US2656147A - Cooling of gas turbine rotors - Google Patents

Cooling of gas turbine rotors Download PDF

Info

Publication number
US2656147A
US2656147A US778618A US77861847A US2656147A US 2656147 A US2656147 A US 2656147A US 778618 A US778618 A US 778618A US 77861847 A US77861847 A US 77861847A US 2656147 A US2656147 A US 2656147A
Authority
US
United States
Prior art keywords
rotor
cooling
blade
blades
gas turbine
Prior art date
Legal status (The legal status is an assumption and is not a legal conclusion. Google has not performed a legal analysis and makes no representation as to the accuracy of the status listed.)
Expired - Lifetime
Application number
US778618A
Inventor
Brownhill Frank Denison
Wolff Paul Heinz Walter
Current Assignee (The listed assignees may be inaccurate. Google has not performed a legal analysis and makes no representation or warranty as to the accuracy of the list.)
English Electric Co Ltd
Original Assignee
English Electric Co Ltd
Priority date (The priority date is an assumption and is not a legal conclusion. Google has not performed a legal analysis and makes no representation as to the accuracy of the date listed.)
Filing date
Publication date
Application filed by English Electric Co Ltd filed Critical English Electric Co Ltd
Application granted granted Critical
Publication of US2656147A publication Critical patent/US2656147A/en
Anticipated expiration legal-status Critical
Expired - Lifetime legal-status Critical Current

Links

Images

Classifications

    • F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
    • F01—MACHINES OR ENGINES IN GENERAL; ENGINE PLANTS IN GENERAL; STEAM ENGINES
    • F01D—NON-POSITIVE DISPLACEMENT MACHINES OR ENGINES, e.g. STEAM TURBINES
    • F01D5/00—Blades; Blade-carrying members; Heating, heat-insulating, cooling or antivibration means on the blades or the members
    • F01D5/02—Blade-carrying members, e.g. rotors
    • F01D5/08—Heating, heat-insulating or cooling means
    • F01D5/081—Cooling fluid being directed on the side of the rotor disc or at the roots of the blades
    • F01D5/084—Cooling fluid being directed on the side of the rotor disc or at the roots of the blades the fluid circulating at the periphery of a multistage rotor, e.g. of drum type
    • F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
    • F01—MACHINES OR ENGINES IN GENERAL; ENGINE PLANTS IN GENERAL; STEAM ENGINES
    • F01D—NON-POSITIVE DISPLACEMENT MACHINES OR ENGINES, e.g. STEAM TURBINES
    • F01D11/00—Preventing or minimising internal leakage of working-fluid, e.g. between stages
    • F01D11/001—Preventing or minimising internal leakage of working-fluid, e.g. between stages for sealing space between stator blade and rotor
    • F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
    • F05—INDEXING SCHEMES RELATING TO ENGINES OR PUMPS IN VARIOUS SUBCLASSES OF CLASSES F01-F04
    • F05D—INDEXING SCHEME FOR ASPECTS RELATING TO NON-POSITIVE-DISPLACEMENT MACHINES OR ENGINES, GAS-TURBINES OR JET-PROPULSION PLANTS
    • F05D2240/00—Components
    • F05D2240/80—Platforms for stationary or moving blades
    • F05D2240/81—Cooled platforms

Definitions

  • the invention relates to cooling of the rotors of multi-stage gas turbines and has the object to provide a gas turbine rotor construction in which the only surfaces in contact with the hot working fluid are those of the blades proper and the circumference of a drum-shaped portion intermediate adjacent rows of rotor blades.
  • the rotor discs that carry the blades are intended to work at considerably lower temperatures and are required to be free from thermal stresses such as would occur in a disc which is heated at the rim but remains comparatively cold at the centre.
  • the main body of the rotor and the outer portion of the rotor which comprises the turbine blades, a series of blade platforms at the inner extremity of the blade profiles and a series of independent platforms located between the blade platforms of adjacent blade rings at the same diameter as the said blade platforms enclose between them a passage for a continuous and undiluted stream of cooling air forming an insulating wall between said main body and outer portion of the rotor, the only connection through the insulating stream of air being the lower portions of the blades which, between the said blade platform and the serrated roots by means of which they are secured into the body of the rotor, are reduced in cross-sectional area and made of sufiicient length to offer a substantial resistance to the flow of heat between the said outer hot portion of the rotor and the said main body, and the only connection between the said independent platforms and the said body of the rotor being through locating arrangements of relatively small area capable of offering considerable resistance to the flow of heat.
  • the invention is applicable both to axially inserted blades and to circumferential fixtures, both types having preferably serrated fixings of the so-called fir tree type.
  • a reduced neck is formed between the blade root platform and the serrated fir tree stub so as to provide a narrow neck for the fiow of heat and at the same time a recess through which the cooling medium can flow.
  • the cooling medium is preferably a small percentage of air tapped oif from the turbine compressor and emerges into the bladed zone of the turbine on the leading edge side of the rotor blades of each stage so as to form a comparatively cool boundary layer on the blade root before mixing with the combustion gases.
  • high pressure cooling air is introduced into the insulating space 12 Claims. (01. 253-3915) i at one of the stages at the low pressure end of the rotor only, preferably the lowest stage, and then passes along substantially the whole length of the rotor in the insulating space through passages formed between adjacent blade stubs and between the reduced area portions of the blade root, these passages being shaped and proportioned so as to ensure a, high degree of heat abstraction from the blade roots and from the main body of the rotor. After passing along substantially the whole length of the rotor the cooling air isthen discharged into the main work ing fluid on the upstream side of one of the early stages of rotor blades, preferably the first stage.
  • This embodiment permits the economical use of larger total quantities of cooling air since the cooling air is returned to do useful work in the turbine at a pressure equal to that of the main working fluid after passing through the first stage nozzles.
  • This together with the fact that the whole of the cooling air passes through each row of blade roots permits the use of compara tively large passages between them and thu min; imizes the danger of the passages becoming choked by foreign matter and at the same time promotes more efiicient cooling due to theincreased surface area.
  • V A further advantage of this embodiment is that heat abstracted by the cooling air from the low pressure portions of the turbine is returned to the cycle at a higher pressure and therefore with higher availability for conversion into work thus, making the cycle to some extent regenerative and increasing its efiiciency.
  • Fig. 1 is a side elevation in section of part of three consecutive rotor discs having axially inserted blades with series flow cooling means ar-.. ranged between adjacent discs, 1
  • Fig. 2 is part 'of a front elevation of a rotor disc and intermediate pieces corresponding to Fig. 1, g Fig. 3 is a side elevation in section of part of two consecutive rotor discs having axially inserted blades with parallel flow cooling means'ar.
  • the radial wall 32 forms a lip that leaves a comparatively narrow gap 36 between its end and the leading edge of the platform 2% of the adjacent blade 22.
  • the other end of the bridge plate 31 has a radially inward pointing extension serrated-with the same fir tree profile as the blade attachment stubs 22a and of an axial dimension slightly smaller than the aforesaid circumferential groove 39b in the rotor disc.
  • a rib 34 joins the platform 31, the serrated extension 35 and the wall portion 32.
  • Adjacent platform sections 21 may be joined either by means of a spigotted joint 2162 (Fig.
  • the aforesaid separate intermediate pieces 21 are dropped with their serrated radial extensions 35 into the said circumferential groove 39a of the disc and then shifted in an axial direction until the projection 33 of the radial wall portion 32 on the other side of the bridge piece -31 engages the recess 26 between the flanges lel streams in a direction opposite to the flow of the combustion gases.
  • These streams mix in another annular cavity 3
  • the cooling medium emerges from this second annular cavity 3
  • , M respectively, are preferably oblique, i. e.
  • the path of air (Figs. 4a and 4b) in this embodiment is preferably radially outward in the cavity between two adjacent discs 4 4 and 4
  • the air flows then radially inward through a recess left in the radial faces of adjacent blade attachment stubs 42a until it reaches another circumferential channel left at 43 between the inner end of the blade attachment stubs 42a. and the bottom of the circumferential groove lla left in the rim of the disc.
  • are again provided with spigotted flanges 56a, 591), respectively, which after assembly are welded with one another by the seam 59.
  • is connected with the source of cooling medium, preferably unburnt air of a pressure at least equal to the highest pressure occurring in the gas turbine cycle.
  • is provided on the downstream side of disc 5
  • Channels 53 and 54 are formed under the inner ends and under the platforms 52b of the blades, corresponding to the channels'f's, 24, respectively, of the embodiment of Fig. 3.
  • the leading edges 52d of blades 52 have a spigotted joint with the flange 5!] of the intermediate platform pieces 51 which correspond otherwise substantially to the intermediate platform pieces 21 of Figs. 3 and 3b, in that they engage with their extensions 63 the grooves 56 between the circumferential projections 55, 55a of the discs 5
  • the radial wall portions 62 of the intermediate platform pieces 5'! have wide openings 52a.
  • this arrangement provides a series flow of the cooling medium from space 59 at the downstream side of the lowest stage through the channels 53, 5d holes 62a under the platforms 61 of the intermediate pieces 51 through the channels 53, 54 of the next higher stage, and so on, until the cooling medium is lasso, 147
  • Inxa multistage 'gas.turbineaaseset forthsin claim 14 317 least one. of. the said intermediateplat- .formxpiecescomprising two parts.:of-. a'.total: axial dimension. smaller than the 'distancebetween adiacent 'dlSCS'.S0"8.-S 'to beinserted iradially and :then shifted 1 axially into engagement with the .respective "adjacent disc, and wedge rmembers :drivenrradially between xthe-zsaidi partsro'f platformpieces and subsequently .welded .to timer-said parts.
  • cylindrical flanges on the said discs forming inner drums the outside radius of which is less than the radius defining the bottoms'of the recesses in the said outer portions, at least one of the said flanges being perforated by holes adapted to admit metered quantities of cooling medium from the space within the said cylindrical flanges to the said insulating spaces arranged in separate groups of the outer rotor portion, each group comprising at least one row of the said blade roots, blade platforms and intermediate platforms and'including the said passages for the cooling medium, the said cooling medium being discharged thereafter into the main Working fluid stream passing the blades proper of the gas turbine.
  • the said means for the supply of coolingmedium being connected to the insulating space on the downstream side of the lowest pressure stage of the rotor, the said means passing the said cooling medium consecutively through the 10 insulating spaces and passages 01' all stages of the said rotor, and eventually discharging it into the said main working fluid stream at the upstream side of the rotor blades of the highest pressure stage.
  • a hollow central portion of the rotor in supply connection on the one hand with a source of coolin medium, and on the other hand with the said insulating space of at least one rotor stage.

Landscapes

  • Engineering & Computer Science (AREA)
  • Mechanical Engineering (AREA)
  • General Engineering & Computer Science (AREA)
  • Turbine Rotor Nozzle Sealing (AREA)

Description

Oct. 20, 1953 F. D. BROWNHILL ET AL COOLING OF GAS TURBINE ROTORS 8 Sheets-Sheet 2 Filed 001;. 8, 1947 FIG. 2.
mwaw
CoMPmsssoR Oct. 20, 1953 F. D. BROWNHILL ETAL COOLING OF GAS TURBINE ROTORS v Filed 001'.- 8, 1947 8 Sheets-Sheet 3 Oct. 20, 19,53 F. D. BR OWNH ILL ET AL 2,656,147
COOLING OF GAS TURBINE ROTORS F iled Oct. 8, 1947 8 Sheets-Sheet 4 9, 1953 F. D. BROWNHILL ET AL 2,656,147
I COOLING OF GAS TURBINE ROTORS Filed Oct. 8, 1947 8 Sheets-Sheet 5 Inventors: yrmw 2 F. D. BROWNHILL ETAL 47 COOLING oF'cAs TURBINE ROTORS Oct. 20, 1953 8 Sheets-Sheet 7 Filed 001;. 8, 1947 1953 F. D. BROWNHILL ET AL 6,
COOLING 0F GAS TURBINE RoToR Patented Oct. 20, 1953 COOLING OF GAS TURBINE ROTORS Frank Denison Brownhill and Paul Heinz Walter Wolfl", Rugby, England, assignors to The English Electric Company Limited, London, England, a British company Application October 8, 1947, Serial No. 778,618 In Great Britain October 9, 1946 The invention relates to cooling of the rotors of multi-stage gas turbines and has the object to provide a gas turbine rotor construction in which the only surfaces in contact with the hot working fluid are those of the blades proper and the circumference of a drum-shaped portion intermediate adjacent rows of rotor blades. The rotor discs that carry the blades are intended to work at considerably lower temperatures and are required to be free from thermal stresses such as would occur in a disc which is heated at the rim but remains comparatively cold at the centre.
According to the invention, in a multi-stage gas turbine the main body of the rotor and the outer portion of the rotor, which comprises the turbine blades, a series of blade platforms at the inner extremity of the blade profiles and a series of independent platforms located between the blade platforms of adjacent blade rings at the same diameter as the said blade platforms enclose between them a passage for a continuous and undiluted stream of cooling air forming an insulating wall between said main body and outer portion of the rotor, the only connection through the insulating stream of air being the lower portions of the blades which, between the said blade platform and the serrated roots by means of which they are secured into the body of the rotor, are reduced in cross-sectional area and made of sufiicient length to offer a substantial resistance to the flow of heat between the said outer hot portion of the rotor and the said main body, and the only connection between the said independent platforms and the said body of the rotor being through locating arrangements of relatively small area capable of offering considerable resistance to the flow of heat.
The invention is applicable both to axially inserted blades and to circumferential fixtures, both types having preferably serrated fixings of the so-called fir tree type. A reduced neck is formed between the blade root platform and the serrated fir tree stub so as to provide a narrow neck for the fiow of heat and at the same time a recess through which the cooling medium can flow. The cooling medium is preferably a small percentage of air tapped oif from the turbine compressor and emerges into the bladed zone of the turbine on the leading edge side of the rotor blades of each stage so as to form a comparatively cool boundary layer on the blade root before mixing with the combustion gases.
- In another form of embodiment high pressure cooling air is introduced into the insulating space 12 Claims. (01. 253-3915) i at one of the stages at the low pressure end of the rotor only, preferably the lowest stage, and then passes along substantially the whole length of the rotor in the insulating space through passages formed between adjacent blade stubs and between the reduced area portions of the blade root, these passages being shaped and proportioned so as to ensure a, high degree of heat abstraction from the blade roots and from the main body of the rotor. After passing along substantially the whole length of the rotor the cooling air isthen discharged into the main work ing fluid on the upstream side of one of the early stages of rotor blades, preferably the first stage.
This embodiment permits the economical use of larger total quantities of cooling air since the cooling air is returned to do useful work in the turbine at a pressure equal to that of the main working fluid after passing through the first stage nozzles. This together with the fact that the whole of the cooling air passes through each row of blade roots permits the use of compara tively large passages between them and thu min; imizes the danger of the passages becoming choked by foreign matter and at the same time promotes more efiicient cooling due to theincreased surface area.
V A further advantage of this embodiment is that heat abstracted by the cooling air from the low pressure portions of the turbine is returned to the cycle at a higher pressure and therefore with higher availability for conversion into work thus, making the cycle to some extent regenerative and increasing its efiiciency.
In order to be better understood and readily: carried into effect, the invention is illustrated by way of example in the accompanying drawings of which:
Fig. 1 is a side elevation in section of part of three consecutive rotor discs having axially inserted blades with series flow cooling means ar-.. ranged between adjacent discs, 1
Fig. 2 is part 'of a front elevation of a rotor disc and intermediate pieces corresponding to Fig. 1, g Fig. 3 is a side elevation in section of part of two consecutive rotor discs having axially inserted blades with parallel flow cooling means'ar.
Fig. 2,
nular space between themselves and the flanges a and 20b of the adjacent discs 2|, 2|. end of that bridge portion 3'! on the side of the The radial wall 32 forms a lip that leaves a comparatively narrow gap 36 between its end and the leading edge of the platform 2% of the adjacent blade 22. The other end of the bridge plate 31 has a radially inward pointing extension serrated-with the same fir tree profile as the blade attachment stubs 22a and of an axial dimension slightly smaller than the aforesaid circumferential groove 39b in the rotor disc. A rib 34 joins the platform 31, the serrated extension 35 and the wall portion 32. Adjacent platform sections 21 may be joined either by means of a spigotted joint 2162 (Fig. 3d) or by means of an inserted metal strip 2le (Fig. 36) forming a seal between the -annular space 28 and the stream of combustion gases passing the blades 22, 22', which will remain substantially airtight and will at the same time permit expansion of the said platform sections.
Before the assembly of the blades with the rotor the aforesaid separate intermediate pieces 21 are dropped with their serrated radial extensions 35 into the said circumferential groove 39a of the disc and then shifted in an axial direction until the projection 33 of the radial wall portion 32 on the other side of the bridge piece -31 engages the recess 26 between the flanges lel streams in a direction opposite to the flow of the combustion gases. These streams mix in another annular cavity 3| formed between the rim of each disc and a ring-shaped wall formed by the wall portions 32 of all of the separate intermediate pieces 21. The cooling medium emerges from this second annular cavity 3| through a circumferential gap 36 between the aforesaid lip :portions 39 of the intermediate pieces 2? and the :leading edges of the blade platforms 22b so as .to form a comparatively cool boundary layer covering the blade platforms before mixing with the hot combustion gases.
, Referring now to Figs. 4 to 4c showing circumferential fixtures, the intermediate pieces 41, bridging the space between two adjacent discs 4|, M respectively, are preferably oblique, i. e.
approximately of parallelogram shape in plan view. These pieces di are again secured by the engagement of circumferential recesses with circumferential flanges 45, a, 4552) respectively on .oppositefaces of adjacent discs. The last sec- 'tion lla (Fig. 4, right hand side) to be inserted is again of an axial dimension which is shorter than the distance between adjacent discs M, 31 so that it can be inserted radially and shifted axially to engagement with the one 4| of the adjacent discs, the remaining gap being closed by the radial insertion of a recessed piece 55 which can also be dropped in radially and shifted axially until it engages the other adjacent disc il", the remaining gap between these two por- :tions being filled by a wedge-shaped member iii, the two portions 46a, 56 and the said wedge- 6 shaped member 5i being eventually welded toether at 52.
The path of air (Figs. 4a and 4b) in this embodiment is preferably radially outward in the cavity between two adjacent discs 4 4 and 4|, 4| respectively, then at 46 in an axial direction between two adjacent intermediate pieces 41 into a circumferential space Ma left between the platform 4212 of the blades and the outer circumference of the rim of the disc. The air flows then radially inward through a recess left in the radial faces of adjacent blade attachment stubs 42a until it reaches another circumferential channel left at 43 between the inner end of the blade attachment stubs 42a. and the bottom of the circumferential groove lla left in the rim of the disc. From there the cooling air moves again radially outward in a recess left between the next pair of adjacent blade attachment stubs 420, into a circumferential channel 44b left between the outer circumference of the rim of the disc and the platform 42b of the blades, and from there in a substantially radial direction through a hole 420 arranged in the leading edge side of the blade root platform 321) so that the emerging air again forms a boundary layer of comparatively low temperature along the platforms of the rotor blades before mixing with the combustion gases.
Referring now to Fig. 5 the adjacent rotor discs 5| and 5| are again provided with spigotted flanges 56a, 591), respectively, which after assembly are welded with one another by the seam 59. The space thus enclosed between the extreme low pressure stage disc 51 and the adjacent disc 5| is connected with the source of cooling medium, preferably unburnt air of a pressure at least equal to the highest pressure occurring in the gas turbine cycle. A flange 7| is provided on the downstream side of disc 5| and is equipped with an inward axial ledge 12. Pieces in similar in shape to the intermediate platform pieces 2'! of Figs. 3 and 321 but generally shorter in the axial dimension are adapted to be inserted radially with their serrated inward extension 75 into the circumferential recess 59a of the disc 5| and shifted axially into the position shown in J'ig. 5 in which they engage with their extensions 73 the ledge 72 of the flange 1|. In this position they are held by the axial insertion from the left of the blades 52 having correspondingly serrated roots. The pieces 19 enclose between themselves and the disc 5| an air space 69 communicating with the space between the discs 5| and 5| by ducts 68.
Channels 53 and 54 are formed under the inner ends and under the platforms 52b of the blades, corresponding to the channels'f's, 24, respectively, of the embodiment of Fig. 3. The leading edges 52d of blades 52 have a spigotted joint with the flange 5!] of the intermediate platform pieces 51 which correspond otherwise substantially to the intermediate platform pieces 21 of Figs. 3 and 3b, in that they engage with their extensions 63 the grooves 56 between the circumferential projections 55, 55a of the discs 5|, 5| etc. The radial wall portions 62 of the intermediate platform pieces 5'! have wide openings 52a.
As will be seen from Fig. 5, this arrangement provides a series flow of the cooling medium from space 59 at the downstream side of the lowest stage through the channels 53, 5d holes 62a under the platforms 61 of the intermediate pieces 51 through the channels 53, 54 of the next higher stage, and so on, until the cooling medium is lasso, 147
channels 153, 5.4 :of f'liheli highest zpressureustage ihomever ercorresponding. otherwise ;tov zthe end zrpieces l0: atithe; low; ressure. end.
zlnrallithe .various embodiments::referred: to *hereabove, severe temperatureegradients inthe rroton can be -avoided'by applyingrcooling medhim ratzthee blade; roots.
Z'I'hustthe.rotorisrkeptrcool enoughtto minimise -creep in thezblade-attachment stubs, even when msmgwcomparatively low grade material L'for :the motor. Creep in the :blade .rootsthemselves 'is ssalsoz-rkept" to (:2, :min-imum. By controlling :the irpresslrre. of therxcooling: medium, any: inward I flow not the hot gasesis. prevented. As. stated aboveca aboundary-layerof cool. airrover'. the outside ofthe blades): and of "the: intermediate platforms; is set by the escaping; cooling; medium.
.The arim of. the :rotor discs ;of:.. any ".of the serfrhodiments xm-ay be rdivided into. segments by .YfidifilslSIOtS 1(notrshown) :in a way known in 2itselt,.:in order to.:allow anindependent thermal .7
.iexpansionofithat portionrwithout setting upondniy highicomp-ressivestresses.
' iiByitheaforesaid improvements .quick starting and changes of load and temperature ofthegas :turbine are: permissible.
:ZIhe 'construction :described: is "applicable also toztaperech rotors. Solid: forged. orbuilt up. rotors .maythe: used=and,: aseshown in the. embodiments described, :with .reierence to v:Figs. i3 and 15 the welding seamscof .built up; rotorszican be kept out =otrcontactnvith therstreami;of:comhustion gases, :a design which; isapplicable also sto other emi-hodim-ents.
The-method or securing the intermediate pieces adjacent blade rows.
What- We claim-asour' invention-and "desire to secure by' Letters Patentv is:
1. In "a multistage gas turbinea rotor coin- :prisingseveral rotor discsintegral with one another and each'coinprising a mainbody, proflled projections. of a cross sectional area restricted ascompared with that of'the said main :body extending from-its circumference, and an outer ;.portion comprising a. rowof blades. each blade-having an airfoileportion, a platform portion'and a root section of a crosssectional' area comparatively restricted adjacent. the said.platform portion and comparatively largerin. a portion profiled complementary to .and. engaging with the saidproflled projectionsofithe main "body; and separate platforms arranged, between the platform portions of adjacent rows of blades -:at" the same diameter asthe said platform portions, complementary coolant conducting .pas-
sagewaysformed in said profiled projections and said Tootsections'the-said root sections, plat- "form *portions and separate platforms forming enclosed passages between "them "and 'the said main body, means connecting said passages and passageways for conducting acontinuous stream of cooling air forming a heat insulating layer between the-said-main body andthe said outer "portiom the said root sections-forming the only integral connection between thesaid airfoil porad'apted' to be axially I inserted into :theesaid recessed and serrated projectionya circumferential :groove in the said projectionand an :opposi'te groove in each. blade root, theseegrooves' forming together a circumferential ductgthe said intermediate platforms having equallyserrated projections pointing inwardsof =an axialedimension I shorter than the. said a groove -.in *the rotor, radially inserted-into f the said: groove and axially moved into engagement with the: serrated 11'6- cesses of the rotor. and held-.therebythe serrated :bla'd'e roots, the .total axialilengthiof :the said serrated projections of thenintermediate :platforms plus the serrated blade rootsi-being .=substantially equal to 'theaXiaILIengthaof-"the F581?- rated recesses in the said projection: of. the:
rotor body.
3'. "In a multistage gas turbine asiset claim 1 -'cy1in'drical "flanges .on 517119 :si'des of a the said discs, adjacent flanges bemg-weldedstozgethen'the said intermediate:pIatformsrseparat ing the said welded flanges from the streamtof hot workingfluidipas'sing through the said-trotor :blades.
14. ?In :a :multistage gas turbine :a motor. :as claimed in 'claim 11 wherein the 1 said separate latforms' are anchoredini thesaidprojections of the main bodyz'andj form with; the adjacent rotor discs substantially closed i'drums exposed iinside to :the cooling; airflowing through -:thezrsaid ;pas-
sages.
5. Inxa multistage 'gas.turbineaaseset forthsin claim 14 317 least one. of. the said intermediateplat- .formxpiecescomprising two parts.:of-. a'.total: axial dimension. smaller than the 'distancebetween adiacent 'dlSCS'.S0"8.-S 'to beinserted iradially and :then shifted 1 axially into engagement with the .respective "adjacent disc, and wedge rmembers :drivenrradially between xthe-zsaidi partsro'f platformpieces and subsequently .welded .to timer-said parts.
6.. In: a multistage gas turbinezas; set. forth: in ciaim 4::meanstfor .circulatingca gaseousaicool'ing zmediumisuchr as compressed air-sthroughnthe said "passages and past: the. inside of the said intermediate platforms.
'i. .Inamultistage-gasxturbine aszset forth: in
:claim .6, the said tblades .tbeing iinserted axially into the said; prOjECti'OIillO'f the 'main'. rotortbody, "andthe said meansiforrcirculating thezcooling :medium directing the same; firstly axially; the
direction. of the main working :fluid -stream through. the said passages :under the inner ends of the blade roots of arotor. stage, then deflecting thesaid :cooling medium rradially. outwards, and
.directing the sai'd'cooli-ng' medium in a direction opposed to' that of the main working fluid stream through the said passages underthe'blade platin: the directionyiopposite :to: that :of the main working fluid; stream: both: through.' the saidma'ssages under theinnerblade root endsof a'rrotor stage and through theasaid passages .underith'e blade: platforms of; the: saidrotorstage, and eventually discharging-the:coolingmediumxatthe upstream side of the. rotor-blades voffthezsaid rotor s age.
9. In a multistage gas turbine as set forth in claim 6 cylindrical flanges on the said discs forming inner drums the outside radius of which is less than the radius defining the bottoms'of the recesses in the said outer portions, at least one of the said flanges being perforated by holes adapted to admit metered quantities of cooling medium from the space within the said cylindrical flanges to the said insulating spaces arranged in separate groups of the outer rotor portion, each group comprising at least one row of the said blade roots, blade platforms and intermediate platforms and'including the said passages for the cooling medium, the said cooling medium being discharged thereafter into the main Working fluid stream passing the blades proper of the gas turbine.
10. In a multistage gas turbine as set forth in claim 9 means for connecting the said insulating pressure stage, and the said cooling medium being discharged into the said main working fluid stream at the upstream side of the rotor blades of a higher pressure stage.
11. In a multistage gas turbine as set forth in claim 10 the said means for the supply of coolingmedium being connected to the insulating space on the downstream side of the lowest pressure stage of the rotor, the said means passing the said cooling medium consecutively through the 10 insulating spaces and passages 01' all stages of the said rotor, and eventually discharging it into the said main working fluid stream at the upstream side of the rotor blades of the highest pressure stage.
12. In a multistage gas turbine as set forth in claim 10 a hollow central portion of the rotor in supply connection on the one hand with a source of coolin medium, and on the other hand with the said insulating space of at least one rotor stage.
FRANK DENISON BROWNHILL. PAUL I-IEINZ WALTER WOLFF.
References Cited in the file of this patent UNITED STATES PATENTS Number Name Date 2,141,401 Martinka Dec. 27, 1938 2,149,510 Darrieus Mar. 7, 1939 2,213,940 Jendrassik Sept. 3, 1940 2,241,782 Jendrassik May 13, 1941 2,440,069 Bloomberg Apr. 20, 1948 2,452,782 McLeod et a1. Nov. 2, 1948 2,489,683 Stalker Nov. 29, 1949 FOREIGN PATENTS Number Country Date 452,412 Great Britain Aug. 24, 1936 543,985 Great Britain Mar. 23, 1942 578,191 Great Britain June 19, 1946 579.316 Great Britain July 31, 1946
US778618A 1946-10-09 1947-10-08 Cooling of gas turbine rotors Expired - Lifetime US2656147A (en)

Applications Claiming Priority (1)

Application Number Priority Date Filing Date Title
GB278723X 1946-10-09

Publications (1)

Publication Number Publication Date
US2656147A true US2656147A (en) 1953-10-20

Family

ID=10266651

Family Applications (1)

Application Number Title Priority Date Filing Date
US778618A Expired - Lifetime US2656147A (en) 1946-10-09 1947-10-08 Cooling of gas turbine rotors

Country Status (4)

Country Link
US (1) US2656147A (en)
CH (1) CH278723A (en)
DE (1) DE971297C (en)
GB (1) GB612097A (en)

Cited By (92)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
US2807434A (en) * 1952-04-22 1957-09-24 Gen Motors Corp Turbine rotor assembly
US2858103A (en) * 1956-03-26 1958-10-28 Westinghouse Electric Corp Gas turbine apparatus
DE1043718B (en) * 1956-07-31 1958-11-13 Maschf Augsburg Nuernberg Ag Axially loaded turbine runner with cooling by a gaseous coolant, especially for gas turbines
US2931624A (en) * 1957-05-08 1960-04-05 Orenda Engines Ltd Gas turbine blade
US2972470A (en) * 1958-11-03 1961-02-21 Gen Motors Corp Turbine construction
US2974925A (en) * 1957-02-11 1961-03-14 John C Freche External liquid-spray cooling of turbine blades
US2996280A (en) * 1959-04-07 1961-08-15 Iii John A Wilson Heat shield
US3034763A (en) * 1959-08-20 1962-05-15 United Aircraft Corp Rotor construction
US3139310A (en) * 1961-12-29 1964-06-30 Svenska Flaektfabriken Ab Arrangement in axial fans for the transport of dust commingled gases
US3356340A (en) * 1965-03-15 1967-12-05 Gen Electric Turbine rotor constructions
US3437313A (en) * 1966-05-18 1969-04-08 Bristol Siddeley Engines Ltd Gas turbine blade cooling
US3689177A (en) * 1971-04-19 1972-09-05 Gen Electric Blade constraining structure
US3730644A (en) * 1969-06-26 1973-05-01 Rolls Royce Gas turbine engine
US4035102A (en) * 1975-04-01 1977-07-12 Kraftwerk Union Aktiengesellschaft Gas turbine of disc-type construction
US4484858A (en) * 1981-12-03 1984-11-27 Hitachi, Ltd. Turbine rotor with means for preventing air leaks through outward end of spacer
FR2557205A1 (en) * 1983-12-22 1985-06-28 United Technologies Corp ROTOR WITH DOUBLE PASS COOLING OF THE HEELS OF THE AUBES
US4536129A (en) * 1984-06-15 1985-08-20 United Technologies Corporation Turbine blade with disk rim shield
US4551063A (en) * 1983-03-18 1985-11-05 Kraftwerke Union Ag Medium-pressure steam turbine
US4582467A (en) * 1983-12-22 1986-04-15 United Technologies Corporation Two stage rotor assembly with improved coolant flow
US4645424A (en) * 1984-07-23 1987-02-24 United Technologies Corporation Rotating seal for gas turbine engine
US4648793A (en) * 1985-05-31 1987-03-10 General Electric Company Turbine wheel key and keyway ventilation
US4659289A (en) * 1984-07-23 1987-04-21 United Technologies Corporation Turbine side plate assembly
US4730982A (en) * 1986-06-18 1988-03-15 Societe Nationale D'etude Et De Construction De Moteurs D'aviation "S.N.E.C.M.A." Assembly for controlling the flow of cooling air in an engine turbine
US4795307A (en) * 1986-02-28 1989-01-03 Mtu Motoren- Und Turbinen-Union Munchen Gmbh Method and apparatus for optimizing the vane clearance in a multi-stage axial flow compressor of a gas turbine
US5193982A (en) * 1991-07-17 1993-03-16 Societe Nationale D'etude Et De Construction De Moteurs D'aviation S.N.E.C.M.A. Separate inter-blade platform for a bladed rotor disk
US5271718A (en) * 1992-08-11 1993-12-21 General Electric Company Lightweight platform blade
US5507620A (en) * 1993-07-17 1996-04-16 Abb Management Ag Gas turbine with cooled rotor
US5525032A (en) * 1994-04-02 1996-06-11 Abb Management Ag Process for the operation of a fluid flow engine
EP0735238A1 (en) * 1995-03-31 1996-10-02 General Electric Company Closed or open circuit cooling of turbine rotor components
US5611669A (en) * 1994-09-27 1997-03-18 Eupopean Gas Turbines Limited Turbines with platforms between stages
EP0921273A1 (en) * 1997-06-11 1999-06-09 Mitsubishi Heavy Industries, Ltd. Rotor for gas turbines
US5993154A (en) * 1996-11-21 1999-11-30 Asea Brown Boveri Ag Welded rotor of a turbo-engine
US6457935B1 (en) * 2000-06-15 2002-10-01 Snecma Moteurs System for ventilating a pair of juxtaposed vane platforms
US20040035118A1 (en) * 2002-08-20 2004-02-26 Alm Development, Inc. Blade cooling in a gas turbine engine
US20040081556A1 (en) * 2002-10-24 2004-04-29 Andre Chevrefils Blade passive cooling feature
US20040165983A1 (en) * 2003-02-26 2004-08-26 Rolls-Royce Plc Damper seal
US20040253113A1 (en) * 2003-06-16 2004-12-16 Snecma Moteurs Retention capacity of a blade having an asymmetrical hammerhead fastener, with the help of platform stiffeners
US20050058545A1 (en) * 2003-09-12 2005-03-17 Siemens Westinghouse Power Corporation Turbine blade platform cooling system
US20050111970A1 (en) * 2003-11-26 2005-05-26 Gabriel Suciu Turbine durm rotor for a turbine engine
GB2409240A (en) * 2003-12-18 2005-06-22 Rolls Royce Plc Cooling arrangement
US20050180847A1 (en) * 2004-02-14 2005-08-18 Alstom Technology Ltd Rotor
US20060024164A1 (en) * 2004-07-30 2006-02-02 Keith Sean R Method and apparatus for cooling gas turbine engine rotor blades
JP2007332973A (en) * 2007-08-24 2007-12-27 Mitsubishi Heavy Ind Ltd Gas turbine
US20080240927A1 (en) * 2006-10-16 2008-10-02 Katharina Bergander Turbine blade for a turbine with a cooling medium passage
US20100290922A1 (en) * 2008-02-27 2010-11-18 Mitsubisihi Heavy Industries, Ltd Turbine disk and gas turbine
US20110016884A1 (en) * 2008-03-28 2011-01-27 Mitsubishi Heavy Industries, Ltd. Cooling passage cover, manufacturing method of the cover, and gas turbine
US20110061395A1 (en) * 2009-09-13 2011-03-17 Kendrick Donald W Method of fuel staging in combustion apparatus
FR2954797A1 (en) * 2009-12-29 2011-07-01 Snecma Low pressure turbine rotor for two-shaft gas turbine engine of aircraft, has elastic sealing unit fixed on each of blade roots and projected with respect to rear transverse face of disk at level of joints so as to be in contact with flange
US20110164982A1 (en) * 2010-01-06 2011-07-07 General Electric Company Apparatus and method for a low distortion weld for rotors
US20120027584A1 (en) * 2010-08-02 2012-02-02 General Electric Company Turbine seal system
US20120134778A1 (en) * 2010-11-29 2012-05-31 Alexander Anatolievich Khanin Axial flow gas turbine
WO2012076588A1 (en) * 2010-12-09 2012-06-14 Alstom Technology Ltd Fluid flow machine especially gas turbine penetrated axially by a hot gas stream
US20120148405A1 (en) * 2010-12-13 2012-06-14 General Electric Company Cooling circuit for a drum rotor
US20120321441A1 (en) * 2011-06-20 2012-12-20 Kenneth Moore Ventilated compressor rotor for a turbine engine and a turbine engine incorporating same
US20130017095A1 (en) * 2011-07-12 2013-01-17 Ching-Pang Lee Flow directing member for gas turbine engine
US20130108413A1 (en) * 2011-10-28 2013-05-02 Gabriel L. Suciu Secondary flow arrangement for slotted rotor
US20130108468A1 (en) * 2011-10-28 2013-05-02 Gabriel L. Suciu Spoked spacer for a gas turbine engine
RU2481481C2 (en) * 2007-07-06 2013-05-10 Снекма Air supply device for ventilation of blades of low pressure turbine of gas turbine engine; rotor of gas turbine engine, and gas turbine engine
US20130236289A1 (en) * 2012-03-12 2013-09-12 General Electric Company Turbine interstage seal system
US20130264779A1 (en) * 2012-04-10 2013-10-10 General Electric Company Segmented interstage seal system
US20130302143A1 (en) * 2010-12-14 2013-11-14 Rolls-Royce Deutschland Ltd & Co Kg Cooling device for a jet engine
US20140069101A1 (en) * 2012-09-13 2014-03-13 General Electric Company Compressor fairing segment
US20140212270A1 (en) * 2012-12-27 2014-07-31 United Technologies Corporation Gas turbine engine component having suction side cutback opening
US20140301851A1 (en) * 2013-04-08 2014-10-09 Alstom Technology Ltd Rotor
US20140334929A1 (en) * 2013-05-13 2014-11-13 General Electric Company Compressor rotor heat shield
US20140363307A1 (en) * 2013-06-05 2014-12-11 Siemens Aktiengesellschaft Rotor disc with fluid removal channels to enhance life of spindle bolt
US20150114001A1 (en) * 2013-10-28 2015-04-30 General Electric Company Sealing component for reducing secondary airflow in a turbine system
US9115587B2 (en) 2012-08-22 2015-08-25 Siemens Energy, Inc. Cooling air configuration in a gas turbine engine
US9145772B2 (en) 2012-01-31 2015-09-29 United Technologies Corporation Compressor disk bleed air scallops
JP2015200242A (en) * 2014-04-09 2015-11-12 株式会社東芝 Axial flow turbine
US20160053688A1 (en) * 2014-08-20 2016-02-25 United Technologies Corporation Gas turbine rotors
US20160153302A1 (en) * 2014-12-01 2016-06-02 General Electric Company Turbine wheel cover-plate mounted gas turbine interstage seal
WO2016087153A1 (en) * 2014-12-04 2016-06-09 Siemens Aktiengesellschaft Rotor, axial compressor, installation method
US20160186591A1 (en) * 2014-12-31 2016-06-30 General Electric Company Flowpath boundary and rotor assemblies in gas turbines
US20160186593A1 (en) * 2014-12-31 2016-06-30 General Electric Company Flowpath boundary and rotor assemblies in gas turbines
US20160230773A1 (en) * 2013-07-17 2016-08-11 Siemens Aktiengesellschaft Rotor for a thermal turbomachine
EP2484866A3 (en) * 2011-02-03 2017-03-15 General Electric Company Cross-over purge flow system for a turbomachine wheel member
EP2586969A3 (en) * 2011-10-28 2017-05-03 United Technologies Corporation Spoked Rotor for a Gas Turbine Engine
EP2586968A3 (en) * 2011-10-28 2017-05-03 United Technologies Corporation Secondary flow arrangement for slotted rotor
US20170211590A1 (en) * 2016-01-27 2017-07-27 General Electric Company Compressor Aft Rotor Rim Cooling for High OPR (T3) Engine
CN107002690A (en) * 2014-10-15 2017-08-01 赛峰航空器发动机 Runner assembly for the turbogenerator including self-supporting rotor case
US9790792B2 (en) 2011-10-28 2017-10-17 United Technologies Corporation Asymmetrically slotted rotor for a gas turbine engine
US9938831B2 (en) 2011-10-28 2018-04-10 United Technologies Corporation Spoked rotor for a gas turbine engine
US20180128114A1 (en) * 2016-11-10 2018-05-10 Doosan Heavy Industries & Construction Co., Ltd Structure for cooling rotor of turbomachine, rotor and turbomachine having the same
EP3447244A1 (en) * 2017-08-23 2019-02-27 Siemens Aktiengesellschaft Turbine rotor assembly with lap joints between the rotor discs for torque transmission
US10337345B2 (en) 2015-02-20 2019-07-02 General Electric Company Bucket mounted multi-stage turbine interstage seal and method of assembly
US10837288B2 (en) 2014-09-17 2020-11-17 Raytheon Technologies Corporation Secondary flowpath system for a gas turbine engine
US11041396B2 (en) * 2016-10-06 2021-06-22 Raytheon Technologies Corporation Axial-radial cooling slots on inner air seal
US11098604B2 (en) * 2016-10-06 2021-08-24 Raytheon Technologies Corporation Radial-axial cooling slots
US11215056B2 (en) 2020-04-09 2022-01-04 Raytheon Technologies Corporation Thermally isolated rotor systems and methods
US11339662B2 (en) * 2018-08-02 2022-05-24 Siemens Energy Global GmbH & Co. KG Rotor comprising a rotor component arranged between two rotor disks
US20240084708A1 (en) * 2016-02-05 2024-03-14 Siemens Energy Global GmbH & Co. KG Rotor comprising a rotor component arranged between two rotor discs

Families Citing this family (20)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
NL72215C (en) * 1949-04-29
BE501031A (en) * 1950-02-03
US2773667A (en) * 1950-02-08 1956-12-11 Gen Motors Corp Turbine rotor sealing ring
US2791091A (en) * 1950-05-15 1957-05-07 Gen Motors Corp Power plant cooling and thrust balancing systems
US2751189A (en) * 1950-09-08 1956-06-19 United Aircraft Corp Blade fastening means
GB699582A (en) * 1950-11-14 1953-11-11 Rolls Royce Improvements in or relating to gas-turbine engines
US2669383A (en) * 1951-02-06 1954-02-16 A V Roe Canada Ltd Rotor blade
GB742241A (en) * 1951-02-15 1955-12-21 Power Jets Res & Dev Ltd Improvements in the cooling of turbines
DE1075380B (en) * 1952-05-22 1960-02-11 Siemens-Schuckertwcrkc Aktiengesellschaft, Berlin und Erlangen Liquid-cooled rotor for gas turbines made up of disks and rings
BE530262A (en) * 1953-07-11
GB789197A (en) * 1956-01-06 1958-01-15 British Thomson Houston Co Ltd Improvements in cooling systems for high temperature turbines
DE1076446B (en) * 1957-10-25 1960-02-25 Siemens Ag Device for blade cooling in gas turbines
DE1258661C2 (en) * 1964-10-28 1973-12-13 GAS TURBINE WITH INLET SIDE SEAL OF THE COOLING AIR SUPPLY
FR1514932A (en) * 1965-06-24 1968-03-01 Snecma Counter-rotating twin rotor axial compressor
CA1209482A (en) * 1983-12-22 1986-08-12 Douglas L. Kisling Two stage rotor assembly with improved coolant flow
KR100389990B1 (en) * 1995-04-06 2003-11-17 가부시끼가이샤 히다치 세이사꾸쇼 Gas turbine
GB2307520B (en) * 1995-11-14 1999-07-07 Rolls Royce Plc A gas turbine engine
JP3758792B2 (en) * 1997-02-25 2006-03-22 三菱重工業株式会社 Gas turbine rotor platform cooling mechanism
DE19940556B4 (en) * 1999-08-26 2012-02-02 Alstom Device for cooling guide vanes or rotor blades in a gas turbine
DE19950109A1 (en) * 1999-10-18 2001-04-19 Asea Brown Boveri Rotor for a gas turbine

Citations (11)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
GB452412A (en) * 1933-11-25 1936-08-24 Michael Martinka Improvements relating to apparatus for cooling the rotors of turbines
US2141401A (en) * 1936-07-01 1938-12-27 Martinka Michael Gas turbine
US2149510A (en) * 1934-01-29 1939-03-07 Cem Comp Electro Mec Method and means for preventing deterioration of turbo-machines
US2213940A (en) * 1937-07-07 1940-09-03 Jendrassik George Rotor for gas turbines and rotary compressors
US2241782A (en) * 1937-07-07 1941-05-13 Jendrassik George Gas turbine
GB543985A (en) * 1939-09-25 1942-03-23 Sulzer Ag Improvements in or relating to rotors for turbines
GB578191A (en) * 1941-11-21 1946-06-19 Frank Bernard Halford Improvements in or relating to turbines
GB579316A (en) * 1941-05-07 1946-07-31 Hayne Constant Improvements in gas turbines, axial flow or turbine type gas compressors and the like machines
US2440069A (en) * 1944-08-26 1948-04-20 Gen Electric High-temperature elastic fluid turbine
US2452782A (en) * 1945-01-16 1948-11-02 Power Jets Res & Dev Ltd Construction of rotors for compressors and like machines
US2489683A (en) * 1943-11-19 1949-11-29 Edward A Stalker Turbine

Family Cites Families (5)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
DE708758C (en) * 1933-11-25 1941-07-28 Michael Martinka Dipl Ing Cooling device for turbine runners
CH179552A (en) * 1933-11-25 1935-09-15 Martinka Michael Device for cooling the impeller of a combustion turbine.
DE718939C (en) * 1936-08-09 1942-03-25 Rheinmetall Borsig Ag Gas or steam turbine
DE665762C (en) * 1936-09-12 1938-10-03 Rheinmetall Borsig Akt Ges Wer Device for cooling turbines, in particular gas turbines
DE733048C (en) * 1937-07-07 1943-03-18 Talalmanykifejlesztoe Es Ertek Gas turbine

Patent Citations (11)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
GB452412A (en) * 1933-11-25 1936-08-24 Michael Martinka Improvements relating to apparatus for cooling the rotors of turbines
US2149510A (en) * 1934-01-29 1939-03-07 Cem Comp Electro Mec Method and means for preventing deterioration of turbo-machines
US2141401A (en) * 1936-07-01 1938-12-27 Martinka Michael Gas turbine
US2213940A (en) * 1937-07-07 1940-09-03 Jendrassik George Rotor for gas turbines and rotary compressors
US2241782A (en) * 1937-07-07 1941-05-13 Jendrassik George Gas turbine
GB543985A (en) * 1939-09-25 1942-03-23 Sulzer Ag Improvements in or relating to rotors for turbines
GB579316A (en) * 1941-05-07 1946-07-31 Hayne Constant Improvements in gas turbines, axial flow or turbine type gas compressors and the like machines
GB578191A (en) * 1941-11-21 1946-06-19 Frank Bernard Halford Improvements in or relating to turbines
US2489683A (en) * 1943-11-19 1949-11-29 Edward A Stalker Turbine
US2440069A (en) * 1944-08-26 1948-04-20 Gen Electric High-temperature elastic fluid turbine
US2452782A (en) * 1945-01-16 1948-11-02 Power Jets Res & Dev Ltd Construction of rotors for compressors and like machines

Cited By (158)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
US2807434A (en) * 1952-04-22 1957-09-24 Gen Motors Corp Turbine rotor assembly
US2858103A (en) * 1956-03-26 1958-10-28 Westinghouse Electric Corp Gas turbine apparatus
DE1043718B (en) * 1956-07-31 1958-11-13 Maschf Augsburg Nuernberg Ag Axially loaded turbine runner with cooling by a gaseous coolant, especially for gas turbines
US2974925A (en) * 1957-02-11 1961-03-14 John C Freche External liquid-spray cooling of turbine blades
US2931624A (en) * 1957-05-08 1960-04-05 Orenda Engines Ltd Gas turbine blade
US2972470A (en) * 1958-11-03 1961-02-21 Gen Motors Corp Turbine construction
US2996280A (en) * 1959-04-07 1961-08-15 Iii John A Wilson Heat shield
US3034763A (en) * 1959-08-20 1962-05-15 United Aircraft Corp Rotor construction
US3139310A (en) * 1961-12-29 1964-06-30 Svenska Flaektfabriken Ab Arrangement in axial fans for the transport of dust commingled gases
US3356340A (en) * 1965-03-15 1967-12-05 Gen Electric Turbine rotor constructions
US3437313A (en) * 1966-05-18 1969-04-08 Bristol Siddeley Engines Ltd Gas turbine blade cooling
US3730644A (en) * 1969-06-26 1973-05-01 Rolls Royce Gas turbine engine
US3689177A (en) * 1971-04-19 1972-09-05 Gen Electric Blade constraining structure
US4035102A (en) * 1975-04-01 1977-07-12 Kraftwerk Union Aktiengesellschaft Gas turbine of disc-type construction
US4484858A (en) * 1981-12-03 1984-11-27 Hitachi, Ltd. Turbine rotor with means for preventing air leaks through outward end of spacer
US4551063A (en) * 1983-03-18 1985-11-05 Kraftwerke Union Ag Medium-pressure steam turbine
FR2557205A1 (en) * 1983-12-22 1985-06-28 United Technologies Corp ROTOR WITH DOUBLE PASS COOLING OF THE HEELS OF THE AUBES
US4582467A (en) * 1983-12-22 1986-04-15 United Technologies Corporation Two stage rotor assembly with improved coolant flow
US4536129A (en) * 1984-06-15 1985-08-20 United Technologies Corporation Turbine blade with disk rim shield
US4645424A (en) * 1984-07-23 1987-02-24 United Technologies Corporation Rotating seal for gas turbine engine
US4659289A (en) * 1984-07-23 1987-04-21 United Technologies Corporation Turbine side plate assembly
US4648793A (en) * 1985-05-31 1987-03-10 General Electric Company Turbine wheel key and keyway ventilation
US4795307A (en) * 1986-02-28 1989-01-03 Mtu Motoren- Und Turbinen-Union Munchen Gmbh Method and apparatus for optimizing the vane clearance in a multi-stage axial flow compressor of a gas turbine
US4730982A (en) * 1986-06-18 1988-03-15 Societe Nationale D'etude Et De Construction De Moteurs D'aviation "S.N.E.C.M.A." Assembly for controlling the flow of cooling air in an engine turbine
US5193982A (en) * 1991-07-17 1993-03-16 Societe Nationale D'etude Et De Construction De Moteurs D'aviation S.N.E.C.M.A. Separate inter-blade platform for a bladed rotor disk
US5271718A (en) * 1992-08-11 1993-12-21 General Electric Company Lightweight platform blade
US5507620A (en) * 1993-07-17 1996-04-16 Abb Management Ag Gas turbine with cooled rotor
US5525032A (en) * 1994-04-02 1996-06-11 Abb Management Ag Process for the operation of a fluid flow engine
US5611669A (en) * 1994-09-27 1997-03-18 Eupopean Gas Turbines Limited Turbines with platforms between stages
EP0735238A1 (en) * 1995-03-31 1996-10-02 General Electric Company Closed or open circuit cooling of turbine rotor components
US5993154A (en) * 1996-11-21 1999-11-30 Asea Brown Boveri Ag Welded rotor of a turbo-engine
EP0921273A1 (en) * 1997-06-11 1999-06-09 Mitsubishi Heavy Industries, Ltd. Rotor for gas turbines
EP0921273A4 (en) * 1997-06-11 2001-01-24 Mitsubishi Heavy Ind Ltd Rotor for gas turbines
US6457935B1 (en) * 2000-06-15 2002-10-01 Snecma Moteurs System for ventilating a pair of juxtaposed vane platforms
US6817190B2 (en) * 2002-08-20 2004-11-16 Alm Development, Inc. Blade cooling in a gas turbine engine
WO2005001260A2 (en) * 2002-08-20 2005-01-06 Alm Development, Inc. Blade cooling in a gas turbine engine
US20040035118A1 (en) * 2002-08-20 2004-02-26 Alm Development, Inc. Blade cooling in a gas turbine engine
WO2005001260A3 (en) * 2002-08-20 2005-04-21 Alm Dev Inc Blade cooling in a gas turbine engine
US20040081556A1 (en) * 2002-10-24 2004-04-29 Andre Chevrefils Blade passive cooling feature
US6832893B2 (en) * 2002-10-24 2004-12-21 Pratt & Whitney Canada Corp. Blade passive cooling feature
US20040165983A1 (en) * 2003-02-26 2004-08-26 Rolls-Royce Plc Damper seal
US7021898B2 (en) * 2003-02-26 2006-04-04 Rolls-Royce Plc Damper seal
US20040253113A1 (en) * 2003-06-16 2004-12-16 Snecma Moteurs Retention capacity of a blade having an asymmetrical hammerhead fastener, with the help of platform stiffeners
EP1489266A1 (en) * 2003-06-16 2004-12-22 Snecma Moteurs Retention of a blade with an asymmetric hammerhead foot using platform stiffeners
FR2856105A1 (en) * 2003-06-16 2004-12-17 Snecma Moteurs IMPROVING THE RETENTION CAPACITY OF A VANE WITH A DISSYMMETRIC HAMMER ATTACHMENT USING PLATFORM STRAINERS
US7080974B2 (en) 2003-06-16 2006-07-25 Snecma Moteurs Retention capacity of a blade having an asymmetrical hammerhead fastener, with the help of platform stiffeners
US20050058545A1 (en) * 2003-09-12 2005-03-17 Siemens Westinghouse Power Corporation Turbine blade platform cooling system
US6945749B2 (en) * 2003-09-12 2005-09-20 Siemens Westinghouse Power Corporation Turbine blade platform cooling system
US20050111970A1 (en) * 2003-11-26 2005-05-26 Gabriel Suciu Turbine durm rotor for a turbine engine
US7128535B2 (en) * 2003-11-26 2006-10-31 United Technologies Corporation Turbine drum rotor for a turbine engine
US20050232751A1 (en) * 2003-12-18 2005-10-20 Townes Roderick M Cooling arrangement
GB2409240B (en) * 2003-12-18 2007-04-11 Rolls Royce Plc A gas turbine rotor
US7207776B2 (en) 2003-12-18 2007-04-24 Rolls-Royce Plc Cooling arrangement
GB2409240A (en) * 2003-12-18 2005-06-22 Rolls Royce Plc Cooling arrangement
US7476078B2 (en) * 2004-02-14 2009-01-13 Alstom Technology Ltd Rotor with core surrounded by shielding rings
US20050180847A1 (en) * 2004-02-14 2005-08-18 Alstom Technology Ltd Rotor
US20060269403A9 (en) * 2004-02-14 2006-11-30 Alstom Technology Ltd Rotor
US20060024164A1 (en) * 2004-07-30 2006-02-02 Keith Sean R Method and apparatus for cooling gas turbine engine rotor blades
US7131817B2 (en) * 2004-07-30 2006-11-07 General Electric Company Method and apparatus for cooling gas turbine engine rotor blades
US20080240927A1 (en) * 2006-10-16 2008-10-02 Katharina Bergander Turbine blade for a turbine with a cooling medium passage
US8021118B2 (en) * 2006-10-16 2011-09-20 Siemens Aktiengesellschaft Turbine blade for a turbine with a cooling medium passage
RU2481481C2 (en) * 2007-07-06 2013-05-10 Снекма Air supply device for ventilation of blades of low pressure turbine of gas turbine engine; rotor of gas turbine engine, and gas turbine engine
JP4616869B2 (en) * 2007-08-24 2011-01-19 三菱重工業株式会社 gas turbine
JP2007332973A (en) * 2007-08-24 2007-12-27 Mitsubishi Heavy Ind Ltd Gas turbine
US20100290922A1 (en) * 2008-02-27 2010-11-18 Mitsubisihi Heavy Industries, Ltd Turbine disk and gas turbine
US8770919B2 (en) * 2008-02-27 2014-07-08 Mitsubishi Heavy Industries, Ltd. Turbine disk and gas turbine
US20110016884A1 (en) * 2008-03-28 2011-01-27 Mitsubishi Heavy Industries, Ltd. Cooling passage cover, manufacturing method of the cover, and gas turbine
CN101970802A (en) * 2008-03-28 2011-02-09 三菱重工业株式会社 Cover for cooling passage, method of manufacturing the cover, and gas turbine
US8387401B2 (en) * 2008-03-28 2013-03-05 Mitsubishi Heavy Industries, Ltd. Cooling passage cover, manufacturing method of the cover, and gas turbine
CN101970802B (en) * 2008-03-28 2013-11-06 三菱重工业株式会社 Cover for cooling passage, method of manufacturing the cover, and gas turbine
US20110061395A1 (en) * 2009-09-13 2011-03-17 Kendrick Donald W Method of fuel staging in combustion apparatus
US8689561B2 (en) 2009-09-13 2014-04-08 Donald W. Kendrick Vortex premixer for combustion apparatus
US8689562B2 (en) 2009-09-13 2014-04-08 Donald W. Kendrick Combustion cavity layouts for fuel staging in trapped vortex combustors
US20110061391A1 (en) * 2009-09-13 2011-03-17 Kendrick Donald W Vortex premixer for combustion apparatus
US8549862B2 (en) 2009-09-13 2013-10-08 Lean Flame, Inc. Method of fuel staging in combustion apparatus
US20110061390A1 (en) * 2009-09-13 2011-03-17 Kendrick Donald W Inlet premixer for combustion apparatus
US20110061392A1 (en) * 2009-09-13 2011-03-17 Kendrick Donald W Combustion cavity layouts for fuel staging in trapped vortex combustors
FR2954797A1 (en) * 2009-12-29 2011-07-01 Snecma Low pressure turbine rotor for two-shaft gas turbine engine of aircraft, has elastic sealing unit fixed on each of blade roots and projected with respect to rear transverse face of disk at level of joints so as to be in contact with flange
US20110164982A1 (en) * 2010-01-06 2011-07-07 General Electric Company Apparatus and method for a low distortion weld for rotors
US20120027584A1 (en) * 2010-08-02 2012-02-02 General Electric Company Turbine seal system
US8511976B2 (en) * 2010-08-02 2013-08-20 General Electric Company Turbine seal system
US20120134778A1 (en) * 2010-11-29 2012-05-31 Alexander Anatolievich Khanin Axial flow gas turbine
US8932007B2 (en) * 2010-11-29 2015-01-13 Alstom Technology Ltd. Axial flow gas turbine
JP2013545926A (en) * 2010-12-09 2013-12-26 アルストム テクノロジー リミテッド Fluid flow devices, in particular gas turbines penetrated axially by hot gas flow
RU2548226C2 (en) * 2010-12-09 2015-04-20 Альстом Текнолоджи Лтд Fluid medium flow unit, in particular, turbine with axially passing heated gas flow
WO2012076588A1 (en) * 2010-12-09 2012-06-14 Alstom Technology Ltd Fluid flow machine especially gas turbine penetrated axially by a hot gas stream
CN103249916B (en) * 2010-12-09 2016-01-20 阿尔斯通技术有限公司 By the fluid stream machine that hot air flow passes axially through, especially gas turbine
US9657641B2 (en) 2010-12-09 2017-05-23 General Electric Company Fluid flow machine especially gas turbine penetrated axially by a hot gas stream
AU2011340576B2 (en) * 2010-12-09 2015-09-24 General Electric Technology Gmbh Fluid flow machine especially gas turbine penetrated axially by a hot gas stream
CN103249916A (en) * 2010-12-09 2013-08-14 阿尔斯通技术有限公司 Fluid flow machine especially gas turbine penetrated axially by a hot gas stream
RU2578016C2 (en) * 2010-12-13 2016-03-20 Дженерал Электрик Компани Cooling circuit for rotor drum
US8662826B2 (en) * 2010-12-13 2014-03-04 General Electric Company Cooling circuit for a drum rotor
JP2012127338A (en) * 2010-12-13 2012-07-05 General Electric Co <Ge> Cooling circuit for drum rotor
US20120148405A1 (en) * 2010-12-13 2012-06-14 General Electric Company Cooling circuit for a drum rotor
US20130302143A1 (en) * 2010-12-14 2013-11-14 Rolls-Royce Deutschland Ltd & Co Kg Cooling device for a jet engine
US9657592B2 (en) * 2010-12-14 2017-05-23 Rolls-Royce Deutschland Ltd & Co Kg Cooling device for a jet engine
EP2484866A3 (en) * 2011-02-03 2017-03-15 General Electric Company Cross-over purge flow system for a turbomachine wheel member
CN102840144A (en) * 2011-06-20 2012-12-26 通用电气公司 Ventilated compressor rotor and a turbine engine having the same
US20120321441A1 (en) * 2011-06-20 2012-12-20 Kenneth Moore Ventilated compressor rotor for a turbine engine and a turbine engine incorporating same
US8721291B2 (en) * 2011-07-12 2014-05-13 Siemens Energy, Inc. Flow directing member for gas turbine engine
US20130017095A1 (en) * 2011-07-12 2013-01-17 Ching-Pang Lee Flow directing member for gas turbine engine
EP2586970A3 (en) * 2011-10-28 2017-05-24 United Technologies Corporation Spoked spacer for a gas turbine engine
EP2586968A3 (en) * 2011-10-28 2017-05-03 United Technologies Corporation Secondary flow arrangement for slotted rotor
US8944762B2 (en) * 2011-10-28 2015-02-03 United Technologies Corporation Spoked spacer for a gas turbine engine
US8961132B2 (en) * 2011-10-28 2015-02-24 United Technologies Corporation Secondary flow arrangement for slotted rotor
US9790792B2 (en) 2011-10-28 2017-10-17 United Technologies Corporation Asymmetrically slotted rotor for a gas turbine engine
US10760423B2 (en) * 2011-10-28 2020-09-01 Raytheon Technologies Corporation Spoked rotor for a gas turbine engine
US20130108413A1 (en) * 2011-10-28 2013-05-02 Gabriel L. Suciu Secondary flow arrangement for slotted rotor
US9938831B2 (en) 2011-10-28 2018-04-10 United Technologies Corporation Spoked rotor for a gas turbine engine
EP2586969A3 (en) * 2011-10-28 2017-05-03 United Technologies Corporation Spoked Rotor for a Gas Turbine Engine
EP2586971A3 (en) * 2011-10-28 2017-05-24 United Technologies Corporation A spacer, a rotor, a spool and a method of orienting a rotor stack load path
US20180223668A1 (en) * 2011-10-28 2018-08-09 United Technologies Corporation Spoked rotor for a gas turbine engine
US20130108468A1 (en) * 2011-10-28 2013-05-02 Gabriel L. Suciu Spoked spacer for a gas turbine engine
US9145772B2 (en) 2012-01-31 2015-09-29 United Technologies Corporation Compressor disk bleed air scallops
US20130236289A1 (en) * 2012-03-12 2013-09-12 General Electric Company Turbine interstage seal system
US9540940B2 (en) * 2012-03-12 2017-01-10 General Electric Company Turbine interstage seal system
US20130264779A1 (en) * 2012-04-10 2013-10-10 General Electric Company Segmented interstage seal system
US9115587B2 (en) 2012-08-22 2015-08-25 Siemens Energy, Inc. Cooling air configuration in a gas turbine engine
US9528376B2 (en) * 2012-09-13 2016-12-27 General Electric Company Compressor fairing segment
US20140069101A1 (en) * 2012-09-13 2014-03-13 General Electric Company Compressor fairing segment
US9790801B2 (en) * 2012-12-27 2017-10-17 United Technologies Corporation Gas turbine engine component having suction side cutback opening
US20140212270A1 (en) * 2012-12-27 2014-07-31 United Technologies Corporation Gas turbine engine component having suction side cutback opening
US20140301851A1 (en) * 2013-04-08 2014-10-09 Alstom Technology Ltd Rotor
US9441639B2 (en) * 2013-05-13 2016-09-13 General Electric Company Compressor rotor heat shield
US20140334929A1 (en) * 2013-05-13 2014-11-13 General Electric Company Compressor rotor heat shield
US20140363307A1 (en) * 2013-06-05 2014-12-11 Siemens Aktiengesellschaft Rotor disc with fluid removal channels to enhance life of spindle bolt
US9951621B2 (en) * 2013-06-05 2018-04-24 Siemens Aktiengesellschaft Rotor disc with fluid removal channels to enhance life of spindle bolt
US10233757B2 (en) * 2013-07-17 2019-03-19 Siemens Aktiengesellschaft Rotor for a thermal turbomachine
US10107103B2 (en) 2013-07-17 2018-10-23 Siemens Aktiengesellschaft Rotor for a thermal turbomachine
US20160230773A1 (en) * 2013-07-17 2016-08-11 Siemens Aktiengesellschaft Rotor for a thermal turbomachine
US20150114001A1 (en) * 2013-10-28 2015-04-30 General Electric Company Sealing component for reducing secondary airflow in a turbine system
US9404376B2 (en) * 2013-10-28 2016-08-02 General Electric Company Sealing component for reducing secondary airflow in a turbine system
JP2015200242A (en) * 2014-04-09 2015-11-12 株式会社東芝 Axial flow turbine
US10006364B2 (en) * 2014-08-20 2018-06-26 United Technologies Corporation Gas turbine rotors
US20160053688A1 (en) * 2014-08-20 2016-02-25 United Technologies Corporation Gas turbine rotors
US10837288B2 (en) 2014-09-17 2020-11-17 Raytheon Technologies Corporation Secondary flowpath system for a gas turbine engine
CN107002690A (en) * 2014-10-15 2017-08-01 赛峰航空器发动机 Runner assembly for the turbogenerator including self-supporting rotor case
US20170226861A1 (en) * 2014-10-15 2017-08-10 Safran Aircraft Engines Rotary assembly for a turbine engine comprising a self-supported rotor collar
US10662793B2 (en) * 2014-12-01 2020-05-26 General Electric Company Turbine wheel cover-plate mounted gas turbine interstage seal
US20160153302A1 (en) * 2014-12-01 2016-06-02 General Electric Company Turbine wheel cover-plate mounted gas turbine interstage seal
WO2016087153A1 (en) * 2014-12-04 2016-06-09 Siemens Aktiengesellschaft Rotor, axial compressor, installation method
CN107002493A (en) * 2014-12-04 2017-08-01 西门子公司 Rotor, Axial Flow Compressor, the method for installation
US10830253B2 (en) 2014-12-04 2020-11-10 Siemens Aktiengesellschaft Rotor, axial compressor, installation method
CN107002493B (en) * 2014-12-04 2019-10-08 西门子公司 Rotor, Axial Flow Compressor, the method for installation
US20160186591A1 (en) * 2014-12-31 2016-06-30 General Electric Company Flowpath boundary and rotor assemblies in gas turbines
US9664058B2 (en) * 2014-12-31 2017-05-30 General Electric Company Flowpath boundary and rotor assemblies in gas turbines
US20160186593A1 (en) * 2014-12-31 2016-06-30 General Electric Company Flowpath boundary and rotor assemblies in gas turbines
US10337345B2 (en) 2015-02-20 2019-07-02 General Electric Company Bucket mounted multi-stage turbine interstage seal and method of assembly
US10612383B2 (en) * 2016-01-27 2020-04-07 General Electric Company Compressor aft rotor rim cooling for high OPR (T3) engine
US20170211590A1 (en) * 2016-01-27 2017-07-27 General Electric Company Compressor Aft Rotor Rim Cooling for High OPR (T3) Engine
US20240084708A1 (en) * 2016-02-05 2024-03-14 Siemens Energy Global GmbH & Co. KG Rotor comprising a rotor component arranged between two rotor discs
US11041396B2 (en) * 2016-10-06 2021-06-22 Raytheon Technologies Corporation Axial-radial cooling slots on inner air seal
US11098604B2 (en) * 2016-10-06 2021-08-24 Raytheon Technologies Corporation Radial-axial cooling slots
US20180128114A1 (en) * 2016-11-10 2018-05-10 Doosan Heavy Industries & Construction Co., Ltd Structure for cooling rotor of turbomachine, rotor and turbomachine having the same
US10837290B2 (en) * 2016-11-10 2020-11-17 DOOSAN Heavy Industries Construction Co., LTD Structure for cooling rotor of turbomachine, rotor and turbomachine having the same
EP3447244A1 (en) * 2017-08-23 2019-02-27 Siemens Aktiengesellschaft Turbine rotor assembly with lap joints between the rotor discs for torque transmission
US11339662B2 (en) * 2018-08-02 2022-05-24 Siemens Energy Global GmbH & Co. KG Rotor comprising a rotor component arranged between two rotor disks
US11215056B2 (en) 2020-04-09 2022-01-04 Raytheon Technologies Corporation Thermally isolated rotor systems and methods

Also Published As

Publication number Publication date
GB612097A (en) 1948-11-08
CH278723A (en) 1951-10-31
DE971297C (en) 1959-01-08

Similar Documents

Publication Publication Date Title
US2656147A (en) Cooling of gas turbine rotors
US3670497A (en) Combustion chamber support
US7819629B2 (en) Blade for a gas turbine
US3742706A (en) Dual flow cooled turbine arrangement for gas turbine engines
US8376697B2 (en) Gas turbine sealing apparatus
US4127358A (en) Blade or vane for a gas turbine engine
US4180373A (en) Turbine blade
US7704048B2 (en) Turbine airfoil with controlled area cooling arrangement
US2268464A (en) Combustion chamber
US4292008A (en) Gas turbine cooling systems
US5288207A (en) Internally cooled turbine airfoil
US7967567B2 (en) Multi-pass cooling for turbine airfoils
KR100364183B1 (en) Gas turbine blade with a cooled platform
US2548485A (en) Combustion chamber lining
US4173120A (en) Turbine nozzle and rotor cooling systems
US4582467A (en) Two stage rotor assembly with improved coolant flow
US2858103A (en) Gas turbine apparatus
US2603453A (en) Cooling means for turbines
US3443791A (en) Turbine vane assembly
US2440069A (en) High-temperature elastic fluid turbine
US2823008A (en) Rotors for fluid flow machines such as turbines
US20130011238A1 (en) Cooled ring segment
US2405190A (en) Fluid turbine
US2945671A (en) Bladed rotor constructions for fluid machines
US2807434A (en) Turbine rotor assembly