BACKGROUND OF THE INVENTION 1. Field of the Invention This invention pertains to gas turbine engines and pertains more particularly to a low cost, improved combustor liner and cooling therefor.

2. Description of the Prior Art The combustor of a gas turbine typically includes a combustor liner which encloses the hot combustion zone. Cooling of the combustor liner is critical for engine operation. The combustor liner is typically cooled either by film cooling or effusion cooling. Such cooling techniques are inherently based upon penetration of the cooling air through the combustor wall. Certain applications however are not amenable to effusion cooling or film cooling except by construction of an expensive combustor liner.

SUMMARY OF THE INVENTION It is an important object of the present invention to provide an improved combustor liner which can be inexpensively manufactured. More specifically, it is an object of the present invention to provide a cast combustor liner, and to provide an inexpensive manner of convectively cooling the cast combustor liner.

More particularly, the present invention contemplates a cast combustor liner which has an exterior surface exposed to cooling airflow received from the compressor section of the gas turbine engine for convective cooling of the liner, along with a plurality of turbulators integrally cast upon the liner's exterior surface for increasing the heat transfer coefficient of the combustor liner.

These and other objects and advantages of the present invention are specifically set forth in or will become apparent from the following detailed description of the invention when read in conjunction with the accompanying drawings.

BRIEF DESCRIPTION OF THE DRAWINGS In the drawings: Fig. 1 is a partial elevational cross-sectional view of a portion of a gas turbine engine utilizing the present invention : Fig. 2 is a perspective of the combustor casing and associated assembly; Fig. 3 is an enlarged elevational cross-sectional view of the combustor liner; Fig. 4 is a plan cross-sectional view of a portion of the combustor; Fig. 5 is an elevation of the exterior surface of the combustor liner; and Fig. 6 is an enlarged cross-section of the combustor liner showing details of the turbulators.

DETAILED DESCRIPTION OF THE INVENTION Referring now more particularly to the drawings. a gas turbine engine 10 conventionally includes a combustor section 12 driven by a turbine section 14 through axially extending shaft 16. Highly compressed air 17 is delivered from combustor section 12 to a combustor generally denoted by the numeral 18, and hot gases exit the compressor 18 to drive the turbine section 14.

The particular combustor 18 illustrated is of a type wherein airflow to the combustor zone of the combustor is modulated or adjusted for pollution control purposes. This type of arrangement is illustrated and described in detail in U. S.

Patents and 5,481,866 of R. J. Mowill, as well as copending U. S. Patent Application Serial No. 08/66,393 filed November 7,1997 entitled"Combustion Dilution Bypass System"of James Lenertz, et al, and commonly assigned herewith.

To the extent necessary for a complete understanding of the present invention, these patents and application are incorporated herein by reference.

Briefly, this type of combustor 18 includes a combustor liner 20. described in greater detail below, of annular configuration having toroidal cross-section.

Combustor liner 20 encloses the internal combustor zone 22 of the combustor. A fuel/air mixture is admitted into the combustor zone 22 through a pair of venturi's 24, one of which is shown in Figs. 2 and 4. Compressed air 17 from the compressor section 12 is routed inside combustor casing 19 and about the exterior of combustor liner 20, then through a duct 26 to be controlled by a three-way valve 28. A portion of the air is bypassed through a combustor bypass 30, while the remaining compressed air flow is directed through duct 32 into the venturi 24. Fuel is admitted through fuel nozzle 34, and intimate mixing thereof occurs within the venturi 24 before subsequent

introduction into the combustion zone 22 through opening 25. It will be noted that a three-way valve and associated ducting is included for the other venturi, not shown.

Combustion occurs within zone 22 and passes through the combustor exit 36 to be delivered to the power turbine section 14.

The bypass air going through bypass duct 30 is collected in a chamber 38 and then passes through a plurality of bypass tubes 40 for reintroduction into the main combustor gas flow adjacent the exit 36, as described in greater detail below.

The improved combustor liner 20 of the present invention is illustrated in greater detail in Fig. 3,5 and 6. Importantly, combustor 20 is of low cost construction inasmuch as it is a cast structure. Preferably, it is a multiple part casting. In the embodiment illustrated, liner 20 is a two-part casting which mating part line lips 42, 44. The casting parts are, of course, rigidly and sealingly intersecured to one another.

The relatively complex geometry of the annularly shaped, irregularly toroidal cross- section of combustor liner 20 may readily be produced via casting processes.

Integrally cast on the exterior surface of the combustor liner 20 are a plurality of upstanding ribs or turbulators 46. In the arrangement illustrated the turbulators are disposed approximately perpendicular to the central axis 48 of the gas turbine engine such that the airflow from the compressor section 17 crosses the pattern of the turbulators 46. If there is no swirl in the airflow 17 exiting the compressor section, the turbulators will be at about 90 degrees to the airflow direction. Preferably the turbulators are arranged between 45 degrees and 90 degrees to the direction of airflow 17. In the preferred pattern illustrated, each of the turbulators 46 extend in a circle about the exterior surface of toroidally shaped combustor, thus presenting a pattern of

circular ringlets. This pattern is readily amenable to casting processes with simple tooling.

Turbulators 46 are sized and configured to enhance the heat transfer coefficient of combustor liner 20 by periodically tripping or interrupting the boundary layer of the airflow 17 from the compressor section. Adjacent turbulators are spaced from one another a sufficient distance to permit reattachment and regrowth of the boundary layer before encountering the next turbulator 46. This periodic reattachment and regrowth of the boundary liner results in a periodic cooling air heat transfer coefficient increase. The net effect is that the average heat transfer coefficient is increased on the exterior surface of combustor liner 20.

The turbulators 46 are of quite small height, however, just enough to cause tripping of the boundary layer but without introducing any significant pressure drop to the compressed airflow 17 flowing there across. That is, the turbulators 46 are not included for the purpose of increasing the cooling surface area of the exterior of liner 20, as such would introduce a pressure drop thereacross. Rather, instead of increasing the cooling surface area, the turbulators increase the cooling heat transfer coefficient through the periodic tripping, reattachment and regrowth of the boundary layer.

In a preferred arrangement the ratio of the width, W, of the turbulators 46 to their height, H, is between about approximately 1 to 2. While the cast combustor liner 20 is of a very thin wall construction with a thickness, T, the height, H, of the turbulators is significantly smaller than this thickness, T. The distance, D, between adjacent turbulators is also carefully controlled to permit the reattachment and regrowth of the tripped boundary layer in the space between adjacent turbulators 46.

Preferably, the distance, D, is between about 5 and 20 times the width, W. An

example of a size for the turbulators which has been found to be highly useful in increasing the heat transfer coefficient is a turbulator where W equals about 0.07 inches, H equals about 0.07 inches, and D equals about 0.7 inches. Preferably, the outer corners of the turbulators are relatively sharp. For casting cost reduction purposes, the outer corners can be slightly without impairing the effectiveness of operation of the turbulators.

Adjacent the combustor exit 36, the combustor liner 20 includes a plurality of dilution orifices 50 drilled therethrough. The bypass airflow from bypass tubes 40 passes through the dilution orifices 50 for reintroduction into the gas flow proceeding to the turbine section of the engine. The dilution orifices 50 serve to provide effusion cooling for the combustor liner 20 in the narrow-necked exit area 36.

It will be apparent that various other patterns for the ribbed turbulators may be utilized effectively. For example, the turbulators may be arranged in a spiral like pattern across the toroidally shaped exterior surface, swirling around the liner exterior in a screw thread fashion. Further, in certain instances the ribs may be intermittent in length rather than continuous.

Various modifications and alterations to the foregoing detailed description of the preferred arrangement of the invention will be apparent to those skilled in the art.

Accordingly, the foregoing should be considered as exemplary and not as limiting to the scope and spirit of the invention as set forth in the appended claims.

Having described the invention with sufficient clarity that those skilled in the art may make and use it, what is claimed is: