US3043561A - Turbine rotor ventilation system - Google Patents

Turbine rotor ventilation system Download PDF

Info

Publication number
US3043561A
US3043561A US783376A US78337658A US3043561A US 3043561 A US3043561 A US 3043561A US 783376 A US783376 A US 783376A US 78337658 A US78337658 A US 78337658A US 3043561 A US3043561 A US 3043561A
Authority
US
United States
Prior art keywords
rotor
cooling
wheel
bucket
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
US783376A
Other languages
English (en)
Inventor
Jr George W Scheper
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.)
General Electric Co
Original Assignee
General Electric Co
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 General Electric Co filed Critical General Electric Co
Priority to US783376A priority Critical patent/US3043561A/en
Priority to CH8225859A priority patent/CH378100A/de
Priority to GB44099/59A priority patent/GB926160A/en
Application granted granted Critical
Publication of US3043561A publication Critical patent/US3043561A/en
Anticipated expiration legal-status Critical
Expired - Lifetime legal-status Critical Current

Links

Images

Classifications

    • F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
    • F02—COMBUSTION ENGINES; HOT-GAS OR COMBUSTION-PRODUCT ENGINE PLANTS
    • F02C—GAS-TURBINE PLANTS; AIR INTAKES FOR JET-PROPULSION PLANTS; CONTROLLING FUEL SUPPLY IN AIR-BREATHING JET-PROPULSION PLANTS
    • F02C7/00—Features, components parts, details or accessories, not provided for in, or of interest apart form groups F02C1/00 - F02C6/00; Air intakes for jet-propulsion plants
    • F02C7/12—Cooling of plants
    • 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
    • F01D5/082—Cooling fluid being directed on the side of the rotor disc or at the roots of the blades on the side of the rotor disc
    • 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
    • F01D5/085—Heating, heat-insulating or cooling means cooling fluid circulating inside the rotor
    • F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
    • F02—COMBUSTION ENGINES; HOT-GAS OR COMBUSTION-PRODUCT ENGINE PLANTS
    • F02C—GAS-TURBINE PLANTS; AIR INTAKES FOR JET-PROPULSION PLANTS; CONTROLLING FUEL SUPPLY IN AIR-BREATHING JET-PROPULSION PLANTS
    • F02C7/00—Features, components parts, details or accessories, not provided for in, or of interest apart form groups F02C1/00 - F02C6/00; Air intakes for jet-propulsion plants
    • F02C7/12—Cooling of plants
    • F02C7/16—Cooling of plants characterised by cooling medium
    • F02C7/18—Cooling of plants characterised by cooling medium the medium being gaseous, e.g. air

Definitions

  • the rotor of a gas turbine powerplant having an output on the order of 13,000 H.P. may Afor instance be about 510 inches in diameter, designed to have a normal rated speed of 4,800 r.p.m.
  • the tlrst stage buckets of such a rotor may operate at an average temperature on the order of 1,200 F., and the rim of the first stage wheel may attain temperatures on the order off 1,000 F. Because of the enormous centrifugal forces generated at these speeds, complicated by the thermal stresses added by the high temperatures and the substantial temperature gradients set up by the cooling systems employed to protect the turbine lfrom excessive temperatures, the rotor must be fabricated of carefully chosen high temperature alloy materials.
  • an object of the present invention is to provide an improved system for preheating the high temperature alloy rotor of a gas turbine power-plant during the starting cycle so that the alloy material is quickly brought above the transition temperature, so that its full normal srength will be lavailable before the rotor is brought up to rated speed.
  • Another object is to provide an improved cooling system for the high temperature alloy rotor ofA a gas turbine power-plant, of great simplicity from the standpoint of changes in the turbine structure required, while entailing a minimum expenditure of pressure energy of the cooling lluid by optimum utilization of a minimum quantity of coolant flow.
  • FIG. 1 is a partial longitudinal sectional view of a gas turbine Powerplant having Va rotor preheating and cooling system incorporating the invention
  • FIG. 2 is an enlarged detail view of a portion of FIG, 1;
  • FIG. 3 is a detail sectional view taken at the plane 3 3 in FIG. 2;
  • FIG. 4 is Aa detail view taken at the plane 4-4 in FIG. 2;
  • FIG. 5 is a velocity ldiagram for the invention.
  • the invention is practiced by taking a small quantity of preheating and' cooling air from the discharge passage of the compressor of the gas turbine powerplant, admitting it to a central chamber in the high temperature rotor by way of a circumferential row of nozzles disposed in the rotor at an angle so as to extract thermal energy from the cooling air and at the same time impart rotational energy to the rotor, conducting the coolant Huid through cooling passages in the high temperature rotor, and discharging it by way of a plurality of nozzles which are also directed rearwardly so as to extract additional thermal energy from the coolant and impart additional rotational energy to the rotor.
  • FIG. 1 illustrates the invention as applied to a gas turbine powerplant comprisingra multi-stage axial ilowcom pressor 1, a two-stage axial ow turbine rotor 2, and a combustion system represented by the single combustion chamber or combustor 3.
  • the axial Ilow compressor 1 may be of any suitable construction, and may for instance have on the order of 16 stages, and a discharge pressure in the neighborhood of pounds per square inch, absolute.
  • the annular discharge passage 1a delivers high pressure air past a circumferential row of supporting struts 1b into an annular air supply passage 3a dened lbetween the inner liner 3b ⁇ and the outer cylindrical housing 3c of the combustor 3, yas indicated by the flow arrows in FIG. l.
  • the combustor 3 is a cylindrical or cantype combustor, for instance of the type described in Patent 2,601,000, issued in the name of A. I.
  • the casing of the turbine includes an outer casing mem-ber 4 provided with a suitable air or water cooling system (not shown) and supporting a segmental shroud ring 4a 4which surrounds with a small clearance the circumferential row of buckets 5 on the rst stage turbine wheel 2a.
  • Casing 4 also supports an intermediate row of stationary nozzle 'blades 6 which discharge-the motive fluid to the second stage buckets 7.
  • Spent motive' fluid is discharged through an annular expanding dilfusing passage 8.
  • the downstream end of casing 4 supports a segmental shroud ring 4b dening appropriate close clearances with the tips of the buckets 7.
  • the outer wall structure of the discharge casing 8 is illustrated -at 8a as having an upstream end ange 8b appropriately secured to the downstream flange 4c of the casing 4.
  • the rotor structure to which the present invention particularly relates, includes a cylindrical member 9, having at one end a flange 9a sceured to the abuttingend disk 1d of the axial ilow compressor 1.
  • the opposite end of the connecting cylinder 9 has a ange 9b secured to an abutting ange 2c of the first stage bucket-wheel 2a.
  • the periphery of the flange 9b of the cylindrical member 9 and the abutting liange 2c of the bucket wheel 2a defines a plurality of circumferential grooves, the annular lands between which form a labyrinth seal with a segmental sealing member aoaaeei supported ina circumferential groove formed ⁇ by a ring member 11 connected at one end to a flange y12a of a cylindrical member 12 supported from the inner ends of the struts 1b.
  • the other end portion of the seal support member 11 supports the inner periphery of the high temperature nozzle ring lassembly 3g ⁇ (FIG. 1).
  • the outer periphery of the high temperature nozzle ring is connected 14e extending across the exhaust gas discharge .passage 8, with their outer ends supported in the outer casing memberSav.
  • the bucket-wheels 2a, 2lb are fabricated separately and bolted together v by means of abutting circumferential flange portions 2e, 2f.
  • the periphery of these anges deiine annular labyrinth sealing teeth cooperating with a segmental sealing ring member 15 carried in a casing 15a supported from the inner ends of the intermediate nozzle blades 6.
  • the special arrangement for circulating preheating and cooling air through the high temperature turbine rotor includes the following.
  • v The cylindrical linner wall y12b of the compressor discharge passage 1a defines an annular clearance space at 12e through which high pressure air is admitted to the 'annular chamber 9c dened between the walls 12, 12b and the connecting cylinder member 9. Adjacent its righthand end, the cylinder 9 defines a circumferential row of v nozzles 9d. As illustrated more particularly in FIG. 3,
  • these nozzles are holes drilled and reamed to have a slight inward taper so as to define a contracting' nozzle directing a jet inwardly and with a substantial tangential absolute velocity component.
  • the Vaxis of the nozzle defines approximately a 250 angle with a radial line.y
  • a compressor last stage rotor discharge pressure of approximately 90 p.'s.i.a. maintained in the chamber 9c, there will be a pressure drop across the nozzles 9d with the downstream pressure inside cylinder 9 maintained at approximately 45 p.s.i.
  • the inwardly directed jets impart rotational energy to the cylinder 9.
  • the Work energy thus extracted ⁇ from the kcooling air reduces its temperature Y sub stanti ally.
  • nozzles 9d there are three of the nozzles 9d, ⁇ each of a minimum diameter of .80 inch, and the cylinder 9 has an approximate diameter of 26 inches and is 2 inches thick, in the powenplant shown in the drawings.
  • cooling air flows to the right through a central axial passage 2g in the hub 2n of the irst stage bucket-wheel 2a.
  • the rst and second stage bucket-wheels 2a, 2b have hub portions 2n, 2p and web portions 2q, 2r spaced axially Vto deiine a radially extending annular cooling air passage 2h.
  • the abutting annular hub flange portions 2e, 2f have an intertting rabbet portion at 2j for maintaining accurate concentricity.
  • This abutting ange portion is provided with a plurality of radially extending slots 2k which admit theycooling air flow to axial holes 2m spaced around the flange 2e.
  • the cooling air is discharged through a circumferential row of nozzles 17b ldetined in special members 17 the shape of vwhich is shown more particularly in FIG. 4.
  • the nozzle member indicated generally as 17 in the drawings, comprises a trapezoidal shaped block, the beveled end portion of which substantially abuts a mating beveled end of the next adjacent nozzle block, as shown at 17d in FIG. 4.
  • the right-hand half of the :block 17 defines a ⁇ circular recess 17a communicating with the air inlet port 2m, and discharges through a nozzle 5171;, the latter preferably formed as a simple drilled and reamed passage because of the ditliculty of machining a contracting tapered nozzle in this location.
  • the axis of the nozzle 17h is at an angle of approximately 30 to a tangent to the periphery of the flange 2e.
  • FIG. 5 A typical vector diagram for the nozzle of FIG. 4 may be seen in FIG. 5, it being understood that the same type of analysis applies to the nozzle of FIG. 3.
  • Vector OW represents the absolute velocity of the nozzles in a tangential'direction as determined by their distance from the rotor axis ⁇ and the rotor rpm.
  • Vector WR is the exit velocity of the gas relative to the nozzle.
  • the resultant OR is the absolute velocity of the gas as it leaves the nozzle. The change between the tangential component of absolute gas velocity OR and the tangential absolute velocity component of the gas entering the nozzle will impar-t rotational energy to the rotor. rl'he energy thus lost by the gas serves to reduce its temperature.
  • a second function of the segmental blocks 17 is to serve as the head of a bolt member, shown in dotted lines at 17C in FIG. 4, which bolt passes through the flanges 2e, 2f to hold the bucket-wheels together.
  • the coolingV air in the annular space 15b escapes by two paths.
  • the first is radially outward along the web ⁇ 2q of the Wheel 2a and past annular sealing rings identified 15C, 15d in FIG.V1.
  • the second path of escape for the cooling air is lthrough the segmental packing 15, the rate of such flow being determined by the packing clearance.
  • This coolant passes radially outward along the upstream face 2t of the second stage wheel 2b, past annular sealing ring 15e.
  • the downstream surface of second stage wheel 2b is cooled by a portion of the outer casing cooling air which enters chamber 14a through hollow struts 14C.
  • the high pressure of the cooling air flow is utilized to impart rotational energy to the rotor, and this work energy extracted from the cooling ilow reduces its tempera-ture substantially so as to improve its cooling capacity.
  • a rate of cooling air flow on the order of 2 pounds per second, a total temperature drop on the order of 90 7F. is obtained from' the two sets of nozzles, and at-the same time approximately 60 H.P. of mechanical energy is imparted to the rotor.
  • the cooling function may be performed with a minimum quantity of coolant extracted from the compressor discharge ilow, and with minimum structural complications.
  • a second important advantage of the invention lies in the fact that the cooling air flow described above may also be employed to preheat the rotor during the starting cycle. As described above, it is important to quickly bring the high temperature alloy rotor up past its transition temperature, above which it develops its optimum strength characteristics.' With a ferritic high temperature alloy such as, for instance, one composed of 12% chronium, 1% molybdenum, 1% tungsten, 0.25% vanadium, and the balance iron, this transition temperature may be on the order of 130 F.
  • the preheating arrangement serves to avoid problems which would otherwise arise from stressing the wheel highly while it is still below its transition temperature, with further benefits derived from the standpoint of the transient thermal stresses created in the respective hub and rim portions.
  • the invention provides means for performing effectively the turbine rotor cooling function in normal operation, as well as providing simple means for preheating the rotor to attain optimum strength qualities in the high temperature alloy during the starting cycle and minimize the thermal stresses created.
  • a source of elastic liluid under pressure having an initial absolute tangential velocity component
  • a turbine rotor having at least one turbine bucket-wheel lfor converting iuid pressure energy torotational energy
  • said turbine rotor Ihaving a cylindrical rotor member dening a first central chamber disposed to supply cooling fluid to the bucket-wheel, a wall member spaced radiallly from said cylindrical rotor member to define a second coolant fluid supply chamber, passage means admitting uid under pressure from said source to the second coolant fluid supply chamber, rst nozzle means disposed in said cylindrical rotor member and directed backwardly with respect to the direction of rotation of the rotor and so constructed and arranged as to have a first tangential absolute velocity component at turbine rated speed and to receive iiuid from the second supply chamber and to discharge jets of Ifluid having a second tangential velocity component relative to said trst nozzle means into said
  • a turbine rotor cooling arrangment for va gas turbine powerplant a compressor supplying air under pressure, a turbine rotor having at least one turbine bucket-wheel ⁇ for converting iluid pressure energy to rotational energy, said turbine rotor including a cylindrical mem-ber 4defining a iirst central chamber disposed to supply cooling iluid to the bucket-wheel, a Wall member spaced radially from said cylindrical rotor member to dene a second coolant fluid supply chamber, passage means admitting air under pressure from the discharge passage of the compressor to said second chamber with an initial tangential absolute velocity component, rst nozzle means disposed in said cylindrical rotor member and directed backwardly with respect to the direction of rotation of the rotor ⁇ and so constructed and arranged as to have a first tangential absolute Velocity component at turbine rated speed and to receive fluid from said second chamber and arranged to ndischarge jets of coolant having a second tangential velocity component relative to said rst nozzle means into
  • a bucket-wheel having hub and web portions and a circumferential row of blade members, the hub portion defining a central axial passage, a cylindrical rotor member projecting axially from the upstream side of said hub portion ⁇ and defining a central chamber communicating with the axial passage in the bucket wheel hub, a source of elastic iluid under pressure having an initial tangential absolute velocity component, said cylindrical rotor member defining a first plurality of circumferential- 1y spaced nozzle means directed backwardly with respect to the direction of rotation of the rotor and so constructed and arranged as to have a first tangential yabsolute velocity component at turbine rated speed and to direct jets of cooling ⁇ fluid having a second tangential velocity corn- 7 5 ponent relative to said first no'zzle means into said central chamber with a third resultant'tangential absolute veloc ity component, the downstream side of the bucket-wheel hub portion including arcircumferentia-l portion supporting
  • SIA two-stage high tempera-ture turbine rotor comprising first and second Vstage bucket-wheels fabricated separately and havng abutting circumferential hu-b flange portions, means securing said hub flange portions together, the first stage bucket-wheel having a central axial passage for cooling fluid, a cylindrical rotor member disposed at the upstream side of the first stage wheel and extending axially to define a central coolant supply chamber, said cylindrical rotor member having a plurality of circumferentially disposed first nozzle means having a first tangential absolute velocity component at turbine rated speed for directing jets of cooling fluid having a second tangential velocity component relative to said firs-t nozzle means inwardly into said central chamber with a resultant third tangential absolute velocity component backwardly relative to the direction of rotation, said hub flange securing means comprising a plurality of fastener members disposed circumferentially around said abutting hub flange portions of the bucket wheels, each of said fasten

Landscapes

  • Engineering & Computer Science (AREA)
  • Mechanical Engineering (AREA)
  • General Engineering & Computer Science (AREA)
  • Chemical & Material Sciences (AREA)
  • Combustion & Propulsion (AREA)
  • Turbine Rotor Nozzle Sealing (AREA)
US783376A 1958-12-29 1958-12-29 Turbine rotor ventilation system Expired - Lifetime US3043561A (en)

Priority Applications (3)

Application Number Priority Date Filing Date Title
US783376A US3043561A (en) 1958-12-29 1958-12-29 Turbine rotor ventilation system
CH8225859A CH378100A (de) 1958-12-29 1959-12-22 Turbinenrotorkühlung bei einem Gasturbinentriebwerk
GB44099/59A GB926160A (en) 1958-12-29 1959-12-29 A gas turbine with preheating and cooling of the turbine rotor

Applications Claiming Priority (1)

Application Number Priority Date Filing Date Title
US783376A US3043561A (en) 1958-12-29 1958-12-29 Turbine rotor ventilation system

Publications (1)

Publication Number Publication Date
US3043561A true US3043561A (en) 1962-07-10

Family

ID=25129059

Family Applications (1)

Application Number Title Priority Date Filing Date
US783376A Expired - Lifetime US3043561A (en) 1958-12-29 1958-12-29 Turbine rotor ventilation system

Country Status (3)

Country Link
US (1) US3043561A (de)
CH (1) CH378100A (de)
GB (1) GB926160A (de)

Cited By (28)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
US3230710A (en) * 1962-12-24 1966-01-25 Garrett Corp Gas turbine
US3437313A (en) * 1966-05-18 1969-04-08 Bristol Siddeley Engines Ltd Gas turbine blade cooling
US3631672A (en) * 1969-08-04 1972-01-04 Gen Electric Eductor cooled gas turbine casing
US3632221A (en) * 1970-08-03 1972-01-04 Gen Electric Gas turbine engine cooling system incorporating a vortex shaft valve
US3748056A (en) * 1971-02-09 1973-07-24 Nissan Motor Turbine blade cooling
US3904307A (en) * 1974-04-10 1975-09-09 United Technologies Corp Gas generator turbine cooling scheme
US4008977A (en) * 1975-09-19 1977-02-22 United Technologies Corporation Compressor bleed system
FR2449789A1 (fr) * 1979-02-26 1980-09-19 Gen Electric Turbomachine a structure de refroidissement de joint d'etancheite perfectionnee
FR2469555A1 (fr) * 1979-11-14 1981-05-22 United Technologies Corp Systeme de purge de compresseur pour le refroidissement des elements d'une section de turbine et le reglage de l'espace libre du joint lenticulaire
US4453888A (en) * 1981-04-01 1984-06-12 United Technologies Corporation Nozzle for a coolable rotor blade
US4543038A (en) * 1982-03-08 1985-09-24 The Garrett Corporation Sealing apparatus and method and machinery utilizing same
EP0188910A1 (de) * 1984-12-21 1986-07-30 AlliedSignal Inc. Turbinenschaufelkühlung
US4759688A (en) * 1986-12-16 1988-07-26 Allied-Signal Inc. Cooling flow side entry for cooled turbine blading
FR2614654A1 (fr) * 1987-04-29 1988-11-04 Snecma Disque de compresseur axial de turbomachine a prelevement d'air centripete
EP0916808A3 (de) * 1997-11-05 2000-01-12 Rolls-Royce Plc Turbine
US20030101730A1 (en) * 2001-12-05 2003-06-05 Stefan Hein Vortex reducer in the high-pressure compressor of a gas turbine
EP1445421A1 (de) * 2003-02-06 2004-08-11 Snecma Moteurs Vorrichtung für die Belüftung eines Rotor einer Hochdruckturbine
RU2261350C2 (ru) * 2003-08-26 2005-09-27 Открытое акционерное общество "Авиадвигатель" Турбина газотурбинного двигателя
FR2881794A1 (fr) * 2005-02-09 2006-08-11 Snecma Moteurs Sa Turbomachine pourvue d'un moyen de reduction du bruit
US20070189890A1 (en) * 2006-02-15 2007-08-16 Snowsill Guy D Gas turbine engine rotor ventilation arrangement
US20110250057A1 (en) * 2010-04-12 2011-10-13 Laurello Vincent P Radial pre-swirl assembly and cooling fluid metering structure for a gas turbine engine
RU2443882C1 (ru) * 2010-08-23 2012-02-27 Открытое акционерное общество "Авиадвигатель" Газотурбинный двигатель
US20120227414A1 (en) * 2011-03-08 2012-09-13 Rolls-Royce Plc Gas turbine engine swirled cooling air
US20130051974A1 (en) * 2011-08-25 2013-02-28 Honeywell International Inc. Gas turbine engines and methods for cooling components thereof with mid-impeller bleed cooling air
CN103046964A (zh) * 2012-06-27 2013-04-17 北京航空航天大学 一种基于主动温度梯度控制应力的航空发动机涡轮盘
CN106640212A (zh) * 2016-11-04 2017-05-10 北京航空航天大学 一种航空燃气涡轮发动机高压涡轮盘腔冷却空气斜向预旋进气喷嘴
US20170234135A1 (en) * 2014-08-29 2017-08-17 Mitsubishi Hitachi Power Systems, Ltd. Gas turbine
CN108884714A (zh) * 2016-03-16 2018-11-23 赛峰飞机发动机公司 包括通风间隔件的涡轮转子

Families Citing this family (3)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
DE2633291C3 (de) * 1976-07-23 1981-05-14 Kraftwerk Union AG, 4330 Mülheim Gasturbinenanlage mit Kühlung durch zwei unabhängige Kühlluftströme
DE3014279A1 (de) * 1980-04-15 1981-10-22 M.A.N. Maschinenfabrik Augsburg-Nürnberg AG, 4200 Oberhausen Einrichtung zur kuehlung des inneren einer gasturbine
DE19733148C1 (de) * 1997-07-31 1998-11-12 Siemens Ag Kühlluftverteilung in einer Turbinenstufe einer Gasturbine

Citations (5)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
US2369795A (en) * 1941-11-17 1945-02-20 Andre P E Planiol Gaseous fluid turbine or the like
US2632626A (en) * 1947-02-12 1953-03-24 United Aircraft Corp Dirt trap for turbine cooling air
US2639579A (en) * 1949-06-21 1953-05-26 Hartford Nat Bank & Trust Co Turbojet engine having tail pipe ejector to induce flow of cooling air
US2680001A (en) * 1950-11-13 1954-06-01 United Aircraft Corp Arrangement for cooling turbine bearings
US2858101A (en) * 1954-01-28 1958-10-28 Gen Electric Cooling of turbine wheels

Patent Citations (5)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
US2369795A (en) * 1941-11-17 1945-02-20 Andre P E Planiol Gaseous fluid turbine or the like
US2632626A (en) * 1947-02-12 1953-03-24 United Aircraft Corp Dirt trap for turbine cooling air
US2639579A (en) * 1949-06-21 1953-05-26 Hartford Nat Bank & Trust Co Turbojet engine having tail pipe ejector to induce flow of cooling air
US2680001A (en) * 1950-11-13 1954-06-01 United Aircraft Corp Arrangement for cooling turbine bearings
US2858101A (en) * 1954-01-28 1958-10-28 Gen Electric Cooling of turbine wheels

Cited By (40)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
US3230710A (en) * 1962-12-24 1966-01-25 Garrett Corp Gas turbine
US3437313A (en) * 1966-05-18 1969-04-08 Bristol Siddeley Engines Ltd Gas turbine blade cooling
US3631672A (en) * 1969-08-04 1972-01-04 Gen Electric Eductor cooled gas turbine casing
US3632221A (en) * 1970-08-03 1972-01-04 Gen Electric Gas turbine engine cooling system incorporating a vortex shaft valve
US3748056A (en) * 1971-02-09 1973-07-24 Nissan Motor Turbine blade cooling
US3904307A (en) * 1974-04-10 1975-09-09 United Technologies Corp Gas generator turbine cooling scheme
US4008977A (en) * 1975-09-19 1977-02-22 United Technologies Corporation Compressor bleed system
FR2324874A1 (fr) * 1975-09-19 1977-04-15 United Technologies Corp Dispositif pour soutirer de l'air d'un compresseur
FR2449789A1 (fr) * 1979-02-26 1980-09-19 Gen Electric Turbomachine a structure de refroidissement de joint d'etancheite perfectionnee
FR2469555A1 (fr) * 1979-11-14 1981-05-22 United Technologies Corp Systeme de purge de compresseur pour le refroidissement des elements d'une section de turbine et le reglage de l'espace libre du joint lenticulaire
US4453888A (en) * 1981-04-01 1984-06-12 United Technologies Corporation Nozzle for a coolable rotor blade
US4543038A (en) * 1982-03-08 1985-09-24 The Garrett Corporation Sealing apparatus and method and machinery utilizing same
EP0188910A1 (de) * 1984-12-21 1986-07-30 AlliedSignal Inc. Turbinenschaufelkühlung
US4759688A (en) * 1986-12-16 1988-07-26 Allied-Signal Inc. Cooling flow side entry for cooled turbine blading
FR2614654A1 (fr) * 1987-04-29 1988-11-04 Snecma Disque de compresseur axial de turbomachine a prelevement d'air centripete
EP0916808A3 (de) * 1997-11-05 2000-01-12 Rolls-Royce Plc Turbine
US20030101730A1 (en) * 2001-12-05 2003-06-05 Stefan Hein Vortex reducer in the high-pressure compressor of a gas turbine
US7159402B2 (en) * 2001-12-05 2007-01-09 Rolls-Royce Deutschland Ltd & Co Kg Vortex reducer in the high-pressure compressor of a gas turbine
US6916151B2 (en) 2003-02-06 2005-07-12 Snecma Moteurs Ventilation device for a high pressure turbine rotor of a turbomachine
FR2851010A1 (fr) * 2003-02-06 2004-08-13 Snecma Moteurs Dispositif de ventilation d'un rotor de turbine a haute pression d'une turbomachine
EP1445421A1 (de) * 2003-02-06 2004-08-11 Snecma Moteurs Vorrichtung für die Belüftung eines Rotor einer Hochdruckturbine
RU2330976C2 (ru) * 2003-02-06 2008-08-10 Снекма Мотёр Устройство вентиляции ротора турбины высокого давления турбомашины
US20040219008A1 (en) * 2003-02-06 2004-11-04 Snecma Moteurs Ventilation device for a high pressure turbine rotor of a turbomachine
RU2261350C2 (ru) * 2003-08-26 2005-09-27 Открытое акционерное общество "Авиадвигатель" Турбина газотурбинного двигателя
FR2881794A1 (fr) * 2005-02-09 2006-08-11 Snecma Moteurs Sa Turbomachine pourvue d'un moyen de reduction du bruit
US20070189890A1 (en) * 2006-02-15 2007-08-16 Snowsill Guy D Gas turbine engine rotor ventilation arrangement
US7775764B2 (en) * 2006-02-15 2010-08-17 Rolls-Royce Plc Gas turbine engine rotor ventilation arrangement
US8677766B2 (en) * 2010-04-12 2014-03-25 Siemens Energy, Inc. Radial pre-swirl assembly and cooling fluid metering structure for a gas turbine engine
US20110250057A1 (en) * 2010-04-12 2011-10-13 Laurello Vincent P Radial pre-swirl assembly and cooling fluid metering structure for a gas turbine engine
RU2443882C1 (ru) * 2010-08-23 2012-02-27 Открытое акционерное общество "Авиадвигатель" Газотурбинный двигатель
US20120227414A1 (en) * 2011-03-08 2012-09-13 Rolls-Royce Plc Gas turbine engine swirled cooling air
US8555654B2 (en) * 2011-03-08 2013-10-15 Rolls-Royce, Plc Gas turbine engine swirled cooling air
US20130051974A1 (en) * 2011-08-25 2013-02-28 Honeywell International Inc. Gas turbine engines and methods for cooling components thereof with mid-impeller bleed cooling air
CN103046964A (zh) * 2012-06-27 2013-04-17 北京航空航天大学 一种基于主动温度梯度控制应力的航空发动机涡轮盘
US20170234135A1 (en) * 2014-08-29 2017-08-17 Mitsubishi Hitachi Power Systems, Ltd. Gas turbine
US11248466B2 (en) * 2014-08-29 2022-02-15 Mitsubishi Power, Ltd. Gas turbine
CN108884714A (zh) * 2016-03-16 2018-11-23 赛峰飞机发动机公司 包括通风间隔件的涡轮转子
CN108884714B (zh) * 2016-03-16 2021-08-31 赛峰飞机发动机公司 包括通风间隔件的涡轮转子
CN106640212A (zh) * 2016-11-04 2017-05-10 北京航空航天大学 一种航空燃气涡轮发动机高压涡轮盘腔冷却空气斜向预旋进气喷嘴
CN106640212B (zh) * 2016-11-04 2019-07-02 北京航空航天大学 一种航空燃气涡轮发动机高压涡轮盘腔冷却空气斜向预旋进气喷嘴

Also Published As

Publication number Publication date
GB926160A (en) 1963-05-15
CH378100A (de) 1964-05-31

Similar Documents

Publication Publication Date Title
US3043561A (en) Turbine rotor ventilation system
US2080425A (en) Turbine
US2591399A (en) Power plant frame structure having air-cooling means for turbine rotors and exhaust frame struts
US3647313A (en) Gas turbine engines with compressor rotor cooling
US2584899A (en) Construction of stator elements of turbines, compressors, or like machines
US3011762A (en) Turbines and in particular gas turbines
US2479573A (en) Gas turbine power plant
CA2254885C (en) Turbomachine rotor cooling
US6382903B1 (en) Rotor bore and turbine rotor wheel/spacer heat exchange flow circuit
JP3939762B2 (ja) タービン機械
JP3219930B2 (ja) ガスターボグループ
US3703808A (en) Turbine blade tip cooling air expander
US3453825A (en) Gas turbine engine having turbine discs with reduced temperature differential
US2603453A (en) Cooling means for turbines
JPS6143201A (ja) ガスタービンエンジンの圧縮機用の羽根及びパツキングの隙間最適化装置
EP0909878B1 (de) Gasturbine
GB712051A (en) Improvements in or relating to axial-flow fluid machines
US2552239A (en) Turbine rotor cooling arrangement
US3756019A (en) Gas turbine blade arrangement
US2788951A (en) Cooling of turbine rotors
US2467818A (en) High-temperature turbine casing arrangement
US2411124A (en) Internal-combustion turbine plant
US2888240A (en) Fluid cooled barrel cylinder for turbines
US2382564A (en) Turbine system
US6435812B1 (en) Bore tube assembly for steam cooling a turbine rotor