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Publication numberUS2658338 A
Publication typeGrant
Publication dateNov 10, 1953
Filing dateFeb 21, 1948
Priority dateSep 6, 1946
Publication numberUS 2658338 A, US 2658338A, US-A-2658338, US2658338 A, US2658338A
InventorsLeduc Rene
Original AssigneeLeduc Rene
Export CitationBiBTeX, EndNote, RefMan
External Links: USPTO, USPTO Assignment, Espacenet
Gas turbine housing
US 2658338 A
Abstract  available in
Images(1)
Previous page
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Claims  available in
Description  (OCR text may contain errors)

NOV- 10, LEDUC GAS TURBINE HOUSiNG Filed Feb. 21, 1948 INVENTOR RENE Lzvuc.

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Patented Nov. 10, 1953 GAS TURBINE HOUSING Ren Leduc, Toulouse, France Application February 21, 1948, Serial No. 10,059 In France September 6, 1946 Section 1, Public Law 690, August 8, 1946 Patent expires September 6, 1966 2 Claims.

This invention relates to gas-turbine engines of the kind in which a centrifugal compressor is directly coupled to a turbine and a fixed combustion chamber is located intermediate in the flow path between the compressor outlet and the turbine inlet, and especially to gas-turbine engines of this kind designed to run at very high speeds of the order of 30,000 R. P. M.

An object of the invention is an arrangement of the compressor-turbine assembly by which air for cooling the turbine blades, the rim of the turbine-wheel and the inner wall of the turbineexhaust annulus is delivered directly from the impeller of the compressor into the turbine annulus, the impeller being closely adjacent, and optionally integral with, the turbine wheel.

Another object of the invention is the provision of a special form of combustion chamber which ensures adequate turbulence and the formation of a layer of cool air between the outer wall of the combustion chamber and the flame region.

The preferred combustion-chamber arrangement comprises a single annular combustion chamber in the form of a solid of revolution struck about the compressor-turbine axis, of generally toroidal shape. Within this is preferably mounted an internal partition of similar general shape to the outer wall of the combustion chamber and spaced at a substantially constant distance from the outer wall; this partition extends to the entry of the combustion chamber and serves to divide the entering air into two streams of which the outer stream flows between the partition and the outer wall which it follows without separation and serves to cool the wall and partition, while the inner stream supplies the primary combustion air.

The said partition is preferably supported on stays which converge inwardly towards the compressor-turbine axis, on which the partition has at least limited freedom to move axially of the stays, thus permitting the partition to expand freely. The convergent stays may advantageously be constituted by the fuel-burners themselves.

Another object of the invention is the provision of improved diffuser means between the compressor outlets and the combustion chamber inlets.

Preferably, the compressor outlets deliver into a number of generally trumpet-shaped diffuser passages, which may be pierced in the solid metal of a diffuser housing, the expansion-cone portion 'of each diffuser-passage having a straight-line axis; and the section of the expansion-cone portion, which is preferably of circular section; increases in the down-stream direction in accordance with the low,

where w is the area of an orthogonal cross-section of the expansion-cone, s is the axial coordinate of the section, U the peripheral length of the section, M the local Mach number of the fluid stream, and k a coefficient depending mainly on the physical characteristics of the fluid.

The invention will be more fully understood from the following description with reference to the accompanying drawing of a gas-turbine engine in accordance therewith, which is given by way of example only and not of limitation, the scope of the invention being defined in the hereto annexed claims. In the drawing,

Figure 1 is an axial half-section of the gasturbine engine.

Figure 2 is a detail section along the line 22 of Figure 1.

Figure 3 is an axial section of two adjacent difiuser ducts corresponding approximately to the line 3-3 of Figure 1.

The gas-turbine engine shown in Figure 1 comprises a centrifugal impeller I mounted on an impeller boss 5, which is keyed to a turbine rotor 5 carrying blades 1. The rotor assembly I, 5, 6, I, is mounted on a shaft I6 supported by the stationary structure 2 of the engine in bearings I1, I 8, the rotor assembly being overhung on the end. of shaft [6 which projects beyond the bearing 3.

A second built-up fixed structure 8 forms the impeller casing and the turbine casing, and encloses an annular combustion chamber 20 of generally toroidal shape, bounded partly by a thin wall I I supported from the outer wall of the structure 8 and partly by an inner wall 8b of the latter structure, which together define a solid of revolution about the axis X-X of the shaft 16.

The fixed structures 8 and 2 define between them an annular air entry through which the air enters, as shown by the large arrow in Figure 1, to reach the eye of the impeller which is provided with two rows of curved inclined rotary guide vanes 3 mounted in the impeller boss 5.

An internal wall 8a of the structure 8 separates the outlet of the combustion chamber or turbine nozzle ring provided with stationary blading Ia from the peripheral part of the impeller cham- ,ber occupied by the tips of the impeller blades 1.

Between the wall 8a and the rim of the turbine rotor 6 is a small gap through which a part of 3 the air delivered by the impeller escapes directly into the turbine annulus, as shown by the arrow 11. This air serves to cool the roots of the turbine blades 1, the rim of the turbine rotor 6 and the internal wall of the turbine exhaust duct indicated at], to ,4

The nan of the ai'i' b'si tlg'e impeller follows the course indicated by arrow f2 and enters a series of diffuser ducts 4 formed in the mass of the intermediate wall 80 of the structure 8.

Within the combustion chamber 20 is an internal partition in which is of the same gerii-ai'roiin as and at a substantially constant distance from the outer wall I I. Supported in the outer wall hi the combustion chamber casing 8 3 e a number of fuel burners 5| which project into't e lgri chamber converging towards the axis X- of the engine. The burners l5 also act as stays supporting the combustion chamber wall II and ae tinon JJL-whicnare free to slide in. the axial 43991191 9f the bu ners 15. thus prov din amp freedom of expansion effects, g g

The partition l0 extends into the mouth of an annular diffuser outlet defined by an annular hood 2| and serves to divide the air-stream f: is- 1.P ,d fi sm n two stream ff ff. =91; .rhi hfim ants: f'.= P s es e wee rm yo r!"sediheauienwq o he 'c bus p h r d. e ie vi o tih na e a l t amends, 9Q QmJwwP m aa a e of col vnsiii 9 i .w ll 9? thec b t si wbs iw iqh. on rea hi heis bine a nu i9 @001 t e t ps a ihei rbinab d s 1 and we! a .T. me sian c! iif'z. 9f w ake 7 'r s iilia m s fiqn .sha b pr vid t m am u iin,a i,. pr,ilie. bs fn rs a d th form 0! fife om q n a bsraa i s mist ia P i i n th t .t'iiqminarys is given sufficient turbulence to ensure efficient combusi C e n th h s xi aa s b bs 'the cdn'ibus'tion'chamber follow paths of curved .9 m QP -PI B-R W1 I i t in; f 9 b ii9 ha 1t2 ,Ii n x q no thiscircular axis with the'plalri e of Figure 1 is moi ne ninan s 'bfi e' i i' rou d ng anther/constitute the gener tors or riyperboioia whichjis an'u doveionomo surfaee 'orrgiomtion, The l nen-1o: Figure 1 is the projectio for one 'ofjthese' duct axes o 'thefplane :or the figure, one a is'the' development of aicoiilcalsurface ,withfaiis X-X, whose gne'rotpris'tlie lin s-a or Figural. Th'e'lin'es A-A and a'fflyfot 3 are'the projections 'of the annular 'cluct axe on this conical "surf ace.

Irwin be se n that the diffuser ducts are trumpet-shaped andjthe expansion-cone parts thereof are of circular'cross-section, asshown at s's'in m l r i. .v

, 'jl 'he areas andper ipheries of these sections are related bythelawpreviously 'stated, and the cirjjcularsections of'the diffuser expansion cones aremerged into the'fsquare sections of thediffuserentryindlcatedat ll'in'Fig'ur'e 3 by means nyi mpcqnn i w a. a

I'o safeguard the shaft it while passing through critical: whirlingfspeeds' in running up to its operns s e fwhw i m b as h i 304000 ;R, 'l. ;'M,,"aclearance bearing [9 is interposed fietween'the bearings none la. This'beari'ng normally has a generous clearance from the shaft which is therefore supported only by the bearings [1, I8. As the shaft approaches a critical speed, however, it will deflect and come into contact with the bearing IS; the support thus provided will immediately approximately double the 'oi'rmoai speed. Ifthe speed continues to increase. the original critical is therefore passed through harmlessly and the shaft realigns itself at speeds between the (original) critical speed for two-point support and the temporary critical speed for three-point support; but once the shaft is ieallg'hedit ceases to make contact with the bearing 19 and consequently the former critical mood for two-point support is restored, the higher speed which is only critical for three-point sqenorti oing no longer critical in the final condi "ion. The engine may, therefore, be safely run up to its full operating speed which is normally well above the critical speed for two-point support of the shaft.

roman:

1 in agas turbine, in combination, an impeller an outer periphery and being mounted for rotation about itsaxis; a plurality of turbine iblades'also mounted 'for rotation about said impeller axis and being located beyond said outer periphery of said impeller and adjacent thereto; and a housing a central axis coincidl'fik with said impeller aic'is'a'nd said lin'p'ller and turbine blades for'iotation therein, said housin]; being formed with air inlet leading'to said impeller, with an combustion chamber located about said central ails thereof and with annular 'p alss'afg'e leading from said combustion to said turbine blades, and

housing having an wouno'rtion about said central 'a'iiis thereof and said impeller from said combustion chamber,'said wall portion having outer 'annula'r forming one side of said -annular'passage-leading from said combustion chamber to said turbine blades and sai d wallportion being formed'with a plurality of separated diffuserpassages extending therethrough each having one open end "facsaid impeller and an opposite open end facing sri5 =..m u iqe cha fl PM: being evenly distributed in said wall-portion and n extending along an axis which is skewed with respect to said central axis, and each of said passages increasing in cross-section from sa id one open end thereof to said opposite open end thereogwhereby air passing throughsaid-air inlet to said impeller must pass through said diffuser passages before entering said combustion chamber and whereby said air after-enteringsaid combustion chamber flows through said annular passage to said turbine blades, said diffuser passagesdecreasing the speed of the air flowing therethrough and increasingthe pressure thereof. a 2. In a gas turbine, in combination an impeller having v an outer periphery and being :rnounted for rotation about its axis; a=plurality of turbine blades also mounted forrotation about said impeller axis an'd-being' located beyond said outer peripheryof said impeller and adjacent thereto; a housing havinga centralaxis coinciding with; said impeller axis and encasing said impeller and turbine blades for rotation therein, said housing being formed with'an air inlet leadto i mpeller, with "an annular combustion chamber located about said central axis "thereoffand with an passage T leading from said "combustion chamber tosaid turbine blades, and Said an armular wan 5 portion located about said central axis thereof and separating said impeller from said combustion chamber, said wall portion having an outer annular surface forming one side of said annular passage leading from said combustion chamber to said turbine blades and said wall portion being formed with a plurality of separated diffuser passages extending therethrough and each having one open end facing said impeller and an opposite open end facing said combustion chamber, said diffuser passages being evenly distributed in said wall portion and each extending along an axis which is skewed with respect to said central axis, and each of said diffuser passages increasing in cross-section from said one open end thereof to said opposite open end thereof, whereby air passing through said air inlet to said impeller must pass through said diffuser passages before entering said combustion chamber and whereby said air after entering said combustion chamber flows through said annular passage to said turbine blades, said diffuser passages decreasing the speed of the air flowing therethrough and increasing the pressure thereof; and a plurality of stationary blades mounted in said annular passage leading from 6 said combustion chamber to said turbine blades, extending across the same and being located adjacent to said turbine blades.

RENE LEDUo.

References Cited in the file of this patent UNITED STATES PATENTS Number Name Date 1,926,225 Birmann Sept. 12, 1933 1,959,703 Birmann May 22, 1934 2,080,425 Lysholm May 18, 1937 2,114,285 Berger Apr. 19, 1938 2,157,002 Moss May 2, 1939 2,256,198 Hahn Sept. 16, 1941 2,272,676 Leduc Feb. 10, 1942 2,408,743 Elliott Oct. 8, 1946 2,489,692 Whittle Nov. 29, 1949 2,556,161 Bailey et a1 June 12, 1951 2,595,505 Bachle May 6, 1952 2,609,141 Aue Sept. 2, 1952 FOREIGN PATENTS Number Country Date 441,683 Germany Mar. 8, 1927

Patent Citations
Cited PatentFiling datePublication dateApplicantTitle
US1926225 *Sep 12, 1930Sep 12, 1933Rudolph BirmannTurbo compressor
US1959703 *Jan 26, 1932May 22, 1934Rudolph BirmannBlading for centrifugal impellers or turbines
US2080425 *Feb 9, 1934May 18, 1937Milo AbTurbine
US2114285 *Nov 28, 1936Apr 19, 1938Berger Adolph LDiffuser for centrifugal compressors
US2151002 *Aug 14, 1937Mar 21, 1939Lewis GompersPipe for smoking tobacco
US2256198 *May 31, 1939Sep 16, 1941Ernst HeinkelAircraft power plant
US2272676 *Dec 23, 1938Feb 10, 1942Rene LeducContinuous flow gas turbine
US2408743 *Oct 7, 1943Oct 8, 1946Elliott Albert GeorgeJet-propulsion apparatus for aircraft
US2489692 *Aug 24, 1945Nov 29, 1949Power Jets Res Dev LtdCompressor
US2556161 *Mar 20, 1945Jun 12, 1951Power Jets Res & Dev LtdGas diffusers for air supplied to combustion chambers
US2595505 *Apr 20, 1946May 6, 1952Continental Aviat & EngineerinCoaxial combustion products generator, turbine, and compressor
US2609141 *Oct 2, 1945Sep 2, 1952Sulzer AgCentrifugal compressor
DE441683C *Dec 1, 1922Mar 8, 1927Lorenzen G M B H CVerfahren zum Betrieb von Gasturbinen
Referenced by
Citing PatentFiling datePublication dateApplicantTitle
US2759662 *Apr 26, 1950Aug 21, 1956Carrier CorpCentrifugal compressors
US2801519 *Feb 17, 1951Aug 6, 1957Garrett CorpGas turbine motor scroll structure
US2855754 *Dec 31, 1953Oct 14, 1958Giannottl Hugo VGas turbine with combustion chamber of the toroidal flow type and integral regenerator
US3150823 *Feb 7, 1963Sep 29, 1964Ass Elect IndDiffusers
US3365892 *Aug 10, 1965Jan 30, 1968Derderian GeorgeTurbomachine
US3604818 *Dec 10, 1969Sep 14, 1971Avco CorpCentrifugal compressor diffuser
US3709629 *May 26, 1970Jan 9, 1973Traut EIntegrated flow gas turbine
US4070824 *Mar 12, 1975Jan 31, 1978Traut Earl WIntegrated flow turbine engine
US4151709 *Nov 26, 1976May 1, 1979Avco CorporationGas turbine engines with toroidal combustors
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US4487548 *May 19, 1983Dec 11, 1984Chandler Evans Inc.Centrifugal main fuel pump having starting element
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Classifications
U.S. Classification60/805, 415/211.1, 416/95, 415/178, 415/143
International ClassificationF02C7/04, F23R3/52, F02C3/045, F04D25/04, F02C6/20, F04D29/44
Cooperative ClassificationF02C7/04, F04D25/045, F23R3/52, F04D29/441, F02C6/20, F02C3/045
European ClassificationF02C7/04, F02C6/20, F23R3/52, F04D29/44C, F02C3/045, F04D25/04B