Patent Application
20170107830
This disclosure relates to gas turbine engines, and more particularly to thermal management of turbine components of gas turbine engines.
Gas turbines hot section components, in particular turbine vanes and blades in the turbine section of the gas turbine are configured for use within particular temperature ranges. Such components often rely on cooling airflow to maintain turbine components within this particular temperature range. For example, stationary turbine vanes often have internal passages for cooling airflow to flow through, and additionally may have openings in an outer surface of the vane for cooling airflow to exit the interior of the vane structure and form a cooling film of air over the outer surface to provide the necessary thermal conditioning. Other components of the turbine often also require such thermal conditioning to reduce thermal gradients that would otherwise be present in the structure and which are generally undesirable. Thus, ways to increase thermal conditioning capability in the turbine are desired.
In one embodiment, a blade assembly for a gas turbine engine includes a blade and a blade platform secured to the blade. The blade extends radially outwardly from the blade platform. The blade platform includes at least one platform airflow passage located therein. A gusset extends from the blade to the blade platform. The gusset includes a gusset airflow passage fluidly connected to the platform airflow passage to convey an airflow to the platform airflow passage.
Additionally or alternatively, in this or other embodiments a blade airflow passage is positioned at the blade, the gusset airflow passage connecting the blade airflow passage to the platform airflow passage.
Additionally or alternatively, in this or other embodiments the platform airflow passage is configured to convey the airflow toward a leading edge of the blade platform from a platform airflow passage entrance.
Additionally or alternatively, in this or other embodiments the gusset is located at about midchord of the blade.
Additionally or alternatively, in this or other embodiments the gusset is located at a pressure side of the blade.
Additionally or alternatively, in this or other embodiments the gusset airflow passage is formed integral with the platform airflow passage.
Additionally or alternatively, in this or other embodiments the gusset airflow passage is formed by casting.
In another embodiment, a rotor assembly for a gas turbine engine includes a rotor disc and a plurality of rotor blades extending radially outwardly from the rotor disc. Each rotor blade includes a blade and a blade platform secured to the blade. The blade extends radially outwardly from the blade platform. The blade platform includes at least one platform airflow passage located therein. A gusset extends from the blade to the blade platform. The gusset includes a gusset airflow passage fluidly connected to the platform airflow passage to convey an airflow to the platform airflow passage.
Additionally or alternatively, in this or other embodiments a blade airflow passage is located at the blade. The gusset airflow passage connects the blade airflow passage to the platform airflow passage.
Additionally or alternatively, in this or other embodiments the platform airflow passage is configured to convey the airflow toward a leading edge of the blade platform from a platform airflow passage entrance.
Additionally or alternatively, in this or other embodiments the gusset is located at about midchord of the blade.
Additionally or alternatively, in this or other embodiments the gusset is located at a pressure side of the blade.
Additionally or alternatively, in this or other embodiments the gusset airflow passage is formed integral with the platform airflow passage.
Additionally or alternatively, in this or other embodiments the rotor assembly is a turbine rotor assembly.
In yet another embodiment, a gas turbine engine includes a combustor and a plurality of gas turbine engine components located in fluid communication with the combustor. The gas turbine engine component includes an airfoil portion and a platform secured to the airfoil portion. The platform includes at least one platform airflow passage positioned therein. A gusset extends from the airfoil portion to the platform. The gusset includes a gusset airflow passage fluidly connected to the platform airflow passage to convey an airflow to the platform airflow passage.
Additionally or alternatively, in this or other embodiments an airflow passage is located at the airfoil portion, the gusset airflow passage connecting the airfoil portion airflow passage to the platform airflow passage.
Additionally or alternatively, in this or other embodiments the platform airflow passage is configured to convey the airflow toward a leading edge of the platform from a platform airflow passage entrance.
Additionally or alternatively, in this or other embodiments the gusset is located at about midchord of the airfoil portion.
Additionally or alternatively, in this or other embodiments the gusset is located at a pressure side of the airfoil portion.
Additionally or alternatively, in this or other embodiments the gusset airflow passage is formed integral with the platform airflow passage.
The subject matter which is regarded as the present disclosure is particularly pointed out and distinctly claimed in the claims at the conclusion of the specification. The foregoing and other features, and advantages of the present disclosure are apparent from the following detailed description taken in conjunction with the accompanying drawings in which:
The gas turbine engine 10 further comprises a turbine section 20 for extracting energy from the combustion gases. Fuel is injected into the combustor 18 of the gas turbine engine 10 for mixing with the compressed air from the compressor 16 and ignition of the resultant mixture. The fan 12, compressor 16, combustor 18, and turbine 20 are typically all concentric about a common central longitudinal axis of the gas turbine engine 10.
The gas turbine engine 10 may further comprise a low pressure compressor located upstream of a high pressure compressor and a high pressure turbine located upstream of a low pressure turbine. For example, the compressor 16 may be a multi-stage compressor 16 that has a low-pressure compressor and a high-pressure compressor and the turbine 20 may be a multistage turbine 20 that has a high-pressure turbine and a low-pressure turbine. In one embodiment, the low-pressure compressor is connected to the low-pressure turbine and the high pressure compressor is connected to the high-pressure turbine.
The turbine 20 includes one or more sets, or stages, of fixed turbine vanes 22 and turbine rotors 24, each turbine rotor 24 including a plurality of turbine blades 26 (shown in
Referring now to
Referring now to
Referring now to
While the present disclosure has been described in detail in connection with only a limited number of embodiments, it should be readily understood that the present disclosure is not limited to such disclosed embodiments. Rather, the present disclosure can be modified to incorporate any number of variations, alterations, substitutions or equivalent arrangements not heretofore described, but which are commensurate with the spirit and scope of the present disclosure. Additionally, while various embodiments of the present disclosure have been described, it is to be understood that aspects of the present disclosure may include only some of the described embodiments. Accordingly, the present disclosure is not to be seen as limited by the foregoing description, but is only limited by the scope of the appended claims.
