Segregated cooling air passages for turbine vane

An airfoil for a gas turbine engine includes a cavity including an internal surface of an outer wall. A baffle is disposed within the cavity and spaced apart from the internal surface. A partition is disposed between the baffle and the internal surface to divide a space between the baffle and the internal surface into at least a first passage and a second passage. A gas turbine engine is also disclosed.

BACKGROUND

Components within a turbine section include features for cooling air onto surfaces exposed to the gas flow and maintain temperatures within acceptable limits. Air for cooling is provided at pressures above those generated by the gas flow through a core flow path to prevent ingestion into interior cavities of components. Increasing airflow for cooling purposes can reduce overall engine operating efficiencies.

Turbine engine manufacturers continue to seek further improvements to engine performance including improvements to thermal, transfer and propulsive efficiencies.

SUMMARY

In a featured embodiment, an airfoil for a gas turbine engine includes a cavity including an internal surface of an outer wall. A baffle is disposed within the cavity and spaced apart from the internal surface. A partition is disposed between the baffle and the internal surface to divide a space between the baffle and the internal surface into at least a first passage and a second passage.

In another embodiment according to the previous embodiment, the partition extends from a root of the airfoil to a tip. Airflow within the first passage is not communicated to the second passage.

In another embodiment according to any of the previous embodiments, the baffle includes a plurality of impingement openings communicating airflow into the first passage. The impingement openings direct airflow from the baffle against the internal surface.

In another embodiment according to any of the previous embodiments, the outer wall includes a plurality of film cooling holes communicating air from the first passage to an outer surface of the outer wall.

In another embodiment according to any of the previous embodiments, the baffle includes at least one bleed air opening communicating airflow into the second passage.

In another embodiment according to any of the previous embodiments, the second passage includes flow disrupting features.

In another embodiment according to any of the previous embodiments, the flow disrupting features include one of a pedestal extending from the internal surface into the second passage and a pedestal extending from the baffle into the second passage.

In another embodiment according to any of the previous embodiments, the baffle is without impingement cooling openings that communicate airflow into the second passage.

In another embodiment according to any of the previous embodiments, the partition includes a first partition and a second partition dividing space between the baffle and the internal surface into the first passage and the second passage and a third passage.

In another embodiment according to any of the previous embodiments, the cavity is defined at a leading edge of the airfoil and the outer wall includes a portion of the leading edge of the airfoil.

In another embodiment according to any of the previous embodiments, the second passage is disposed on one of a suction side and a pressure side of the airfoil.

In another embodiment according to any of the previous embodiments, the baffle includes an interior space in communication with a source of airflow.

In another embodiment according to any of the previous embodiments, the airfoil is part of a turbine vane.

In another featured embodiment, a gas turbine engine includes a turbine section. At least one vane within the turbine section includes a cavity at a leading edge portion. The cavity includes an internal surface of an outer wall, a baffle disposed within the cavity and spaced apart from the internal surface and a partition disposed between the baffle and the internal surface to divide a space between the baffle and the internal surface into at least a first passage and a second passage. The second passage is disposed on one of suction side and a pressure side of the vane.

In another embodiment according to the previous embodiments, the partition extends from a root of the vane to a tip such that airflow the first passage is not communicated to the second passage.

In another embodiment according to any of the previous embodiments, the baffle includes a plurality of impingement openings communicating airflow into the first passage. The impingement openings direct airflow from the baffle against the internal surface.

In another embodiment according to any of the previous embodiments, the outer wall includes a plurality of film cooling holes communicating air from the first passage to an outer surface of the outer wall.

In another embodiment according to any of the previous embodiments, the baffle includes at least one bleed air opening communicating airflow into the second passage and at least one of a pedestal extending from the internal surface into the second passage and a pedestal extending from the baffle into the second passage.

In another embodiment according to any of the previous embodiments, the second passage is disposed on one of the suction side and pressure side of the vane.

Although the different examples have the specific components shown in the illustrations, embodiments of this invention are not limited to those particular combinations. It is possible to use some of the components or features from one of the examples in combination with features or components from another one of the examples.

DETAILED DESCRIPTION

A mid-turbine frame58of the engine static structure36is arranged generally between the high pressure turbine54and the low pressure turbine46. The mid-turbine frame58further supports bearing systems38in the turbine section28as well as setting airflow entering the low pressure turbine46.

The disclosed gas turbine engine20in one example is a high-bypass geared aircraft engine. In a further example, the gas turbine engine20includes a bypass ratio greater than about six (6), with an example embodiment being greater than about ten (10). In another disclosed embodiment the gas turbine engine20includes a bypass ratio greater than six (6) and less than fifteen (15) The example geared architecture48is an epicyclical gear train, such as a planetary gear system, star gear system or other known gear system, with a gear reduction ratio of greater than about 2.3. The gear ratio may be less than about 5.0 to provide the desired speed of the fan relative to rotational speed of the turbine section28.

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 example turbine engine20includes vanes62disposed within the turbine section28. The vanes62includes features for cooling so that it may operate with a predefined operational life within the extreme temperatures.

Referring toFIGS. 2 and 3with continued reference toFIG. 1, the example turbine vane62includes a plurality of film cooling holes78through which a cooling film airflow82is communicated to maintain the vane62within a predefined temperature range. The predefined temperature range corresponds to material capabilities and the amount of cooling air provided. The disclosed example turbine vane62includes an airfoil portion66with a leading edge68, a trailing edge70, a pressure side74and a suction side72. The vane62extends radially away from the engine axis A from a platform64to a tip76. It should be appreciated that although a vane62is disclosed and explained by way of example that turbine rotors and other airfoils that require cooling would benefit from the disclosures herein.

The turbine vane62includes internal cavities84and86. The cavity84is proximate the leading edge68of the vane62and includes a baffle88. An inner rib114divides the internal cavities84and86. The trailing edge includes the cavity86that includes a trailing edge baffle108. In this example, the baffle88within the leading edge cavity84is a separate structure supported within the cavity84and spaced apart from an inner surface92of the cavity84. The baffle88receives cooling air80within an inner space98and disperses that cooling air against the inner surface92of a wall90. The wall90includes the inner surface92and an outer surface104. The baffle88includes a plurality of impingement openings96through which air flows as schematically indicated at118to impinge against the inner surface92of the wall90. Impingement of cooling air flow118cools against the inner surface92of the wall90to keep it within a desired temperature range.

The amount of cooling air required to cool the vane62is metered by the cooling holes78in fluidic connection to the cooling air supply source98. Moreover, the airflow into the baffle and against the inner surface92must maintain a pressure above that encountered outside of the vane62so that the exhaust gas flow is not ingested into the cavity84. The pressure differential between the core airflow outside of the vanes62and the cooling airflow within the vane62is referred to as back flow margin (BFM). In some locations on the vane62, the BFM is relatively small and therefore variations in pressure along the surface of the vane62may vary such that cooling airflow is required to be increased to maintain higher pressure within the vane62. Increased airflow bleed from other parts of the engine can reduce overall engine operating efficiency. The example vane62includes features for maintaining BFM within acceptable predefined ranges while reducing the amount of airflow required to maintain the vane62within desired temperature ranges.

The example baffle88defines a flow passage between the inner surface92and the baffle88that increases heat transfer coefficients without significantly impacting or requiring complex assembly or manufacturing techniques. The external pressures on the turbine vane62differ between the suction side72, pressure side74, leading edge68and trailing edge70. A lower external pressure is typically encountered on the suction side72relative to pressures at the leading edge68, pressure side74and at the trailing edge70. The lower pressures along the suction side72enable more airflow through openings compared to airflow through similar sized openings in other higher pressure regions such as at the leading edge68. The increased flow on one side of the vane62as compared to flows in other regions of the vane62can reduce efficiency.

A partition94is provided between the baffle88and the inner surface92to define a first passage110proximate the leading edge68and a second passage112proximate the pressure side74. The partition94may be part of the vane62, the baffle88or a separate part. In this example the partition94is formed as part of the vane62. The partition94extends along the entire length of the cavity84. The division of the passage between the baffle88and the inner surface of the vane62provides localized control of internal pressures that are tailored to compensate for variations in BFM along the outer surfaces of the vane62. Pressure within the first passage110is different than pressure within the second passage112.

In this example, the second passage112does not include impingement cooling holes but instead includes a bleed cooling opening100in the baffle88. The bleed cooling opening100is substantially larger than the film cooling holes102fluidically connected to passage112as to minimize pressure fluctuations from the source air inside the baffle98. The second passage112includes a plurality of flow disrupting features106and exhausts air through an opening102in the wall90.

In the disclosed example, the flow disrupting features extend from the inner surface92towards the baffle88to create turbulent flow that improves heat transfer.—In the second passage112, heat transfer is provided by flow through the passage112along the inner surface92of the wall90and not impingement against the inner surface92. Airflow, schematically shown at116into the second passage112flows through the bleed air opening100across the flow disrupting features106and out the opening102. Airflow exhausted through the opening102flows along an outer surface104of the vane62to provide film cooling function.

Airflow through the impingement openings96into the first passage110impinges on the inner surface92of the wall90to cool the wall90. Air then flows through the film cooling holes78that injects the cooling air into the gas flow path along the outer surface104of the vane62.

The partition94enables the baffle88to be utilized to provide different flow characteristics for localized portions of the vane62to compensate for the variations in BFM between the suction side72, leading edge78and pressure side74. The pressures and air flows in the first passage110can be controlled independent of air flow and pressures in the second passage112by sizing the impingement openings96. The pressures and air flows in the second passage112can be controlled independent of air flow and pressures in the first passage110by sizing the bleed air opening100, as well as the size, shape and number of flow disrupting features106.

Referring toFIG. 4, another example turbine vane120is schematically illustrated and includes a partition122on the suction side72. The partition122divides the space between the baffle88and an inner surface92into a first passage124and a second passage126. The first passage124extends along the leading edge68around to the pressure side74. The second passage126is disposed along the suction side72. The second passage is provided cooling air116through a bleed air opening100. The bleed air opening100supplies air into the second passage126where it flows over flow disrupting features106and then out an opening102into the gas flow path along the suction side72.

Air flow is communicated through impingement openings96against the inner surface92within the first passage124. The first passage124utilizes impingement flow through openings96to cool the wall90within the leading edge68along the cavity84to the pressure side74. Along the leading edge68, and pressure side74, the plurality of impingement openings96provide a cooling impingement flow against the inner surface92to cool the wall90. The cooling air flow then flows out of the film cooling air passages78into the gas stream and along the outer surface of the leading edge68and the pressure side74of the vane.

Referring toFIG. 5, another example turbine vane130is schematically shown and includes a first partition122and a second partition94to divide the space between the baffle88and the inner surface92into three passages. In this example, the first passage132is disposed along the leading edge68. A second passage134is disposed along the suction side72and a third passage136is disposed along the pressure side74. Each of the second passage134and the third passage136includes a plurality of flow disrupting features106and are fed air through bleed air openings100. Heat transfer within the second passage134and the third passage136is accomplished through flow along the inner surface92and against the flow disrupting features106.

The example flow disrupting features106may be of any structure understood to disrupt flow and increase surface area to improve heat transfer. In one example embodiment, the flow disrupting features106disrupt flow and increase surface area to improve heat transfer. In one example embodiment, the flow disrupting features106are pedestals that extend from the inner surface92towards the baffle88. In another example, the pedestals are part of the baffle88and extend outward toward the inner surface92. Moreover, instead of the disclosed example flow pedestals other shapes such as ribs, dimples or other shapes that improve heat transfer could be utilized and are within the contemplation of this disclosure.

The disclosed examples show flow entering the secondary passages close to the structural rib114through the bleed hole100, and continuing over the flow obstructing features towards the leading edge68. The bleed hole100may be located near the flow separator122, and cooling flow may instead flow towards the trailing edge70, being exhausted into the gaspath through cooling holes78located in close proximity to the structural rib114.

The first passage132is cooled by way of an impingement flow through impingement openings96defined within the baffle88. As with the other example embodiments, cooling air flow through the openings96impinge on the inner surface92of the wall90proximate the leading edge68to maintain the wall90within a desired temperature range. The impingement flow then flows out the film cooling hole78along the outer surface of the vane130to provide a film cooling function.

Accordingly, the example vane62includes a partition94that divides the space between the inner surface92and the baffle88into different and separate passages to tailor pressure and air flow to specific regions to maintain a predefined BFM and temperatures.