US3866417A - Gas turbine engine augmenter liner coolant flow control system - Google Patents

Gas turbine engine augmenter liner coolant flow control system Download PDF

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Publication number
US3866417A
US3866417A US331078A US33107873A US3866417A US 3866417 A US3866417 A US 3866417A US 331078 A US331078 A US 331078A US 33107873 A US33107873 A US 33107873A US 3866417 A US3866417 A US 3866417A
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United States
Prior art keywords
liner
cooling
gas turbine
turbine engine
chambers
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Expired - Lifetime
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US331078A
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English (en)
Inventor
David A Velegol
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General Electric Co
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General Electric Co
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Filing date
Publication date
Application filed by General Electric Co filed Critical General Electric Co
Priority to US331078A priority Critical patent/US3866417A/en
Priority to CA191,342A priority patent/CA995472A/en
Priority to GB448374A priority patent/GB1458054A/en
Priority to DE19742405840 priority patent/DE2405840A1/de
Priority to IT20314/74A priority patent/IT1006315B/it
Priority to JP49015533A priority patent/JPS49112017A/ja
Priority to BE140711A priority patent/BE810794A/xx
Priority to FR7404254A priority patent/FR2217547A1/fr
Application granted granted Critical
Publication of US3866417A publication Critical patent/US3866417A/en
Anticipated expiration legal-status Critical
Expired - Lifetime legal-status Critical Current

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    • F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
    • F02—COMBUSTION ENGINES; HOT-GAS OR COMBUSTION-PRODUCT ENGINE PLANTS
    • F02K—JET-PROPULSION PLANTS
    • F02K1/00—Plants characterised by the form or arrangement of the jet pipe or nozzle; Jet pipes or nozzles peculiar thereto
    • F02K1/78—Other construction of jet pipes
    • F02K1/82—Jet pipe walls, e.g. liners

Definitions

  • the stabilizing and l Appl' 3314,78 support system includes a plurality of stabilizers adapted to mount the cooling liner to the exhaust duct [52] U.S. Cl 60/261, 60/39.32, 60/3966 casing in such a manner as to define the cooling [51 1 Int. Cl F02c 7/20, F03k 3/10 plenum therebetween.
  • the stabilizers are captured on ⁇ 58] Field of Search 60/261, 39.65, 39.66, 39.69, their outer ends by a stabilizer guide which permits 60/3972 R, 270 R, 39.32, 39.31 relative thermal expansion to occur between the cooling liner and the exhaust duct casing, and each of the [56] References Cited stabilizer guides is connected to a positioning band UNITED STATES PATENTS which is adapted to be mounted to the inside of the 2,510,645 6/1950 McMahan 60/3932 exhaust duct Casing' t Prelerred.embdimem the 2,974,486 3/1961 Edwards 60/261 x Positioning band is Pmvded "tegrally formed 3,031,844 5/1962 TOmOlOniUS 60/3931 x flange which defines a restricted ihlel 10 each of the 3,540,216 11/1970 Quillevere et a1 60/3972 R cooling chambers.
  • Gas turbine engines generally comprise a compressor for compressing air flowing through the engine, a combustion system in which high energy fuel is mixed with the compressed air and ignited to form a high energy gas stream, and a turbine which includes a rotor portion operatively connected to the compressor to drive the same.
  • Many modern-day gas turbine engines are of the turbofan type in which a second or low pressure compressor is mounted forwardly of the high pressure compressor and is driven by a second turbine mounted downstream of the first turbine.
  • the low pressure compressor or fan presents an additional stage of compression and, in addition, is normally of a larger diameter than the high pressure compressor.
  • the turbofan engine is therefore capable of flowing a much larger mass of air, thereby greatly increasing the thrust output of the engine.
  • An additional known method of increasing the thrust output of the engine is to provide the engine with an augmenter.
  • additional fuel is injected into an exhaust duct formed downstream of the second turbine and is ignited to provide an additional high energy gas stream, which is ejected through an exhaust nozzle to provide high energy thrust output from the engine.
  • turbofan engines it is also known to mix the fan airflow with core engine airflow prior to supplying the mixed flow with additional fuel by means of the augmenter.
  • the augmenter system is normally located within the exhaust duct of the engine and, in most cases, some means must be provided for protecting the exhaust duct from the extremely high temperatures associated with the gas flow within the augmenter.
  • One common means for providing this protection is to position a cooling liner within the exhaust duct and to pass cooling air between the liner and the exhaust duct. Openings or slots are positioned within the cooling liner such that the cooling air may flow through these openings and form a film of coolant on the inside of the cooling liner.
  • a basic problem confronting the designer of such an augmenter liner consists of the regulation of the coolant flow and the minimization of liner pressure loading.
  • a simple and reliable means is needed to regulate the coolant flow through the openings or slots in the cooling liner.
  • the problem is complicated by the fact that the pressure in the plenum which surrounds the cooling liner is essentially constant along the entire axial length of the liner while the static pressure of the combustion gas inside the liner decreases axially due to the acceleration of the gas as its temperature increases due to operation of the augmenter. This condition not only results in significant pressure differentials across the liner but also in a varying pressure differential from the upstream to the downstream end of such liner.
  • a cooling liner stabilization and mounting system which includes means which divide the coolant plenum surrounding the liner into a number of individual chambers.
  • the flow into each of these chambers is con trolled by means of a flange associated with the mount ing system which acts as a restriction in the flow path and also divides the plenum into the previouslymentioned individual chambers.
  • FIG. 1 is a schematic, axial, cross-sectional view of a gas turbine engine incorporting the present inventive regulating means
  • FIG. 2 is an enlarged, partial view of the inventive cooling liner stabilization and pressure regulation means of FIG. 1;
  • FIG. 3 is a graphical plot showing the pressure differential along the axial length of a cooling liner incorporating the present inventive means.
  • FIG. 1 a gas turbine engine 10 of the mixed flow turbofan type is shown to include a core engine 12 which includes a fan turbine 14 for driving a plurality of fan blades 15 mounted on a shaft 16.
  • the fan blades are located within an inlet 17 formed by an outer or fan casing 18 which surrounds the entire gas turbine engine 10.
  • the fan casing 18 operates with a core engine casing 20 to define parallel flow paths 22 and 23.
  • Air entering the flow path 23 is compressed by means of a compressor 24 and is mixed with fuel in combustor 26.
  • Fuel is delivered to the combustor 26 by means of a plurality of fuel injection points 27 from fuel tubes 28 which extend through the flow path 22.
  • the resultant high energy gas stream exits the combustor 26 and drives a turbine 30 which, in turn, drives the compressor 24 by means of a shaft 31.
  • air flowing through the outer or fan flow path 22 and air exiting the core engine 12 flow through a mixer 32, which operates to mix the two separate flow paths.
  • the mixed flow path is then acted upon by an augmenter 34, which consists of a plurality of fuel injectors 36.
  • the resultant fuel/air mixture in the augmenter 34 is ignited by means of a suitable igniter (not shown), flows through an exhaust duct 40, and thereafter provides an additional propulsive force by exiting through an exhaust nozzle 42.
  • the exhaust duct is located at the downstream end of the fan casing 18 and is shown in FIG. 1 to include an exhaust duct casing 44 and a cooling air liner which is generally designated by the numeral 46.
  • the cooling liner 46 is spaced radially inwardly from the exhaust duct casing 44 and defines an annular coolant flow path 48 having an inlet 50 formed by a forward lip 52 at the upstream end of the cooling liner 46.
  • the cooling liner 46 includes a plurality of openings or slots 54 adapted to deliver cooling air from the passageway 48 to the inside of the liner 46.
  • the coolant flowing through the open ings 54 provides a film of cool air on the inside of the liner 46 thereby protecting both the liner 46 and the surrounding exhaust duct casing member 44 from the high temperatures associated with the operation of the augmenter 34.
  • a high energy gas stream is generated by the combustor 26 and drives the high pressure turbine 30 and low pressure turbine 14, which, in turn, drive the core engine compressor 24 and the fan 15.
  • Air exiting the low pressure turbine 14 and air flowing through the fan flow path 22 are mixed within the mixer 32 and the mixed flow is delivered to the region of the augmenter 34.
  • a resultant fuel/air mixture generated by the augmenter 34 is ignited to provide an additional propulsive force by exiting through the exhaust nozzle 42.
  • This cooling air thereafter flows through the openings 54 and forms a film on the inside of the cooling liner 46 thereby protecting the liner 46 and the surrounding exhaust duct casing 44 from the high gas temperatures associated with operation of the augmenter 34.
  • gas turbine engine 10 described above is typical of many present-day augmented turbofan engines and has been described solely to place the present invention in proper perspective. As will become clear to those skilled in the art, the present invention will be applicable to other types of gas turbine engines and, therefore, the engine 10 is merely meant to be illustrative.
  • the positioning band 62 includes a plurality of holes 64 which are aligned with threaded openings associated with capture nuts 68 mounted on each of the stabilizer guides 60.
  • a plurality of bolts 70 fit through openings 72 formed within the exhaust duct casing 44 and are threadably received within the nuts 68, thereby firmly attaching the positioning band 62 and thus the cooling liner 46 to the exhaust duct casing 44.
  • the stabilizers 58 are adapted to define radial height 71 of the annular coolant passageway 48.
  • the pressure of the coolant within the passageway 48 is relatively constant along the entire axial length of the liner 46.
  • the static pressure of the combustion gas within the exhaust duct i.e. inside the liner 46, decreases axially due to the acceleration of the gas as its temperature increases due to operation of the augmenter.
  • the pressure differential across the liner 46 increases significantly near the aft end of the liner 46. This condition is shown graphically as a dotted line curve 73 in FIG. 5 wherein pressure differential across the liner is plotted as a function of axial length of the liner.
  • the flange 76 is sized so as to provide a gap 84 between an inner edge 86 of the flange 76 and the outer wall of the cooling liner 46.
  • the gap 84 acts as a restriction in the flow path for the cooling air and acts to control the pressure of the coolant air within the chambers 78, 80 and 82.
  • the gap 84 may be made of various sizes depending upon the desired pressure within the chambers 78, 80 and 82. In certain applications, the gap 84 may be the same for each of the stabilizer assemblies 56, while in other applications the gap size may vary from mounting assembly to mounting assembly. In either case, the gap 84 can be utilized to provide a pressure within the chambers 78, 80 and 82 such that the average pressure differential across the liner is a minimum in each chamber.
  • the flange 76 could be positioned at either end of the positioning band 62.
  • the flange 76 could be formed integrally with, or connected to, the stabilizing guides 60 instead of the positioning band 62.
  • the flange 76 could be connected directly to either the exhaust duct casing 44 or the cooling air liner 46.
  • the stabilizing assemblies 56 are capable of use in combination with more standard mounting techniques such as spacers 86 shown in FIG. 3 near the upstream end of the liner 46. The appended claims are intended to cover these and similar modifications in Applicants broader inventive concept.
  • a gas turbine engine of the type including a compressor, a turbine, a combustion system, an augmenter, an exhaust duct surrounding said augmenter, and a cooling liner positioned within said duct so as to form a cooling plenum therebetween, at least a portion of which liner extends downstream from said augmenter and is adapted to protect said exhaust duct from the high temperature gas generated by said augmenter, the
  • flange means situated within said cooling plenum and dividing said cooling plenum into at least two individual chambers, said flange means occupying a predetermined radial space in said cooling plenum and providing a restricted inlet between said chambers such that the average pressure differentials across the liner in each of said chambers are substantially equal.
  • each of said stabilizer assemblies includes a plurality of stabilizers, a like number of stabilizer guides, one of which partially surrounds each of said stabilizers, and a positioning band surrounding and mounted to each of said stabilizer guides and to said exhaust duct, and said flange means comprises a flange formed integrally with said positioning band and extending toward said cooling liner.
  • cooling liner includes at least two of said stabilizer assemblies such that said plenum is divided into at least three of said chambers.
  • cooling liner includes a plurality of relatively equally sized coolant holes spaced along at least a portion of the axial length thereof.

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  • Engineering & Computer Science (AREA)
  • Chemical & Material Sciences (AREA)
  • Combustion & Propulsion (AREA)
  • Mechanical Engineering (AREA)
  • General Engineering & Computer Science (AREA)
  • Turbine Rotor Nozzle Sealing (AREA)
  • Exhaust Silencers (AREA)
US331078A 1973-02-09 1973-02-09 Gas turbine engine augmenter liner coolant flow control system Expired - Lifetime US3866417A (en)

Priority Applications (8)

Application Number Priority Date Filing Date Title
US331078A US3866417A (en) 1973-02-09 1973-02-09 Gas turbine engine augmenter liner coolant flow control system
CA191,342A CA995472A (en) 1973-02-09 1974-01-30 Gas turbine engine augmenter liner coolant flow control system
GB448374A GB1458054A (en) 1973-02-09 1974-01-31 Gas turbine engines including exhaust reheat combustion equipment
DE19742405840 DE2405840A1 (de) 1973-02-09 1974-02-07 Steuersystem fuer den kuehlstrom fuer die verkleidung des nachbrenners eines gasturbinen-triebwerkes
IT20314/74A IT1006315B (it) 1973-02-09 1974-02-08 Sistema di controllo di portata refrigerante per camicia di po tenziatore di turbomotore a gas
JP49015533A JPS49112017A (enrdf_load_stackoverflow) 1973-02-09 1974-02-08
BE140711A BE810794A (fr) 1973-02-09 1974-02-08 Moteur a turbine a gaz a dispositif de post-combustion
FR7404254A FR2217547A1 (enrdf_load_stackoverflow) 1973-02-09 1974-02-08

Applications Claiming Priority (1)

Application Number Priority Date Filing Date Title
US331078A US3866417A (en) 1973-02-09 1973-02-09 Gas turbine engine augmenter liner coolant flow control system

Publications (1)

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US3866417A true US3866417A (en) 1975-02-18

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US331078A Expired - Lifetime US3866417A (en) 1973-02-09 1973-02-09 Gas turbine engine augmenter liner coolant flow control system

Country Status (8)

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US (1) US3866417A (enrdf_load_stackoverflow)
JP (1) JPS49112017A (enrdf_load_stackoverflow)
BE (1) BE810794A (enrdf_load_stackoverflow)
CA (1) CA995472A (enrdf_load_stackoverflow)
DE (1) DE2405840A1 (enrdf_load_stackoverflow)
FR (1) FR2217547A1 (enrdf_load_stackoverflow)
GB (1) GB1458054A (enrdf_load_stackoverflow)
IT (1) IT1006315B (enrdf_load_stackoverflow)

Cited By (25)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
US4072008A (en) * 1976-05-04 1978-02-07 General Electric Company Variable area bypass injector system
US4461146A (en) * 1982-10-22 1984-07-24 The United States Of America As Represented By The Secretary Of The Navy Mixed flow swirl augmentor for turbofan engine
DE3606286A1 (de) * 1985-03-04 1986-09-04 General Electric Co., Schenectady, N.Y. Verfahren und einrichtung zum steuern des kuehlmittelstroemungsflusses in einer nachbrennerauskleidung
US4706453A (en) * 1986-11-12 1987-11-17 General Motors Corporation Support and seal assembly
US4718230A (en) * 1986-11-10 1988-01-12 United Technologies Corporation Augmentor liner construction
US4773593A (en) * 1987-05-04 1988-09-27 United Technologies Corporation Coolable thin metal sheet
US4813229A (en) * 1985-03-04 1989-03-21 General Electric Company Method for controlling augmentor liner coolant flow pressure in a mixed flow, variable cycle gas
US4833881A (en) * 1984-12-17 1989-05-30 General Electric Company Gas turbine engine augmentor
US4958489A (en) * 1985-03-04 1990-09-25 General Electric Company Means for controlling augmentor liner coolant flow pressure in a mixed flow, variable cycle gas turbine engine
US4961312A (en) * 1985-03-04 1990-10-09 General Electric Company Method for controlling augmentor liner coolant flow pressure in a mixed flow, variable cycle gas turbine engine
US5157917A (en) * 1991-05-20 1992-10-27 United Technologies Corporation Gas turbine engine cooling air flow
US5201887A (en) * 1991-11-26 1993-04-13 United Technologies Corporation Damper for augmentor liners
DE4338745A1 (de) * 1993-11-12 1995-05-18 Abb Management Ag Vorrichtung zur Wärmeabschirmung des Rotors in Gasturbinen
US5465572A (en) * 1991-03-11 1995-11-14 General Electric Company Multi-hole film cooled afterburner cumbustor liner
US5697213A (en) * 1995-12-05 1997-12-16 Brewer; Keith S. Serviceable liner for gas turbine engine
US5720434A (en) * 1991-11-05 1998-02-24 General Electric Company Cooling apparatus for aircraft gas turbine engine exhaust nozzles
US20070157621A1 (en) * 2006-01-06 2007-07-12 General Electric Company Exhaust dust flow splitter system
US20080110176A1 (en) * 2006-04-28 2008-05-15 Snecma Turbojet engine comprising an afterburner duct cooled by a variable-throughput ventilation stream
US20090136342A1 (en) * 2007-05-24 2009-05-28 Rolls-Royce Plc Duct installation
US20100242494A1 (en) * 2009-03-24 2010-09-30 Rolls-Royce Plc Casing arrangement
US20130336759A1 (en) * 2012-06-14 2013-12-19 Joseph T. Christians Turbomachine flow control device
US20140109593A1 (en) * 2012-10-22 2014-04-24 United Technologies Corporation Coil spring hanger for exhaust duct liner
US20160032863A1 (en) * 2013-04-15 2016-02-04 Aircelle Nozzle for an aircraft turboprop engine with an unducted fan
US10385868B2 (en) * 2016-07-05 2019-08-20 General Electric Company Strut assembly for an aircraft engine
US20210310377A1 (en) * 2020-04-01 2021-10-07 General Electric Company Liner Support System

Families Citing this family (2)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
US4920742A (en) * 1988-05-31 1990-05-01 General Electric Company Heat shield for gas turbine engine frame
FR3100284B1 (fr) * 2019-08-30 2021-12-03 Safran Aircraft Engines Couple volet convergent-volet divergent pour tuyère de turboréacteur à géométrie variable dont les volets comprennent chacun un conduit de circulation d’air de refroidissement

Citations (6)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
US2510645A (en) * 1946-10-26 1950-06-06 Gen Electric Air nozzle and porting for combustion chamber liners
US2974486A (en) * 1958-03-27 1961-03-14 United Aircraft Corp Afterburner mixture and flame control baffle
US3031844A (en) * 1960-08-12 1962-05-01 William A Tomolonius Split combustion liner
US3540216A (en) * 1967-01-23 1970-11-17 Snecma Two-flow gas turbine jet engine
US3570241A (en) * 1968-08-02 1971-03-16 Rolls Royce Flame tube for combustion chamber of a gas turbine engine
US3712062A (en) * 1968-04-17 1973-01-23 Gen Electric Cooled augmentor liner

Patent Citations (6)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
US2510645A (en) * 1946-10-26 1950-06-06 Gen Electric Air nozzle and porting for combustion chamber liners
US2974486A (en) * 1958-03-27 1961-03-14 United Aircraft Corp Afterburner mixture and flame control baffle
US3031844A (en) * 1960-08-12 1962-05-01 William A Tomolonius Split combustion liner
US3540216A (en) * 1967-01-23 1970-11-17 Snecma Two-flow gas turbine jet engine
US3712062A (en) * 1968-04-17 1973-01-23 Gen Electric Cooled augmentor liner
US3570241A (en) * 1968-08-02 1971-03-16 Rolls Royce Flame tube for combustion chamber of a gas turbine engine

Cited By (38)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
US4072008A (en) * 1976-05-04 1978-02-07 General Electric Company Variable area bypass injector system
US4461146A (en) * 1982-10-22 1984-07-24 The United States Of America As Represented By The Secretary Of The Navy Mixed flow swirl augmentor for turbofan engine
US4833881A (en) * 1984-12-17 1989-05-30 General Electric Company Gas turbine engine augmentor
US4961312A (en) * 1985-03-04 1990-10-09 General Electric Company Method for controlling augmentor liner coolant flow pressure in a mixed flow, variable cycle gas turbine engine
US4813229A (en) * 1985-03-04 1989-03-21 General Electric Company Method for controlling augmentor liner coolant flow pressure in a mixed flow, variable cycle gas
US4958489A (en) * 1985-03-04 1990-09-25 General Electric Company Means for controlling augmentor liner coolant flow pressure in a mixed flow, variable cycle gas turbine engine
DE3606286A1 (de) * 1985-03-04 1986-09-04 General Electric Co., Schenectady, N.Y. Verfahren und einrichtung zum steuern des kuehlmittelstroemungsflusses in einer nachbrennerauskleidung
US4718230A (en) * 1986-11-10 1988-01-12 United Technologies Corporation Augmentor liner construction
US4706453A (en) * 1986-11-12 1987-11-17 General Motors Corporation Support and seal assembly
US4773593A (en) * 1987-05-04 1988-09-27 United Technologies Corporation Coolable thin metal sheet
US5465572A (en) * 1991-03-11 1995-11-14 General Electric Company Multi-hole film cooled afterburner cumbustor liner
US5483794A (en) * 1991-03-11 1996-01-16 General Electric Company Multi-hole film cooled afterburner combustor liner
US5157917A (en) * 1991-05-20 1992-10-27 United Technologies Corporation Gas turbine engine cooling air flow
US5775589A (en) * 1991-11-05 1998-07-07 General Electric Company Cooling apparatus for aircraft gas turbine engine exhaust nozzles
US5720434A (en) * 1991-11-05 1998-02-24 General Electric Company Cooling apparatus for aircraft gas turbine engine exhaust nozzles
US5201887A (en) * 1991-11-26 1993-04-13 United Technologies Corporation Damper for augmentor liners
DE4338745B4 (de) * 1993-11-12 2005-05-19 Alstom Vorrichtung zur Wärmeabschirmung des Rotors in Gasturbinen
DE4338745A1 (de) * 1993-11-12 1995-05-18 Abb Management Ag Vorrichtung zur Wärmeabschirmung des Rotors in Gasturbinen
US5697213A (en) * 1995-12-05 1997-12-16 Brewer; Keith S. Serviceable liner for gas turbine engine
US5704208A (en) * 1995-12-05 1998-01-06 Brewer; Keith S. Serviceable liner for gas turbine engine
US7966823B2 (en) 2006-01-06 2011-06-28 General Electric Company Exhaust dust flow splitter system
US20070157621A1 (en) * 2006-01-06 2007-07-12 General Electric Company Exhaust dust flow splitter system
RU2433291C2 (ru) * 2006-04-28 2011-11-10 Снекма Задняя часть турбореактивного двигателя самолета, турбореактивный двигатель и самолет
US7870740B2 (en) * 2006-04-28 2011-01-18 Snecma Turbojet engine comprising an afterburner duct cooled by a variable-throughput ventilation stream
US20080110176A1 (en) * 2006-04-28 2008-05-15 Snecma Turbojet engine comprising an afterburner duct cooled by a variable-throughput ventilation stream
US20090136342A1 (en) * 2007-05-24 2009-05-28 Rolls-Royce Plc Duct installation
US20100242494A1 (en) * 2009-03-24 2010-09-30 Rolls-Royce Plc Casing arrangement
EP2861851A4 (en) * 2012-06-14 2015-08-05 United Technologies Corp RIVER CONTROL DEVICE FOR A TURBOMA MACHINE
US20130336759A1 (en) * 2012-06-14 2013-12-19 Joseph T. Christians Turbomachine flow control device
US10253651B2 (en) * 2012-06-14 2019-04-09 United Technologies Corporation Turbomachine flow control device
US20140109593A1 (en) * 2012-10-22 2014-04-24 United Technologies Corporation Coil spring hanger for exhaust duct liner
US10054080B2 (en) * 2012-10-22 2018-08-21 United Technologies Corporation Coil spring hanger for exhaust duct liner
US10125723B1 (en) 2012-10-22 2018-11-13 United Technologies Corporation Coil spring hanger for exhaust duct liner
US20160032863A1 (en) * 2013-04-15 2016-02-04 Aircelle Nozzle for an aircraft turboprop engine with an unducted fan
US10113506B2 (en) * 2013-04-15 2018-10-30 Aircelle Nozzle for an aircraft turboprop engine with an unducted fan
US10385868B2 (en) * 2016-07-05 2019-08-20 General Electric Company Strut assembly for an aircraft engine
US20210310377A1 (en) * 2020-04-01 2021-10-07 General Electric Company Liner Support System
US11905843B2 (en) * 2020-04-01 2024-02-20 General Electric Company Liner support system

Also Published As

Publication number Publication date
DE2405840A1 (de) 1974-08-15
CA995472A (en) 1976-08-24
BE810794A (fr) 1974-05-29
FR2217547A1 (enrdf_load_stackoverflow) 1974-09-06
JPS49112017A (enrdf_load_stackoverflow) 1974-10-25
IT1006315B (it) 1976-09-30
GB1458054A (en) 1976-12-08

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