Patent Application
20100003142
This disclosure relates to a supplemental radial cooling passage for an airfoil.
Turbine blades are utilized in gas turbine engines. As known, a turbine blade typically includes a platform having a root on one side and an airfoil extending from the platform opposite the root. The root is secured to a turbine rotor. Cooling circuits are formed within the airfoil to circulate cooling fluid, such as air. Typically, multiple relatively large cooling channels extend radially from the root toward a tip of the airfoil. Air flows through the channels and cools the airfoil, which is relatively hot during operation of the gas turbine engine.
Some advanced cooling designs use one or more radial cooling passages that extend from the root toward the tip. Typically, the cooling passages are arranged between the cooling channels and an exterior surface of the airfoil. The cooling passages provide extremely high convective cooling.
The Assignee of the present disclosure has discovered that in some cooling designs the airfoil is overcooled at the base of the airfoil near the platform. It is believed that strong secondary flows, particularly on the suction side, force the migration of relatively cool fluid off the end wall and onto the suction side of the blade. This results in relatively low external gas temperatures. Internally, the coolant temperature is relatively cool as it has just entered the blade. The high heat transfer coefficients provided by the cooling passage in this region are undesirable as it causes overcooling of the external surface and premature heating of the coolant air.
Tapered radial cooling passages have been used. However, in one arrangement, the wall adjacent to the suction side exterior surface is tapered as it extends towards the tip. This configuration undesirably results in increased cooling near the platform as compared to near the tip due to the larger convection surface near the platform.
In another arrangement in which the cross-sectional area of the cooling passage remains relatively constant cooling fluid, Mach numbers also remain relatively constant resulting in uniform heat transfer rates within the passage. Coolant fluid entering the airfoil at low temperature and increases in temperature as it moves through the cooling passage. External three-dimensional flows and non-uniform gas temperature profiles cause temperatures and heat transfer rates to be typically lower near the inner and outer radii of the airfoil. This external heat load, combined with the cool coolant fluid near the inlet to the airfoil cause the external surface to be overcooled.
What is needed is a radial cooling passage that provides desired cooling of the airfoil.
A turbine engine airfoil is disclosed that includes an airfoil structure having an exterior surface and an end portion. A cooling passage extends a length radially within the structure in a direction toward the end portion. The cooling passage provides a convection surface along the length adjacent to the exterior surface. The convection surface includes a generally uniform width along the length. The cooling passage has generally decreasing cross-sectional areas along the length in the direction. The width and the cross-sectional areas are generally perpendicular to the length.
The cooling passage is provided by a core structure that extends from a first end to a second end along the length. The core structure includes a side having a generally uniform width along the length. The core structure includes a first thickness at the first end providing with the width a first area that is greater than a second area, which is provided by the width and a second thickness at the second end. Accordingly, a radial cooling passage provides desired cooling of the airfoil.
These and other features of the disclosure can be best understood from the following specification and drawings, the following of which is a brief description.
The turbine section 11 includes alternating rows of blades 20 and static airfoils or vanes 19. It should be understood that
An example blade 20 is shown in
The airfoil 34 includes an exterior surface 58 extending in a chord-wise direction C from a leading edge 38 to a trailing edge 40. The airfoil 34 extends between pressure and suction sides 42, 44 in a airfoil thickness direction T, which is generally perpendicular to the chord-wise direction C. The airfoil 34 extends from the platform 32 in a radial direction R to an end portion or tip 33. Cooling holes 48 are typically provided on the leading edge 38 and various other locations on the airfoil 34 (not shown).
Referring to
Current advanced cooling designs incorporate supplemental cooling passages arranged between the exterior surface 58 and one or more of the cooling channels 50, 52, 54. In the example disclosed, a radially extending cooling passage 56 is provided in a wall 60 between the exterior surface 58 and the cooling channels 50, 52, 54 at the suction side 44. First and second wall portions 68, 70 are provided on either side of the radial cooling passage 56 respectively adjacent to the exterior surface 58 and the cooling channel 52. However, it should be understood that the example cooling passages can be provided at other locations within the airfoil. For example, the disclosed cooling passage 56 can also be provided on the pressure side (shown) and leading edge (not shown).
As shown in
In one example, the cooling channels 50, 52, 54 are provided by ceramic cores during a casting process, as known. The radial cooling passages 56 are provided by a refractory metal core 74 (
Referring to
The reduction in the cross-sectional area increases the Mach number as the coolant moves to the end of the coolant passage. The increase in Mach number in turn allows the heat transfer coefficient near the exit of the passage to be higher than near the inlet. The heat transfer coefficients in the region of the blade 20 near the platform 32 is reduced. This allows the designer to maintain a uniform value (or adjust to the most desirable value) based upon the product of h*(ΔT) resulting in a uniformly cooled blade, where h is the convection heat transfer coefficient and ΔT is the temperature gradient.
Although example embodiments have been disclosed, a worker of ordinary skill in this art would recognize that certain modifications would come within the scope of the claims. For that reason, the following claims should be studied to determine their true scope and content.
