Patent Description:
Gas turbine engines are rotary-type combustion turbine engines built around a power core made up of a compressor, combustor and turbine, arranged in flow series with an upstream inlet and downstream exhaust. The compressor compresses air from the inlet, which is mixed with fuel in the combustor and ignited to generate hot combustion gas. The turbine extracts energy from the expanding combustion gas, and drives the compressor via a common shaft. Energy is delivered in the form of rotational energy in the shaft, reactive thrust from the exhaust, or both.

The individual compressor and turbine sections in each spool are subdivided into a number of stages, which are formed of alternating rows of rotor blade and stator vane airfoils. The airfoils are shaped to turn, accelerate and compress the working fluid flow, or to generate lift for conversion to rotational energy in the turbine.

The compressor section and the turbine section each have airfoils including rotating blades and stationary vanes. It may be desirable to provide a cooling (or heating in the case of the compressor section) airflow through the airfoils due to the relatively great temperatures at which they are operated. In that regard, the airfoils may include exterior walls along with internal ribs or walls that form internal air passages through which a cooling airflow may flow. Because the exterior walls are exposed to relatively hot gaspath air, they may experience greater thermal expansion than the internal ribs or walls. Such difference in thermal expansion undesirably results in compressive and tensile stress experienced between the exterior walls and the internal ribs or walls.

<CIT> discloses a gas turbine engine airfoil including an airfoil outer wall having widthwise spaced apart pressure and suction sidewall sections extending chordally between leading and trailing edges of the airfoil and extending longitudinally from a base to a tip.

<CIT> discloses a turbine component with an airfoil portion having an external suction side and pressure side wall enclosing a central cavity, the cavity being portioned into a leading edge and a trailing edge region by at least one longitudinally extending first web connecting the suction side wall with the pressure side wall.

<CIT> discloses an airfoil blade comprising an internal serpentine coolant circuit having a last downstream passageway bounded by four monolithic inner walls which are monolithic with at least a portion of the outer walls.

<CIT> discloses a turbine blade having an airfoil defined by a concave shaped pressure side out wall and a convex shaped suction side outer wall. The turbine blade includes a rib configuration that partitions the chamber of the airfoil into a plurality of radially extending flow passages.

According to a first aspect, there is provided an airfoil for a gas turbine engine according to claim <NUM>.

The plurality of cooling passages may comprise at least one leading edge cooling passage arranged along the leading edge of the airfoil.

The plurality of cooling passages may comprise at least one trailing edge cooling passages arranged along the trailing edge of the airfoil.

The camber line may define a camber line height Hc and the pressure side main body cooling passage may have a passage height from a base to an apex defined as <NUM> · HC ≤ H<NUM> ≤ <NUM> · HC.

The camber line may not intersect any of the suction side main body cooling passages.

The camber line may define a camber line height Hc and the suction side main body cooling passage may have a passage height from a base to an apex defined as <NUM> · HC ≤ H<NUM> ≤ <NUM> · HC.

The airfoil may comprise a first heat transfer augmentation feature formed on the pressure side exterior wall within the pressure side main body cooling passage and a second heat transfer augmentation feature formed on the suction side exterior wall within the suction side main body cooling passage.

Each of the first and second heat transfer augmentation features may have a height between <NUM> (<NUM> inches) and <NUM> (<NUM> inches).

The first heat transfer augmentation feature may have a first height E<NUM> and the pressure side main body cooling passage has a first passage height H<NUM>, wherein: <MAT>.

The second heat transfer augmentation feature may have a second height E<NUM> and the suction side main body cooling passage has a second passage height H<NUM>, wherein: <MAT>.

Each of the first and second heat transfer augmentation features may comprise at least one of a skewed trip strip, a chevron trip strip, a hemispherical protrusion, and a pin fin.

A cross-sectional area of the at least one pressure side main body cooling passage in a flow direction through the at least one pressure side main body cooling passage may be greater than a cross-sectional area of the at least one suction side main body cooling passage in a flow direction through the at least one suction side main body cooling passage.

According to a second aspect there is provided a gas turbine engine according to claim <NUM>.

It should be understood, however, the following description and drawings are intended to be illustrative and explanatory in nature and non-limiting.

The following descriptions should be considered exemplary. With reference to the accompanying drawings, like elements are numbered alike. The foregoing and other features, and advantages of the present disclosure are apparent from the following detailed description taken in conjunction with the accompanying drawings in which like elements may be numbered alike and:.

With reference to <FIG>, as used herein, "aft" refers to the direction associated with the tail (e.g., the back end) of an aircraft, or generally, to the direction of exhaust of the gas turbine engine (to the right in <FIG>). The term "forward" refers to the direction associated with the nose (e.g., the front end) of an aircraft, or generally, to the direction of flight or motion (to the left in <FIG>). An axial direction A is along an engine central longitudinal axis Ax (left and right on <FIG>). Further, radially inward refers to a negative radial direction relative to the engine axis Ax and radially outward refers to a positive radial direction (radial being up and down in the cross-section of the page of <FIG>). A circumferential direction C is a direction relative to the engine axis Ax (e.g., a direction of rotation of components of the engine; in <FIG>, circumferential is a direction into and out of the page, when offset from the engine axis Ax). An A-R-C axis is shown throughout the drawings to illustrate the relative position of various components.

The exemplary engine <NUM> generally includes a low speed spool <NUM> and a high speed spool <NUM> mounted for rotation about the engine central longitudinal axis Ax relative to an engine static structure <NUM> via several bearing systems <NUM>.

The inner shaft <NUM> and the outer shaft <NUM> are concentric and rotate via bearing systems <NUM> about the engine central longitudinal axis Ax which is collinear with their longitudinal axes.

In a further example, the engine <NUM> bypass ratio is greater than about six (<NUM>), optionally greater than about ten (<NUM>), the geared architecture <NUM> is an epicyclic gear train, such as a planetary gear system or other gear system, with a gear reduction ratio of greater than about <NUM> and the low pressure turbine <NUM> has a pressure ratio that is greater than about five. The engine <NUM> bypass ratio may be greater than about ten (<NUM>:<NUM>), the fan diameter is significantly larger than that of the low pressure compressor <NUM>, and the low pressure turbine <NUM> has a pressure ratio that is greater than about five <NUM>:<NUM>. The geared architecture <NUM> may be an epicycle gear train, such as a planetary gear system or other gear system, with a gear reduction ratio of greater than about <NUM>:<NUM>. It should be understood, however, that the above parameters are only examples of a geared architecture engine and that the present disclosure is applicable to other gas turbine engines including direct drive turbofans.

The fan section <NUM> of the engine <NUM> is designed for a particular flight conditiontypically cruise at about <NUM> Mach and about <NUM>,<NUM> feet (<NUM>,<NUM> meters). The low fan pressure ratio as disclosed herein may be less than about <NUM>. The "Low corrected fan tip speed" as disclosed herein may be less than about <NUM> ft/second (<NUM>/sec).

Although the gas turbine engine <NUM> is depicted as a turbofan, it should be understood that the concepts described herein are not limited to use with the described configuration, as the teachings may be applied to other types of engines such as, but not limited to, turbojets, turboshafts, etc..

<FIG> is a schematic view of a portion of the turbine section <NUM>. Turbine section <NUM> includes a plurality of airfoils <NUM>, <NUM> including, for example, one or more blades and vanes. The airfoils <NUM>, <NUM> may be hollow bodies with internal cavities or cooling passages defining a number of channels, hereinafter airfoil cooling passages, formed therein and extending from an inner diameter <NUM> to an outer diameter <NUM>, or vice-versa. The airfoil cooling passages may be separated by partitions within the airfoils <NUM>, <NUM> that may extend either from the inner diameter <NUM> or the outer diameter <NUM> of the airfoil <NUM>, <NUM>. The partitions may extend the entire length of the component. The partitions may extend for a portion of the length of the airfoil <NUM>, <NUM>, but may stop or end prior to forming a complete wall within the airfoil <NUM>, <NUM>. Thus, each of the airfoil cores may be fluidly connected and form a fluid path within the respective airfoil <NUM>, <NUM>. The airfoils <NUM>, <NUM> may include platforms <NUM> located proximal to the inner diameter <NUM> thereof. Located below the platforms <NUM> (e.g., radially inward with respect to the engine axis Ax) may be airflow ports and/or bleed orifices that enable air to bleed from the internal cooling passages of the airfoils <NUM>, <NUM>. A root of the airfoil may connect to or be part of the platform <NUM>.

The turbine section <NUM> is housed within a case <NUM>, which may have multiple parts (e.g., turbine case, diffuser case, etc.). In various locations, components, such as seals, may be positioned between airfoils <NUM>, <NUM> and the case <NUM>. For example, as shown in <FIG>, blade outer air seals <NUM> (hereafter "BOAS") are located radially outward from the blade <NUM>. As will be appreciated by those of skill in the art, the BOAS <NUM> may include BOAS supports that are configured to fixedly connect or attach the BOAS <NUM> to the case <NUM> (e.g., the BOAS supports may be located between the BOAS <NUM> and the case <NUM>). As shown in <FIG>, the case <NUM> includes a plurality of case hooks <NUM> that engage with BOAS hooks <NUM> to secure the BOAS <NUM> between the case <NUM> and a tip of the airfoil <NUM>.

Referring now to <FIG>, an airfoil <NUM> which is not covered by the claims is shown. The airfoil <NUM> may be a blade employed in a turbine or compressor section of a gas turbine engine. The airfoil <NUM> has a pressure side exterior wall <NUM> and a suction side exterior wall <NUM>. The pressure side exterior wall <NUM> may receive a hot airflow from a combustor section of the gas turbine engine. In that regard, the pressure side exterior wall <NUM> may be exposed to greater pressure than the suction side exterior wall <NUM> during operation of the gas turbine engine. The hot airflow may cause the airfoil <NUM> to rotate about the engine axis Ax, as will be appreciated by those of skill in the art. The airfoil <NUM> includes a leading edge <NUM> and a trailing edge <NUM>. The leading edge <NUM> may be located axially forward of the trailing edge <NUM> and may receive the hot airflow prior to the trailing edge <NUM>.

The airfoil <NUM>, as shown, includes interior ribs <NUM> that define multiple air passages <NUM> therebetween. Further, at least one of the air passages <NUM> may also be defined by the pressure side exterior wall <NUM> and/or the suction side exterior wall <NUM>, as illustratively shown. The interior ribs <NUM> may be arranged into sets of ribs, with a set of first interior ribs <NUM> oriented in a first direction and a set of second interior ribs <NUM> oriented in a second direction that may differ from the first direction. The interior ribs <NUM> may define multiple air passages <NUM> within the airfoil <NUM>. The multiple air passages <NUM> may receive a cooling airflow to reduce a temperature of the airfoil <NUM>. Each of the interior ribs of the set of first interior ribs <NUM> may be oriented at an angle <NUM> relative to the each of the ribs of the set of second interior ribs <NUM>. The angle <NUM> may be between <NUM>° and <NUM>°. Each of the interior ribs <NUM> may contact at least one of the pressure side exterior wall <NUM> or the suction side exterior wall <NUM> and the interior ribs <NUM> may not extend all the way to the opposing pressure side or suction side exterior wall <NUM>, <NUM>. As such, the interior ribs <NUM> may create triangular passages adjacent to only one of the pressure side exterior wall <NUM> or suction side exterior wall <NUM>. Each of the interior ribs <NUM> may extend from the pressure side exterior wall <NUM> to the suction side exterior wall <NUM>. In that regard, the interior ribs <NUM> may form a modified truss structure that defines the multiple air passages <NUM> (as illustratively shown in <FIG>) including a first plurality of triangular air passages <NUM> (pressure side), a second plurality of triangular air passages <NUM> (suction side), and a plurality of internal air passages <NUM>. As shown in <FIG>, the internal air passages <NUM> are diamond shaped. Some of the interior ribs <NUM> may be arranged to form one or more leading edge cooling passages including a leading edge feed cooling passage <NUM> and a leading edge cooling passage <NUM>, as shown in <FIG>. The interior ribs <NUM> may further form one or more trailing edge cooling passages including a trailing edge cooling passage <NUM>, as shown in <FIG>.

Interior ribs of the first set of interior ribs <NUM> and the ribs of the second set of interior ribs <NUM> may be oriented such that the angle <NUM> that is formed between the respective ribs may vary between <NUM>° and <NUM>°. Interior ribs of each set of interior ribs <NUM>, <NUM> intersect and bisect the airfoil <NUM> at a location that is approximate the mean camber line, located between the airfoil pressure side exterior wall <NUM> and suction side exterior wall <NUM>. The interior ribs <NUM> have partial rib segments (of the sets of ribs <NUM>, <NUM> which generally fully extend between the pressure side exterior wall <NUM> and suction side exterior wall <NUM>) that partially extend to a location approximate the mean camber line.

The multiple air passages <NUM> may be oriented in such a way as to segregate the cooling flows into different regions. For example, the first plurality of triangular air passages <NUM> may transport a pressure side cooling airflow, and the second plurality of triangular air passages <NUM> may transport a suction side cooling airflow. The internal air passages <NUM> may function as tip feed passages to transport cooling air to an inner diameter or an exterior diameter extent of the airfoil <NUM> (e.g., to the tip). Because the internal air passages <NUM> are bordered by the interior ribs <NUM> only, instead of the pressure side exterior wall <NUM> or the suction side exterior wall <NUM>, the cooling airflow traveling through the internal air passages <NUM> remains relatively cool. In that regard, the internal air passages <NUM> may provide relatively cool air to the inner diameter or the exterior diameter extent of the airfoil <NUM>.

As shown, the internal passage may be used to provide resupply cooling air flow, through one or more resupply flow apertures <NUM>, to either, or at least one of the first plurality of triangular air passages <NUM> and/or at least one of the second triangular air passages <NUM>. The resupply flow apertures <NUM>, as shown, emanate from the internal air passages <NUM> and provide a fluidic connection through which relatively higher pressure and lower temperature cooling air may be provided to the respective first and second plurality of triangular passages <NUM>, <NUM>. The resupply of higher pressure, colder cooling air from the internal air passages <NUM> may be required to mitigate internal flow separation that may occur in the triangular air passages <NUM>, <NUM> due to Coriolis forces that occur in rotating air passages. In addition to mitigating adverse internal convective heat transfer consequences related to rotating passages, the resupply flow apertures <NUM> emanating from the internal air passages <NUM> may also be necessary to mitigate excessive cooling air heat pickup and/or high pressure losses that may be incurred in respective triangular air passages <NUM>, <NUM>.

It will be appreciated by those of skill in the art that the location of the resupply flow apertures <NUM> shown in the illustrative figures are for illustrative purposes and are not limiting in any way. That is, any combination, orientation, and selection of connected passages by use of resupply flow apertures may be used and/or optimized based on the local external heat flux, cooling flow, pressure loss, and cooling air temperature heat pickup in order achieve local and overall component thermal cooling effectiveness and durability life requirements, without departing from the scope of the present disclosure.

Further, as shown, film cooling hole apertures <NUM> may be formed to emanate from any of the internal cooling passages <NUM>, <NUM>, <NUM> to expel air to an exterior of the airfoil <NUM>. In some such configurations, it may be necessary to incorporate the resupply flow apertures <NUM>, fed from the internal air passages <NUM> to respective triangular passages <NUM>, <NUM> to ensure adequate pressure ratio and back flow margin is maintained across the film cooling hole apertures <NUM> in order to achieve local film cooling effectiveness and thermal cooling performance requirements.

The leading edge feed cooling passage <NUM> and the leading edge cooling passage <NUM> may be configured to transport a leading edge cooling airflow. In some configurations, an airflow from the leading edge feed cooling passage <NUM> into the leading edge cooling passage <NUM> may be an impinging flow. Further, one or more film cooling hole apertures <NUM> may be located on the leading edge <NUM> such that a film layer may be formed on the exterior surface of the airfoil <NUM>, as will be appreciated by those of skill in the art. The trailing edge cooling passage <NUM> may be arranged to transport a trailing edge cooling airflow. The trailing edge cooling airflow may exit the airfoil <NUM> through one or more trailing edge cooling exits <NUM>, such as holes, slots, etc., as will be appreciated by those of skill in the art.

With respect to the interior cavities (i.e., between the leading edge <NUM>, <NUM> and trailing edge <NUM> cavities) are the geometric shaped first plurality of triangular air passages <NUM>, the second plurality of triangular air passages <NUM>, and the plurality of internal air passages <NUM>. The first plurality of triangular air passages <NUM> may each be bordered by a combination of one or more of the interior ribs <NUM> and the pressure side exterior wall <NUM>. For example, the first plurality of triangular air passages <NUM> may include a first triangular air passage <NUM>. The first triangular air passage <NUM> may have a first wall that is defined by a first interior rib <NUM>, a second wall that is defined by a second interior rib <NUM>, and a third wall that is defined by the pressure side exterior wall <NUM>.

Similarly, the second plurality of triangular air passages <NUM> may each be bordered by a combination of one or more of the interior ribs <NUM> and the suction side exterior wall <NUM>. For example, the second plurality of triangular air passages <NUM> may include a second triangular air passage <NUM>. The second triangular air passage <NUM> may have a first wall that is defined by a third rib <NUM>, a second wall that is defined by a fourth rib <NUM>, and a third wall that is defined by the suction side exterior wall <NUM>.

The internal air passages <NUM> may be bordered entirely by three or more ribs of the interior ribs <NUM>. For example, the internal air passages <NUM> may include a first internal air passage <NUM> that is bordered entirely by interior ribs <NUM>. In this configuration, the first internal air passage <NUM> has four sides, each side defined by a portion of the first interior rib <NUM>, the second interior rib <NUM>, the third rib <NUM>, and the fourth rib <NUM>.

One or more of the interior ribs <NUM> may define openings between adjacent air passages <NUM>. For example, shown proximate the leading edge <NUM>, an interior rib <NUM> may define a cooling flow aperture opening <NUM> between the leading edge feed cooling passage <NUM> and the leading edge cooling passage <NUM>. The cooling flow aperture opening <NUM> may allow air to transfer between the leading edge feed cooling passage <NUM> and the leading edge cooling passage <NUM>, as described above. The cooling flow aperture opening <NUM> may be one or more (e.g., an array) of impingement holes between the leading edge feed cooling passage <NUM> and the leading edge cooling passage <NUM>. Likewise, one or more of the internal air passages <NUM> defined by the interior ribs <NUM> may include cooling holes, bleed holes, transfer holes, impingement holes, etc. For example, the pressure side exterior wall <NUM> may include the trailing edge cooling exits <NUM> designed to facilitate movement of the cooling airflow from the trailing edge cooling passage <NUM> to the pressure side exterior wall <NUM> in order to cool the pressure side exterior wall <NUM>. Additional holes or apertures may be arranged in or through one or more of the interior ribs <NUM> and/or on the pressure side <NUM> or the suction side <NUM> of the airfoil to provide a desired cooling scheme, as will be appreciated by those of skill in the art. Further, the internal cooling passages (those not at the leading or trailing edge) may each be fluidly separated or separate from each of the other internal cooling passages.

In order to achieve the target oxidation and thermal mechanical fatigue lives in modern engines with high gaspath temperatures and low cooling flow allotments, a cooling scheme is needed that utilizes the cooling air effectively to meet the oxidation lives and still provide the flexibility that the airfoils need to meet the thermal mechanical fatigue lives. Some airfoils are configured to provide sufficient heat transfer on both the pressure and suction side exterior walls of a blade by utilizing a cold internal wall that is parallel to the exterior wall(s) and ribs that are perpendicular to the exterior wall (e.g., circumferentially extending ribs). A disadvantage of this type of geometric arrangement is the high compressive strains that are induced in the hot exterior wall due to the relatively large differential in absolute operating metal temperature that exists between the cold internal wall and the hot exterior wall. The metal temperature difference between the cold internal wall and the hot exterior wall adversely impacts the relative rate of thermal expansion. In this sense, the stiff cold internal wall thereby constrains the expansion of the hot exterior wall, increasing both compressive stresses and strains. Such high compressive strains reduce the thermal mechanical fatigue capability of the hot exterior wall, resulting in premature crack initiation and accelerated crack propagation.

To address this, a truss-configuration, for example as shown and described above, provides for increases to the thermal mechanical fatigue life by replacing the cold parallel internal wall and perpendicular ribs that cause the high compressive strains with ribs that intersect in an x-shape or truss arrangement. Such blades are also shown and described in <CIT>. The x-shape arrangement of the ribs can provide for a compliant structure that allows the exterior walls of the airfoil to expand without incurring significant compressive load stresses and strains due to the constrains associated with cold internal walls.

The intersection of internal rib geometries may form a non-optimal triangular passage aspect ratio of the suction side cooling passages. In this sense, the suction side triangular cooling passages may exhibit reduced convective cooling characteristics due to the adverse pressure gradients and passage vortices generated by the Coriolis forces that occur in rotating air passages along the "leading" internal wall surfaces (i.e., the suction side internal surfaces) immediately adjacent the hot exterior suction side airfoil wall. Similarly, the intersection of internal rib geometries may form a non-optimal triangular cooling passage aspect ratio of the pressure side cooling passages. In so doing, the pressure side triangular cooling passages may not leverage enhanced internal
convective cooling characteristics due to the favorable pressure gradients and passage vortices generated by the Coriolis forces that occur in rotating air passages along the "trailing" internal wall surfaces (i.e., the pressure side internal surfaces) immediately adjacent the hot exterior pressure side airfoil wall.

Rather than utilizing the x-shape arrangement shown in <FIG>, a Y-shaped arrangement of ribs arranged in opposing orientations may be employed to create relatively small suction side passages, relatively large pressure side passages, and some main body passages isolated from the gaspath. As used herein, the terms large and small with respect to the passages (or cavities) refers to a cross-sectional area in a flow direction through the respective passage (or cavity). That is, the flow area through which a cooling flow passes through a passage in a radial or flow direction through the respective passage.

For example, turning now to <FIG>, a schematic illustration of an airfoil <NUM> in accordance with an embodiment of the present invention is shown. The airfoil <NUM> has a leading edge <NUM>, a trailing edge <NUM>, a pressure side exterior wall <NUM>, and a suction side exterior wall <NUM>. The pressure side exterior wall <NUM> defines a pressure side 406a of the airfoil <NUM> that is exposed to a hot gaspath during operation of a gas turbine engine and the suction side exterior wall <NUM> defines a suction side 408a that is exposed to the hot gaspath during operation. The airfoil <NUM> includes a plurality of internal cavities or cooling passages that are configured to cooling the material of the airfoil <NUM> during operation. The internal passages can include, for example, a leading edge cooling passage <NUM>, a leading edge feed passage <NUM>, a trailing edge cooling passage <NUM>, and a plurality of main body cooling passage <NUM>, <NUM>, <NUM>. The main body cooling passages <NUM>, <NUM>, <NUM> include, as shown, pressure side main body cooling passages <NUM> (arranged along the pressure side <NUM> of the airfoil <NUM>), suction side main body cooling passage <NUM> (arranged along the suction side <NUM> of the airfoil <NUM>), and isolated main body passages <NUM> (arranged internally within the airfoil <NUM> and not exposed to an exterior side wall of the airfoil <NUM>).

As shown, the pressure side main body cooling passages <NUM> have relatively large triangular shapes in cross-section (relative to a flow direction through the respective cooling passages). In contrast, the suction side main body cooling passages <NUM> have relatively small triangular shapes in cross-section. The isolated main body cooling passages <NUM> have generally diamond shapes in cross-section. The isolated main body cooling passages <NUM> are isolated from and not exposed to exterior hot gaspath surfaces of the airfoil <NUM>.

The geometries are defined by ribs within the airfoil <NUM>. As shown, first interior ribs <NUM> extend a full width of the airfoil <NUM> in a circumferential direction. That is, the first interior ribs <NUM> extend fully between the pressure side exterior wall <NUM> and the suction side exterior wall <NUM>. Second interior ribs <NUM> extend a partial width of the airfoil <NUM> in a circumferential direction. That is, the second interior ribs <NUM> extend between a side wall <NUM>, <NUM> and one of the first interior ribs <NUM>. As shown in <FIG>, the second interior ribs <NUM> extend from the suction side exterior wall <NUM> and intersect with a first interior rib <NUM>, with the first interior ribs <NUM> extending fully between the pressure side exterior wall <NUM> and the suction side exterior wall <NUM>.

As shown, the pressure side main body cooling passages <NUM> are defined between two first interior ribs <NUM> and a portion of the pressure side exterior wall <NUM>. The triangular geometry of the pressure side main body cooling passages <NUM> is achieved by the two first interior ribs <NUM> starting at separate locations in a chordwise direction (i.e., in a direction from the leading edge <NUM> to the trailing edge <NUM>) along the pressure side exterior wall <NUM> and converging to substantially the same location along the suction side exterior wall <NUM>. The triangular geometry of the suction side main body cooling passages <NUM> is achieved by the arrangement of a second interior rib <NUM> extending from the suction side exterior wall <NUM> and intersecting with a portion of a first interior rib <NUM>, rather than intersecting with the opposing pressure side exterior wall <NUM>. The diamond geometry of the isolated main body cooling passages <NUM> is defined by portion of two first interior ribs <NUM> and two second interior ribs <NUM>, as illustratively shown.

The dimensions of the various interior main body cooling passages <NUM>, <NUM> may be defined based on a relationship relative to a portion of a first interior rib <NUM> that defines a wall of the respective main body cooling passage <NUM>, <NUM>. In such configurations, for the triangular shape cooling passages shown in <FIG>, an interior wall (in the circumferential direction) of a main body cooling passage <NUM>, <NUM> may be defined as a percentage of a length of a respective first interior rib <NUM> (extending between the pressure side exterior wall <NUM> and the suction side exterior wall <NUM>) that defines a wall or surface of a respective main body cooling passage <NUM>, <NUM>. For example, a length of wall (in the circumferential direction) of the pressure side main body cooling passage <NUM> may range between <NUM>%-<NUM>% of the total length of a respective first interior rib <NUM>. In contrast, length of wall (in the circumferential direction) of the suction side main body cooling passages <NUM> may range between <NUM>%-<NUM>% of the total length of a respective first interior rib <NUM>. Said differently, a wall of the suction side main body cooling passages <NUM> may be defined by a location where a respective second interior rib <NUM> intersects a respective first interior rib <NUM> at a location ranging between <NUM>%-<NUM>% of the length of the first interior rib <NUM>.

The first interior ribs <NUM> and the second interior ribs <NUM>, as arranged as shown in <FIG>, form a generally y-shape arrangement. This is in contrast to the generally x-shape arrangement shown in <FIG>. The y-shape arrangement enables or provides for the pressure side main body cooling passages <NUM> to have larger cross-sectional areas (relative to a flow direction through the respective cooling passages) than the suction side main body cooling passages <NUM>.

Each of the main body cooling passages <NUM>, <NUM>, <NUM> may be fluidly separated from each adjacent cooling passage. A cooling flow may enter each main body cooling passage <NUM>, <NUM>, <NUM> at a root of the airfoil <NUM> and the cooling flow may flow radially outward from the root toward the tip, as will be appreciated by those of skill in the art. The cooling flow may exit the main body cooling passages <NUM>, <NUM>, <NUM> through tip purge holes located at a tip of the airfoil. The cooling flow may exit the main body cooling passages <NUM>, <NUM>, <NUM> into a tip flag cooling passage, as will be appreciated by those of skill in the art. The cooling flow may exit the main body cooling passages <NUM>, <NUM>, <NUM> and may be directed toward and out a trailing edge flag exit slot, as will be appreciated by those of skill in the art. Further, the cooling flow may exit the main body cooling passages <NUM>, <NUM>, <NUM> through one or more film holes located on/in the respective pressure side exterior wall <NUM> or suction side exterior wall <NUM>.

As shown, the internal passage may be used to provide resupply cooling air flow through one or more resupply flow apertures <NUM> between the pressure side main body cooling passages <NUM>, the suction side main body cooling passages <NUM>, and the isolated main body cooling passages <NUM>. The resupply flow apertures <NUM>, as shown, can emanate from the isolated main body cooling passages <NUM> to one or more of the pressure side main body cooling passages <NUM> and/or the suction side main body cooling passages <NUM>. Similarly, the resupply flow apertures <NUM> can emanate from a pressure side main body cooling passage <NUM> to supply cooling air to a suction side main body cooling passage <NUM>. Such fluid connections can enable relatively higher pressure and lower temperature cooling air to be provided to the suction side main body cooling passages <NUM>. The resupply of higher pressure, colder cooling air from the internal main body cooling passages <NUM> and/or the larger pressure side main body cooling passages <NUM> may be required to mitigate internal flow separation that may occur in the relatively smaller suction side main body cooling passages air passages <NUM> due to Coriolis forces that occur in rotating air passages. In addition to mitigating adverse internal convective heat transfer consequences related to rotating passages, the resupply flow apertures <NUM> emanating from the internal main body cooling passages <NUM> may also be necessary to mitigate excessive cooling air heat pickup and/or high pressure losses that may be incurred in both the pressure side main body cooling passages <NUM> and the suction side main body cooling passages <NUM>.

Further, as shown, film cooling hole apertures <NUM> may be incorporated and emanate from any of the internal cooling passages <NUM>, <NUM>, <NUM>, <NUM>, <NUM> to expel air to an exterior of the airfoil <NUM>. In some such configurations, it may be necessary to incorporate the resupply flow apertures <NUM>, fed from the internal main body cooling passages <NUM> and/or the pressure side main body cooling passages <NUM> to the suction side main body cooling passages <NUM> to ensure adequate pressure ratio and back flow margin is maintained across the film cooling hole apertures <NUM> in order to achieve local film cooling effectiveness and thermal cooling performance requirements.

Turning now to <FIG>, a schematic illustration of an airfoil <NUM> in accordance with an embodiment of the present invention is shown. The airfoil <NUM> may be substantially similar to that shown and described with respect to <FIG>, and thus similar features may not be described or labeled again. The airfoil <NUM> includes a plurality of pressure side main body cooling passages <NUM> arranged along an exterior pressure side wall <NUM> of the airfoil <NUM> and a plurality of suction side main body cooling passages <NUM> arranged along an exterior suction side wall <NUM> of the airfoil <NUM>. The pressure side and suction side main body cooling passages <NUM>, <NUM> are substantially triangular in shape, as illustratively shown. As described above, the pressure side main body cooling passages <NUM> are relatively larger than the suction side main body cooling passages <NUM>. Further, as shown (but not labeled), the airfoil <NUM> includes one or more internal isolated main body cooling passages, leading edge cooling passages, and trailing edge cooling passages.

As shown in <FIG>, the airfoil <NUM> (and airfoils in general) have or define a camber line <NUM>. The camber line <NUM> is a line drawn from a leading edge <NUM> to a trailing edge <NUM> of the airfoil <NUM>, with the camber line <NUM> being equidistant from the exterior pressure side wall <NUM> and the exterior suction side wall <NUM> of the airfoil <NUM>. That is, the camber line <NUM> is midsurface on the airfoil <NUM> or equidistant from the exterior pressure side wall <NUM> and the exterior suction side wall <NUM>. Accordingly, along the camber line <NUM>, at any given axial location, a camber line height Hc is the same dimension or distance from both the pressure side and the suction side. The pressure and suction main body cooling passages <NUM>, <NUM> are defined relative to the camber line <NUM>. For example, as shown in <FIG>, the pressure side main body cooling passages <NUM> extend across the camber line <NUM> in a direction between the exterior pressure side wall <NUM> and the exterior suction side wall <NUM>. Stated another way, the camber line <NUM> passes through the interior of the pressure side main body cooling passages <NUM>. In contrast, the suction side main body cooling passages <NUM> do not extend across the camber line <NUM>. That is, the camber line <NUM> does not pass through an interior of the suction side main body cooling passages <NUM>.

As illustratively labeled, the pressure side main body cooling passages <NUM> have a first passage height H<NUM> that is defined as a maximum height of the cooling passage from a hot wall side (i.e., along the pressure side exterior wall <NUM>) to an apex of the same cooling passage (e.g., proximate the opposing suction side exterior wall <NUM>). As used herein, the term apex of a cooling passage refers to a portion of a cooling passage defined by the junction of two ribs (e.g., ribs <NUM>, <NUM> shown in <FIG>). The base of the cooling passage is a side of the respective cooling passage along a hot wall of the airfoil. Stated another way, the base may be defined as an axial length of the cooling passage along a wall thereof and the apex is a point or location where two ribs that define the cooling passage intersect with or extend from a side wall of the airfoil. The suction side main body cooling passages <NUM> have a second passage height H<NUM> that is defined as a maximum height of the cooling passage from a hot wall side (i.e., along the suction side exterior wall <NUM>) to an apex of the same cooling passage (e.g., at the junction of two ribs). In some configurations, the heights of the respective pressure and suction side main body cooling passages may be defined relative to the camber line height He. For example, the pressure side main body cooling passages <NUM> may have a passage height defined as <NUM> · HC ≤ H<NUM> ≤ <NUM> · HC. Further, the suction side main body cooling passages <NUM> may have a passage height defined as <NUM> · HC ≤ H<NUM> ≤ <NUM> · HC.

It will be appreciated that the surfaces of walls or ribs that define the main body cooling passages described herein can include one or more heat transfer augmentation features. Heat transfer augmentation features can include, without limitation, normal trip strips, chevron trip strips, angled trip strips, pin fins, hemispherical protrusions, etc. In some configurations such heat transfer augmentation features may be arranged on the hot surfaces of the respective cooling passages (i.e., on the pressure or suction side exterior walls that define, in part, surfaces of the main body cooling passages).

For example, turning now to <FIG>, a schematic illustration of an airfoil <NUM> in accordance with an embodiment of the present invention is shown. The airfoil <NUM> may be substantially similar to that shown and described with respect to <FIG>, and thus similar features may not be described or labeled again. The airfoil <NUM> includes a plurality of suction side main body cooling passages <NUM> arranged along a pressure side exterior wall <NUM> of the airfoil <NUM> and a plurality of suction side main body cooling passages <NUM> arranged along a suction side exterior wall <NUM> of the airfoil <NUM>. As described above, the pressure side main body cooling passages <NUM> are relatively larger than the suction side main body cooling passages <NUM>. Further, as shown (but not labeled), the airfoil <NUM> includes one or more internal isolated main body cooling passages, leading edge cooling passages, and trailing edge cooling passages. The pressure side and suction side main body cooling passages <NUM>, <NUM> are substantially triangular in shape, as illustratively shown.

As shown, the pressure side main body cooling passages <NUM> and the suction side main body cooling passages <NUM> each include respective heat transfer augmentation features <NUM>, <NUM>. First heat transfer augmentation features <NUM> are located within the pressure side main body cooling passages <NUM> and are formed on the pressure side exterior wall <NUM> of the airfoil <NUM>. Second heat transfer augmentation features <NUM> are located within the suction side main body cooling passages <NUM> and are formed on the suction side exterior wall <NUM> of the airfoil <NUM>.

As illustratively labeled, the pressure side main body cooling passages <NUM> have a first passage height H<NUM> that is defined as a maximum height of the cooling passage from a hot wall side (i.e., along the pressure side exterior wall <NUM>) to an apex of the same cooling passage (e.g., proximate the opposing suction side exterior wall <NUM>). As used herein, the term apex of a cooling passage refers to a portion of a cooling passage defined by the junction of two ribs (e.g., ribs <NUM>, <NUM> shown in <FIG>). The base of the cooling passage is a side of the respective cooling passage along a hot wall of the airfoil. Stated another way, the base may be defined as a length of the cooling passage along a wall thereof from a location where two ribs that define the cooling passage intersect with or extend from a side wall of the airfoil. The suction side main body cooling passages <NUM> have a second passage height H<NUM> that is defined as a maximum height of the cooling passage from a hot wall side (i.e., along the suction side exterior wall <NUM>) to an apex of the same cooling passage (e.g., at the junction of two ribs).

In these pressure and suction main body cooling passages <NUM>, <NUM>, as shown, the heat transfer augmentation features <NUM>, <NUM> are formed on the respective pressure and suction side exterior walls <NUM>, <NUM>. The first heat transfer augmentation features <NUM> have a first height E<NUM>, which may be defined as a depth or length of extension from the respective pressure side exterior wall <NUM> into the pressure side main body cooling passage <NUM>. The second heat transfer augmentation features <NUM> have a second height E<NUM>, which may be defined as a depth or length of extension from the respective suction side exterior wall <NUM> into the suction side main body cooling passage <NUM>. The heights E<NUM>, E<NUM> of the heat transfer augmentation features <NUM>, <NUM> may be between <NUM> inches (~<NUM>) and <NUM> inches (<NUM>). The heat transfer augmentation features <NUM>, <NUM> may be trip strips that are normal to a flow direction, skewed at an angle to the flow direction, or chevron shaped.

The dimensions of the pressure and suction main body cooling passages <NUM>, <NUM> may be defined, in part, based on a relationship between the passage height (H) and the height of the heat transfer augmentation features (E) in the cooling passage. The pressure side main body cooling passages <NUM> (i.e., the relatively larger cooling passages) may be defined by a relationship of <MAT>. The suction side main body cooling passages <NUM> (i.e., the relatively smaller cooling passages) may be defined by a relationship of <MAT>.

It will be appreciated that in the airfoils of <FIG>, the various internal cooling passages can include interconnecting resupply cooling flow apertures and/or film cooling hole apertures to fluidly connect to an exterior of the airfoil body, as described above. As described above, such apertures can be employed to ensure desired cooling of the airfoil. For example, such fluid apertures can be employed to ensure adequate pressure ratio and back flow margin is maintained across film cooling hole apertures in order to achieve local film cooling effectiveness and thermal cooling performance requirements. Further, the resupply flow apertures can provide a fluidic connection through which relatively higher pressure and lower temperature cooling air may be provided to adjacent/connected cooling passages. The resupply of higher pressure, colder cooling air from one cooling passage to another may be required to mitigate internal flow separation that may occur due to Coriolis forces that occur in rotating air passages. In addition, to mitigating adverse internal convective heat transfer consequences related to rotating passages, the resupply flow apertures can be employed to mitigate excessive cooling air heat pickup and/or high pressure losses that may be incurred in some cooling passages.

The formation of such apertures (resupply and/or film cooling) may be aided by the cooling passage geometries described herein. That is, the film holes and hole drill manufacturing capability may be improved through incorporation with airfoils according to the present invention. In conventional rectangular shaped cooling passages the internal ribs or walls of the passages are nearly perpendicular to the local internal and exterior wall surfaces. Thus, during manufacturing of the film cooling apertures, a drilling process may be limited to prevent back strike. However, the use of triangular cooling passages, as shown and described herein, enable improved back strike distance for drill bits. Furthermore, shallower surface angles may be employed, as compared to rectangular passages, due to the triangular passage shapes of the pressure and suction side cooling passages. Additionally, increased wall thickness may be located at the apex of a passages where two internal ribs meet an exterior hot wall. The local increase in metal thickness enables larger film cooling hole diffuser geometries with increased area ratio to be utilized. The increased geometric coverage of the film cooling holes enable higher film cooling effectiveness levels to be achieved, thereby reducing the external heat flux along the exterior airfoil surface, resulting in lower operating metal temperatures and improved durability capability.

Advantageously, embodiments of the present invention are directed to improved cooling schemes for airfoils, and particularly blades, or gas turbine engines. The rotation of the blade sets up a Coriolis effect inside the cooling passages that causes the convective heat transfer to increase along the "trailing" internal wall surface of the plurality of pressure side cooling passages of radially outward flowing cooling passages, while the internal convective heat transfer decreases along the "leading" internal wall surface of the plurality of suction side cooling passages of radially outward flowing cooling passages. The larger (cross-sectional area in a flow direction) the passage, the greater the Coriolis effect. Advantageously, embodiments of the present invention implement relatively larger cooling passages on the pressure side. The large pressure side passages take advantage of the Coriolis effect to significantly increase the heat transfer on the hot pressure side exterior wall and reduce the heat transfer on the cold internal ribs. Meanwhile, the small suction side passages minimize the effect that Coriolis has on the suction side exterior wall. This creates a more uniform temperature gradient throughout the part.

In addition, advantageously, the arrangement of ribs creates some main body isolated cooling passages that are isolated from the gaspath. Because the passages are isolated from the gaspath, the cooling air in these passages does not pick up a lot of heat and can be used to cool the hot tip region of the airfoil.

Further, advantageously, because there are no perpendicular ribs or parallel walls in the airfoil, the airfoil is compliant and can tolerate the thermal expansion of the hot pressure and suction side exterior walls without inducing a significant amount of compressive strain.

As used herein, the term "about" and "substantially" are intended to include the degree of error associated with measurement of the particular quantity based upon the equipment available at the time of filing the application. For example, these terms may include a range of ± <NUM>%, or <NUM>%, or <NUM>% of a given value or other percentage change as will be appreciated by those of skill in the art for the particular measurement and/or dimensions referred to herein.

It should be appreciated that relative positional terms such as "forward," "aft," "upper," "lower," "above," "below," "radial," "axial," "circumferential," and the like are with reference to normal operational attitude and should not be considered otherwise limiting.

Claim 1:
An airfoil (<NUM>; <NUM>; <NUM>) for a gas turbine engine (<NUM>), the airfoil comprising:
a leading edge (<NUM>; <NUM>), a trailing edge (<NUM>; <NUM>), a pressure side exterior wall (<NUM>; <NUM>; <NUM>) extending between the leading edge and the trailing edge and defining a pressure side, and a suction side exterior wall (<NUM>; <NUM>; <NUM>) extending between the leading edge and the trailing edge and defining a suction side, wherein a plurality of cooling passages (<NUM>, <NUM>, <NUM>, <NUM>, <NUM>, <NUM>; <NUM>, <NUM>; <NUM>, <NUM>) are formed within the airfoil;
a plurality of first interior ribs (<NUM>) extending from the pressure side exterior wall to the suction side exterior wall; and
a plurality of second interior ribs (<NUM>) extending from the suction side exterior wall toward the pressure side exterior wall and intersecting with one of the first interior ribs of the plurality of first interior ribs, wherein:
at least one pressure side main body cooling passage (<NUM>; <NUM>; <NUM>) is defined between the pressure side exterior wall and two first interior ribs of the plurality of first interior ribs,
at least one suction side main body cooling passage (<NUM>; <NUM>; <NUM>) is defined between the suction side exterior wall, a first interior rib, and a second interior rib; and
at least one isolated main body cooling passage (<NUM>) is defined between two first interior ribs (<NUM>) and two second interior ribs (<NUM>), characterised in that the at least one isolated main body cooling passage is configured to supply cooling air to a tip of the airfoil;
and the airfoil (<NUM>) defines a camber line (<NUM>) extending from the leading edge (<NUM>) to the trailing edge (<NUM>), wherein the camber line is defined as a line that is equidistant from the pressure side and the suction side, wherein the camber line passes through each pressure side main body cooling passage (<NUM>).