Patent Application
20100236248
The present invention relates generally to gas turbine engines, and, more specifically to combustors therein. In a gas turbine engine, air is pressurized in a compressor and channeled to a combustor, mixed with fuel, and ignited for generating hot combustion gases which flow downstream through one or more turbine stages. In a turbofan engine, a high pressure turbine drives the compressor, and is followed in turn by a low pressure turbine which drives a fan disposed upstream of the compressor.
A typical combustor is annular and axisymmetrical about the longitudinal axial centerline axis of the engine, and includes a radially outer combustion liner and radially inner combustion liner joined at upstream ends thereof to a combustor dome. Mounted in the dome are a plurality of circumferentially spaced apart carburetors each including an air swirler and a center fuel injector. Fuel is mixed with the compressed air from the compressor and ignited for generating the hot combustion gases which flow downstream through the combustor and in turn through the high and low pressure turbines which extract energy therefrom.
A major portion of the compressor air is mixed with the fuel in the combustor for generating the combustion gases. Another portion of the compressor air is channeled externally or outboard of the combustor for use in cooling the combustion liners, while another portion is channeled radially through the combustion liner as a jet of dilution air, which both reduces the temperature of the combustion gases exiting the combustor and controls the circumferential and radial temperature profiles thereof for optimum performance of the turbines.
A combustor is typically cooled by establishing a cooling film of the compressor air in a substantially continuous boundary layer or air blanket along the inner or inboard surfaces of the combustion liners that confine the combustion gases therein. The film cooling layer provides an effective barrier between the metallic combustion liners and the hot combustion gases for protecting the liners against the heat thereof and ensuring a suitable useful life thereof.
In a typical combustor, the film cooling layer is formed in a plurality of axially spaced apart film cooling nuggets which are annular manifolds fed by a plurality of inlet holes, with a downstream extending annular lip which defines a continuous circumferential outlet slot for discharging the cooling air as a film along the hot side of the liners. The rows of nuggets ensure that the film is axially reenergized from row to row for maintaining a suitably thick boundary layer to protect the liners.
In a recent development in combustor design, a multihole film cooled combustor liner eliminates the conventional nuggets and instead uses a substantially uniform thickness, single sheet metal liner with a dense pattern of multiholes to effect film cooling. The individual multiholes are inclined through the liner at a preferred angle of about 20°, with an inlet on the outboard, cold surface of the liner, and an outlet on the inboard, hot surface of the liner spaced axially downstream from the inlet. The diameter of the multiholes is about 20-30 mils (0.51-0.76 mm). This effects a substantially large length to diameter ratio for the multiholes for providing internal convection cooling of the liner therearound. Most significantly, the small inclination angle allows the discharged cooling air to attach along the inboard surface of the liner to establish the cooling film layer which is fed by the multiple rows of the multiholes to achieve a maximum boundary layer thickness, which is reenergized and maintained from row to row in the aft or downstream direction along the combustor liners.
Combustor liner durability in the region of the primary mixing/cooling holes is a concern due to localized hot spots in the vicinity of the mixing holes, which can lead to liner cracking. The hot spots are mainly due to the disturbance to the hot gases by cold jets from the mixing holes leaving the high combustion air in contact with the liner wall. That is, hot combustion gases can be trapped behind cooling jets coming through the mixing holes, thereby causing a temperature increase in the liner near the mixing holes. Such hot spots can result in cracking or other damage to the liner due to thermal fatigue as well as high cycle fatigue (HCF) failures at high frequencies.
In an exemplary embodiment, a combustor liner for a gas turbine combustor includes a cooling hole formed in the liner and a stub secured in the cooling hole. The cooling hole delivers cooling air into a combustion zone of the combustor. The stub is structured to provide added stiffness to an inside edge of the cooling hole.
In another exemplary embodiment, a method of reducing cracking due to thermal fatigue adjacent cooling holes in a gas turbine combustor liner includes a step of securing a stub in the cooling hole, where the stub provides added stiffness to an inside edge of the cooling hole.
In yet another exemplary embodiment, a combustor liner for a gas turbine combustor includes a cooling hole formed in the liner that delivers cooling air into a combustion zone of the combustor. A stub is secured in the cooling hole and includes a plurality of cooling passages disposed substantially surrounding the cooling hole. The cooling passages are angled relative to an axis of the cooling hole in a direction corresponding to a hot gas flow direction through the liner.
With reference to
Still referring to
Hot gases from the combustion section in combustion liner 12 flow therefrom into section 16. There is a transition region indicated generally at 46 in
A problem may occur, however, in that hot combustion gases may be trapped behind cooling jets coming through the cooling holes 34. These hot spots can cause cracking due to thermal fatigue or possibly HCF failures at high frequencies. With reference to
Each stub 50 may include one or a plurality of cooling passages 52 disposed substantially surrounding the cooling hole 34. The cooling passages 52 are preferably oriented at an angle α relative to an axis (represented by arrow 54) of the cooling hole in a direction corresponding to a hot gas flow direction (represented by arrow 56) through the liner 12. That is, as shown in
The addition of stubs or stiffening members to the cooling holes in a combustion liner adds stiffness at the cooling hole edge to reduce cracking due to thermal fatigue. The additional stiffness also provides resistance against HCF failures at high frequencies. The angled cooling passages serve to push the hot gases away from the liner wall, thereby cooling the liner wall and the stub. As a result, durability of the liner can be improved.
While the invention has been described in connection with what is presently considered to be the most practical and preferred embodiments, it is to be understood that the invention is not to be limited to the disclosed embodiments, but on the contrary, is intended to cover various modifications and equivalent arrangements included within the spirit and scope of the appended claims.
