Gas turbine engine airfoil having serpentine fed platform cooling passage

A gas turbine engine airfoil includes a platform, and spaced apart walls that provide an exterior airfoil surface that extends radially from the platform to an end opposite the platform. A serpentine cooling passage is arranged between the walls and has a first passageway that extends from the platform toward the end and a second passageway fluidly connecting to the first passageway and extending from the end toward the platform to an end. A platform cooling passageway is fluidly connected to the end and extends transversely into the platform. A cooling hole fluidly connects the platform cooling passageway to an exterior surface.

BACKGROUND

This disclosure relates to a gas turbine engine airfoil. More particularly, the disclosure relates to a cooling configuration in the airfoil.

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.

Many turbine blades having turns that provide a serpentine shape, which create undesired pressure losses. Some turbine blades use internally cored serpentine cavities to cool the mid-body section of the airfoil between the leading and trailing edges. The cooling flow is fed into a serpentine passage from the root of the blade. Most serpentine configurations use three to five passageways with the last passageway flowing radially outward from the root and finally terminating near the tip. If cooling air does not reach the tip of the last passage, the airfoil could develop a hot spot and burn through. Having the last passageway flow radially outward takes advantage of pumping action from the circumferential forces on the turbine blade, which ensures cooling air reaches the tip of the last passage.

SUMMARY

In one exemplary embodiment, a gas turbine engine airfoil includes a platform, and spaced apart walls that provide an exterior airfoil surface that extends radially from the platform to an end opposite the platform. A serpentine cooling passage is arranged between the walls and has a first passageway that extends from the platform toward the end and a second passageway fluidly connecting to the first passageway and extending from the end toward the platform to an end. A platform cooling passageway is fluidly connected to the end and extends transversely into the platform. A cooling hole fluidly connects the platform cooling passageway to an exterior surface.

In a further embodiment of the above, multiple cooling passages extend radially within the airfoil and are spaced apart from one another in a chord-wise direction.

In a further embodiment of any of the above, the multiple passages include a leading edge passage that is arranged near a leading edge of the exterior airfoil surface.

In a further embodiment of any of the above, the multiple passages include a trailing edge passage that is arranged near a trailing edge of the exterior airfoil surface.

In a further embodiment of any of the above, each of the multiple passages includes discrete inlets that provide cooling flow to the passage.

In a further embodiment of any of the above, the first and second passageways provide an up-pass passageway and a down-pass passageway that form a U-shaped cooling passage.

In a further embodiment of any of the above, the serpentine cooling passage includes a third passageway that is fluidly connected to the second passageway and extends from the platform toward the end.

In a further embodiment of any of the above, the serpentine cooling passage terminates at the end.

In a further embodiment of any of the above, the platform cooling passageway extends along a pressure side of the platform.

In a further embodiment of any of the above, the platform cooling passageway extends along a suction side of the platform.

In a further embodiment of any of the above, the platform cooling passageway extends along a pressure side and a suction side of the platform.

In a further embodiment of any of the above, multiple cooling holes fluidly connect the platform cooling passageway to the exterior surface.

In another exemplary embodiment, a core for a gas turbine engine airfoil includes a serpentine core portion that is configured to provide an inlet that extends to a first passageway. A second passageway is fluidly connected to the first passageway to form a U-shaped cooling passage. A platform core portion is configured to provide a platform cooling passageway that is arranged transverse and connected to the second passageway.

In a further embodiment of the above, the first passageway is an up-pass passageway and the second passageway is a down-pass passageway. The down-pass passageway terminates near the platform cooling passageway.

In a further embodiment of any of the above, the first passageway is an up-pass passageway and the second passageway is a down-pass passageway. The platform cooling passageway is generally normal to the down-pass passageway.

In a further embodiment of any of the above, the platform core portion extends in the opposite directions from the serpentine core portion.

In another exemplary embodiment, a method of cooling an airfoil comprising the steps of supplying a cooling fluid to an airfoil in a radial direction toward an end, turning the cooling fluid from the end back toward the root to a region near a platform, conveying the cooling fluid from the region to the platform and exiting the cooling fluid through a cooling hole to an exterior surface.

In a further embodiment of the above, the supplying step includes providing the cooling fluid through multiple discrete inlets to multiple cooling passages.

In a further embodiment of any of the above, the turning step includes flowing the cooling fluid along a U-shaped serpentine cooling passage.

In a further embodiment of any of the above, the conveying step includes conveying the cooling fluid to the platform on opposite sides of an airfoil.

DETAILED DESCRIPTION

The disclosed serpentine cooling passage may be used in various gas turbine engine components. For exemplary purposes, a turbine blade64is described. It should be understood that the cooling passage may also be used in vanes, blade outer air seals, and turbine platforms, for example.

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 exterior airfoil surface extending in a chord-wise direction C from a leading edge82to a trailing edge84. The airfoil78is provided between pressure (typically concave) and suction (typically convex) wall86,88in an airfoil thickness direction T, which is generally perpendicular to the chord-wise direction C. Multiple turbine blades64are arranged circumferentially in a circumferential direction A. The airfoil78extends from the platform76in the radial direction R, or spanwise, to the tip80.

The airfoil78includes multiple cooling passages90provided between the pressure and suction walls86,88. The exterior airfoil surface may include multiple film cooling holes (not shown) in fluid communication with the cooling passage90. Flow through the cooling passage90illustrated inFIG. 2Ais shown in more detail inFIG. 3.

Referring toFIG. 3, a core112is shown in phantom within the turbine blade64. The core112produces correspondingly shaped passages within the turbine blade using known casting techniques. Alternatively, the airfoil64may be constructed using an additive manufacturing technique in which the cooling passages are formed while constructing the blade layer-by-layer.

The turbine blade64includes multiple cooling passages90A,90B,90C. The cooling passage90A corresponds to a leading edge cooling passage, and the cooling passage90C corresponds to a trailing edge cooling passage. In one type of cooling configuration, a serpentine cooling passage90B is provided in the mid-body section of the airfoil78between the leading and trailing edge cooling passages90A,90C, as shown inFIG. 4. In one example, each of the cooling passages90A,90B,90C is fed by discrete inlets92A,92B,92C, respectively, which are joined at a sprue (FIG. 3) for handling during the casting process.

Typically, a serpentine cooling passage90B has at least one up-pass connected to at least one down-pass interconnected to one another by a bend to provide a U-shaped passage, as shown inFIGS. 3 and 5. In the example shown, the cooling passage90B includes a first passageway94extending radially outward from the inlet92B toward an end, in the example, the tip80. A second passageway96is interconnected to the first passageway94at a first bend100and extends radially downward away from the tip80toward the platform76. In one example, a third passageway98is interconnected to the second passageway96at a second bend102and extends radially upward from the platform76toward the tip where the passageway terminates.

The mid-body of the airfoil78may be susceptible to developing a hot spot if the pumping action of the fluid is ineffective. Thus, the disclosed cooling configuration provides a cooling flow exit at a location on the turbine blade64with a low dump pressure, which ensures that the fluid continues to flow through the serpentine cooling passage90B. To this end, a platform passageway104is arranged within the platform76and is fluidly interconnected to the second passageway96at an end106, which is generally arranged near the second bend102in the example. The platform passageway104is generally normal to the second passageway94. At least one cooling hole108fully connects the platform passageway104to an exterior surface110to provide an exit for the cooling flow near the inner gas flow path, which has a relatively low pressure as compared to the fluid pressure at the inlet92B. The cooling holes108may be any suitable shape, for example, slots, circular, non-circular, linear, non-linear and others. The exterior surface110may be provided in on the platform and or blade necks, for example.

The core112includes a serpentine core114providing the first, second and third passageways94,96,98. The core112also includes a platform core116corresponding to the platform passageway104. The serpentine core portion114includes an up-pass portion118and a down-pass portion120that respectively provide the first and second passageways94,96. The platform core portion116is interconnected to the down pass portion120at an intersection122.

In one an example, the platform passageway104is generally perpendicular to the second passageway96. The first, second and third passageways94,96,98extend in a radial direction and the platform passageway104extends in the circumferential direction A. The serpentine cooling passage90B may be provided by any number of passes. For example, two passes are shown inFIG. 5and three passes are shown inFIG. 3.

A platform passageway may be provided on either or both of the pressure and suction side portions of the platform76, as shown inFIG. 6.

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 and other reasons, the following claims should be studied to determine their true scope and content.