The present invention generally relates to a combustion chamber and more specifically, to a can combustion chamber having a uniform effusion cooling method built therein.
Referring to
This classical method for cooling combustion chamber 10 is adequate only for low cycle and low performance engines. Such a method may not be effective in terms of combustion life, cooling efficiency, elimination of carbon build up, maintaining a lower and more uniform temperature, and reducing manufacturing complexity for higher performance engines, such as those used in various Joint Strike Fighter (JSF) aircraft.
Due to non-uniform cooling from conventional louvers 18, 18a, the temperature distribution of the dome 16 and of the chamber walls 14 vary, causing thermal stress and therefore reducing the life of the part. Also, with extended operation, the louvers 18a can deteriorate due to carbon formation and variations in the temperature of the dome 16.
U.S. Pat. No. 6,427,446 discloses an effusion cooling method for the liner walls of a can-annular combustor that has a premixing chamber. The method disclosed in the '446 patent uses a dome plate containing multiple rows of angled film cooling holes, angled in a tangential direction, cold side to hot side, to impart swirl into the airflow. The dome serves as a regulator for controlling the amount of air entering the combustor. This conventional system requires a premixing chamber, into which the film cooling holes may be cut, in order to impart swirl into the airflow before it enters the combustion chamber. This cooling method does not address either cooling the dome 16 or reducing carbon formation.
U.S. Pat. No. 6,546,731 discloses an effusion cooling method using double walls over the entire length of the combustion liner. The outer wall has a plurality of holes therethrough to provide normal impingement to the inner wall to provide cooling through convection. The inner wall also has effusion holes, whereby air can effuse into the combustion chamber. This method requires a double wall system, which may add manufacturing complexity and weight to the engine design.
Accordingly, there is a need for an improved combustion chamber utilizing an improved uniform effusion cooling method, whereby cooling efficiency is maximized while eliminating carbon build up, maintaining a lower and more uniform temperature, and reducing manufacturing complexity of high performance engines.
In one aspect of the present invention, a dome of a combustion chamber of an engine comprises a dome wall having a plurality of effusion holes passing through the dome wall, the effusion holes being uniformly spaced on the surface of the dome, wherein the effusion holes have a density of from about 10 to about 100 holes per square inch of the surface of the dome.
In another aspect of the present invention, a dome for a combustion chamber of an engine comprises a dome wall having a plurality of effusion holes passing through the dome, the effusion holes being uniformly spaced on a surface of the dome with a hole density of from about 10 to about 100 holes per square inch, wherein the effusion holes have a diameter from about 0.010 to about 0.040 inch; a center axis of each the plurality of effusion holes forms a first angle with a tangent to the surface of the dome of from about 7 to about 90 degrees; a centerline of the combustion chamber forms a second angle with the center axis of each of the plurality of effusion holes of from about 0 to about 180 degrees; and a ratio of the length of the combustion chamber to the diameter of the dome is greater than or equal to about 2.
In yet another aspect of the present invention, a combustion chamber for an engine comprises a dome; a can combustion liner having a first end attached to a scroll assembly, and a second end covered by the dome; and a plurality of effusion holes passing through the dome, the effusion holes being uniformly spaced on a surface of the dome with a density of from about 10 to about 100 holes per square inch of the surface of the dome.
In a further aspect of the present invention, a combustion chamber for an aircraft engine comprises a dome; a can combustion liner having a first end attached to a scroll and a second end covered by the dome; and a dome wall having a plurality of effusion holes passing through the dome wall, the effusion holes being uniformly spaced on the dome with a density from about 10 to about 100 holes per square inch of the surface of the dome, wherein the each of the plurality of effusion holes has a diameter from about 0.010 to about 0.040 inch; a center axis of each the plurality of effusion holes forms a first angle, θ, with the surface of the chamber dome of from about 7 to about 90 degrees; a centerline of the combustion chamber forms a second angle, β, with the center axis of each of the plurality of effusion holes of from about 0 to about 180 degrees; and a ratio of the length of the combustion chamber to the diameter of the dome is greater than or equal to about 2.
In another aspect of the present invention, a high performance gas turbine engine comprises a combustion chamber; a dome attached to a first end of the combustion chamber; a scroll assembly attached to a second end of the combustion chamber; and a plurality of effusion holes passing through the dome, the effusion holes being uniformly spaced about the dome with a hole density from about 10 to about 100 holes per square inch, wherein each of the effusion holes has a diameter from about 0.010 to about 0.040 inch; a center axis of each the plurality of effusion holes forms a first angle with the surface of the chamber dome of from about 7 to about 90 degrees; a centerline of the combustion chamber forms a second angle with the center axis of each of the plurality of effusion holes of from about 0 to about 180 degrees; and a ratio of the length of the combustion chamber to the diameter of the dome is greater than or equal to about 2.
In still a further aspect of the present invention, a method for uniformly cooling a dome of a combustion chamber of an engine, comprises a) providing the dome, the dome including a dome wall having a plurality of effusion holes therethrough, the effusion holes being uniformly spaced on a surface of the dome with a hole density from about 10 to about 100 holes per square inch; and b) passing an airflow through the effusion holes into the combustion chamber during operation of the engine.
These and other features, aspects and advantages of the present invention will become better understood with reference to the following drawings, description and claims.
The following detailed description is of the best currently contemplated modes of carrying out the invention. The description is not to be taken in a limiting sense, but is made merely for the purpose of illustrating the general principles of the invention, since the scope of the invention is best defined by the appended claims.
Broadly, the combustion chamber of the present invention may be used in any number of applications where conventional can combustors may be used. These applications include gas turbine engines for aircraft and land-based vehicles, as well as in engines used in generator equipment. The present invention provides a dome for a can combustor having a plurality of effusion holes therein to provide efficient cooling while preventing carbon formation.
In contrast to the present invention, conventional dome cooling designs of the prior art, using dome louvers, for example, may become corroded and/or may allow for ingestion of carbon particles that may build up and eventually separate from the dome. Furthermore, the dome cooling design of the present invention allows for the use of a low profile dome as compared with conventional domes, thereby maximizing liner volume in the constrained combustion envelope while reducing combustor case weight. Additionally, the dome effusion cooling design of the present invention requires the use of minimal or no thermal barrier coating, as compared to conventional designs, in order to minimize thermal variation within the dome and between the dome and the combustor wall.
Referring to
The design of the combustion chamber 34 of the present invention may be used to maximize the volume of the combustion chamber 34 within a limited installation envelope, as may be found on many modern aircraft. Combustion chamber 34 may have a first end 33 attached to the scroll assembly 36 and a second end 35 attached to the dome 32. The length L of the combustion chamber 34 may be defined as the distance from the first end 33 to the second end 35. In one embodiment of the present invention, for example, with reference to
Referring now to
Effusion holes 40 may be created in dome 32 by using various processes, such as a laser drilling operation. Effusion holes 40 allow a cooling air flow to enter the combustion chamber 34. The density of the effusion holes 40 and the size of the effusion holes 40 may vary, for example, according to the operating temperatures of engine 30 and the amount of cooling that is needed, for example, to maintain a particular operating temperature. Typically, there may be from about 10 to about 100 effusion holes 40 per square inch of surface area of dome 32. An exemplary density of effusion holes may be from about 10 to about 100 effusion holes 40 per square inch of surface area of dome 32. Typically, the effusion holes 40 are uniformly spaced on dome 32, as shown in
Referring specifically to
Referring now to
The effusion cooling design of the dome 32 according to the present invention resulted in a considerable improvement in that it provided a lower surface temperature and a more uniform surface temperature distribution. The dome 32 of the present invention gave an average surface temperature of about 1100° F., with a temperature variation of about ±100° F. On the other hand, the conventional dome 16, when used in the same operating environment, gave a non-uniform temperature distribution varying from 540° F. to 1300° F.
Improving the temperature distribution may allow the combustion system to meet a life requirement of at least 10,000 cycles. Furthermore, the design of the present invention reduces the risk of potential carbon formation by eliminating the louvers and flow steps caused by the louvers. To this end, the present invention therefore reduces the risk of carbon deposition on the louvers that may separate and be ingested downstream, thus enhancing the system reliability and durability. The design of the present invention also may reduce manufacturing cost and weight by eliminating the louvers in the combustor dome. The uniform cooling results achieved with the design of the present invention may also reduce, or eliminate, the need for a thermal barrier coating on dome 32, whereas in conventional systems a thermal barrier coating may be required. Further, the design of the present invention allows dome 32 to have a relatively low profile, as compared with conventional dome design, thereby allowing for installation of an engine comprising dome 32 in smaller installation envelopes, for example, in aircraft.
The present invention also relates to a method for uniformly cooling the dome 32 of the combustion chamber 34 of a gas turbine engine, e.g., engine 30. The method may include a step of providing the dome 32, which may include a dome wall 32a having a plurality of effusion holes 40 therethrough. The effusion holes 40 formed by the cutting step may be uniformly spaced on the surface of the dome 32. The effusion holes 40 may be formed by a process such as laser drilling. The effusion holes 40 may have a size, shape, and orientation with respect to the dome surface as described hereinabove. Typically, the effusion holes 40 are cut having a density from about 10 to about 100 holes per square inch of the surface of dome 32. During operation of the engine, air may be passed through the effusion holes 40 into the combustion chamber 34, thereby providing uniform cooling of the dome 32.
It should be understood, of course, that the foregoing relates to exemplary embodiments of the invention and that modifications may be made without departing from the spirit and scope of the invention as set forth in the following claims.
This invention was made with Government support under contract number N00019-02-C-3002, awarded by the U.S. Navy. The Government has certain rights in this invention.