Resupply hole of cooling air into gas turbine blade serpentine passage

An airfoil for a gas turbine engine includes pressure and suction side walls joined to one another at leading and trailing edges and extending in a radial direction. A serpentine cooling passage is provided between the pressure and suction side walls. A passageway adjoins a passage wall and is fluidly interconnected to an upstream turn that has radially spaced apart innermost and outermost contours. The innermost contour is provided at the wall. A resupply hole is configured to exit downstream from the innermost contour.

BACKGROUND

This disclosure relates to a gas turbine engine component, such as an airfoil. More particularly, the disclosure relates to a cooling configuration.

Gas turbine engines typically include a compressor section, a combustor section and a turbine section. During operation, air is pressurized in the compressor section and is mixed with fuel and burned in the combustor section to generate hot combustion gases. The hot combustion gases are communicated through the turbine section, which extracts energy from the hot combustion gases to power the compressor section and other gas turbine engine loads.

Both the compressor and turbine sections may include alternating series of rotating blades and stationary vanes that extend into the core flow path of the gas turbine engine. For example, in the turbine section, turbine blades rotate and extract energy from the hot combustion gases that are communicated along the core flow path of the gas turbine engine. The turbine vanes, which generally do not rotate, guide the airflow and prepare it for the next set of blades.

Gas path temperatures are well above the melting point of hot section components, necessitating cooling component for adequate durability. Many blades and vanes, blade outer air seals, turbine platforms, and other components include internal cooling passages. Multi-pass serpentine cooling passage configurations, which are a common cooling scheme for gas turbine blades, are susceptible to problems stemming from the cooling air heating up as the cooling flow navigates the internal passages. As the cooling air flows through the airfoil, this air cools the metal surfaces of the blade; however, the cooling airfoil becomes increasingly hot and less effective as the air travels down the passageway.

One approach to provide cooler air further down the passageway of a serpentine cooling passage is to provide a resupply hole parallel to a chord-wise direction of the airfoil and tangent to an outer radius of a turn interconnecting parallel passageways of the serpentine cooling passage.

SUMMARY

In one exemplary embodiment, an airfoil for a gas turbine engine includes pressure and suction side walls joined to one another at leading and trailing edges and extending in a radial direction. A serpentine cooling passage is provided between the pressure and suction side walls. A passageway adjoins a passage wall and is fluidly interconnected to an upstream turn that has radially spaced apart innermost and outermost contours. The innermost contour is provided at the wall. A resupply hole is configured to exit downstream from the innermost contour.

In a further embodiment of any of the above, the serpentine cooling passage includes first, second and third passageways. The first and second passageways are separated by a first wall and fluidly interconnected to one another by a first turn. The second and third passageways are separated by a second wall and fluidly interconnected to one another by a second turn. The passageway corresponds to the third passageway. The passage wall corresponding to the second wall and the turn corresponding to the second turn.

In a further embodiment of any of the above, the first and third passageways are configured to carry a cooling fluid radially outward. The second passageway is configured to carry the cooling fluid radially inward.

In a further embodiment of any of the above, a trailing edge cooling passage is fluidly connected to the third passageway and extends to the trailing edge.

In a further embodiment of any of the above, a root includes an inlet passage. The resupply hole fluidly interconnects the inlet passage to the third passageway.

In a further embodiment of any of the above, the inlet passage is fluidly interconnected to the first passageway.

In a further embodiment of any of the above, the inlet passage includes a boss to which the resupply hole is fluidly connected. The boss is provided on one side of the airfoil with respect to a thickness direction.

In a further embodiment of any of the above, the boss is provided on the pressure side.

In a further embodiment of any of the above, the resupply hole is fluidly connected to the inlet passage centrally with respect to the thickness direction.

In a further embodiment of any of the above, the inlet passage includes first and second inlet passages converging to a junction that is configured to fluidly feed the first passageway. The resupply hole is fluidly connected to the junction.

In a further embodiment of any of the above, the resupply hole is arranged at an angle that is less than 90° with respect to the radial direction.

In a further embodiment of any of the above, the angle is in a range of 15°-75°.

In another exemplary embodiment, a gas turbine engine includes a compressor section. A combustor is fluidly connected to the compressor section. A turbine section is fluidly connected to the combustor. The turbine section includes an array of turbine blades. Each turbine blade includes an airfoil that has pressure and suction side walls joined to one another at leading and trailing edges and extending in a radial direction. A serpentine cooling passage is provided between the pressure and suction side walls. A passageway adjoins a passage wall and is fluidly interconnected to an upstream turn that has radially spaced apart innermost and outermost contours. The innermost contour is provided at the wall. A resupply hole is configured to exit downstream from the innermost contour.

In a further embodiment of any of the above, the serpentine cooling passage includes first, second and third passageways. The first and second passageways are separated by a first wall and fluidly interconnected to one another by a first turn. The second and third passageways are separated by a second wall and fluidly interconnected to one another by a second turn. The passageway corresponds to the third passageway. The passage wall corresponds to the second wall. The turn corresponds to the second turn. The first and third passageways are configured to carry a cooling fluid radially outward. The second passageway is configured to carry the cooling fluid radially inward and comprises a trailing edge cooling passage fluidly connected to the third passageway and extending to the trailing edge.

In a further embodiment of any of the above, a root includes an inlet passage. The resupply hole fluidly interconnects the inlet passage to the third passageway. The inlet passage is fluidly interconnected to the first passageway.

In a further embodiment of any of the above, the inlet passage includes a boss to which the resupply hole is fluidly connected. The boss is provided on one side of the airfoil with respect to a thickness direction. The boss is provided on the pressure side.

In a further embodiment of any of the above, the resupply hole is fluidly connected to the inlet passage centrally with respect to the thickness direction.

In a further embodiment of any of the above, the resupply hole is arranged at an angle that is less than 90° with respect to the radial direction.

In a further embodiment of any of the above, the angle is in a range of 15°-75°.

In a further embodiment of any of the above, the blade is a first stage high pressure turbine blade.

DETAILED DESCRIPTION

A mid-turbine frame57of the engine static structure36is arranged generally between the high pressure turbine54and the low pressure turbine46. The mid-turbine frame57further supports bearing systems38in the turbine section28as well as setting airflow entering the low pressure turbine46.

In one disclosed embodiment, the gas turbine engine20includes a bypass ratio greater than about ten (10:1) and the fan diameter is significantly larger than an outer diameter of the low pressure compressor44. It should be understood, however, that the above parameters are only exemplary of one embodiment of a gas turbine engine including a geared architecture and that the present disclosure is applicable to other gas turbine engines.

The disclosed cooling passage may be used in various gas turbine engine components. For exemplary purposes, a turbine blade64is described, which may be provided in a first stage of the high pressure turbine section54. Referring toFIGS. 2A and 2B, a root74of each turbine blade64is mounted to the rotor disk. The turbine blade64includes a platform76, which provides the inner flow path, supported by the root74. An airfoil78extends in a radial direction R from the platform76to a tip80. It should be understood that the turbine blades may be integrally formed with the rotor such that the roots are eliminated. In such a configuration, the platform is provided by the outer diameter of the rotor. The airfoil78provides leading and trailing edges82,84. The tip80is arranged adjacent to a blade outer air seal (not shown).

The airfoil78ofFIG. 2Bsomewhat schematically illustrates an exterior airfoil surface extending in a chord-wise direction C from a leading edge82to a trailing edge84. The airfoil78is provided between spaced apart pressure (typically concave) and suction (typically convex) wall86,88in an airfoil thickness direction T, which is generally perpendicular to the chord-wise direction C. The pressure and suction walls86,88are joined at the leading and trailing edges82,84. Multiple turbine blades64are arranged circumferentially with a stage in a circumferential direction A. The airfoil78extends from the platform76in the radial direction R, or spanwise, to the tip80.

The airfoil78includes a cooling passage configuration90provided between the pressure and suction walls86,88. The cooling passage configuration90, which typically comprises multiple cooling passages, is fed a cooling fluid through the root74from a cooling source100. In one example, the cooling source is provided by a stage of the compressor section24. The exterior airfoil surface may include multiple film cooling holes (not shown) in fluid communication with the cooling passage configuration90.

Referring toFIGS. 3 and 5, multiple inlet passages (numbered1-5inFIG. 3) are provided in the root74to feed the various cooling passages within the blade64. The cooling passage configuration90includes a leading edge cooling passage102fed by an inlet passage at1. Another inlet passage is fed at2to supply cooling fluid to first and second pressure side cooling passages106A,106B. An inlet passage at3feeds cooling fluid to a suction side cooling passage104.

First and second inlet passages110A,110B merge into a junction111to feed cooling fluid to a serpentine cooling passage108, which can be any suitable shape such as S-, U- or L-shaped. The serpentine cooling passage108includes first, second and third passageways112,114,116. A trailing edge cooling passage136fluidly connected to the third passageway116, which carries the cooling fluid toward and out the trailing edge84.

The first and second passageways112,114are separated by a passage wall bridging the pressure and suction side walls86,88in the thickness direction T, i.e., first wall118, that generally extends in the radial direction R. The second and third passageways114,116are separated by a passage wall bridging the pressure and suction side walls86,88in the thickness direction T, i.e., second wall120, that generally extends in the radial direction R. A first turn122fluidly interconnects the first and second passageways112,114, and a second turn124fluidly interconnects the second and third passageways114,116. The first and third passageways112,116carry the cooling fluid in a radial direction toward the tip80(an “up-pass”), and the second passageway114carries cooling fluid from the first passageway112to the third passageway116in the radial direction R from the tip80toward platform76(a “down-pass”). Finally, the third passageway116carries the fluid from the second passage114toward the tip80(an “up-pass”) and out the trailing edge cooling passage126.

As the cooling fluid flows from the inlet passages110A,110B toward the third passageway116, the cooling fluid heats up becoming less effective. The disclosed airfoil includes a resupply hole132that is configured to supply fresh cooling air into the final up-pass, which corresponds to the third passageway116of the serpentine cooling passage108. The resupply hole132is oriented such that the fresh cooling air is biased or carried toward the tip80. Typically cooling fluid from prior art configurations is swept out the trailing edge cooling passage before travelling very far toward the tip, minimizing the cooling effectiveness of the fresh cooling air. In the disclosed embodiment, the resupply hole132is angled toward the tip80and trailing edge84at a relatively shallow angle θ with respect to the radial direction R such that the fresh cooling air tends to travel radially outward to provide cooling to the tip of the blade while encouraging the relatively warmer cooling air to exit the serpentine cooling passage108at an inboard radial span. In one example, the angle θ is less than 90°, in another example, the angle θ is in a range of 15°-75°.

Referring toFIGS. 3, 3A, 5 and 6, the second turn124includes an extrados or outermost contour134and an intrados or innermost contour136. The resupply hole132includes an inlet138provided at the boss128or first passageway112. An outlet140of the resupply hole132is arranged downstream from the innermost contour136. The outlet140is positioned at a radially outward location in the radial direction R with respect to a radially inner edge133of the trailing edge cooling passage126. In the example, the resupply hole132fluidly interconnects an inlet passage (e.g.,110A/110B) to a location at or downstream from the innermost contour136, which ensures that the fresh cooling fluid will encourage cool air to flow radially outward without first being swept out the trailing edge cooling passage126as occurs in prior art cooling arrangements.

As best shown inFIGS. 3, 3A and 6, a boss128is provided by the junction111to provide an inlet138to the resupply hole132on one side of the airfoil with respect to the thickness direction T, for example the pressure side. In this configuration, a portion of the serpentine cooling passage108may include a recess130, best shown inFIG. 3, to accommodate the resupply hole132. An arrangement of core portions providing such a configuration is illustrated inFIG. 3A. A resupply hole core144interconnects an inlet core142to a third passageway core148. The resupply hole core144is arranged to one side of the second passageway core146.

In another example illustrated inFIGS. 4 and 4A, the resupply core hole144may pass through the second passageway core146to locate the resupply hole more centrally with respect to the thickness direction T. In one example, the resupply hole core144is provided by a quartz rod that is supported with respect to an aperture152in the second passageway core146by wax150. This supports the fragile quartz rod during the next step of the airfoil manufacturing process, which typically includes immersing the core assembly in wax within an airfoil mold, as is known in the art.

The tip of rotating gas turbine blades are difficult to cool, particularly in a multi-pass serpentine design. Because the fluid in the serpentine heats up as it continues along the passages, the flow at the end of the final up-pass in these serpentines typically is much less effective at cooling the airfoil than the air at the inlet to the serpentine. Prior art schemes to resupply fresh cooling air to turbine blades introduce the new air at the root of the blade, which it typically gets forced out of the blade at a low radial span of the blade. This may provide a cooling benefit to the lower radial spans of the blade, but very little cooling benefit is provided toward the tip. The disclosed cooling configuration introduces cooling air at a higher radial span in a location where the fresh, cooler flow will tend to provide a cooling benefit toward the outboard radial spans of the airfoil, and very little cooling benefit (if any) toward the root of the blade. Thus, the disclosed cooling configuration locally redistributes additional cooling benefit to the tip of the airfoil, where improved cooling may be more desirable.