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Publication numberUS3877221 A
Publication typeGrant
Publication dateApr 15, 1975
Filing dateAug 27, 1973
Priority dateAug 27, 1973
Also published asCA993668A1, DE2422362A1, DE2422362B2, DE2422362C3
Publication numberUS 3877221 A, US 3877221A, US-A-3877221, US3877221 A, US3877221A
InventorsArthur H Lefebvre, Samuel B Reider, Harold L Stocker, Jerry G Tomlinson
Original AssigneeGen Motors Corp
Export CitationBiBTeX, EndNote, RefMan
External Links: USPTO, USPTO Assignment, Espacenet
Combustion apparatus air supply
US 3877221 A
Abstract
An improvement in diffusers to guide air flow from the compressor of a gas turbine engine into the combustion apparatus, splitting the air between front, outer, and inner walls of the combustion liner. A slot in a wall of the diffuser transfers air from compressor mid-radius to the passage leading to the liner inner wall.
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United States Patent Lefebvre et al.

[451 Apr. 15, 1975 COMBUSTION APPARATUS AIR SUPPLY Inventors: Arthur H. Lefebvre, Bedford,

England; Harold L. Stocker; Samuel B. Reider, both of Indianapolis, Ind.; Jerry G. Tomlinson, Speedway, lnd.

Assignee: General Motors Corporation,

Detroit, Mich.

Filed: Aug. 27, 1973 Appl. N0.: 391,888

US. Cl. 60/39.65 Int. Cl. F02c 3/00 Field of Search 60/3965, 39.66, 39.74 R

References Cited UNITED STATES PATENTS Lauck 60/3965 Kenworthy et al. 60/3965 Keiter et al. 60/39.65

Primary ExaminerC. J. Husar Assistant ExaminerRobert E. Garrett Attorney, Agent, or FirmPaul Fitzpatrick [57] ABSTRACT An improvement in diffusers to guide air flow from the compressor of a gas turbine engine into the combustion apparatus, splitting the air between front, outer, and inner walls of the combustion liner. A slot in a wall of the diffuser transfers air from compressor mid-radius to the passage leading to the liner inner wall.

3 Claims, 2 Drawing Figures COMBUSTION APPARATUS AIR SUPPLY The invention described herein was made in the course of work under a contract with the Department of Defense.

Our invention is directed to combustion apparatus for gas turbine engines. and particularly to an inlet diffuser structure adapted to divide the air flowing to the apparatus between several flow paths. More specifically. our invention resides in diffusing structure which is placed ahead of the combustion zone of an annular gas turbine combustor to divide the air between a portion flowing to the forward wall of the liner. a portion flowing to the radially outer wall. and a portion flowing to the radially inner wall.

Our invention is particularly adapted to maintain the desired division of flow between the several paths leading into the combustion apparatus notwithstanding changes in operating conditions of the engine which alter the pressure distribution profile at the outlet of a compressor which supplies the combustion apparatus with air.

Our apparatus is adapted to be employed at relatively high inlet velocity of the air with a broad operating range and a high pressure recovery for low pressure drop through the combustion apparatus. It is characterized by prevention of localized flow separation and by control of the air so as to maintain a good burner outlet pattern.

The principal objects ofour invention are to improve the operation of gas turbine combustion apparatus. particularly those of the annular type: to improve the inlet diffusing arrangements of such combustion apparatus; and to provide a combustion apparatus which is relatively insensitive to variations in the pressure and velocity profiles of air delivered to the combustion apparatus by a compressor which supplies the air for combustion.

The nature of our invention and its advantages will be clear to those skilled in the art from the succeeding detailed description of the preferred embodiment of the invention and the accompanying drawings.

FIG. 1 is a sectional view of the combustion apparatus of a gas turbine engine taken in a plane containing the axis of the engine and illustrating the environment of our invention.

FIG. 2 is an enlarged view corresponding to FIG. 1 of the diffuser and upstream end of the combustion apparatus.

Referring first to FlG. 1. a typical axial-flow gas turbine engine 2, which is only partly shown. includes. in flow series, an axial-flow compressor 3, combustion apparatus 4. and a turbine 6. Only the discharge end of the compressor is illustrated. and only a portion of the turbine nozzle 7 through which the combustion products flow is illustrated. As is well known. the turbine is connected by a shaft to drive the compressor to force compressed air into the combustion apparatus. Fuel is burned in the air so supplied and the resulting combustion products are fed to the turbine to drive the compressor. Power may be taken off as shaft power or as a pressurized exhaust stream for jet propulsion. The general structure of such engines is well known. and there is no need to describe such in greater detail to explain our invention.

The combustion apparatus 4 which is shown by way of example comprises an outer casing or wall 8 and an inner casing or wall 9. these defining between them an annular space extending from the outlet of the compressor to the inlet of the turbine. Combustion takes place in an annular combustion liner 10 disposed between the outer and inner walls. The combustion liner comprises an outer wall ll and an inner wall 12. these being approximately cylindrical and slightly tapered. An air passage 13 is defined by walls 8 and 11, and an air passage 14 by walls 9 and 12. The liner 10 also comprises a ring-shaped forward wall 15 fixed to and joining the outer and inner walls of the liner. Fuel nozzles 16 are mounted on struts 18 extending through the outer wall 8. through which fuel is supplied to the nozzle. In a particular example. there are sixteen such nozzles disposed equally aroung the axis of the engine.

The downstream ends of the passages 13 and 14 are closed by structures indicated schematically at I) and 20, respectively, which allow some flow of cooling air to the turbine. The details of these structures are immaterial to our invention.

The multistage compressor. so far as illustrated. comprises a final rotor stage 22 and two rows of outlet guide vanes 23. The compressor delivers that air in an axial direction into the combustion apparatus. specifically delivering air through the forward. outer. and inner walls of the liner.

Referring also to FIG. 2, the diffusion and division of the air is effected by a diffuser assembly 24. The diffuser comprises an outer diffuser wall 26 of sheet metal the rear edge of which overlaps and is welded or brazed to the forward portion of the liner outer wall 11. The forward end of outer diffuser wall 26 abuts and is welded to a ring 27 forming the outer boundary of an annular central air entrance 28. The inner boundary of this air entrance is defined by a ring 30. these rings being joined by circumferentially spaced radial struts or spacers 31. The inner diffuser wall 34 comprises a forward section 32 which is attached to the rear edge of the ring 30 and an aft section the rear end of which overlaps and is fixed to the forward end of the liner inner wall 12. Sections 32 and 35 of the inner diffuser wall 34 are overlapped and spaced to define a communicating passage 36, the parts being connected and maintained in proper relation by spacers 38 of any suitable structure distributed around the axis of the diffuser. As will be seen, the walls 26 and 34 define between them a rather large cavity or air space 39 to which air is introduced through a diverging diffuser passage 40 defined between rings 27 and 30. The assembly of rings 27 and 30 may be characterized as a snout 42. It will be noted that this snout projects very close to the outlet of the compressor; that is. the outlet guide vanes 23. and that it serves to split the air discharged by the compressor into three portions, one flowing through the diffusing passage 40, a second flowing through an outer diffusing passage 43 between walls 8 and 26 into passage 13, and a third flowing through an inner diffusing passage 44 between walls 30. 32. and 35 and the wall 9 into passage 14.

The means by which the diffuser 24 and the forward end of liner 10 are supported are not material to the invention. but it may be mentioned that they are attached to circumferentially spaced struts (not illustrated) extending between walls 8 and 9 abreast of snout 42. the forward edge of which is recessed to clear these struts.

The outer and inner walls 11 and 12 of the combustion liner are similar. Each consists of a number of overlapping sections which define narrow gaps at the over-laps between them for flow of air which performs the function of convection cooling of the liner and then flows over the heated surface of the wall for film cooling. Primary combustion air is admitted through holes near the forward end of the liner wall as indicated by the arrows 45. Secondary or dilution air flows through holes in the walls farther downstream as indicated by the arrows 46. Some primary combustion air is admitted through the fuel nozzle 16 to atomize the fuel. Additional air is admitted through the forward wall and flows between the margins of baffles 47 and the outer and inner walls to serve as film cooling air for the up stream end of the liner. The film cooling air so intro ducedmay become primary combustion air and. so far as any of this film cooling air is not combined with the fuel. it ultimately becomes dilution air. So far as our invention is concerned. any suitable structure of the forward. outer. and inner walls of the liner may be employed.

An air seal allowing for some relative motion is provided between fuel nozzle strut 18 and the diffuser outer wall 26. As illustrated in FIG. 2, the strut 18 has a circumferential flange 50 which is disposed with some clearance in a hole 51 in wall 26. A hat-section sheet metal ring or ferrule 52 is assembled around flange 50. The outer margin of the ferrule is slidably retained between the inner surface of wall 26 and a flanged retain ing ring 54 welded to wall 26. This structure minimizes air leakage around strut 18.

Our invention is directed to the diffuser structure 24 which diffuses and divides the air going to the three liner walls. In a typical case. the air may be divided into approximately 44% flowing into passage 13. including about 4% flowing past barrier 19 for turbine cooling;

4071 into passage 14 including some 6% flowing past barrier 20 for turbine cooling; and about l6% through the front liner wall 15.

However. the amount entering through the central air inlet 28 and flowing through diffusing passage 40 is. in this case. approximately 23% of the total air and about 7% of the air is redirected under normal operating-conditions through the air communicating passage 36 into the inner passage 14. The exact proportion of air which is diverted from cavity 39 through the communicating passage to the inner passage 14 varies with the change in outlet velocity and total pressure profiles of the-air entering the combustion apparatus through vanes 23-of the compressor. If the pressure near the inner boundary of the compressor discharge passage decreases the effect of this is to decrease both the quantity and pressure of air flowing through the inner passage 14. while. at the same time it creates an increase in pressure differential between cavity 39 and the exit of the inner diffusing passage 44. as a result of which an increased amount of air flows through the communicating passage 36 from cavity 39 to inner passage 14 thereby increasing both pressure and air flow quantity in passage l4.'ln this manner the communicating passage compensates for variations in compressor discharge velocity and pressure profiles.

In the example shown the diffusing passage 40 is directed much more toward the inner margin of the diffuser'assembly 24 than its'outer margin and thus the air flows more directly toward the communicating passage 36 into inner passage 14. Because of the great enlargement of the passage as it leaves passage 40 and enters cavity 39, the air in cavity 39 is rapidly diffused and thus there are no significant velocity effects to upset distribution of flow through the nozzles 16 and the cooling air inlets between the baffles 47 and the liner walls. Thus. there is a fast-moving stream of air flowing over the outer surface of both walls of the liner and a relatively slower moving stream of air within the cavity 39.

Tests of the structure illustrated and described herein has shown that the combustor is significantly less sensitive to variations in compressor outlet pressure and velocity which occur with changes in operating conditions ofthe engine in which the combustor is used. This is attributable to the provision of the communicating passage 36 from the central passage to the inner passage. which diverts some of the strong mid-radius air stream into the inner passage which is otherwise suppliedfrom near the rotor hub of the compressor. Thus. a good balance of air flow and pressure in the inner and outer passages respectively is achieved. resulting in a desirably uniform temperature distribution in the combustor efflux gases.

The detailed description of the preferred embodiment of the invention for the purpose of explaining the principles thereof is not to be considered as limiting or restricting the invention. since many modifications may 4 be made by the exercise of skill in the art.

We claim:

1. In a gas turbine engine comprising a compressor with an annular axial outlet. outer and inner casing walls defining an annular combustion space connected to the outlet. and an annular combustion liner disposed between the said casing walls. the casing walls having forward portions diverging to define a diffusing zone beginning at the compressor outlet. the combustion liner having an upstream wall and outer and inner walls with openings for entrance of'air into the liner in each of the said liner walls. the improvement comprising an airflow dividing diffuser disposed in the diffusing zone and connected to the upstream end of the liner. the diffuser having a snout providing an annular entrance adjacent the compressor outlet and at a mid-radius of the outlet. an outer diffuser wall extending from the snout to the outer wall of the liner. and an inner diffuser wall extending from the snout to the inner wall of the liner; the said diffuser walls bounding a diffusing flow path to the upstream liner wall and defining with the upstream liner wall an air space. the outer diffuser wall and outer casing wall defining a diffusing flow path to the outer liner wall. the outer diffuser wall blocking flow between the said air space and the outer liner wall. and the inner casing wall defining a diffusing flow path to the inner liner wall; the said diffusing flow paths taking air from the outer. central. and inner annular zones of the compressor outlet. respectively; and the diffuser defining means for diverting a portion of the air from the path leading to the upstream liner wall into the path leading to the inner liner wall responsive to relative pressure conditions in the said paths.

2. In a gas turbine engine comprising a compressor with an annular axial outlet. outer and inner casing walls defining an annular combustion space connected to the outlet. and an annular combustion liner disposed between the said casing walls. the casing walls having forward portions diverging to define a diffusing zone beginning at the compressor outlet. the combustion liner having an upstream wall and outer and inner walls with openings for entrance of air into the liner in each of the said liner walls. the improvement comprising an airflow dividing diffuser disposed in the diffusing zone and connected to the upstream end of the liner. the diffuser having a snout providing an annular entrance adjacent the compressor outlet and at a mid-radius of the outlet. an outer diffuser wall extending from the snout to the outer wall of the liner. and an inner diffuser wall extending from the snout to the inner wall of the liner: the said diffuser walls bounding a diffusing flow path to the upstream liner wall and defining with the upstream liner wall an air space. the outer diffuser wall and outer casing wall defining a diffusing flow path to the outer liner wall. the outer diffuser wall blocking flow between the said air space and the outer liner wall. and the inner casing wall defining a diffusing flow path to the inner liner wall; the said diffusing flow paths taking air from the outer. central. and inner annular zones of the compressor outlet. respectively; and the inner diffuser wall defining circumferentially extending slot means for diverting a portion of the air from the path leading to the upstream liner wall into the path leading to the inner liner wall in response to relative pressure conditions in the said paths.

3. In a gas turbine engine comprising a compressor with an annular axial outlet. outer and inner casing walls defining an annular combustion space connected to the outlet. and an annular combustion liner disposed between the said casing walls. the casing walls having forward portions diverging to define a diffusing zone beginning at the compressor outlet. the combustion liner having an upstream wall and outer and inner walls with openings for entrance of air into the liner in each of the said liner walls. the improvement comprising an airflow dividing diffuser disposed in the diffusing zone and connected to the upstream end of the liner. the diffuser having a snout providing an annular entrance adjacent the compressor outlet and at a mid-radius of the outlet. an outer diffuser wall extending from the snout to the outer wall of the liner. and an inner diffuser wall extending from the snout to the inner wall of the liner; the said diffuser walls bounding a diffusing flow path to the upstream liner wall and defining with the upstream liner wall an air space. the outer diffuser wall and outer casing wall defining a diffusing flow path to the outer liner wall. the outer diffuser wall blocking flow between the said air space and the outer liner wall. and the inner casing wall defining a diffusing flow path to the inner liner wall; the said diffusing flow paths taking .air from the outer. central. and inner annular zones ofthe compressor outlet. respectively; the inner diffuser wall defining circumferentially extending slot.means for diverting a portion of the air from the path leading to the upstream liner wall into the path leading to the inner liner wall in response to relative pressure conditions in the said paths; and the diffusing flow path to the upstream liner wall being disposed to discharge the air into the said air space in a direction more predominantly toward the inner diffuser wall than toward the outer diffuser wall.

Patent Citations
Cited PatentFiling datePublication dateApplicantTitle
US3403510 *Nov 23, 1966Oct 1, 1968United Aircraft CorpRemovable and replaceable fuel nozzle holder assembly for an annular combustion burner
US3589127 *Feb 4, 1969Jun 29, 1971Gen ElectricCombustion apparatus
US3631675 *Sep 11, 1969Jan 4, 1972Gen ElectricCombustor primary air control
Referenced by
Citing PatentFiling datePublication dateApplicantTitle
US4098074 *Jun 1, 1976Jul 4, 1978United Technologies CorporationCombustor diffuser for turbine type power plant and construction thereof
US4458479 *Oct 13, 1981Jul 10, 1984General Motors CorporationDiffuser for gas turbine engine
US4870826 *Jun 8, 1988Oct 3, 1989Societe Nationale D'etude Et De Construction De Moteurs D'aviation (Snecma)Casing for a turbojet engine combustion chamber
US5077967 *Nov 9, 1990Jan 7, 1992General Electric CompanyProfile matched diffuser
US5134855 *Nov 8, 1990Aug 4, 1992Rolls-Royce PlcAir flow diffuser with path splitter to control fluid flow
US5279126 *Dec 18, 1992Jan 18, 1994United Technologies CorporationDiffuser-combustor
US5339622 *Aug 17, 1993Aug 23, 1994Societe Nationale D'etude Et De Construction De Moteurs D'aviation (S.N.E.C.M.A.)Gas turbine engine with improved water ingestion prevention
US5737915 *Feb 9, 1996Apr 14, 1998General Electric Co.Tri-passage diffuser for a gas turbine
US6401447Nov 8, 2000Jun 11, 2002Allison Advanced Development CompanyCombustor apparatus for a gas turbine engine
US7181914Jun 25, 2003Feb 27, 2007Rolls-Royce PlcDiffuser for gas turbine engine
CN100416062CNov 19, 2003Sep 3, 2008通用电气公司Combustion chamber inlet booster with blowing attached layer
EP1426688A1 *Nov 19, 2003Jun 9, 2004General Electric CompanyCombustor inlet diffuser with boundary layer blowing
Classifications
U.S. Classification60/751
International ClassificationF23R3/04, F23R3/02, F23R3/10
Cooperative ClassificationY02T50/675, F23R3/04
European ClassificationF23R3/04