US2594808A - Means for supporting the nozzles of the combustion chambers of internal-combustion turbines - Google Patents
Means for supporting the nozzles of the combustion chambers of internal-combustion turbines Download PDFInfo
- Publication number
- US2594808A US2594808A US13689A US1368948A US2594808A US 2594808 A US2594808 A US 2594808A US 13689 A US13689 A US 13689A US 1368948 A US1368948 A US 1368948A US 2594808 A US2594808 A US 2594808A
- Authority
- US
- United States
- Prior art keywords
- ring
- combustion
- turbine
- apertures
- combustion chambers
- 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
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
- F01D9/00—Stators
- F01D9/02—Nozzles; Nozzle boxes; Stator blades; Guide conduits, e.g. individual nozzles
- F01D9/023—Transition ducts between combustor cans and first stage of the turbine in gas-turbine engines; their cooling or sealings
-
- F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
- F23—COMBUSTION APPARATUS; COMBUSTION PROCESSES
- F23R—GENERATING COMBUSTION PRODUCTS OF HIGH PRESSURE OR HIGH VELOCITY, e.g. GAS-TURBINE COMBUSTION CHAMBERS
- F23R3/00—Continuous combustion chambers using liquid or gaseous fuel
- F23R3/42—Continuous combustion chambers using liquid or gaseous fuel characterised by the arrangement or form of the flame tubes or combustion chambers
- F23R3/60—Support structures; Attaching or mounting means
-
- Y—GENERAL TAGGING OF NEW TECHNOLOGICAL DEVELOPMENTS; GENERAL TAGGING OF CROSS-SECTIONAL TECHNOLOGIES SPANNING OVER SEVERAL SECTIONS OF THE IPC; TECHNICAL SUBJECTS COVERED BY FORMER USPC CROSS-REFERENCE ART COLLECTIONS [XRACs] AND DIGESTS
- Y02—TECHNOLOGIES OR APPLICATIONS FOR MITIGATION OR ADAPTATION AGAINST CLIMATE CHANGE
- Y02T—CLIMATE CHANGE MITIGATION TECHNOLOGIES RELATED TO TRANSPORTATION
- Y02T50/00—Aeronautics or air transport
- Y02T50/60—Efficient propulsion technologies, e.g. for aircraft
Definitions
- cams invention relates to lgas-iturbine-engines kin n which .acompressor delivers comand "the-productsof Tcombustion flow from the combustion chambers through a .ring of nozzle ;guide”vanes or turbine .stator-blading to one or more turbine rotors which drive the compressor :and which the :mainstructure :for supporting jthe turbine rotor assembly comprises the eompressor "casing-and an iin'termediate casing lying within the ring of combustion ohanibersand enclosing the drive shaft between the turbine compressor rotor, and also-in which the turbine stator assembly, including shroud ring or ring ,”stator blading or guide vanes is'suppcrted from theintermedi
- HithertoTit has been the vpractice tolprovide nozzle-boxes .betweenthe combustion chambers and the z-inlet to the turbine, the nozzle-boxes being'circular at their forward or .inlet endswhere they receive the combustion products from he' combustion equipment-land.substantiallyrec- V itangulariat their rearor outlet ends where they direct the ⁇ combustion, products on to the nozzle guide van assembly; the outlet-ends-of the nozz'les were iarranged side -by-side to form a complete annular ring corresponding to theannulus' :of'the nozzle guide vane assembly.
- This construction comprises a ring having a: series of apertures each arranged to receive and support the outlet end oi a combustion chamber, the apertured ring being secured I casing and to the outer ring of the nozzle-guide-vane assembly so that strucsolely through the ringQi
- the ring conv an apertured ring for receiving and supp tural loads are tra-nsmi ed from one tothe provides a wall of ahQusing 'whichisuppor .Tth'e nozzle-boxes.
- This apertured ring ' has b'eeni'firoduced bya casting operation.
- This invention seeks to provide an improved construction of such an aperturedring which is relatively simple to manufacture and which, while retaining the desirable features of a'castccnstruction and a predetermined over-answer the outlet ends of the combustionchambefsiof a gas-turbine engine of the hind desbrib' f the form of a casting Which'h'as beeiij achine'd substantially over its wholesurface scia'sta have maximum thickness between the apertures and to taper in thicknessito'vvards its edges.
- one surface of the ring is machined to the form ofa part of a sphere having its centre on the axis of the ring 'rein'ote from the plane thereof and the othersurface ofj'th'e ring is machined by a turning operation so that the ring has a substantially triangular crosssection between the apertures.
- Reces'se's in be formed in the ring between the apertures so that the ring has a web-like formation between the apertures.
- the weight of the ring may be kept to a minimum-having regard to aprodetermined overall strength thereon n
- socket members are secured to the ring engage in the apertures to guide and support the outlet ends of the combustion chambers,- and, where the ring is formed with a part-spherical surface,- the socket members are preferably each formed with a flange having a spherical surface to co-operate with thespherical surfaceof the ring. 7
- the invention is particularly applicable to the type of gas-'turbine-engine in which: the com-lous tion chambers are disposed in a ring around the engine with their axes lyingin the surface or a cone the apex of which is'beyond the outlet ends of the combustion chambers, and in thiscas the "rearwardly by the the direction of flight
- the com-lous tion chambers are disposed in a ring around the engine with their axes lyingin the surface or a cone the apex of which is'beyond the outlet ends of the combustion chambers, and in thiscas the "rearwardly by the the direction of flight
- centre of the spherical surface of the ring will be the apex of the cone.
- a plurality of combustion chambers l2 receive air from the compressor H) by ducts Na and liquid fuel is burnt in these chambers.
- the products of combustion are delivered to the turbine H and then pass to atmosphere by way of an exhaust assembly l3.
- the exhaust gases are directed exhaust assembly relative to of the aeroplane for propulsion purposes.
- theturbine shaft Hi supports the turbine disc it? which carries blades IS.
- the shaft M is supported by a bearing H carried by a structure 13 extending from an intermediate casing 29 which is bolted to the compressor casing Hi.
- the turbine rotor assembly is supported from the compressor casing It, the intermediate cas ing [9 and the structure 13 constituting a backbone structure for the engine.
- the exhaust assembly I3 is bolted at 2! to a turbine shroud ring 22 which in turn is bolted at 23 to an outer ring 26 between which and an inner ring 20 extend the fixed guide vanes 25 of the turbine.
- a ring 26 forming one part of a labyrinth seal. which is itself bolted at 21 to a tubular member 28.
- a nozzle-box assembly is located between the ou tlet ends of the combustion chambers i2 and the guide vanes -25.
- the nozzle-box assembly comprises a ring-like casting 29 having an apertured forwardly-facing wall 2% by which the outlet ends of the combustion chambers l2 are sup ported and a series of nozzle-boxes 3b which are supported at their inlet ends by the casting 29 to receive the combustion gases from the combustion chambers and which are supported at their outlet ends in between a flanged ring SI and the tubular member 28.
- the nozzle -boxes are circular at their inlet ends and are substantially rectangular at their outlet ends "to abut'laterally to form a substantially anguide vanes 25.
- the casting 29' is formed with an outer periphral wall 2% which extends rearwardly and um wardly to be bolted at 32 to the flange 32a on.
- blades 25 are arranged to be located circumfer- 24 but to have radial freedom with respect to the ring so that any loads on the rings 29, 2d are not imparted to the blades.
- the nozzle-boxes 30 are provided at their inlet and outlet ends with springy flanges 35 by which they engage with the casting 29 and the parts 3
- the combustion chambers l2 which are disposed in a ring around the shaft I4, have their inlet ends at a greater radius than their outlet ends and the axes of the combustion chambers lie in the surfaceof a cone having its'apex 36 on the shaft axis and the apertured wall 29a of the casting 25 is arranged to be substantially normal to the axes of the combustion'chambers.
- the casting 29 is formed with a series of circular apertures 31 which are machined to have a cylindrical periphery and therefore to' lie'substantially in a plane which is normal to-the axis of the combustion chamber associated with it.
- the wall 29a is machined on its inner surface 38 to be part of a sphere having its centre at the apex 3 6 of the cone in the surface of which the axes lie.
- the outer surface 3'3 of the wall 29a isf also machined by a simple turning operation about the axis oi the ring 29 so that the wall 29a thereof has .a thickness predetermined in' relation to the part-spherical surface 38. It is arranged that the minimum thickness is obtained in the areas lying beyond the apertures 31, that is in the peripheral areas oft-he wall 2% and that the thickness is increased in the region between the apertures in order to increase the strength of the ring where it is weakened by the presence of the apertures $1.
- the wall 25a has its maximum thickness between the apertures 3i and tapers smoothly towards its peripheral edges so as to have a substantially triangular radial cross-section between the apertures.
- Each socket-member M is formed with a peripheral flange 44 having one surface of spherical form to-Jco-operate with the surface 38' of the wall 29c; and the socket-members All are-secured in position by studs .5 extending through the wall' 29a into the flanges M.
- Flats 86 are machined on the front surface 39 of the wall to provide a seating for-the heads of the studs 45.
- the outer peripheral Wall 2% and the inner peripheral flange 290 may also be machined to a desired thickness by a simple turning operation about the ring axis.
- An axial-flow gas-turbine comprising a turbine rotor; a stator casing enclosing the rotor; a
- socket members are formed with a flange having a surface machined to be part of a sphere to co-operate with the spherical surface of the ring.
- An axial-flow gas-turbine comprising a turbine rotor; a stator casing enclosing the rotor; a driving shaft on said turbine rotor; a plurality of combustion chambers disposed in a ring around the said shaft; an intermediate casing surrounding the shaft within the ring of combustion chambers; a ring-like casing member connected to the intermediate casing and to the stator casing to transmit structural loads therebetween, said ring-like casing member having one machined surface formed as a portion of a sphere whose centre lies on the axis of the turbine, and having formed therein a plurality of circumferentially spaced apertures, one aperture 'for each combustion chamber, said ring-like casing member face opposite to said first machined surface, said second surface being of substantially non-uniform convex contour and cooperating with said first machined surface to define areas of greater thickness of the ring-like casing member between said apertures and of relatively less thickness in the peripheral areas of the ring-like casing memher
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)
- Control Of Turbines (AREA)
Applications Claiming Priority (1)
Application Number | Priority Date | Filing Date | Title |
---|---|---|---|
GB268285X | 1947-03-14 |
Publications (1)
Publication Number | Publication Date |
---|---|
US2594808A true US2594808A (en) | 1952-04-29 |
Family
ID=10249898
Family Applications (1)
Application Number | Title | Priority Date | Filing Date |
---|---|---|---|
US13689A Expired - Lifetime US2594808A (en) | 1947-03-14 | 1948-03-08 | Means for supporting the nozzles of the combustion chambers of internal-combustion turbines |
Country Status (5)
Country | Link |
---|---|
US (1) | US2594808A (ko) |
BE (1) | BE487768A (ko) |
CH (1) | CH268285A (ko) |
FR (1) | FR963256A (ko) |
GB (1) | GB622384A (ko) |
Cited By (15)
Publication number | Priority date | Publication date | Assignee | Title |
---|---|---|---|---|
US2702454A (en) * | 1951-06-07 | 1955-02-22 | United Aircraft Corp | Transition piece providing a connection between the combustion chambers and the turbine nozzle in gas turbine power plants |
US2836959A (en) * | 1950-01-11 | 1958-06-03 | Gen Motors Corp | Gas turbine engine supporting frame |
US3086363A (en) * | 1960-07-22 | 1963-04-23 | United Aircraft Corp | Annular transition duct |
US3116604A (en) * | 1962-05-17 | 1964-01-07 | Gen Electric | Engine exhaust system |
US3750398A (en) * | 1971-05-17 | 1973-08-07 | Westinghouse Electric Corp | Static seal structure |
US4573315A (en) * | 1984-05-15 | 1986-03-04 | A/S Kongsberg Vapenfabrikk | Low pressure loss, convectively gas-cooled inlet manifold for high temperature radial turbine |
US5414999A (en) * | 1993-11-05 | 1995-05-16 | General Electric Company | Integral aft frame mount for a gas turbine combustor transition piece |
US5761898A (en) * | 1994-12-20 | 1998-06-09 | General Electric Co. | Transition piece external frame support |
US6116013A (en) * | 1998-01-02 | 2000-09-12 | Siemens Westinghouse Power Corporation | Bolted gas turbine combustor transition coupling |
US20040134199A1 (en) * | 2003-01-15 | 2004-07-15 | Manteiga John A | Methods and apparatus for controlling engine clearance closures |
WO2009103636A1 (de) * | 2008-02-20 | 2009-08-27 | Alstom Technology Ltd. | Thermische maschine |
US20100180605A1 (en) * | 2009-01-22 | 2010-07-22 | Siemens Energy, Inc. | Structural Attachment System for Transition Duct Outlet |
US20120159954A1 (en) * | 2010-12-21 | 2012-06-28 | Shoko Ito | Transition piece and gas turbine |
US20120291451A1 (en) * | 2011-05-20 | 2012-11-22 | Frank Moehrle | Structural frame for gas turbine combustion cap assembly |
US9404421B2 (en) | 2014-01-23 | 2016-08-02 | Siemens Energy, Inc. | Structural support bracket for gas flow path |
Families Citing this family (1)
Publication number | Priority date | Publication date | Assignee | Title |
---|---|---|---|---|
WO1998016764A1 (en) * | 1996-10-16 | 1998-04-23 | Siemens Westinghouse Power Corporation | Brush seal for gas turbine combustor-transition interface |
Citations (5)
Publication number | Priority date | Publication date | Assignee | Title |
---|---|---|---|---|
US1794454A (en) * | 1928-09-08 | 1931-03-03 | Huron Ind Inc | Sealing ring for rotating machines or cylinders |
US1852279A (en) * | 1929-05-16 | 1932-04-05 | Superheater Co Ltd | Means for securing tubes to headers |
US2445114A (en) * | 1948-07-13 | Arrangement of jet propulsion | ||
US2479573A (en) * | 1943-10-20 | 1949-08-23 | Gen Electric | Gas turbine power plant |
US2494821A (en) * | 1946-03-25 | 1950-01-17 | Rolls Royce | Means for supporting the nozzles of the combustion chambers of internal-combustion turbines |
-
0
- BE BE487768D patent/BE487768A/xx unknown
- FR FR963256D patent/FR963256A/fr not_active Expired
-
1947
- 1947-03-14 GB GB7219/47A patent/GB622384A/en not_active Expired
-
1948
- 1948-03-08 US US13689A patent/US2594808A/en not_active Expired - Lifetime
- 1948-03-12 CH CH268285D patent/CH268285A/de unknown
Patent Citations (5)
Publication number | Priority date | Publication date | Assignee | Title |
---|---|---|---|---|
US2445114A (en) * | 1948-07-13 | Arrangement of jet propulsion | ||
US1794454A (en) * | 1928-09-08 | 1931-03-03 | Huron Ind Inc | Sealing ring for rotating machines or cylinders |
US1852279A (en) * | 1929-05-16 | 1932-04-05 | Superheater Co Ltd | Means for securing tubes to headers |
US2479573A (en) * | 1943-10-20 | 1949-08-23 | Gen Electric | Gas turbine power plant |
US2494821A (en) * | 1946-03-25 | 1950-01-17 | Rolls Royce | Means for supporting the nozzles of the combustion chambers of internal-combustion turbines |
Cited By (21)
Publication number | Priority date | Publication date | Assignee | Title |
---|---|---|---|---|
US2836959A (en) * | 1950-01-11 | 1958-06-03 | Gen Motors Corp | Gas turbine engine supporting frame |
US2702454A (en) * | 1951-06-07 | 1955-02-22 | United Aircraft Corp | Transition piece providing a connection between the combustion chambers and the turbine nozzle in gas turbine power plants |
US3086363A (en) * | 1960-07-22 | 1963-04-23 | United Aircraft Corp | Annular transition duct |
US3116604A (en) * | 1962-05-17 | 1964-01-07 | Gen Electric | Engine exhaust system |
US3750398A (en) * | 1971-05-17 | 1973-08-07 | Westinghouse Electric Corp | Static seal structure |
US4573315A (en) * | 1984-05-15 | 1986-03-04 | A/S Kongsberg Vapenfabrikk | Low pressure loss, convectively gas-cooled inlet manifold for high temperature radial turbine |
US5414999A (en) * | 1993-11-05 | 1995-05-16 | General Electric Company | Integral aft frame mount for a gas turbine combustor transition piece |
US5761898A (en) * | 1994-12-20 | 1998-06-09 | General Electric Co. | Transition piece external frame support |
US6116013A (en) * | 1998-01-02 | 2000-09-12 | Siemens Westinghouse Power Corporation | Bolted gas turbine combustor transition coupling |
US6886343B2 (en) * | 2003-01-15 | 2005-05-03 | General Electric Company | Methods and apparatus for controlling engine clearance closures |
US20040134199A1 (en) * | 2003-01-15 | 2004-07-15 | Manteiga John A | Methods and apparatus for controlling engine clearance closures |
WO2009103636A1 (de) * | 2008-02-20 | 2009-08-27 | Alstom Technology Ltd. | Thermische maschine |
US20110113790A1 (en) * | 2008-02-20 | 2011-05-19 | Alstom Technology Ltd | Thermal machine |
US8272220B2 (en) | 2008-02-20 | 2012-09-25 | Alstom Technology Ltd | Impingement cooling plate for a hot gas duct of a thermal machine |
US20100180605A1 (en) * | 2009-01-22 | 2010-07-22 | Siemens Energy, Inc. | Structural Attachment System for Transition Duct Outlet |
US8616007B2 (en) * | 2009-01-22 | 2013-12-31 | Siemens Energy, Inc. | Structural attachment system for transition duct outlet |
US20120159954A1 (en) * | 2010-12-21 | 2012-06-28 | Shoko Ito | Transition piece and gas turbine |
US9200526B2 (en) * | 2010-12-21 | 2015-12-01 | Kabushiki Kaisha Toshiba | Transition piece between combustor liner and gas turbine |
US20120291451A1 (en) * | 2011-05-20 | 2012-11-22 | Frank Moehrle | Structural frame for gas turbine combustion cap assembly |
US8938976B2 (en) * | 2011-05-20 | 2015-01-27 | Siemens Energy, Inc. | Structural frame for gas turbine combustion cap assembly |
US9404421B2 (en) | 2014-01-23 | 2016-08-02 | Siemens Energy, Inc. | Structural support bracket for gas flow path |
Also Published As
Publication number | Publication date |
---|---|
BE487768A (ko) | |
FR963256A (ko) | 1950-07-05 |
GB622384A (en) | 1949-05-02 |
CH268285A (de) | 1950-05-15 |
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