Turbine airfoil with near wall cooling

None.

None.

1. Field of the Invention

The present invention relates generally to a gas turbine engine, and more specifically to an air cooled turbine airfoil with relatively low cooling flow volume.

2. Description of the Related Art including information disclosed under 37 CFR 1.97 and 1.98

In a gas turbine engine, airfoils such as stator vanes (sometimes referred to as guide nozzles) and rotor blades are exposed to extremely high temperature gas flows in order to produce a high efficient conversion of the combustion gas. In order to allow for higher temperature, the airfoils exposed to the higher temperatures require internal and external cooling. Elaborate internal cooling circuits have been designed to produce convection cooling for the airfoils, and film cooling holes are used to provide a layer of film cooling air over certain external surfaces of the airfoil to limit exposure to the hot gas flow. An airfoil with many film cooling holes requires a larger volume of cooling air flow because much of the cooling air is discharged out from the airfoil before passing through much of the airfoil internal cooling circuit.

In future industrial gas turbine engines, large airfoils are proposed in which the latter stages will also require some sort of internal cooling in order to increase the useful life of the airfoils. In recent engines, typically only the first and second stages of the turbine require internal cooling while the last two stages (third and fourth) are not cooled at all.

It is an object of the present invention to provide for a turbine airfoil with a high level of cooling while using a low amount of cooling air.

It is another object of the present invention to provide for a turbine airfoil with near wall cooling.

It is another object of the present invention to provide for a turbine airfoil with an internal cooling circuit that can be tailored to the external airfoil heat load on individual sections of the airfoil.

A turbine airfoil, such as a stator vane or a rotor blade, used in a gas turbine engine, the airfoil having a main support spar extending from the root and platform section that has a general airfoil cross sectional shape and forms a main cooling air supply chamber near the leading edge and a series of alternating cooling air impingement chambers and cooling air collecting chambers extending along the airfoil to the trailing edge region. A thin thermal skin is bonded to the outside surface of the spar to form the airfoil outer surface and to define a plurality of impingement chambers formed between the spar and the thermal skin. Cooling air supplied to the cooling air supply chamber flows through impingement holes into a series of impinge cut chambers spaced along the leading edge region and into a cooling air collector chamber, and then through impingement holes into an impingement chamber on the suction side wall. This series of impingement cooling followed by collector chamber cooling is repeated in an alternating process from the pressure side wall to the suction side wall until the cooling air is discharged into a trailing edge region collector chamber, where the cooling air is then discharged out through trailing edge exit holes. In another embodiment, one or more rows of film cooling holes can be connected to selected ones of the collector chambers to discharge some of the cooling air through film holes to provide a layer of film cooling air onto the external airfoil surface.

FIG. 1 shows a cross section view of the near wall cooling circuit of the present invention.

FIG. 2 shows a detailed view of the thin thermal skin with micro pin fins on the inner surface used in the airfoil of the present invention.

FIG. 3 shows a isometric view of the inner and outer surfaces of the thermal skin in the leading edge region of the airfoil of FIG. 1 .

FIG. 4 shows the backside surface of the thermal skin with the pattern of micro pin fins.

FIG. 5 shows an isometric view of a turbine blade with an opening to show several of the collector chambers with impingement holes arranged along the passage.

FIG. 6 shows a second embodiment of the near wall cooling circuit of the present invention with film cooling holes.

The low flow near wall cooling circuit of the present invention is disclosed for use in a turbine blade. However, the cooling circuit of the present invention can also be used in a stator vane. FIG. 1 shows a cross section view of the airfoil with the low flow near wall cooling circuit that include a main spar 11 that provides support for the blade and extends from the root. The main spar and root can be cast as a single piece or the spar can be secured to the root to form an integral piece. The main spar is formed of a number of ribs extending from the pressure side to the suction side of the airfoil to form collector chamber 17 , and a number of slanted outer walls that define pressure side impingement chambers 16 and suction side impingement chambers 18 . The main spar 11 also forms a cooling air supply chamber 12 and four impingement chambers 13 - 15 that extend along the leading edge of the airfoil.

A thin thermal skin 22 is bonded to the outer surface of the spar to form the outer airfoil surface. The thermal skin can be bonded to the spar using a transient liquid phase (TLP) bonding process. A thermal barrier coating (TBC) 23 is applied over the thermal skin 22 . FIG. 2 shows a detailed view of the thermal skin 22 and TBC 23 with a number of micro pin fins 24 fanned along the inner surface of the thermal skin 22 . The thermal skin 22 and the spar 11 form the leading edge impingement chambers 13 - 15 and the pressure side and suction side impingement chambers 16 and 18 that extend along the airfoil walls as seen in FIG. 1 . The thin thermal skin includes micro pin fins arranged along the inner surface as shown in FIGS. 3 and 4 to enhance the impingement cooling of the cooling air discharged from the impingement holes 18 in the spar 11 . The thin thermal skin has a thickness in the range of 0.010 inches to 0.030 inches and the micro pin fins have a diameter and a height of about the same order as the thermal skin thickness. The density of the micro pin fins 24 can be in the range of 50 to 75 percent.

The collector chambers 17 are formed within the spar 11 and extend from the leading edge to the trailing edge of the airfoil. Impingement holes 21 formed in the main spar 11 connect the collector chambers with the impingement chambers 13 - 15 and 16 and 18 . As seen in FIG. 1 , the last leading edge impingement chamber 15 is connected to a pressure side impingement chamber 16 and the collector chamber 17 through impingement holes 21 , which connects to a suction side impingement chamber 18 through impingement holes 21 , and then into an adjacent collector chamber 17 through impingement holes 21 . Spent cooling air from the impingement chambers 18 flow into the collection chambers 17 through return holes 20 . This alternating repetition of collector chamber 17 , then pressure side impingement chamber 16 , then collector chamber 17 , then suction side impingement chamber 18 repeats along the airfoil until the last collector chamber 25 position adjacent to the trailing edge of the airfoil. The last or trailing edge collector chamber 25 is connected to a row of trailing edge exit cooling holes that discharge the spent cooling air from the airfoil. Cooling holes that discharge cooling air for impingement against the airfoil wall are impingement holes 21 while the cooling holes that return spent cooling air into the collector chambers are return holes 20 .

The pressure and suction side collector chambers 16 and 18 and leading edge collector chambers 13 - 15 extend along the airfoil from the root to the tip as seen in the opening of the blade in FIG. 5 . The spar includes a row of impingement holes 21 and return holes 20 extending along the spar wall to provide impingement cooling along the length of the thermal skin and return holes for the spent impingement air that forms the particular collector chamber.

The operation of the low flow near wall cooling circuit in FIG. 1 is described below. Pressurized cooling air is supplied to the supply chamber 12 and flows into the first leading edge impingement chamber 13 through a row of impingement holes 21 to provide impingement cooling on the backside surface of the thermal skin 22 in that chamber. The cooling air then flows into the second leading edge impingement chamber 14 through another row of impingement holes 21 to provide impingement cooling to the backside surface of the thermal skin. This repeats through the third impingement chamber 15 to provide impingement cooling of the remaining section of the leading edge. This cooling air then flows into the first pressure side impingement chamber 16 and then into the first collector chamber 17 adjacent to the cooling air supply chamber 12 through a row of return holes 20 .

From the first collector chamber 17 , the cooling air flows into the first suction side impingement chamber 18 through a row of impingement holes 21 to provide impingement cooling to the backside surface of the thermal skin. From the first suction side impingement chamber 18 , the cooling air flows into the next collector chamber 17 through the return holes 20 and then into the second pressure side impingement chamber 16 through a row of impingement holes 21 to provide impingement cooling to the backside surface of the thermal skin 22 of the pressure side wall. This cooling air flow—from pressure side impingement chamber 16 into a collector chamber 17 and then into the suction side impingement chamber 18 —is repeated until the cooling air flows through a row of return holes 20 and into the last or trailing edge collector chamber 25 . From here the spent cooling air flows through a row of exit holes 26 along the trailing edge and out from the airfoil. In the embodiment of FIG. 1 , all of the cooling air supplied to the cooling air supply chamber 12 flows through the airfoil cooling circuit and then eventually out through the exit cooling holes 26 along the trailing edge. Thus, near wall cooling of the entire airfoil surface is performed with a low volume of cooling air since none of the cooling air is discharged through film holes before reaching the exit cooling holes.

Also, the cooling air supplied to the cooling air supply chamber 12 first flows through impingement chambers that extend along the leading edge region of the airfoil where the heat load is the highest. Thus, the relatively cool cooling an is first used to provide cooling to the hottest section. The micro pin fins on the backside surface of the thin thermal skin in each of the impingement chambers increase the heat transfer rate. In the first embodiment of FIG. 1 , a number of rows of film cooling holes can be used to form a showerhead arrangement that connect to the leading edge impingement chambers 13 - 16 and discharge a layer of film cooling air can be used if warranted.

FIG. 6 shows a second embodiment of the near wall cooling circuit of the present invention where some of the collector chambers 17 include a row of film cooling holes 31 that open onto the airfoil surface as film cooling holes 32 . As seen in FIG. 6 , the first collector chamber discharges film cooling air onto the pressure side wall and the next adjacent collector chamber discharges film cooling air onto the suction side wall. This repeats for the next three collector chambers. The number of collector chambers with film cooling holes and the location of the film cooling holes will depend on the location required for film cooling of the airfoil outer surfaces. In the FIG. 6 embodiment, some of the cooling an collected in the collector chamber 17 will flow out through the film holes 32 while the remaining cooling air flows into the next impingement chamber 17 .

In either embodiment above, the main spar can be cast and the impingement holes and the film holes (if used) can be drilled into the spar after the casting process. The thin thermal skin 22 can then be bonded to the spar and the TBC applied over the surface. The film holes can be drilled after the TBC is applied of the film holes can be covered while the TBC is applied and then uncovered to leave the holes open on the TBC surface.

The impingement holes 21 of each of the collector and impingement chambers can be sized to regulate the amount of impingement cooling produced depending upon the temperature and pressure profile of the airfoil. Also, the use of total cooling for repeating impingement cooling process generates extremely high turbulence for a fixed amount of coolant flow and therefore creates a high value of internal heat transfer coefficient. This yields a higher internal convection cooling effectiveness than the prior art single pass impingement cooling design. The end result of the low flow near wall cooling design of the present invention is to achieve a balance between a longer airfoil life and a reduced cooling flow amount.

Instead of the micro pin fins on the backside surface of the thermal skin, concaved shaped dimples can be used. The thermal skin can be a different material than the spar or can be the same material. also, the thermal skin can be one piece that extends around the leading edge on both sides of the airfoil ending at the trailing edge, or can be formed from multiple pieces all bonded to the spar. The thermal skin can be a high temperature resistant material in a thin sheet form in order to produce very high levels of near wall cooling. The micro pin fins 24 can be formed by photo etching or chemical etching onto the backside of the thermal skin. A low conductivity TBC material can be used on the thermal skin external surface to provide further reduction of the heat flux onto the airfoil external wall.