US2584899A - Construction of stator elements of turbines, compressors, or like machines - Google Patents

Construction of stator elements of turbines, compressors, or like machines Download PDF

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Publication number
US2584899A
US2584899A US641146A US64114646A US2584899A US 2584899 A US2584899 A US 2584899A US 641146 A US641146 A US 641146A US 64114646 A US64114646 A US 64114646A US 2584899 A US2584899 A US 2584899A
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Prior art keywords
ring
rotor
stator
turbine
shroud
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US641146A
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English (en)
Inventor
Mcleod Roderick Cristall
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Power Jets Research and Development Ltd
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Power Jets Research and Development Ltd
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Priority claimed from GB1855/45A external-priority patent/GB597165A/en
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    • F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
    • F01—MACHINES OR ENGINES IN GENERAL; ENGINE PLANTS IN GENERAL; STEAM ENGINES
    • F01D—NON-POSITIVE DISPLACEMENT MACHINES OR ENGINES, e.g. STEAM TURBINES
    • F01D11/00—Preventing or minimising internal leakage of working-fluid, e.g. between stages
    • F01D11/08—Preventing or minimising internal leakage of working-fluid, e.g. between stages for sealing space between rotor blade tips and stator
    • F01D11/14—Adjusting or regulating tip-clearance, i.e. distance between rotor-blade tips and stator casing
    • F01D11/20—Actively adjusting tip-clearance
    • F01D11/24—Actively adjusting tip-clearance by selectively cooling-heating stator or rotor components
    • F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
    • F01—MACHINES OR ENGINES IN GENERAL; ENGINE PLANTS IN GENERAL; STEAM ENGINES
    • F01D—NON-POSITIVE DISPLACEMENT MACHINES OR ENGINES, e.g. STEAM TURBINES
    • F01D11/00—Preventing or minimising internal leakage of working-fluid, e.g. between stages
    • F01D11/08—Preventing or minimising internal leakage of working-fluid, e.g. between stages for sealing space between rotor blade tips and stator
    • F01D11/14—Adjusting or regulating tip-clearance, i.e. distance between rotor-blade tips and stator casing
    • F01D11/16—Adjusting or regulating tip-clearance, i.e. distance between rotor-blade tips and stator casing by self-adjusting means
    • F01D11/18—Adjusting or regulating tip-clearance, i.e. distance between rotor-blade tips and stator casing by self-adjusting means using stator or rotor components with predetermined thermal response, e.g. selective insulation, thermal inertia, differential expansion
    • 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
    • F01D25/00—Component parts, details, or accessories, not provided for in, or of interest apart from, other groups
    • F01D25/08—Cooling; Heating; Heat-insulation
    • F01D25/12—Cooling
    • 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
    • F01D25/00—Component parts, details, or accessories, not provided for in, or of interest apart from, other groups
    • F01D25/08—Cooling; Heating; Heat-insulation
    • F01D25/14—Casings modified therefor
    • 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
    • F01D25/00—Component parts, details, or accessories, not provided for in, or of interest apart from, other groups
    • F01D25/24—Casings; Casing parts, e.g. diaphragms, casing fastenings
    • F01D25/246—Fastening of diaphragms or stator-rings
    • 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
    • F01D25/00—Component parts, details, or accessories, not provided for in, or of interest apart from, other groups
    • F01D25/24—Casings; Casing parts, e.g. diaphragms, casing fastenings
    • F01D25/26—Double casings; Measures against temperature strain in casings
    • 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
    • F01D9/04—Nozzles; Nozzle boxes; Stator blades; Guide conduits, e.g. individual nozzles forming ring or sector
    • F01D9/041—Nozzles; Nozzle boxes; Stator blades; Guide conduits, e.g. individual nozzles forming ring or sector using blades
    • 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
    • F01D9/04—Nozzles; Nozzle boxes; Stator blades; Guide conduits, e.g. individual nozzles forming ring or sector
    • F01D9/042—Nozzles; Nozzle boxes; Stator blades; Guide conduits, e.g. individual nozzles forming ring or sector fixing blades to stators
    • 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
    • F02C7/185—Cooling means for reducing the temperature of the cooling air or gas
    • 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

  • This invention relates to the construction of stator elements of turbines, compressors or like machines and has for its principal object to provide a construction of working fluid channel for the stator portions of such machines which will simplify the problem of maintaining at reasonably small value the nominal clearance between a wall of such a channel and an element which has to be contained with a clearance therein, notwithstanding that the parts are subject to expansion and contraction due to temperature variation.
  • the invention is deemed to be particularly applicable to, and has special value in machines in which a large temperature variation is experienced, and thus has particular application to internal combustion gas turbines operating at high temperatures, for example axial flow gas turbines such as are used in aircraft power plants, in which the problem of maintaining a small clearance between the working fluid channel and moving parts therein arises in its most acute form.
  • the invention has been evolved primarily to meet this problem as it arises in turbines of that type in the case of the clearance between the tips of the turbine rotor blades and their stator shroud, and will be described in relation to that particular case without prejudice to the generality of the scope of the appended claims.
  • stator construction presents a smaller mass than the rotor and thus heats more rapidly, with the result that the stator diameter and therefore the blade tip clearance increases on starting, and emciency falls until the rotor becomes heated.
  • this difllculty cannot be met by a preliminary warming-up as with steam turbines; on the contrary, it is in the nature of the machine that a rapid temperature rise is inevitable.
  • the stator being more favourably situated for cooling as well as having a smaller mass, cools more rapidly than the rotor so that unless adequate tip clearance is allowed in the first place, the stator will shrink on to the blades, causing serious damage.
  • a further object of the invention is to enable the stator bladlng, whether as a whole or as individual blades or pluralities of blades, to be re- 2 moved or replaced simply and easily.
  • Yet another object is the provision of a stator construction especially well adapted to the requirement 0f multistage machines such that the stator as a whole can be built up in the course of assembly of a complete machine.
  • a turbine, compressor, or like machine especially an internal combustion turbine operating at high temperatures, having a stator working fluid channel containing an element which has to have a clearance therein, both the channel and the element being required to be subject to thermal radial expansion and contraction, wherein a boundary of said channel is defined by shrouding embodying expansion joints effective in the peripheral direction such as will allow thermal expansion and contraction of the material of the shroud without a corresponding variation in the radius thereof, said shrouding being supported in relation to the channel assembly by means which control the radius of the shrouding independently of the expansion and contraction characteristic of the latter, so that at any time said radius depends upon the condition of said supporting means and not upon the temperature of the shroud.
  • the shrouding is constituted by a high temperature inner wall carried by a low temperature support made as a rigid unit which is screened from the channel by and operates within a smaller temperature range than the shrouding, the mounting of the inner wall on its support being such that the radius of the former is determined by the condition of the latter.
  • the high temperature inner wall will be subject only to the same variations in diameter as the low temperature outer casing, so that the diameter of said inner wall may be maintained constant to a degree determined by the limits of temperature variation permitted to the outer casing, and it becomes possible, by allowing a cold blade tip clearance calculated only on the estimated expansion of the rotor, to eliminate the danger of the shroud ring shrinking into contact with the blades when cooling. It may be, however, in some cases that such clearance may still exceed an acceptable value when the turbine is being started from cold, due to the fact that the rotor heats relatively slowly. To meet this difficulty the rotor, according to a further feature of the invention, is also cooled so that the tip clearance at starting may the more closely approach that existing under running conditions.
  • the low temperature outer casing of the stator assembly is built up of a plurality of ring units (which may be continuous peripherally or may be built up in segments) interattached axially; these ring units by their juxtaposition locate the roots of stator blades in internal grooves presented by the assembly of rings; the high temperature inner wall of the stator assembly is built up (in the case of a multistage turbine) of a plurality of ring units; and the high temperature inner wall and low temperature outer casing, or their component ring units, have mutually interengaging formations by which the inner wall is rigidly locked to the outer casing.
  • the invention further provides for cooling the latter.
  • the said inner wall is mounted on the outer casing in radially spaced relationship thereto and provision is made for a circulation of collant to take place through the cavity so provided.
  • the means employed for cooling the turbine rotor preferably includes ports and guides coupled to a source of compressed air and adapted to supply such air to and conduct it along one or more surfaces of the rotor, the preferred arrangement being one in which only a thin layer of rapidly changed air passes over the surface to be cooled.
  • Figure 2 is an enlarged axial section of the turbine incorporated in said plant.
  • Figure 3 is a detail of the inner high temperature shroud construction.
  • the power plant in which the invention has been illustrated comprises a multistage axial flow, centrifugal discharge compressor I of which only the final stages are shown, said compressor supplying air under pressure to an annular aircasing or combustion chamber 2 defined by inner and outer walls 3, 4, symmetrically surrounding a shaft 5 by which the compressor I is driven from a turbine generally indicated as 8.
  • the atmosphere of the chamber I enclosed by the inner wall 3 of the aircasing 2 is maintained underpressure by supplying compressed air through a cooler 8 and a pipe connection 9 from the aircasing 2.
  • the latter contains, peripherally spaced around its annulus, combustion elements comprising fuel nozzles or burners III directed into flame tubes I I, which discharge the working gases at the inlet nozzle I2 of the turbine.
  • the outlets from the flame tubes II are peripherally flattened to form seg-.
  • the preferred application of the invention is to a turbine which, as illustrated, has a working fluid passage that expands in the downstream direction.
  • the turbine nozzle and stator blade assembly are preferably built up by an assembly of ring elements of appropriately varying diameter which are mutually interengaging and supporting and constitute together the high temperature inner wall and low temperature outer casing of the invention, and are assembled by being entered from the wider end of the passage as the assembly progresses, each successive element to be so entered retaining the one before it.
  • the turbine nozzle and stator blade assembly is constructed as follows (see more particularly Figure 2)
  • the first element of the rigid outer casing which also forms the outer wall of the turbine nozzle ring, is a first casing ring I3, being the casing ring of least diameter, which is attached to the engine structure by bolts or the like and comprises, extending radially inwardy and outwardly at its forward end, a flange I4 and at its rearward edge internally at rebated groove and also at its rearward edge an external radially extending flange I5 for the attachment of the second casing ring.
  • the nozzle ring stator blading comprises individual blades I6 each having a platform I! at its root, of rectangular or rhomboidal form.
  • a rib or key I8 which lies in a corresponding axially directed or skewed groove formed in the ring I3 and this rib at its rear end has a tooth I9 to engage the rebated groove of the ring.
  • Each blade is entered from the rear until the front edge of its platform is adjacent to the first mentioned flange I4 of the ring I3 and its tooth I9 is engaged; its rib I8 is now in the complementary groove in the ring l3.
  • the second casing ring 20 is now positioned against the first ring I3 by a flange 2I whose inner end abuts against the rear edges of the teeth I9 of the blades I5, locking them in their rebated groove.
  • at its outer part is extended forwardly at 22 to form an angle by which the outer edge of the flange I5 of the ring I3 is located, and also to meet the outer part of the flange I4, to which it is bolted and with which and the first ring I3 it defines an annular air chamber 23 to which cooling air is supplied from the cooler 8 by way of a connection 24 to an inlet 25 associated with the chamber 23.
  • the latter has outlets to the spaces between the ribs I8 of the blades It.
  • This second casing ring 28 is of frusto-conical form and at its rear edge presents a radially outwardly extending flange 26 interengagement with the rear edges of the blade platforms I! already in place, through a stepped formation 21a.
  • the shroud formed by the ring of sections 21 is spaced from the outer casing ring to define therewith an air cavity 30, having an inlet port 3
  • the radial flange 32 also has a peripheral forwardly directed flange engaging a corresponding peripheral groove in the second casing ring 20, thereby, with the stepped formation 28, supporting the shroud 21 in radially fixed relationship to the ring 20.
  • the third low temperature casing ring 35 in the main, is again of frusto-conical form and has set back slightly from its forward edge a radially outwardly extending flange 36 at whose periphery is a forwardly extending annular flange 31 which, with the flanges 36, 26 defines an annular chamber of generally rectangular channelled section in which are accommodated the roots 3% of a row of stator blades 39.
  • the roots 38 are located in the channel formation by bolts 40 engaging axial grooves therein, which bolts also serve to secure together the flanges 36, 26.
  • the roots 38 are further positioned by the interengagement therewith of an axially raised rib on the flange 36. This rib also, with the edge of the ring 35, defines an air cavity 4
  • the roots 38 are themselves recessed to form coolant passages, as indicated at 42 and in dotted lines.
  • the rear edge of the third casing ring 35 has a radial flange 43 surrounding an undercut groove, similarly to the rear edge of the second ring 20, the said groove receiving a forwardly directed flange 44 at the rear edge of a second high temperature inner shroud ring 45 which is similar in its construction to the first shroud ring 21 and is supported by the interengagement of its forward edge with the blade root 38.
  • a coolant space 45 is left between the second shroud ring 45 and the third casing ring 35, the space having an inlet port 41 by which air is received from the passages of the blade root 38, and an outlet port 48.
  • the rear face of the second shroud ring 45 has attached to it an extension piece 49, which is of similar construction to the shroud ring 45 in that it is in peripherally discontinuous sections interconnected by expansion joints.
  • Said extension piece 49 forms a liner for the exhaust duct wall 50, which has a flange and bolt attachment with the flange 43 of the third casing ring.
  • the tips of the blades 16 are also formed with platforms 56 which abut peripherally to form a virtually continuous floor for the nozzle passage and also enclose beneath them a chamber 51 for cooling air, which may enter by way of an inlet at 58 between the aircasing structure and platforms 56 from the main air stream emerging from the air-casing, and be discharged at 59 to the slight gap between the rotor and the rear edge of the nozzle ring 5
  • the inner high temperature shrouds 21, 45 by virtue of their peripherally effective expansion joints are capable of maintaining a constant diameter notwithstanding temperature variations; further, by virtue of their attachment to the low temperature outer casing rings 13, 20, 35, their actual diameter at any time is determined by the diameter of said casing rings; and finally that by reason of the circulation of cooling air from the chamber 23 through the cavities between the shroud rings and outer casing rings the temperature variation of the latter and the consequential variation in their diameter, and therefore, also, in that of the shroud rings can be kept within quite narrow limits, thus enabling a smaller tip clearance to be left between the rotor blades and shrouding when cold.
  • , are reinforced by an annular disc 66 having at its inner edge an enlarged ring formation 6'! which engages the inner faces of annular ribs 68 on the rotors 60, 6
  • the disc 66 At its outer edge the disc 66 has a flanged rim 10 which occupies the axial gap between blade mountings in the rims of the rotors 60, 6
  • the disc 66 is peripherally interrupted by short radial slots (not seen in the drawings) which, for the purpose of preventing leakage of cooling air direct from the chambers 13 to the working gas stream, instead of through or under the blade roots, are constructed so that under working conditions they are closed at their radially outer ends. This may be achieved by making the slots as narrow saw cuts whose edges are burred over to close the slots by a hammering operation prior to the final machining of the rim 10.
  • the weak flanges thus formed are crushed at the first expansion of the rim so that the slots, though open when the disc 66 is cold, are closed at the working temperature. It is desirable, in order to maintain a balanced flow of air past the blade roots in the rotors 50, 6
  • the forward face of the rotor 60 is screened by an air guide 14 which is mounted on a fixed part of the bearing 15 of the shaft and defines a narrow chamber extending along the greater part of the radius of the rotor 50.
  • a similar air guide 16 is formed by a part of the exhaust structure adjacent to the rear face of the rotor 6
  • the shaft 5 is ported to receive compressed air from the chamber 1, and this passes by way of outlet ports 80 to the passages 64 in the hub of the rotor GI and thence up the inner faces of the two rotors and the rear face of the rotor, eventually leaking into the working fluid as before.
  • a turbine, compressor. or like machine especially an internal combustion turbine operating at high temperatures, having radially inner and outer structures defining inner and outer boundary walls of an annular axial flow working fluid channel, and an element in said channel which has to have a radial clearance from at least one wall thereof, both the walls and the element being required to be subject to thermal radial expansion and contraction; at least one of said boundary walls comprising a shroud ring exposed to the fluid in and forming a high temperature boundary surface of said channel, and a lower temperature shroud ring-supporting structure screened from the channel by said shroud ring, said shroud ring comprising a plurality of peripherally discontinuous sections having slots at their axially extending edges and mounted on said supporting structure in fixed radial relation thereto so that the radius of the ring is determined thereby, and said sections having between the expansion joints which close the spaces between successive sections to provide for its peripheral continuity and are effectlve;to accommodate expansion and contraction of said sections in the peripheral direction, each of said expansion joints being formed by a
  • An axial flow turbine, compressor, or like machine especially an internal combustion turbine operating at high temperatures, having a bladed axial flow rotor defining an inner boundary wall of an annular axial flow working fluid channel; an outer annular structure enclosing said blading and defining an outer boundary wall of said channel; said outer structure comprising a radially inner shroud ring exposed to the fluid in and forming the outer high temperature boundary surface of said channel, and a lower temperature shroud ring-supporting structure screened from the channel by said shroud ring, said shroud ring comprising a plurality of peripherally discontinuous sections having slots at their axially extending edges and mounted on said supporting structure in fixed radial relation thereto so that the radius of the ring i determined thereby, and said sections having between them expansion joints which close the spaces between successive sections to provide for it peripheral continuity and are effective to accommodate expansion and contraction of said sections in the peripheral direction, each of said expansion Joints being formed by a strip element which engages in slots in axial edges of
  • An axial flow turbine, compressor, or like machine especially an internal combustion turbine operating at high temperatures, having a bladed axial flow rotor defining an inner boundary wall of an annular axial flow working fluid channel; an outer annular structure enclosing said blading and defining an outer boundary wall of said channel; said outer structure comprising a radially inner shroud ring exposed to the fluid in and forming the outer high temperature boundary surface of said channel, and a lower temperature shroud ring-supporting structure screened from the channel by said shroud rin said shroud ring being mounted on said supporting structure in fixed radial relation thereto so that its radius is determined thereby, and expansion gaps being embodied in the shroud ring which accommodate its expansion and contraction in the peripheral direction; said lower temperature shroud-supporting structure being built up of a plurality of ring elements inter-attached in axial succession to form a rigid unit and having blading mounted thereon to extend radially inwardly into the working fluid channel through said inner high
  • said further structure being separably attached to the downstream end of said extension; and a shroud ring surrounding but not substantially exceeding in axial extent said rotor blading, said shroud rin extending from the downstream side of said nozzle blading to the downstream end of said extension and comprising a plurality of separate arcuate elements free to expand peripherally and being supported in fixed concentric relation to said stationary structure and extension by annular seating means at its upstream and downstream edges effective to afford uniform peripheral distribution of such support while allowing freedom of expansion to said arcuate elements in the peripheral direction; said shroud ring being radially spaced from said extension and defining therewith an enclosed chamber.
  • An axial flow turbine, compressor, or like machine especially an internal combustion turbine operating at high temperatures, having at least one annulus of stator blading and at least one annulus of rotor blading downstream thereof mounted on a radially inner rotor; stationary structure axially substantially coextensive with and annularly surrounding said one annulus of stator blading and said one annulus of rotor blading, said stator blading being supported by attachment to said stationary structure, further stationary structure surrounding the downstream continuation of the working fluid channel through said blading and attached to said first mentioned structure; a shroud ring spanning axially the region from the outlet of said stator blading to the downstream edge of said first mentioned structure, said shroud ring forming the immediate boundary of the working fluid channel in that region and surrounding said rotor blading; said shroud ring comprising a plurality of separate arcuate elements between which are expansion gaps and each having slots at their axially extending edges and being supported'in fixed con
  • An axial flow turbine, compressor, or like machine especially an internal combustion turbine operating at high temperatures, having at least one annulus of stator blading and at least one annulus of rotor blading downstream thereof mounted on a radially inner rotor; stationary structure axially substantiall coextensive with and annularly surrounding said one annulus of stator blading and said one annulus of rotor blading, said stator blading being supported by attachment to said stationary structure; further stationary structure surrounding the downstream continuation of the working fluid channel through said blading and attached to said first mentioned structure; a shroud ring spanning axially the region from the outlet of said stator blading to the downstream edge of said first mentioned structure, said shroud ring forming the immediate boundary of the working fluid channel in that region and surrounding said rotor blading; and said shroud ring comprising a plurality of separate arcuate elements each being supported in fixed concentric spaced radial relation to said first mentioned structure by annular seating
  • a multistage axial flow turbine. compressor. or like machine, especially an internal combustion turbine operating at high temperatures, comprising at least two stages of rotor blading mounted on a radially inner rotor to define the inner boundary of a working fluid passage through the blading; stator blading extending radially inwardly between said stages of rotor blading to constitute an intermediate stator stage in relation thereto, stationary structure annularly surrounding said rotor blading and comprising in respect of two successive stages, each consisting of one stage of stator and rotor blading, successive ring units each substantially coextensive in the axial direction with one set of stator and rotor blading.
  • each shroud ring comprising a plurality of separate arcuate elements between which are expansion means and being supported in fixed concentric radial relation to said stationary structure by annular seating means at the upstream and downstream edges of the ring effective to afford uniform peripheral distribution of such support while allowing freedom of expansion to the elements of the ring in the peripheral direction.
  • An axial-flow gaseous-fluid turbine especially a gas-turbine operating at high temperatures, comprising a rotor incorporating at least one ring of rotor blades, a stator structure enclosing the rotor, incorporating at least one ring of stator blades upstream of the rotor blades, and, within said stator structure, a stationary shroudring constituted by a ring of separate, short, axially straight arcuate elements enveloping said ring of rotor blades as an outer boundary of the gas path through the turbine blading and spanning only the one ring of rotor blades, the arcuate elements being separated by expansionpermitting means, each edge of each arcuate ele ment being held on the stator structure by a ture into which fits an axially-extending part of the other of said arcuate element and stator struca,se4,soo
  • a turbine according to claim 9 wherein the expansion permitting means between two adjacent arcuate elements is constituted by the axially extending edge of one element overlapping the corresponding edge of the other element.
  • a turbine according to claim 9 wherein the expansion permitting means between two adjacent arcuate elements is constituted by the axially extending edge or one element entering a groove in the corresponding edge of its neighbor.

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  • Engineering & Computer Science (AREA)
  • Mechanical Engineering (AREA)
  • General Engineering & Computer Science (AREA)
  • Chemical & Material Sciences (AREA)
  • Combustion & Propulsion (AREA)
  • Turbine Rotor Nozzle Sealing (AREA)
US641146A 1945-01-23 1946-01-14 Construction of stator elements of turbines, compressors, or like machines Expired - Lifetime US2584899A (en)

Applications Claiming Priority (1)

Application Number Priority Date Filing Date Title
GB1855/45A GB597165A (en) 1945-01-23 Improvements relating to the construction of stator elements of turbines, compressors or like machines

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US2584899A true US2584899A (en) 1952-02-05

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US (1) US2584899A (de)
BE (1) BE463344A (de)
CH (1) CH279730A (de)
DE (1) DE859089C (de)
NL (1) NL70901C (de)

Cited By (55)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
US2638743A (en) * 1949-04-29 1953-05-19 Ruston & Hornsby Ltd Construction of turbine-inlet and stator elements of gas turbines
US2672013A (en) * 1950-06-30 1954-03-16 Curtiss Wright Corp Gas turbine cooling system
US2702687A (en) * 1950-11-13 1955-02-22 United Aircraft Corp Rotor construction
US2710523A (en) * 1951-09-27 1955-06-14 A V Roe Canada Ltd Gas turbine tail cone
US2738920A (en) * 1950-12-23 1956-03-20 Gen Motors Corp Gas turbine engine with thrust balancing coupling
US2763462A (en) * 1950-01-11 1956-09-18 Gen Motors Corp Turbine casing construction
US2791090A (en) * 1952-08-05 1957-05-07 Bristol Aeroplane Co Ltd Improved cooling and lubricating arrangement for bearings of a gas turbine engine
US2791091A (en) * 1950-05-15 1957-05-07 Gen Motors Corp Power plant cooling and thrust balancing systems
US2793832A (en) * 1952-04-30 1957-05-28 Gen Motors Corp Means for cooling stator vane assemblies
US2837892A (en) * 1949-10-08 1958-06-10 Joseph E Van Acker Combustion gas turbine assembly embodying a plurality of like turbine members
US2858101A (en) * 1954-01-28 1958-10-28 Gen Electric Cooling of turbine wheels
US2859935A (en) * 1951-02-15 1958-11-11 Power Jets Res & Dev Ltd Cooling of turbines
US2859934A (en) * 1953-07-29 1958-11-11 Havilland Engine Co Ltd Gas turbines
US2863634A (en) * 1954-12-16 1958-12-09 Napier & Son Ltd Shroud ring construction for turbines and compressors
US2873909A (en) * 1954-10-26 1959-02-17 Svenska Rotor Maskiner Ab Rotary devices and casing structures therefor
US2917276A (en) * 1955-02-28 1959-12-15 Orenda Engines Ltd Segmented stator ring assembly
US2919104A (en) * 1953-12-02 1959-12-29 Napier & Son Ltd Interstage seals and cooling means in axial flow turbines
US3042364A (en) * 1960-12-12 1962-07-03 Gen Electric Sealing mechanism
US3056582A (en) * 1960-08-26 1962-10-02 Gen Electric Turbine stator construction
US3083532A (en) * 1953-09-07 1963-04-02 Rolls Royce Gas turbine engine with air-cooling means and means to control the temperature of cooling air by liquid injection
US3204406A (en) * 1960-04-04 1965-09-07 Ford Motor Co Cooling system for a re-expansion gas turbine engine
US3321910A (en) * 1964-09-04 1967-05-30 Rolls Royce Gas turbine lubrication
US3408044A (en) * 1965-07-23 1968-10-29 Bbc Brown Boveri & Cie Combustion gas turbine with cooled guide vane support structure
US3768817A (en) * 1972-04-27 1973-10-30 Westinghouse Electric Corp Static seal for a gas turbine
FR2450345A1 (fr) * 1979-02-28 1980-09-26 Mtu Muenchen Gmbh Dispositif pour reduire au minimum et maintenir constants des jeux existant dans les turbines axiales, notamment turbomachines a gaz
FR2450344A1 (fr) * 1979-02-28 1980-09-26 Mtu Muenchen Gmbh Dispositif pour reduire au minimum et maintenir constants les jeux a la crete des aubes existants dans les turbines axiales, notamment pour turbomachines a gaz
US4465429A (en) * 1982-02-01 1984-08-14 Westinghouse Electric Corp. Steam turbine with superheated blade disc cavities
EP0134186A1 (de) * 1983-08-01 1985-03-13 United Technologies Corporation Statoranordnung für eine Turbine
US4648791A (en) * 1984-06-30 1987-03-10 Bbc Brown, Boveri & Company, Limited Rotor, consisting essentially of a disc requiring cooling and of a drum
US4721433A (en) * 1985-12-19 1988-01-26 United Technologies Corporation Coolable stator structure for a gas turbine engine
US5394687A (en) * 1993-12-03 1995-03-07 The United States Of America As Represented By The Department Of Energy Gas turbine vane cooling system
US5482431A (en) * 1992-02-04 1996-01-09 Bayerische Motoren Werke Ag Arrangement for supplying cooling air to a turbine casing of an aircraft gas turbine
EP0919700A1 (de) * 1997-06-19 1999-06-02 Mitsubishi Heavy Industries, Ltd. Vorrichtung zum dichten der leitschaufeln von gasturbinen
WO1999054609A1 (en) * 1998-04-21 1999-10-28 Pratt & Whitney Canada Corp. Turbine engine with cooled p3 air to impeller rear cavity
US6382903B1 (en) 1999-03-03 2002-05-07 General Electric Company Rotor bore and turbine rotor wheel/spacer heat exchange flow circuit
US6578363B2 (en) * 2001-03-05 2003-06-17 Mitsubishi Heavy Industries, Ltd. Air-cooled gas turbine exhaust casing
US6691503B2 (en) * 2001-03-26 2004-02-17 Siemens Aktiengesellschaft Gas turbine having first and second combustion chambers and cooling system
EP1398474A2 (de) * 2002-08-15 2004-03-17 General Electric Company Zapfluft-Gehäuse für einen Verdichter
US20070003410A1 (en) * 2005-06-23 2007-01-04 Siemens Westinghouse Power Corporation Turbine blade tip clearance control
EP1744016A1 (de) * 2005-07-11 2007-01-17 Siemens Aktiengesellschaft Heissgasführendes Gehäuseelement, Wellenschutzmantel und Gasturbinenanlage
US20070243811A1 (en) * 2006-03-27 2007-10-18 Pratt & Whitney Canada Corp. Ejector controlled twin air source gas turbine pressurizing air system
US20090304502A1 (en) * 2008-05-23 2009-12-10 Honeywell International Inc. Pre-diffuser for centrifugal compressor
EP2159384A1 (de) * 2008-08-27 2010-03-03 Siemens Aktiengesellschaft Leitschaufelträger für eine Gasturbine
US20100236249A1 (en) * 2009-03-20 2010-09-23 General Electric Company Systems and Methods for Reintroducing Gas Turbine Combustion Bypass Flow
US20100247292A1 (en) * 2009-03-30 2010-09-30 General Electric Company System and Method of Cooling Turbine Airfoils with Sequestered Carbon Dioxide
US20120167595A1 (en) * 2010-12-30 2012-07-05 Nathan Wesley Ottow Gas turbine engine with secondary air flow circuit
US20130192252A1 (en) * 2012-01-31 2013-08-01 William A. ACKERMANN Gas turbine engine buffer system
US20140260292A1 (en) * 2011-10-24 2014-09-18 Siemens Aktiengesellschaft Gas turbine and method for guiding compressed fluid in a gas turbine
US20160061210A1 (en) * 2013-05-30 2016-03-03 Mitsubishi Heavy Industries, Ltd. Turbo compressor and turbo chiller using same
US20170138209A1 (en) * 2015-08-07 2017-05-18 MTU Aero Engines AG Device and method for influencing the temperatures in inner ring segments of a gas turbine
US10018116B2 (en) 2012-01-31 2018-07-10 United Technologies Corporation Gas turbine engine buffer system providing zoned ventilation
US10316854B2 (en) * 2016-03-31 2019-06-11 Rolls-Royce Plc Shaft and a turbomachine
US20190368381A1 (en) * 2018-05-30 2019-12-05 General Electric Company Combustion System Deflection Mitigation Structure
US10502135B2 (en) 2012-01-31 2019-12-10 United Technologies Corporation Buffer system for communicating one or more buffer supply airs throughout a gas turbine engine
US11499479B2 (en) * 2017-08-31 2022-11-15 General Electric Company Air delivery system for a gas turbine engine

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GB1086432A (en) * 1965-09-21 1967-10-11 Bristol Siddeley Engines Ltd Gas turbine engines
DE4304989A1 (de) * 1993-02-18 1994-08-25 Abb Management Ag Verfahren zur Kühlung einer Gasturbinenanlage

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US2241782A (en) * 1937-07-07 1941-05-13 Jendrassik George Gas turbine
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Cited By (79)

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US2638743A (en) * 1949-04-29 1953-05-19 Ruston & Hornsby Ltd Construction of turbine-inlet and stator elements of gas turbines
US2837892A (en) * 1949-10-08 1958-06-10 Joseph E Van Acker Combustion gas turbine assembly embodying a plurality of like turbine members
US2763462A (en) * 1950-01-11 1956-09-18 Gen Motors Corp Turbine casing construction
US2791091A (en) * 1950-05-15 1957-05-07 Gen Motors Corp Power plant cooling and thrust balancing systems
US2672013A (en) * 1950-06-30 1954-03-16 Curtiss Wright Corp Gas turbine cooling system
US2702687A (en) * 1950-11-13 1955-02-22 United Aircraft Corp Rotor construction
US2738920A (en) * 1950-12-23 1956-03-20 Gen Motors Corp Gas turbine engine with thrust balancing coupling
US2859935A (en) * 1951-02-15 1958-11-11 Power Jets Res & Dev Ltd Cooling of turbines
US2710523A (en) * 1951-09-27 1955-06-14 A V Roe Canada Ltd Gas turbine tail cone
US2793832A (en) * 1952-04-30 1957-05-28 Gen Motors Corp Means for cooling stator vane assemblies
US2791090A (en) * 1952-08-05 1957-05-07 Bristol Aeroplane Co Ltd Improved cooling and lubricating arrangement for bearings of a gas turbine engine
US2859934A (en) * 1953-07-29 1958-11-11 Havilland Engine Co Ltd Gas turbines
US3083532A (en) * 1953-09-07 1963-04-02 Rolls Royce Gas turbine engine with air-cooling means and means to control the temperature of cooling air by liquid injection
US2919104A (en) * 1953-12-02 1959-12-29 Napier & Son Ltd Interstage seals and cooling means in axial flow turbines
US2858101A (en) * 1954-01-28 1958-10-28 Gen Electric Cooling of turbine wheels
US2873909A (en) * 1954-10-26 1959-02-17 Svenska Rotor Maskiner Ab Rotary devices and casing structures therefor
US2863634A (en) * 1954-12-16 1958-12-09 Napier & Son Ltd Shroud ring construction for turbines and compressors
US2917276A (en) * 1955-02-28 1959-12-15 Orenda Engines Ltd Segmented stator ring assembly
US3204406A (en) * 1960-04-04 1965-09-07 Ford Motor Co Cooling system for a re-expansion gas turbine engine
US3056582A (en) * 1960-08-26 1962-10-02 Gen Electric Turbine stator construction
US3042364A (en) * 1960-12-12 1962-07-03 Gen Electric Sealing mechanism
US3321910A (en) * 1964-09-04 1967-05-30 Rolls Royce Gas turbine lubrication
US3408044A (en) * 1965-07-23 1968-10-29 Bbc Brown Boveri & Cie Combustion gas turbine with cooled guide vane support structure
US3768817A (en) * 1972-04-27 1973-10-30 Westinghouse Electric Corp Static seal for a gas turbine
FR2450345A1 (fr) * 1979-02-28 1980-09-26 Mtu Muenchen Gmbh Dispositif pour reduire au minimum et maintenir constants des jeux existant dans les turbines axiales, notamment turbomachines a gaz
FR2450344A1 (fr) * 1979-02-28 1980-09-26 Mtu Muenchen Gmbh Dispositif pour reduire au minimum et maintenir constants les jeux a la crete des aubes existants dans les turbines axiales, notamment pour turbomachines a gaz
US4465429A (en) * 1982-02-01 1984-08-14 Westinghouse Electric Corp. Steam turbine with superheated blade disc cavities
EP0134186A1 (de) * 1983-08-01 1985-03-13 United Technologies Corporation Statoranordnung für eine Turbine
US4648791A (en) * 1984-06-30 1987-03-10 Bbc Brown, Boveri & Company, Limited Rotor, consisting essentially of a disc requiring cooling and of a drum
US4721433A (en) * 1985-12-19 1988-01-26 United Technologies Corporation Coolable stator structure for a gas turbine engine
US5482431A (en) * 1992-02-04 1996-01-09 Bayerische Motoren Werke Ag Arrangement for supplying cooling air to a turbine casing of an aircraft gas turbine
US5394687A (en) * 1993-12-03 1995-03-07 The United States Of America As Represented By The Department Of Energy Gas turbine vane cooling system
EP0919700A1 (de) * 1997-06-19 1999-06-02 Mitsubishi Heavy Industries, Ltd. Vorrichtung zum dichten der leitschaufeln von gasturbinen
EP0919700A4 (de) * 1997-06-19 2000-12-13 Mitsubishi Heavy Ind Ltd Vorrichtung zum dichten der leitschaufeln von gasturbinen
US6217279B1 (en) 1997-06-19 2001-04-17 Mitsubishi Heavy Industries, Ltd. Device for sealing gas turbine stator blades
WO1999054609A1 (en) * 1998-04-21 1999-10-28 Pratt & Whitney Canada Corp. Turbine engine with cooled p3 air to impeller rear cavity
US6035627A (en) * 1998-04-21 2000-03-14 Pratt & Whitney Canada Inc. Turbine engine with cooled P3 air to impeller rear cavity
US6382903B1 (en) 1999-03-03 2002-05-07 General Electric Company Rotor bore and turbine rotor wheel/spacer heat exchange flow circuit
US6578363B2 (en) * 2001-03-05 2003-06-17 Mitsubishi Heavy Industries, Ltd. Air-cooled gas turbine exhaust casing
US6691503B2 (en) * 2001-03-26 2004-02-17 Siemens Aktiengesellschaft Gas turbine having first and second combustion chambers and cooling system
EP1398474A2 (de) * 2002-08-15 2004-03-17 General Electric Company Zapfluft-Gehäuse für einen Verdichter
EP1398474A3 (de) * 2002-08-15 2005-01-26 General Electric Company Zapfluft-Gehäuse für einen Verdichter
US7708518B2 (en) * 2005-06-23 2010-05-04 Siemens Energy, Inc. Turbine blade tip clearance control
US20070003410A1 (en) * 2005-06-23 2007-01-04 Siemens Westinghouse Power Corporation Turbine blade tip clearance control
EP1744016A1 (de) * 2005-07-11 2007-01-17 Siemens Aktiengesellschaft Heissgasführendes Gehäuseelement, Wellenschutzmantel und Gasturbinenanlage
WO2007006680A3 (de) * 2005-07-11 2007-04-26 Siemens Ag HEIßGASFÜHRENDES GEHÄUSEELEMENT, WELLENSCHUTZMANTEL UND GASTURBINENANLAGE
US20090035124A1 (en) * 2005-07-11 2009-02-05 Bohrenkaemper Gerhard Hot-Gas-Ducting Housing Element, Protective Shaft Jacket and Gas Turbine System
US8147179B2 (en) 2005-07-11 2012-04-03 Siemens Aktiengesellschaft Hot-gas-ducting housing element, protective shaft jacket and gas turbine system
US8302407B2 (en) 2006-03-27 2012-11-06 Pratt & Whitney Canada Corp. Ejector controlled twin air source gas turbine pressurizing air system
US7861536B2 (en) * 2006-03-27 2011-01-04 Pratt & Whitney Canada Corp. Ejector controlled twin air source gas turbine pressurizing air system
US20070243811A1 (en) * 2006-03-27 2007-10-18 Pratt & Whitney Canada Corp. Ejector controlled twin air source gas turbine pressurizing air system
US20110067413A1 (en) * 2006-03-27 2011-03-24 Pratt & Whitney Canada Corp. Ejector controlled twin air source gas turbine pressurizing air system
US8438854B2 (en) * 2008-05-23 2013-05-14 Honeywell International Inc. Pre-diffuser for centrifugal compressor
US20090304502A1 (en) * 2008-05-23 2009-12-10 Honeywell International Inc. Pre-diffuser for centrifugal compressor
EP2159384A1 (de) * 2008-08-27 2010-03-03 Siemens Aktiengesellschaft Leitschaufelträger für eine Gasturbine
US8281601B2 (en) * 2009-03-20 2012-10-09 General Electric Company Systems and methods for reintroducing gas turbine combustion bypass flow
US20100236249A1 (en) * 2009-03-20 2010-09-23 General Electric Company Systems and Methods for Reintroducing Gas Turbine Combustion Bypass Flow
US20100247292A1 (en) * 2009-03-30 2010-09-30 General Electric Company System and Method of Cooling Turbine Airfoils with Sequestered Carbon Dioxide
US8631639B2 (en) * 2009-03-30 2014-01-21 General Electric Company System and method of cooling turbine airfoils with sequestered carbon dioxide
US20120167595A1 (en) * 2010-12-30 2012-07-05 Nathan Wesley Ottow Gas turbine engine with secondary air flow circuit
US9228497B2 (en) * 2010-12-30 2016-01-05 Rolls-Royce Corporation Gas turbine engine with secondary air flow circuit
US9745894B2 (en) * 2011-10-24 2017-08-29 Siemens Aktiengesellschaft Compressor air provided to combustion chamber plenum and turbine guide vane
US20140260292A1 (en) * 2011-10-24 2014-09-18 Siemens Aktiengesellschaft Gas turbine and method for guiding compressed fluid in a gas turbine
US20130192252A1 (en) * 2012-01-31 2013-08-01 William A. ACKERMANN Gas turbine engine buffer system
US10415468B2 (en) * 2012-01-31 2019-09-17 United Technologies Corporation Gas turbine engine buffer system
US11560839B2 (en) 2012-01-31 2023-01-24 Raytheon Technologies Corporation Gas turbine engine buffer system
US11286852B2 (en) 2012-01-31 2022-03-29 Raytheon Technologies Corporation Gas turbine engine buffer system
WO2013154630A1 (en) * 2012-01-31 2013-10-17 United Technologies Corporation Gas turbine engine buffer system
US10018116B2 (en) 2012-01-31 2018-07-10 United Technologies Corporation Gas turbine engine buffer system providing zoned ventilation
US10502135B2 (en) 2012-01-31 2019-12-10 United Technologies Corporation Buffer system for communicating one or more buffer supply airs throughout a gas turbine engine
EP2809918A4 (de) * 2012-01-31 2015-09-16 United Technologies Corp Puffersystem für gasturbinenmotor
US10858951B2 (en) * 2013-05-30 2020-12-08 Mitsubishi Heavy Industries Thermal Systems, Ltd. Turbo compressor and turbo chiller using same
US20160061210A1 (en) * 2013-05-30 2016-03-03 Mitsubishi Heavy Industries, Ltd. Turbo compressor and turbo chiller using same
US10590788B2 (en) * 2015-08-07 2020-03-17 MTU Aero Engines AG Device and method for influencing the temperatures in inner ring segments of a gas turbine
US20170138209A1 (en) * 2015-08-07 2017-05-18 MTU Aero Engines AG Device and method for influencing the temperatures in inner ring segments of a gas turbine
US10316854B2 (en) * 2016-03-31 2019-06-11 Rolls-Royce Plc Shaft and a turbomachine
US11499479B2 (en) * 2017-08-31 2022-11-15 General Electric Company Air delivery system for a gas turbine engine
US20190368381A1 (en) * 2018-05-30 2019-12-05 General Electric Company Combustion System Deflection Mitigation Structure
CN110552747A (zh) * 2018-05-30 2019-12-10 通用电气公司 燃烧系统偏转减轻结构

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CH279730A (de) 1951-12-15
DE859089C (de) 1952-12-11
NL70901C (de)
BE463344A (de)

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