine nozzle guide vane segment 20 in the axial direction 40 a first turbine guide vane ring that is arranged down
( shroud segment). The finger 920 of the groove - finger stream of the combustion chamber, wherein the first
connection 95 is formed by a wall of the first radial wall area turbine guide vane ring includes a turbine nozzle
92 that projects axially into the groove and extends radially guide vane segment, wherein the turbine nozzle
between its inner and outer wall . guide vane segment includes :
Through the groove - finger connection 95 , the guide vane 45 at least one guide vane;
support ring 9 is fixed to the turbine casing 51 in the radial an outer platform ;
direction . In this manner, a radial fixation of the guide vane an inner platform ; and
support ring 9 to the turbine casing 51 is provided, whereby a guide vane hook ;
radial relative movements between the turbine nozzle guide an outer housing , including the combustion chamber
vane segment 20 and the structure adjoining thereto in the 50 casing and the turbine casing , wherein the guide van
axial direction are minimized . hook is positioned in the axial groove of the turbine
FIG . 5 shows a further exemplary embodiment that differs casing to fix the turbine nozzle guide vane segment in
from the exemplary embodiment of FIG . 4 with respect to a radial direction with respect to the rotational axis , and
the embodiment of the wall areas 91 , 92 of the guide vane wherein occurring radial loads are transferred into the
support ring 9. While in the exemplary embodiment of FIG . 55 outer housing ;
4 these wall areas 91 , 92 are formed in a slit - like manner, a first turbine rotor blade ring arranged downstream from
they are not provided with slits in the exemplary embodi the first turbine guide vane ring and including a rotor
ment of FIG . 5. Further, a groove - finger connection corre blade; and
sponding to the groove - finger connection 95 of FIG . 4 is not a structure arranged axially downstream from the first
present. Small radial relative movements between the guide 60 turbine guide vane ring, wherein the structure is an
vane segments 20 and the structures axially adjoining outer boundary of a main flow path of the rotor blade ,
thereat, such as structures 75 , 76 , are permitted. and wherein the structure is fixed in the radial direction
FIG . 6 shows a further exemplary embodiment, in which via a portion of the structure being inserted between the
the turbine nozzle guide vane segments 20 are fixed at the two side walls of the axial groove .
turbine casing 51 not directly, but rather via an intermediate 65 2. The gas turbine according to claim 1 , further compris
structure 9. The intermediate structure 9 comprises a guide ing an inner casing of the combustion chamber, wherein the
vane support ring 9 corresponding to FIG . 4. In addition , a turbine nozzle guide vane segment is supported in an axial
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13 14
direction at the outer housing as well as at the inner casing, a turbine nozzle guide van segment, and wherein
wherein occurring axial loads are transferred into the outer the turbine nozzle guide vane segment includes a
housing and the inner casing . guide vane hook ;
3. The gas turbine according to claim 2 , wherein the inner an outer housing including the turbine casing ;
platform includes an inner platform wall that extends 5 a first turbine rotor blade ring arranged downstream
inwards in the radial direction and comprises an inner wall from the first turbine guide vane ring and including
sealing edge extending in a circumferential direction with a rotor blade ; and
respect to the rotational axis , wherein the inner casing a structure arranged axially downstream from the first
includes a radially extending recess, wherein the inner turbine guide vane ring , wherein the structure is an
platform wall is arranged inside the recess so as to be 10 outer boundary of a main flow path of the rotor
displaceable therein in the radial direction, wherein the blade ; and
recess forms an axial stop collar of the inner casing for the
inner sealing edge of the inner platform wall , and thus for the positioning the guide vane hook in the axial groove of the
turbine nozzle guide vane segment. turbine casing to fix the turbine nozzle guide vane
4. The gas turbine according to claim 1 , wherein , at the 15 segment in a radial direction with respect to the rota
outer platform , the turbine nozzle guide vane segment is tional axis ;
axially supported at the structure . fixing the turbine nozzle guide vane segment in a
5. The gas turbine according to claim 1 , wherein , at the circumferential direction with respect to the rota
outer platform , the turbine nozzle guide vane segment forms tional axis;
a radially extending outer platform wall that comprises an 20 fixing the turbine nozzle guide vane segment in the axial
axially supported sealing edge that extends in a circumfer direction at the outer housing and at the inner casing ;
ential direction with respect to the rotational axis . and
6. The gas turbine according to claim 1 , wherein the fixing the structure in the radial direction via a portion of
turbine nozzle guide vane segment is fixed at the outer the structure being inserted between the two side walls
housing in the radial direction and in a circumferential 25 of the axial groove.
direction with respect to the rotational axis, and wherein 11. The gas turbine according to claim 1 , wherein the
occurring radial and tangential loads are transferred into the structure is a shroud segment.
outer housing.
7. The gas turbine according to claim 1 , wherein the outer 12. The gas turbine according to claim 1 , wherein the first
housing is formed by the turbine casing and the combustion 30 turbine guide vane ring includes a plurality of turbine nozzle
chamber casing, which are connected to one another. guide vane segments , wherein the structure includes struc
8. The gas turbine according to claim 1 , wherein the ture hooks , and wherein the guide vane hooks and the
turbine nozzle guide vane segment is fixed directly the structure hooks alternately sit in the axial groove of the
outer housing in the radial direction. turbine housing
9. The gas turbine according to claim 1 , wherein the rotor 35 13. The gas turbine according to claim 1 , wherein the
blade is formed without a tip shroud. structure includes an abradable material facing the main
10. A method for attaching a turbine nozzle guide vane flow path .
segment of a gas turbine, comprising: 14. The gas turbine according to claim 1 , wherein the
providing the gas turbine, including: turbine guide vane segment is fixed in a circumferential
a combustion chamber including an inner casing ; 40
direction with respect to the rotational axis via at least one
a high -pressure turbine including: pin .
a turbine casing including an axial groove with 15. The gas turbine according to claim 1 , wherein the
respect to a rotational axis of the gas turbine; the structure is supported by a side surface of the turbine hook
axial groove being defined by two opposing side in a circumferential direction with respect to the rotational
walls and a connecting base surface and 45
axis and the structure is fixed in the circumferential direc
a first turbine guide vane ring that is arranged tion .
downstream of the combustion chamber, and
wherein the first turbine guide vane ring includes