US2594808A

US2594808A - Means for supporting the nozzles of the combustion chambers of internal-combustion turbines

Info

Publication number: US2594808A
Application number: US13689A
Authority: US
Inventors: Rubbra Arthur Alexander
Original assignee: Rolls Royce PLC
Current assignee: Rolls Royce PLC
Priority date: 1947-03-14
Filing date: 1948-03-08
Publication date: 1952-04-29
Anticipated expiration: 1969-04-29
Also published as: BE487768A; CH268285A; FR963256A; GB622384A

Description

April 29, 1952 RUBBRA 2,594,808

. MEANS FOR SUPPORTING THE NOZZLES v0F THE COMBUSTION CHAMBERS 0F INTERNAL-COMBUSTION TURBINES Filed March 8, 1948 5 Sheets-Sheet l April 29, 1952 A. A. RUBBRA 2,594,808 MEANS FOR SUPPORTING THE NOZZLES OF THE COMBUSTION INTERNAL-COMBUSTION TURBINES Filed March 8, l9

CHAMBERS OF 48 3 Sheets-Sheet 2 'll-l MM MJBBM 7% #77 April 29, 1952 r v A. A. RUBBRA 2,594,808 MEANS FOR SUPPORTING m: NOZZLES OF THE COMBUSTION CHAMBERS OF INTERNAL-COMBUSTION TURBINES Filed March 8, 1948 3 Sheets-Sheet 3 of the pressed air-to-a plurality: of combustion chambers vrotor 101' rotors and the to the intermediate Patented Apr. 29, 1952 PATENT OFFICE 2,594,808 'Moms'roascrrommo-Tun NozzL-cs on 'THE U'QMBUS'IYION CHADIBERS OF IN TER- WhL-ooMBUsTIoN rURBINEs -ArthucAlexanderIRubbra,

v Littleover, Derby, England, 'ass'ignor to Rolls-R oyce Limited, Derby,

IEnglan'fl,ja'British company Application March 8, 1948;Sei-ialNoJ 13Q689 :sImGrea-t Britain "comers. (o1. 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 theintermediate casing.

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.

Inrsuchjrknown constructions the nozzles were welded to'gether'along radially abutting cages at their outlet ends and further the inner and outer diametral edgesat the outlet ends were welded to-inner and outer support rings. With "such a construction, it has been the practice to "support the inner ring from the intermediate casing referred to, whilst the turbine shroud ring or rings (together with "thejet pipe) were -supported from the outer ring. As a consequence the weight of the jet pipe was taken through the welded nozzles and also, where the nozzle guide vanes were rigidly secured by welding or otherwise to their inner and outer shroud rings, through itheguide vanes.

In the specification of'U. SsPatent No.2,494;82'1 granted January 17, 1950, there is described a construction in which the nozzle-boxes and nozzle guide-vanes are relieved of structural loading and which permits of these components being so supported that they can expand freely within the 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 ringQiThe 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.

In one constructionof apierturedwring according to this invention, 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.

By the invention, the weight of the ring may be kept to a minimum-having regard to aprodetermined overall strength thereon n According to another feature of this nvention, 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 Referring now e bolted to the inner ring 20 nular outlet to the end to the tubular member .23. applied to'the turbine shroud ring 22 by the exh'aust assembly 3, for example due to its weight,

centre of the spherical surface of the ring will be the apex of the cone.

There will now be described a gas-turbinecngine of which the nozzle-box assembly comprises one construction of apertured ring according to this invention, the description making reference to the accompanying drawings, in

view of the asingle-stage turbine generally indicated at H. a

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.

to Figure 2, it will be seen that 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. In this way 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. a

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. It will be understood that 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.

the ring 3! and the outer ring 14 of the guide vane assembly. and with an axial, forwardly-di rected flange 290 at its inner periphery which is bolted at 3-3 at its front end to the intermediatecasing 1'9" and structure 18 and at 34 at itsrear In this way loads 23 and inner ring guide-vane assembly to the casting 29 which transmits them to the intermediate casing [9. These parts constitute the main load carrying structure of the engine and the tubular member 20 which support the guide blades 25 are substantially unstressed. The guide are transmitted through the outer ring 2 of the entially by the outer rin combustion chamber the apertures, the

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| and 28 so that the nozzle-boxes are also substantially free from loads in the main engine structure.

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.

Referring now to Figures3 to 6, which'illustrate the casting 29 in detail, it will be; seen that 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.

F 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. As will be seen more clearly from Figure 5, 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.

In order to avoid giving the ring an excessive weight due to the increase in thickness between ring 29 is preferably cast with recesses it) in the thickened portion so that the metal between each pair-of apertures is in effect constituted by webs (see Figure 6) bounding the apertures circumferentially." It will be appreciated that this web-like formation gives the 'ring 1 and 43 forming guides and sealing surfaces respectively for the outlet ends of the combustion chambers and the inlet ends of the nozzle boxes. 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.

, receive and support the 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.

I claim:

1. 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 with their axes intersecting the turbine axis all at a single point; 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 liesat said point, and having formed therein a plurality of apertures, one aperture for each combustion chamber spaced circumferentially to correspond with the disposition of the outlet ends of the chambers, the angular position of said ring-like casing member being such that the axes of the combustion chambers intersect at said point, said ring-like casing member having a second machined surface opposite to said first machined surface, said second surface being of substantially non-uniform convex contour so that the said two surfaces cooperate to provide the greater thickness of the ring-like casing member in the region lying between said apertures and the less thickness in the peripheral regions of the ring-like casing member, and a plurality of apertured socket members one in each of said apertures to outlet ends of the combustion chambers, said socket members each being secured to said casing member on the spherical side thereof by flange means formed on the socket member.

2. An axial-flow gas turbine as claimed in claim 1 wherein the 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.

3. An axial-flow claim 1 wherein the gas turbine as claimed in cross-section of the portion of the ring-like casing member lying between the apertures, cone whose axis is the axis of the turbine, is of U-shaped form.

4. 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, and. a plurality of socket members one in each of said apertures to receive and support the outlet ends of the combustion chambers, said ARTHUR ALEXANDER RUBBRA.

REFERENCES CITED The following references are of record in the file of this patent:

UNITED STATES PATENTS on a plane forming the surface of a having a second machined sur-