Internally cooled turbine blade
An internally cooled turbine blade for use in a gas turbine engine includes a serpentine channel extending within the turbine blade for flowing cooling air through the interior of the blade. The turbine blade includes turbulators positioned within the serpentine channel to cause or increase turbulence of the cooling air flowing through the serpentine channel, increasing the heat transfer between the turbine blade and the cooling air before exiting a plurality of air outlets within the turbine blade.
1. An internally cooled turbine blade for use in a gas turbine engine, the internally cooled turbine blade comprising:
an airfoil with a concave sidewall and a convex sidewall extending spanwise between a platform and a blade tip, and chordwise between a leading edge and a trailing edge;
wherein internal surfaces of the airfoil define an internal cooling passage within the airfoil extending between a cooling air inlet and a plurality of air outlets, the internal cooling passage comprising:
a serpentine channel comprising a leading edge channel, an intermediate channel,
a trailing edge channel, and a plurality of turbulators;
wherein two adjacent rows of the plurality of turbulators are positioned within the trailing edge channel;
wherein:
the leading edge channel is positioned adjacent the leading edge of the airfoil;
the trailing edge channel is positioned adjacent the trailing edge of the airfoil; and
the intermediate channel is positioned between the leading edge channel and the trailing edge channel;
wherein the leading edge channel and the intermediate channel are separated by a first dividing wall that extends from a blade base toward the blade tip and wherein the first dividing wall includes a first radially outer end configured to turn cooling air flowing in the leading edge channel from a platform towards the blade tip into the intermediate channel such that the cooling air flows in the intermediate channel from the blade tip towards the platform, and
wherein the trailing edge channel and the intermediate channel are separated by a second dividing wall that extends from the platform toward the blade tip and wherein the second dividing wall includes a second radially inner end configured to direct cooling air flowing in the intermediate channel from the blade tip to the platform into the trailing edge channel such that the cooling air flows in the trailing edge channel from the platform towards the blade tip.
2. The internally cooled turbine blade of claim 1 , wherein a radially outer end of the second dividing wall is positioned closer to the blade tip than a radially outer end of the first dividing wall.
3. The internally cooled turbine blade of claim 1 , wherein a single row of the plurality of turbulators are positioned within the leading edge channel and the intermediate channel.
4. The internally cooled turbine blade of claim 1 , wherein the turbulators have a chevron shape, and wherein a point of each of the plurality of chevron shaped turbulators points in a direction opposite the flow direction of cooling air flowing through the internal cooling passage.
5. The internally cooled turbine blade of claim 1 , wherein:
a first row of the plurality of turbulators positioned within the trailing edge channel extends at least 75% a distance from the platform towards the blade tip; and
a second row of the plurality of turbulators positioned within the trailing edge channel extends at least 50% the distance from the platform towards the blade tip.
6. The internally cooled turbine blade of claim 1 , wherein the first dividing wall and the second dividing wall gradually decrease in thickness or width as each extends toward the blade tip, such that portions of the first dividing wall and the second dividing wall closer to the root are thicker or than portions of the first dividing wall and the second dividing wall closer to the blade tip.