Patent Grant
7665307
This invention relates to a dual wall combustor for a gas turbine engine. More particularly, this invention relates to a dual wall combustor including a ceramic matrix composite shell that supports a liner assembly.
A combustor for a gas turbine engine includes an outer shell and an inner liner. The inner liner is directly exposed to combustion gases and defines a gas flow path. The inner liner is spaced apart from the outer shell to define an air-cooling passage for cooling and controlling the temperature of the inner liner. Both the inner liner and the outer shell are fabricated from a material capable of withstanding the extreme temperatures generated during the combustion process.
During operation, the inner liner is exposed to thermal gradients caused by the flow and swirl of the fuel air mixture as it is ignited to generate combustion gases. Such differences in temperature cause the thermal gradients within the inner liner. A design concern is providing an inner liner material and configuration that accommodates such gradients. As appreciated, not all materials that perform favorably at high temperatures can also withstand the thermal gradients and the strains produced by such differences in temperature. Disadvantageously, the stress and strains generated in the inner liner by the thermal gradients have complicated the use of many materials capable of withstanding the elevated temperatures produced during combustion.
One example material includes ceramic matrix composites. A ceramic matrix composite includes ceramic fibers interwoven into a sheet that is than impregnated with a material such as Silicon Carbide, Silicon-Nitride or other oxide components that are capable of withstanding elevated temperatures. As appreciated, higher temperatures within a combustor are favorable to provide a more efficient burning of fuel. However, the ceramic matrix composite does not respond favorably to thermal gradients and therefore has not been widely utilized in conventional combustors.
Accordingly, it is desirable to develop a combustor that utilizes the advantageous thermal properties of ceramic matrix materials within a combustor without compromising combustor strength and durability.
An example combustor for a gas turbine engine according to this invention includes an outer shell made of a ceramic matrix composite that supports a plurality of inner heat shields made of a material other than the ceramic matrix composite.
The combustor liner assembly of this invention includes an outer shell made from a ceramic matrix composite. The ceramic matrix composite is a thermally desirable material and provides the requisite thermal insulation between the combustor chamber and other elements within the gas turbine engine. Supported within the outer shell is a plurality of heat shields that are constructed of a material other than the ceramic matrix composite.
The ceramic matrix composite of the outer shell performs optimally at a substantially stable and uniform temperature. However, the ceramic matrix composite does not perform as desired or provide the desired durability when exposed to substantial thermal gradients such as are experienced within a combustor chamber. Therefore, the inner heat shields are fabricated from a material that provides favorable thermal mechanical properties compatible with the thermal gradients generated within a combustor chamber.
The inner heat shield is supported within the outer shell by a plurality of fasteners. The fasteners provide a mechanical coupling between the plurality of heat shields and the outer shell while also providing a thermal de-coupling between the inner heat shields and outer shell. The thermal de-coupling inhibits thermal transfer between the inner heat shields and the outer shell.
A cooling air passage is defined between the plurality of inner heat shields and the outer shell to provide cooling air along the inner heat shields. Cooling air is provided as impingement flow against a cold side of each of the heat shields and also maybe communicated to the hot side surface of the inner heat shields through the plurality of cooling holes.
Accordingly, the combustor liner assembly of this invention provides a structure that utilizes the favorable properties of a ceramic matrix composite material in portions of a combustor that are exposed to substantially uniform temperatures while also accommodating the thermal gradients present within a combustor liner assembly.
These and other features of the present invention can be best understood from the following specification and drawings, the following of which is a brief description.
Referring to
The outer shell 14 is shown in an annular configuration about an axis 19 of the turbine engine 10. The liner assembly 12 includes an outer radial wall 34 and an inner radial wall 32. The outer shell 14 also includes a cowling 30 that is disposed forward of a forward end segment 36. The cowling 30 directs airflow around the combustor 1. The forward end segment 36 provides for the securement of a heat shield 16 on a forward end of the combustor 11. As should be appreciated, the gas turbine engine 10 illustrated in
Referring to
The plurality of heat shields 16 are fastened by way of fasteners 26 to the outer shell 14. The outer shell 14 includes a plurality of openings 25 that correspond to fasteners 26. The outer shell 14 is made of a ceramic matrix composite that provides desirable thermal properties. The ceramic matrix composite may be of any composition known to a worker skilled in the art. For example, the ceramic matrix composite may include a silicon-based composition including silicon carbide, silicon nitride or oxide-based ceramic materials. A worker skilled in the art would understand the composition of the ceramic matrix material favorable for application specific requirements.
The ceramic matrix composite material provides desirable thermal properties, but is not desirable in applications and environments that encounter thermal loading caused by thermal gradients as are present within a combustor. However, although the outer shell 14 of this invention encounters high temperatures, the heating is relatively even such that high amounts of thermal loading are not placed on the ceramic matrix composite material.
The heat shields 16 are supported by the ceramic matrix composite outer shell 14 and are made of material possessing favorable thermal mechanical properties compatible with the high thermal gradients encountered within the combustor assembly 11. The inner heat shields 16 are constructed of a refractory alloy or other advanced alloy composition that is compatible with the ceramic matrix composite of the outer shell 14. A worker skilled in the art would understand and know what materials are chemically and thermally compatible for use with the specific ceramic matrix composite and that also provide the desired thermal mechanical properties.
A plurality of fasteners 26 is utilized to secure the heat shields 16 within the outer shell 14. The fasteners 26 may be separate elements or may be integrally formed with the inner heat shields 16. The configuration of the combustor liner assembly 12 is shown with a convergent portion extending from the forward end segment 36 towards an aft open end 35. The specific shape of the combustor liner assembly 12 is application specific and other configurations and orientations of the combustor liner assembly 12 are within the contemplation of this invention.
Referring to
Each of the fasteners 26 includes a corresponding threaded member 28. The fasteners 26 extend through openings 25 within the outer shell 14 and are secured by the threaded member 28. The fastener 26 shown in
The inner heat shields 16 comprise a plurality of panels that are fit and mounted to the inner surface of the outer shell 14. The inner heat shields 16 are supported within the outer shell 14 and are spaced apart from the outer shell by the tab 24. As appreciated, although a tab 24 is shown other spacers as are understood and within one skilled in the art maybe utilized to define a space between the inner heat shield 16 and the outer shell 14.
Referring to
Referring to
Referring to
A combustor liner assembly 12 according to this invention utilizes the favorable thermal properties of a ceramic matrix composite without exposure to thermal gradients. Attachment of the heat shields 16 to the outer shell 14 through openings in the ceramic matrix composite provides a durable and desirable combination that utilizes thermally and mechanically desirable materials.
The foregoing description is exemplary and not just a material specification. Although a preferred embodiment of this invention has been disclosed, a worker of ordinary skill in this art would recognize that certain modifications would come within the scope of this invention. For that reason, the following claims should be studied to determine the true scope and content of this invention.
This invention was made with Government support under Contract Nos. F33615-98-C-2907 and F33615-01-C-2183 awarded by the United States Air Force. The Government has certain rights in this invention.
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| Number | Date | Country | |
|---|---|---|---|
| 20070144178 A1 | Jun 2007 | US |
