US3581492A

US3581492A - Gas turbine combustor

Info

Publication number: US3581492A
Application number: US839994A
Authority: US
Inventors: Carl T Norgren; William H Roudebush; North Olmsted; Francis M Humenik
Original assignee: National Aeronautics and Space Administration NASA
Current assignee: National Aeronautics and Space Administration NASA
Priority date: 1969-07-08
Filing date: 1969-07-08
Publication date: 1971-06-01
Anticipated expiration: 1988-06-01

Abstract

Maintaining performance in a gas turbine combustor during periods of airflow distortion. A diffuser duct is between a liner forming the combustor and a housing surrounding the liner. Entry ports in the combustor liner have scoops which extend the full height of the diffuser duct to maintain the proper airflow distribution regardless of inlet airflow distortion.

Description

D United States Patent 1 1 3,581,492

[72] Inventors Carl T. Norgren; 56] References Cited William ROUdEbUSh, North Olmsted; UNITED STATES PATENTS [2'] App No $312 Humemk 2,841,958 7/1958 Stokes et al 60/3965 2,993,337 7/1961 Cheeseman 60/39.65 [22] PM 3 338 051 s 1967 Ch b 1 1 60/39 65 [45] Patented June L 1971 am er ameta [73] Assignee The United States of America as represented Primary Examiner-Carroll B. Don'ty, Jr.

by the Administrator of the National AtlarneysN. T. Musial, G. E. Shook and G. T McCoy Aeronautics and Space Administration [54] T EL q 'g ABSTRACT: Maintaining performance in a gas turbine comrawmg bustor during periods of airflow distortion; A difiuser duct is [52] U.S.Cl 60/3936, between a liner forming the combustor and a housing sur- 60/39.65,43l/352 rounding the liner. Entry ports in the combustor liner have [51] lnt.Cl F02c 3/06 scoops which extend the full height of the diffuser duct to [50] Field of Search 60/3936, maintain the proper airflow distribution regardless of inlet air- 39.65;43l/352 flow distortion.

SIDE ENTRY CO-MBUSTOR PATENTEU JUN 1 1971 SHEET 1 BF 3 INVENTORS CARL T. NORGREN WILLIAM H. ROUDEBUSH FRANCIS M. HUMENIK ATTORNEYS PATENIEDJUN 11971 3,581,492

' SHEET 2 OF 3 PRIOR ART COMBUSTOR FIG. 2

SIDE ENTRY COMBUSTOR FIG. 3

INVENTORS CARL T. NORGREN:

WILLIAM H. ROUDEBUSH BY FRANCIS M. HUMENIK ATTORNEYS PATENTED JUN H971 3581.492

SHEET 3 OF 3 FIG. 4

INVENTOR CARL T. NORGREN WILLIAM H. ROUDEBUSH BY FRANCIS M. HUMENIK QV cl 7 JQW 4? J ATTORNEYS GAS TURBINE CDMBUSTOR ORIGIN OF THE INVENTION The invention described herein was made by employees of the United States Government and may be manufactured and used by or for the Government for governmental purposes without the payment of any royalties thereon or therefor.

BACKGROUND OF THE INVENTION This invention is concerned with maintaining good performance in a gas turbine combustor during periods of airflow distortion. The invention is particularly directed to providing the proper airflow through each entry port even if the radial flow profile from the compressor is distorted.

Certain combustion problems have been encountered in gas turbine engines due to flow distortions caused by acceleration, deceleration, off design operation of the engine or local environmental conditions. The problems have become critical in advanced engines which operate near the maximum capability of the materials from which the engine is constructed.

SUMMARY OF THE INVENTION These problems have been solved in a gas turbine com bustor construction in accordance with the present invention wherein air from the compressor is ducted into a single annular diffuser that is adjacent the combustion chamber. The velocity profile across the duct reflects the velocity profile at the compressor exit.

Air is admitted into the combustion chamber through entry ports in a predetermined pattern. A scoop is attached to each entry port, and each scoop extends completely to the full height of the diffuser duct. In this manner the mass flow through the scoop and into the combustion chamber is always proportional to the mass flow from the compressor. Even if the radial flow profile from the compressor becomes inverted, the airflow into each port will remain the same. Thus the combustion chamber will contain the same fuel air distribution which assures that no hot temperature streaks will occur due to local decrease in airflow.

OBJECTS OF THE INVENTION It is, therefore, an object of the present invention to maintain good performance in a gas turbine combustor with no local overheating during periods of flow distortion caused by acceleration, deceleration, off-design operation, and local environmental conditions.

Another object of the invention is to accommodate flow distortions of air to a gas turbine combustor so that the engine can be operated near the maximum capability of the materials.

These and other objects of the invention will be apparent from the specification which follows and from the drawings wherein like numerals are used throughout to identify like parts.

DESCRIPTION OF THE DRAWINGS FIG. I. is an axial quarter section view of a typical gas turbine engine showing the relative positions of its various components;

FIG. 2 is an enlarged sectional view of a prior art combustor;

FIG. 3 is an enlarged sectional view of a combustor constructed in accordance with the present invention;

FIG. 4 is a perspective view of a combustor constructed in accordance with the invention;

FIG. 5 is a perspective view of a portion of a combustor illustrating an alternate embodiment of the invention; and

FIG. 6 is a plan view of still another embodiment of the present invention.

DESCRIPTION OF THE PREFERRED EMBODIMENT Referring now to the drawings there is shown in FIG. I a gas turbine engine 10 having a compressor 12 and a turbine M mounted on a spindle I6. A com bustor 18 enclosed by a housing I9 is positioned between the compressor I2 and the turbine 14. Air from the compressor I2 is heated in the combustor l8 and then directed through the turbine 14 in a manner well known in the art.

One form of combustor has a torus-shaped chamber 20 between the housing I9 and an inner casing 21 which encircles the spindle 16. Fuel for combustion is injected into the chamber 20 through a plurality of nozzles 22. Swirlers 24 may be used with the nozzles 22 for improved combustion.

The prior art combustor 18 is shown in greater detail in FIG. 2. Air from the compressor 12 flows through an inlet passage 26 toward a liner 27. This airflow is divided between an inner duct 28 and an outer duct 30. The duct 28 is in the form of an annulus between the inner casing 21 and the liner 27. The outer duct 30 comprises the annular space between the housing 19 and the liner 27.

A portion of the air in the inlet passage 26 is also directed to the swirlers 24 through snouts 31. The compressed air in the inner duct 28 is admitted into the combustion chamber 20 through

suitable ports

32 and 34. In like manner, compressed air in the outer duct 30 passes through

ports

36 and 38.

The arrows in the inlet passage 26 illustrate a distorted radial air flow profile. The arrows in the

ducts

28 and 30 show the resulting distortions of the radial air profiles in these ducts. As a result of these profile distortions the amount of air entering the combustion chamber 20 through the

ports

32, 34, 36, and 3% will not be the same as with a uniform inlet airflow profile. Consequently, combustion problems occur which may produce local overheating.

A side entry combustor constructed in accordance with the present invention in which air for the combustion process enters substantially from a single annulus is shown in FIG. 3. Here again compressed air from the compressor 12 flows through an inlet passage 26 toward a liner 27. This air then passes into an annular diffuser duct 40 that is adjacent to the combustion chamber 20 between the liner 27 and the housing I9. The radial air profile across the diffuser duct 40 shown by the arrows in FIG. 3 will reflect the radial air profile at the compressor exit in the passage 26.

Air from the diffuser duct 44) is admitted into the combustion chamber 20 in a predetermined entry pattern through primary entry ports 42 and secondary entry ports 44 in the liner 27. Fuel is supplied to the noules 22 from pipes 45. A portion of the air in the inlet passage 26 is directed to the swirlers 24 at the nozzles 22 through inlet diffusers 36.

An important feature of the invention is the provisions of scoops 48 at the primary entry ports 42. The scoops 48 extend to the full height of the diffuser duct 40 from the liner 27 to the housing I9. Each scoop 48 has an open front which faces toward the flowing air. A curved backwall 49 on each scoop 4% extends outward from the downstream edge of each primary entry port 42 the full height of the duct 40 to the housing 19. Vertical sidewalls SI extend from opposite edges of each primary port 42 to the curved backwall 49. If desired, the wall 49 may be a straight plate which extends diagonally from the downstream edge of the port 42 to the housing 19 opposite the upstream edge of the same primary port.

In this manner the mass flow through each scoop 48 and into the combustion chamber 20 is always proportional to the mass flow from the compressor. Even if the radial flow profile shown by the arrows in FIG. 3 from the compressor becomes inverted the airflow into each primary entry port 42 will remain the same. Thus the combustion chamber 20 will have the same fuel air distribution which, in turn, will assure that no hot temperature streaks will occur due to local decreases in airflow.

The operation of the side entry combustor shown in FIG. 3 is illustrated in greater detail in FIG. 4. The combustion chamber 20 encircles the spindle 16 which mounts the compressor at one end and the turbine at the other. Fuel passes through pipes 45 and is injected into the chamber 20 through nozzles as previously described.

Compressed air from the compressor enters the diffuser duct 40 as shown by the arrows A. A portion of this air is intercepted by the scoops 43 at the primary entry port 42. These scoops 48 direct this intercepted air into the combustion chamber 20. Still another portion of the air enters the combustion chamber through the secondary ports 44.

A small amount of air passes through the inlet diffusers 46 into a cooling duct 50 as shown by the arrow B. This cooling air passes through cooling apertures 52 into the combustion chamber 20. The remaining air from the diffusers passes to the swirlers 24 which surround the nozzles 22.

DESCRIPTION OF ALTERNATE EMBODIMENTS In the alternate embodiment shown in FIG. 5 scoops 54 are provided at each of the secondary ports 44. The scoops 54 likewise extend to the full height of the diffuser duct 40.

In the embodiment shown in FIGS. 4 and 5 the secondary ports 44 are staggered relative to the primary entry ports 42. The embodiment shown in FIG. 6 has the primary and secondary ports aligned. Each primary port is provided with a scoop so that is aligned with another scoop 58 on the secondary port. As seen in FIG. 6 the secondary scoop S8 is larger than the primary scoop 56 to obtain the proper proportioning of the air.

The embodiment shown in FIG. 6 also provides a set of secondary ports 60 which are staggered relative to the primary entry ports. This results in combinations of entry ports which may be sized to suit desired combustor performance. When scoops are attached to entry ports, the scoops extend to the full height of the diffuser duct.

While several embodiments of the invention have been shown and described it will be appreciated that various structural modifications may be made without departing from the spirit of the invention or the scope of the subjoined claims. For example, it is contemplated the secondary ports 44 may be staggered with respect to the primary ports 42 as shown in FIG. 5 yet have different size scoops as shown in FIG. 6. Whatever the arrangement all the scoops must extend to the full height of the diffuser duct to obtain at all times correct proportional amount of air which is to be used for combustion purposes. This amount of air is independent of the compressor outlet radial velocity profile.

We claim:

l. A combustor for a gas turbine engine of the type having a compressor and a turbine mounted on a spindle, said combustor comprising a curved liner forming a torus-shaped chamber encircling said spindle between said compressor and said turbine, said liner having a plurality of spaced ports therein, each of said ports having an upstream edge toward said compressor and a downstream edge away from said compressor;

means for injecting fuel into said torus-shaped chamber,

a housing surrounding said liner and spaced therefrom to form an annular diffuser duct, said duct being in communication with said compressor for receiving compressed air therefrom whereby compressed air is admitted to said torus-shaped chamber through said ports, and

a plurality of scoops in said diffuser duct adjacent some of said ports and substantially surrounding said lines for intercepting a portion of the flow of compressed air, each of said scoops having a pair of spaced sidewalls extending from said liner toward said housing, a backwall connecting the outermost edges of each of said pairs of sidewalls whereby the space between the sidewalls at adjacent ports is substantially unobstructed to the flow of compressed air, said backwall extending the full height of said diffuser duct from said downstream edge of said port to said housing substantially opposite the upstream edge of said port, and an open front between said sidewalls extending from said upstream edge of said port to said housinzg whereby the mass airflow through said open front and a acent port into said combustion chamber 18 proportional to the mass flow in said diffuser duct.

2. Apparatus as claimed in claim 1 wherein the ports are in at least two adjacent rows, one of said rows being upstream from the other,

primary entry ports being in said one row, and

secondary entry ports being in said other row.

3. Apparatus as claimed in claim 2 including scoops on the primary entry ports.

4. Apparatus as claimed in claim 2 including scoops on the secondary entry ports.

5. Apparatus as claimed in claim 2 wherein the primary entry ports and secondary entry ports are aligned in the direction of airflow.

6. Apparatus as claimed in claim 2 wherein the primary entry ports and the secondary entry ports are staggered in the direction of the airflow.

7. Apparatus as claimed in claim 1 wherein backwall portion of the scoop comprises a member curving outwardly from said downstream edge of the adjacent port to the housing.

8. Apparatus as claimed in claim 1 wherein the backwall portion of the scoop comprises a substantially straight member extending diagonally outward from said downstream edge of the adjacent port to a housing substantially opposite the upstream edge of said port.

Claims

1. A combustor for a gas turbine engine of the type having a compressor and a turbine mounted on a spindle, said combustor comprising a curved liner forming a torus-shaped chamber encircling said spindle between said compressor and said turbine, said liner having a plurality of spaced ports therein, each of said ports having an upstream edge toward said compressor and a downstream edge away from said compressor; means for injecting fuel into said torus-shaped chamber, a housing surrounding said liner and spaced therefrom to form an annular diffuser duct, said duct being in communication with said compressor for receiving compressed air therefrom whereby compressed air is admitted to said torus-shaped chamber through said ports, and a plurality of scoops in said diffuser duct adjacent some of said ports and substantially surrounding said lines for intercepting a portion of the flow of compressed air, each of said scoops having a pair of spaced sidewalls extending from said liner toward said housing, a backwall connecting the outermost edges of each of said pairs of sidewalls whereby the space between the sidewalls at adjacent ports is substantially unobstructed to the flow of compressed air, said backwall extending the full height of said diffuser duct from said downstream edge of said port to said housing substantially opposite the upstream edge of said port, and an open front between said sidewalls extending from said upstream edge of said port to said housing whereby the mass airflow through said open front and adjacent port into said combustion chamber is proportional to the mass flow in said diffuser duct.

2. Apparatus as claimed in claim 1 wherein the ports are in at least two adjacent rows, one of said rows being upstream from the other, primary entry ports being in said one row, and secondary entry ports being in said other row.

3. Apparatus as claimed in claim 2 including scoops on the primary entry ports.