Gas turbine engine airfoil crossover and pedestal rib cooling arrangement

A gas turbine engine component includes spaced apart walls that provide a cooling passage that extends in a first direction. A cross-over rib joins the walls and extends along the first direction. The cross-over rib has holes that extend in a second direction transverse to the first direction. A row of at least one pedestal joins the walls and extends along the first direction. The row and the cross-over rib overlap one another in the second direction.

BACKGROUND

This disclosure relates to a gas turbine engine airfoil. More particularly, the disclosure relates to a cooling arrangement within a cooling passage of 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 blades and vanes, blade outer air seals, turbine platforms, and other components include internal cooling passages having turns that provide a serpentine shape, which create undesired pressure losses. Some of the cooling passages may include portions having turbulence promoters that enhance the cooling effects of the cooling flow through the cooling passage, in particular, at the trailing edge portion of the airfoil. Typical designs carry either a full rib of cross-overs or pedestals extending radially within the trailing edge portion.

SUMMARY

In one exemplary embodiment, a gas turbine engine component includes spaced apart walls that provide a cooling passage that extends in a first direction. A cross-over rib joins the walls and extends along the first direction. The cross-over rib has holes that extend in a second direction transverse to the first direction. A row of at least one pedestal joins the walls and extends along the first direction. The row and the cross-over rib overlap one another in the second direction.

In a further embodiment of the above, spaced apart walls are provided by pressure and suction walls that form an airfoil. The first direction corresponds to a radial direction along which the airfoil extends from a platform to a tip. The second direction corresponds to a chord-wise direction.

In a further embodiment of any of the above, the row is arranged near the tip, and the cross-over is arranged near the platform.

In a further embodiment of any of the above, the airfoil includes a trailing edge portion. The cross-over rib and the row are arranged in the trailing edge portion.

In a further embodiment of any of the above, the cooling passage includes first and second cooling passageways. The cross-over rib is arranged between the first and second cooling passageways and the holes fluidly connect the first and second passageways.

In a further embodiment of any of the above, the cooling passage is tapered along the first direction from a wide end to a narrow end.

In a further embodiment of any of the above, the row is arranged near the narrow end, and the cross-over rib is arranged near the wide end.

In a further embodiment of any of the above, the cross-over rib and the row are aligned with one another in the second direction to provide a non-contiguous rib in the first direction.

In a further embodiment of any of the above, multiple cross-over ribs are spaced apart from one another in the second direction. The row is aligned in the second direction between the multiple cross-over ribs and spaced apart in the first direction.

In a further embodiment of any of the above, multiple rows are spaced apart from one another in the second direction. The cross-over rib is aligned in the second direction between the multiple rows and spaced apart in the first direction.

In a further embodiment of any of the above, the component is one of an airfoil, a blade outer air seal, a vane, a blade, a platform, a combustor liner and an exhaust liner.

In another exemplary embodiment, a gas turbine engine component includes spaced apart walls that provide a cooling passage that extends in a first direction. A cross-over rib joins the walls and extends along the first direction. The cross-over rib has holes that extend in a second direction transverse to the first direction. A row of at least one pedestal joins the walls and extends along the first direction. The row and the cross-over rib generally are aligned with another in the second direction to provide a rib in the first direction.

In a further embodiment of the above, spaced apart walls are provided by pressure and suction walls that form an airfoil. The first direction corresponds to a radial direction along which the airfoil extends from a platform to a tip. The second direction corresponds to a chord-wise direction.

In a further embodiment of any of the above, the row is arranged near the tip, and the cross-over rib is arranged near the platform.

In a further embodiment of any of the above, the airfoil includes a trailing edge portion. The cross-over rib and the row are arranged in the trailing edge portion.

In a further embodiment of any of the above, the cooling passage includes first and second cooling passageways. The cross-over rib is arranged between the first and second cooling passageways and the holes fluidly connect the first and second passageways.

In a further embodiment of any of the above, the cooling passage is tapered along the first direction from a wide end to a narrow end. The row is arranged near the narrow end, and the cross-over rib is arranged near the wide end.

In a further embodiment of any of the above, multiple cross-over ribs are spaced apart from one another in the second direction. The row is aligned in the second direction between the multiple cross-over ribs and spaced apart in the first direction.

In a further embodiment of any of the above, multiple rows are spaced apart from one another in the second direction. The cross-over rib is aligned in the second direction between the multiple rows and spaced apart in the first direction.

In a further embodiment of any of the above, the component is one of an airfoil, a blade outer air seal, a vane, a blade, a platform, a combustor liner and an exhaust liner.

DETAILED DESCRIPTION

The engine20in one example a high-bypass geared aircraft engine. In a further example, the engine20bypass ratio is greater than about six (6), with an example embodiment being greater than ten (10). The example speed reduction device is a geared architecture48however other speed reducing devices such as fluid or electromechanical devices are also within the contemplation of this disclosure. The example geared architecture48is an epicyclic gear train, such as a star gear system or other gear system, with a gear reduction ratio of greater than about 2.3, or more specifically, a ratio of from about 2.2 to about 4.0. In one disclosed embodiment, the engine20bypass ratio is greater than about ten (10:1), the fan diameter is significantly larger than that of the low pressure compressor44, and the low pressure turbine46has a pressure ratio that is greater than about 5:1. Low pressure turbine46pressure ratio is pressure measured prior to inlet of low pressure turbine46as related to the pressure at the outlet of the low pressure turbine46prior to an exhaust nozzle. It should be understood, however, that the above parameters are only exemplary of one embodiment of a geared architecture engine and that the present invention is applicable to other gas turbine engines including direct drive turbofans.

The disclosed 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 part platforms, combustor liners, and exhaust liners, 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 a cooling passage90provided 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 one portion of the cooling passage90illustrated inFIG. 2Ais shown in more detail inFIGS. 3A and 3B.

Some airfoil designs may result in progressively thinner airfoil cross-sections at the radially outboard spans. These airfoils may be hollow to save weight and to provide cooling passages. It is desirable to structurally support the thin, hollow airfoil sections at the tip while maintaining a consistent load path to the root of the airfoil.

Generally, according to this disclosure a combined “rib” is created that consists of both cross-over ribs100and pedestals104in the internal cooling passage90, as shown inFIG. 3A. The cross-over ribs100and pedestals104line up to form a structural “rib”108while still allowing air to pass through the rib. This arrangement enables thinner airfoil sections at the tip and wider airfoil sections at the platform, for example.

Spaced apart walls92,94provide the cooling passage90that extends in a first direction, which corresponds to the radial direction R in the example. In one example, the walls92,94correspond to the pressure and suction side walls86,88. The cross-over rib100joins the walls92,94and extends along the first direction a first length. The cross-over rib100has holes102extending in a second direction, which corresponds to the chord-wise direction C in the example, which is transverse to the first direction. A row106of pedestals104joins the walls92,94and extends along the first direction a second length. The row106and the cross-over rib100overlap one another in the second direction. In the example shown inFIG. 3A, the cross-over rib100and the row106are aligned with one another in the second direction to provide a non-contiguous rib108in the first direction.

The airfoil78includes a trailing edge portion110, and the cross-over rib100and the row106are arranged in the trailing edge portion110. The cooling passage90includes first and second cooling passageways96,98. The cross-over rib100is arranged between the first and second cooling passageways96,98, and the holes102fluidly connect the first and second passageways96,98.

The cooling passage90is tapered along the first direction from a wide end116to a narrow end114, as shown inFIG. 3B. The row106is arranged near the narrow end114, and the cross-over100rib is arranged near the wide end116. In the example of an airfoil that tapers toward the tip, the row106is arranged near the tip80and the cross-over rib100is arranged near the platform76. In some cases, the order of these features may be reversed with the pedestals occupying the inner diameter portion of the rib and the cross-over ribs occupying the outer portion of the rib.

The configuration is not limited to a single row of pedestals or cross-over ribs. An array of pedestal rows206a,206b,such as that shown in the blade164ofFIG. 4, may be used. In this example, multiple rows206a,206bare spaced apart from one another in the second direction. The cross-over rib200is aligned in the second direction between the multiple rows206a,206band spaced apart in the first direction.

Also multiple cross-over ribs may be used, as shown inFIG. 5. A gas turbine engine component is shown with multiple cross-over ribs300a,300bspaced apart from one another in the second direction. The row306of pedestals is aligned in the second direction between the multiple cross-over ribs300a,300band spaced apart in the first direction.

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.