US2806355A - Axial flow turbine with means for admixing low temperature gas into the high temperature driving gas stream - Google Patents

Axial flow turbine with means for admixing low temperature gas into the high temperature driving gas stream Download PDF

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Publication number
US2806355A
US2806355A US221978A US22197851A US2806355A US 2806355 A US2806355 A US 2806355A US 221978 A US221978 A US 221978A US 22197851 A US22197851 A US 22197851A US 2806355 A US2806355 A US 2806355A
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Prior art keywords
gas
temperature
turbine
high temperature
axial flow
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Expired - Lifetime
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US221978A
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English (en)
Inventor
Schorner Christian
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MAN AG
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MAN Maschinenfabrik Augsburg Nuernberg AG
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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
    • 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
    • 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
    • F02C7/125—Cooling of plants by partial arc admission of the working fluid or by intermittent admission of working and cooling fluid
    • 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
    • F02C7/14—Cooling of plants of fluids in the plant, e.g. lubricant or fuel

Definitions

  • This invention relates to an axial flow turbine for hot gaseous driving agents.
  • the moving blades are the most critical elements in gas turbines. The difficulties in this respect are increased by non-uniformities in the local maximum temperature of the propellant gas occurring in the combustion chamber or in the intermediate heating device. These very peak loads may have a decisive influence upon the service life of the first blade rims and so of the whole turbine.
  • this problem is solved in such a way that the admitted gas is stratified or subdivided in zones of different temperature, in such a way that a circular zone of a suflicient radial extension in the vicinity of the root circle of the blade grid is passed by gas of a lower temperature.
  • the subdivision is effected in such a way that the mixing operation for adjustment of the normal temperature of the gas entering the blade grid is partly carried out only shortly before the inlet guide blade grid of the first rim of moving blades while normally this admixing takes place following the real combustion of the introduced fuel in the combustion chamber or chambers preposed to the turbine.
  • the hot gases are guided in nonscaling sheet metal inserts within the casing in order to permit an optimum utilization of the material required for the design of the casing of the turbine; these sheet metal inserts are practically relieved from the pressure by the compressed air sweeping them and having approximately the same pressure as the working gas, said air moreover holding the pressure-proof walls of the casing at a sufficiently low temperature.
  • the above mentioned last part of the ballast air required within the gas for adjusting the inlet temperature may be used for producing zones of different temperature in the admitted gas.
  • This last part of ballast air may either be taken from the pressure branch of the compressor or, preheated to a higher temperature, from the air heater which is heated by exhaust gas.
  • This air at first surrounding the sheet metal insert filled with hot gas, flows into the annular chamber before the first distributor of the turbine, in such a way that the kinetic energy of the hot working gas is impaired as little as possible by the feeding of the cooling admixture.
  • the last part of the inlet channel for feeding the working gas to the annular space of the blading of the axial flow turbine is corrugated.
  • the working gas is stratified as to the temperature, with a somewhat colder inner annular zone, by which the service life of the moving blading is substantially increased even in case of the desirable high inlet temperature of the gas and in case of a favorable utilization of the material according to the principles of light weight construction.
  • the balance of admixed air blown in may be adapted to the requirements of the construction in various Ways.
  • Either the outer ring zone may be kept on the normal specified temperature of the gas, or, depending on the mechanical and thermal resistance of the blade material and the desired service life, temperatures may be provided for the working gas which are elevated beyond the normay measure, even for this outer zone ofthe annular space of the blading, in which the mechanical tensile stresses on the moving blade are not so high.
  • Fig. 1 is an axial section of a multi-stage axial flow turbine having the invention applied thereto, provided for the supply of the air to be admixed from outside,
  • Fig. 2 is a perspective view of a part of the mixing insert
  • Fig. 3 is an axial section of a multi-stage axial flow turbine in which the air to be. admixed is passed within the turbine casing, V
  • Fig. 4 is an axial section of a multi-stage axial flow turbine in which additional fuel is fed Within the casing of the turbine, and
  • Fig. 5 is a section through a part ofthe guide plate within the casing.
  • a corrugated. covering or bellows shell 5 is arranged on the part of'the. labyrinth insert 4 projecting beyond the hub of the casing, so as to form an inner delimitation of the annular gas feeding space 50 for the first distributor 2a.
  • the bellows so that the depth of the folds is gradually increasing in the flowing direction of the gas, so that a transitional zone of favorable dynamic properties towards the cross sections of the passages before and behind this transitional zone is obtained.
  • the inner part of the bellows shell is connected, throughbranches 6, to the pressure branch of the compressor or another source of supply of a cooling agent to be admixed.
  • This air simultaneously serves as sealing air for the packing of the rotor from the chamber 7 which is supplied through bores 32. That is, part of this air enters the annular space 7 through bores 32 and travels from there via labyrinths 35 into the annular space 36.
  • suitable cut-outs 5a may be arranged on the bellowsshell, along the folds, or at the end thereof, in order to promote the equalization of temperature and kinetic energy. That is, the cutouts 5a effect reduction of undesirably large temperature differences between inner and outer annular zones; The gas jets immerging from out outs 5a are broken up in the stream of hot driving gas with some reduction of temperature.
  • FIG. 2 shows a fragmentary perspective view of the bellows shell 5 withthe cut-outs 5a.
  • Fig. 3 the invention is exemplified by way of a multistage turbine in which for reasons of mechanical strength the casing is constructed as a welded shell type casing.
  • the rotor 8 is of the disc-drum type and bears the movable blades 9.
  • the appertaining distributors 10 including their carrier 11designed as bipartite shellare suspended into .the casing 12, which is also of a bipartite welded type, in conventional manner by means of radial bolts. 13.
  • the rotor shaft is packed at the end faces of the casing by means of labyrinths within divided stuffing box inserts 14 and 15 which aresupplied with sealing or packing air through the hub of the casing.
  • the working gas is guided within the casing 12 which is sound for pressure, in a sheet metal insert 16 which is also of a bipartite type and clamped on the guide blade carrier 11 by means of clamping rings 17 and 17a.
  • Its inlet connectingbranch 18.within the branch 19 of the casing is connected to the preposed combustion chamber (not shown). Since the space between this sheet metal insert 16 and the casing 12 is filled with compressed air of practically equal pressure, it will be sufficient to'provide aJdeSign of this insert consisting of a non-scaling sheet metal which is easily adaptable to the requirements of aerodynamics.
  • the front face ofthis sheet metal insert 16 is provided with radially directed folds 20 in the direction of the radial.
  • a separate insert 24 folded in a corrugated form may be provided on the adjoining cylindrical covering 23 of the packing shell of the packing insert 14, the size and dimensions of this separate insert being influenced by the temperature conditions before the guide blade grid.
  • the cooling ducts of this insert 24 may also be used, through corresponding bores 25, for feeding air to the inner intermediate chamber 26 before the first guide blade disc, so as to reduce the amount of compressed packing air required in the packing insert 14, to cool the same, and to reduce the axial thrust force exerted by the rotor.
  • the operating conditions of the first moving blade rims are similar for a partial turbine of a GT-process located in direct succession to an intermediate preheating device, where the decrease of the temperature of the working gas due to the production of mechanical energy in the preceding turbine is more or less compensated by intermediate superheating; as is well known, this is achieved by after-combustion of a further partial amount of fuel introduced into the excess air carried along by the working gas as a ballast.
  • this is achieved by after-combustion of a further partial amount of fuel introduced into the excess air carried along by the working gas as a ballast.
  • the arrangement according to the present invention is modified in such a way that the rest of the fuel to be introduced, which may be gaseous, is added only within the feeding section of the turbine shortly before the :first distributor; the included areas of the annular inlet space associated to the zone of the root circle are formed with corrugated indentures transversely to the direction of 'flow, e. g., at the border of a sheet metal Y insert of the kind described, between which indentures working gas may flow which has not been fully superheated in the phase of intermediate superheating.
  • Fig. 4 shows the principle according to the present invention applied to the conditions existing behind an intermediate superheating apparatus.
  • the corresponding elements 9-19 are thesame as in the turbine according to Fig. -3 and are denoted by the same reference numerals.
  • the difference consists merely in the fact that instead of air a small share of fuel gas is introduced into the stationary turbine casing 12 where it serves also as a pressure cushion outside of thesheetmetal insert 16 for-the working gas whichhas been heated almost to its full temperature in a preposed re-heating chamber.
  • the required additionalfuel gas is' passed from the branch 21 through an annular space between the casing shell 12 and the insulating cover 28 ofthe sheetmetal insert 16, so 'as to get to similar radial folds 29 (through bores 34 in member 33) atthefront curvature of-this sheet metal insert 16. ;
  • the outer annular zone is heated to the full temperature of the working gas or even to a higher temperature, by means of the-fuel openings 30 provided in the folds.
  • Fig. 1 one radial fold 29 is shown in section. It will be understood that in this case the folds are closed at their ends facing the blading, as shown in Fig. 4.
  • the admixing of both air and fuel gas can be arranged in such a way that deviations of the feeding symmetry owing to the supply through the branch which is usually unilateral, are avoided, and that the admission along the circumference is equalized.
  • a further advantage of the principle according to the present invention resides in the fact that normally the conditions as to complete combustion and temperature at the exit of the combustion chamber, depending on the construction thereof, the fuel, load etc. are never completely equalized throughout the cross section of the flow and that, therefore, it is very useful that the temperature of the working gas admitted to the highly stressed root circle zones of the first moving blade rims is reduced in a reliable manner.
  • the combination which comprises a high temperature inlet channel for supplying said high temperature gas to said turbine blades and having a discharge end spaced from said turbine blades, a separate low temperature inlet channel for supplying said lower temperature gas and terminating adjacent said discharge end of said high temperature inlet channel and radially inwardly thereof, a corrugated partition separating said inlet channels providing alternating adjacent axial passageways of substantial radial extent for guiding said high and lower temperature gases in said channels in heat exchanging relationship therebetween, and a plurality of apertures in said corrugated partition communicating between inner and outer sides thereof for admixing said high and lower temperature gases in varying proportions at varying radial distances prior to impinging on said blades effecting a stratified radially outwardly increasing temperature gradient in said admixed gases a portion of said apertures being positioned in substantially
  • an axial flow gas turbine adapted to be driven by high temperature gas impinging on the turbine blades and having a source of compressed gas at a temperature substantially lower than said high temperature gas
  • the combination which comprises an annular high temperature inlet channel for supplying said high temperature gas to said turbine blades and having a discharge end spaced from said turbine blades, a separate annular low temperature inlet channel radially inward of said high temperature channel for supplying said lower temperature gas, said low temperature inlet channel having a discharge end adjacent said discharge end of said high temperature inlet channel and radially inwardly thereof, a corrugated partition separating said inlet channels at said discharge ends thereof providing axial passageways for guiding said high and lower temperature gases in said channels in heat exchanging relationship therebetween, said corrugated partition at said discharge ends of said channels forming first admixing means for admixing a substantial proportion of said lower temperature gas with radially inward portions of said high temperature gas prior to impinging upon said turbine blades, and second admixing means including a plurality of jet
  • an axial flow gas turbine adapted to be driven by high temperature gas impinging on the turbine blades 7 and having a source of compressed gas at a temperature substantially lower than said high temperature gas
  • the combination which comprises an annular high temperature inlet for supplying said high temperature gas to said turbine blades, an annular lower temperature inlet radially inwardly of said high temperature inlet for supplying said lower temperature gas to said turbine blades, a corrugated partition between said inlets forming axial passages for guiding alternating streams of said high and lower temperature gases toward said turbine blades in heat exchanging relationship therebetween, and admixing means including a plurality of apertures communicating between said high temperature and lower temperature inlets for admixing said lower temperature gas with said high temperature gas in varying proportions over the annular cross section of said inlets forming a concentric temperature stratification in said admixed gases prior to impinging on said turbine blades with a temperature gradient increasing in the radially outward direction.
  • an axial flow gas turbine adapted to be driven by a stream of gas impinging on the turbine blades and having a source of high temperature gas and a source of lower temperature gas
  • the combination which comprises a radially outer inlet channel for said high temperature gas, a radially inward inlet channel for said lower temperature gas, a corrugated partition separating said inlet channels and forming a plurality of axial passages for directing alternate streams of said high and lower temperature gases toward radially inner portions of said turbine blades, and a plurality of apertures in radially outward portions of said corrugated partition and communicating between said inlet channels for admixing said lower temperature gas with portions of said high temperature gas flowing radially outwardly of said passages to provide in said gas stream prior to impinging on said turbine blades a plurality of strata of different temperatures and different proportions of said high and lower temperature gases.

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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)
US221978A 1950-05-09 1951-04-20 Axial flow turbine with means for admixing low temperature gas into the high temperature driving gas stream Expired - Lifetime US2806355A (en)

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DE304143X 1950-05-09

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CH (1) CH304143A (fr)
FR (1) FR1044704A (fr)

Cited By (7)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
US3135496A (en) * 1962-03-02 1964-06-02 Gen Electric Axial flow turbine with radial temperature gradient
US3490747A (en) * 1967-11-29 1970-01-20 Westinghouse Electric Corp Temperature profiling means for turbine inlet
US3652181A (en) * 1970-11-23 1972-03-28 Carl F Wilhelm Jr Cooling sleeve for gas turbine combustor transition member
US4017207A (en) * 1974-11-11 1977-04-12 Rolls-Royce (1971) Limited Gas turbine engine
US4083649A (en) * 1976-05-05 1978-04-11 Carrier Corporation Cooling system for turbomachinery
US4195474A (en) * 1977-10-17 1980-04-01 General Electric Company Liquid-cooled transition member to turbine inlet
US4321006A (en) * 1980-03-05 1982-03-23 Von Ohain Hans J P Gas compression cycle and apparatus therefor

Families Citing this family (1)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
DE1269418B (de) * 1964-08-21 1968-05-30 Gen Motors Corp Gasturbine

Citations (15)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
US1368751A (en) * 1918-11-29 1921-02-15 Auguste C E Rateau Means for cooling turbine-rotors
US1708402A (en) * 1926-09-04 1929-04-09 Holzwarth Gas Turbine Co Turbine blade
FR781057A (fr) * 1934-01-29 1935-05-08 Cem Comp Electro Mec Procédé et dispositif pour protéger contre les hautes températures les organes de turbo-machines plongés dans un fluide chaud en mouvement, en particulier les aubes de turbines à gaz ou à vapeur
CH210655A (de) * 1938-09-16 1940-07-31 Sulzer Ag Axial arbeitende Brennkraftturbine.
US2326072A (en) * 1939-06-28 1943-08-03 Bbc Brown Boveri & Cie Gas turbine plant
US2434134A (en) * 1939-12-19 1948-01-06 Power Jets Res & Dev Ltd Cooling means for internal-combustion turbine wheels of jet propulsion engines
US2435042A (en) * 1942-11-09 1948-01-27 Goetaverken Ab Plural fluid turbine combining impulse and reaction blading
US2445661A (en) * 1941-09-22 1948-07-20 Vickers Electrical Co Ltd Axial flow turbine, compressor and the like
US2483616A (en) * 1947-05-22 1949-10-04 Svenska Flygmotor Aktiebolaget Rotor for multistage turbines or similar machines
US2488867A (en) * 1946-10-02 1949-11-22 Rolls Royce Nozzle-guide-vane assembly for gas turbine engines
USRE23172E (en) * 1940-09-21 1949-11-29 Bochi
US2501633A (en) * 1943-06-28 1950-03-21 Lockheed Aircraft Corp Gas turbine aircraft power plant having ducted propulsive compressor means
US2529946A (en) * 1941-10-30 1950-11-14 Rateau Soc Cooling device for the casings of thermic motors, including gas turbines
US2563269A (en) * 1943-05-22 1951-08-07 Lockheed Aircraft Corp Gas turbine
US2628066A (en) * 1946-10-02 1953-02-10 Rolls Royce Turbine disk

Patent Citations (15)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
US1368751A (en) * 1918-11-29 1921-02-15 Auguste C E Rateau Means for cooling turbine-rotors
US1708402A (en) * 1926-09-04 1929-04-09 Holzwarth Gas Turbine Co Turbine blade
FR781057A (fr) * 1934-01-29 1935-05-08 Cem Comp Electro Mec Procédé et dispositif pour protéger contre les hautes températures les organes de turbo-machines plongés dans un fluide chaud en mouvement, en particulier les aubes de turbines à gaz ou à vapeur
CH210655A (de) * 1938-09-16 1940-07-31 Sulzer Ag Axial arbeitende Brennkraftturbine.
US2326072A (en) * 1939-06-28 1943-08-03 Bbc Brown Boveri & Cie Gas turbine plant
US2434134A (en) * 1939-12-19 1948-01-06 Power Jets Res & Dev Ltd Cooling means for internal-combustion turbine wheels of jet propulsion engines
USRE23172E (en) * 1940-09-21 1949-11-29 Bochi
US2445661A (en) * 1941-09-22 1948-07-20 Vickers Electrical Co Ltd Axial flow turbine, compressor and the like
US2529946A (en) * 1941-10-30 1950-11-14 Rateau Soc Cooling device for the casings of thermic motors, including gas turbines
US2435042A (en) * 1942-11-09 1948-01-27 Goetaverken Ab Plural fluid turbine combining impulse and reaction blading
US2563269A (en) * 1943-05-22 1951-08-07 Lockheed Aircraft Corp Gas turbine
US2501633A (en) * 1943-06-28 1950-03-21 Lockheed Aircraft Corp Gas turbine aircraft power plant having ducted propulsive compressor means
US2488867A (en) * 1946-10-02 1949-11-22 Rolls Royce Nozzle-guide-vane assembly for gas turbine engines
US2628066A (en) * 1946-10-02 1953-02-10 Rolls Royce Turbine disk
US2483616A (en) * 1947-05-22 1949-10-04 Svenska Flygmotor Aktiebolaget Rotor for multistage turbines or similar machines

Cited By (7)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
US3135496A (en) * 1962-03-02 1964-06-02 Gen Electric Axial flow turbine with radial temperature gradient
US3490747A (en) * 1967-11-29 1970-01-20 Westinghouse Electric Corp Temperature profiling means for turbine inlet
US3652181A (en) * 1970-11-23 1972-03-28 Carl F Wilhelm Jr Cooling sleeve for gas turbine combustor transition member
US4017207A (en) * 1974-11-11 1977-04-12 Rolls-Royce (1971) Limited Gas turbine engine
US4083649A (en) * 1976-05-05 1978-04-11 Carrier Corporation Cooling system for turbomachinery
US4195474A (en) * 1977-10-17 1980-04-01 General Electric Company Liquid-cooled transition member to turbine inlet
US4321006A (en) * 1980-03-05 1982-03-23 Von Ohain Hans J P Gas compression cycle and apparatus therefor

Also Published As

Publication number Publication date
CH304143A (de) 1954-12-31
FR1044704A (fr) 1953-11-20

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