Patent Publication Number: US-2016222793-A1

Title: Cooling configuration for engine component

Description:
BACKGROUND 
     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 arrays 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. Stationary vanes, and in particular those located in the turbine section, are cooled to increase performance and service life. One known cooling technique includes providing an internal cooling passageway within the interior of the vane. 
     SUMMARY 
     One exemplary embodiment of this disclosure relates to a gas turbine engine. The engine includes a component having a first wall and a second wall spaced-apart from the first wall. The component further includes a cooling passageway provided in part by a helical wall between the first wall and the second wall. 
     In a further embodiment of any of the above, the second wall is provided about a component axis, and wherein the helical wall extends between the first wall and the second wall in a direction normal to the component axis. 
     In a further embodiment of any of the above, the helical wall is arranged to substantially follow a helical curve around the component axis at an angle inclined relative to a line normal to the component axis. 
     In a further embodiment of any of the above, the cooling passageway is bounded by the helical wall, the first wall, and the second wall. 
     In a further embodiment of any of the above, a distance between the first wall and the second wall varies along the component axis to vary the cross-sectional area of the cooling passageway. 
     In a further embodiment of any of the above, a cross-sectional area of the cooling passageway is greater adjacent a trailing edge of the component than the leading edge of the component. 
     In a further embodiment of any of the above, the component includes a first helical wall and a second helical wall between the first wall and the second wall, the first helical wall spaced-apart from the second helical wall. 
     In a further embodiment of any of the above, the component includes a leading edge cooling passageway provided between a leading edge of the component and a lateral divider adjacent the leading edge. 
     In a further embodiment of any of the above, the component includes a trailing edge passageway provided between a trailing edge of the component and a lateral divider adjacent the trailing edge. 
     In a further embodiment of any of the above, the component is a vane, and wherein the cooling passageway is provided in an airfoil section of the vane. 
     In a further embodiment of any of the above, the component is formed using an additive manufacturing technique. 
     In a further embodiment of any of the above, the first wall is a gas-path wall exposed to a core flow path of the gas turbine engine, and the second wall is a non-gas-path wall. 
     Another exemplary embodiment of this disclosure relates to an engine component. The engine component includes a first wall, a second wall spaced from the first wall, and a cooling passageway provided by a helical wall between the first wall and the second wall. 
     In a further embodiment of any of the above, the helical wall is arranged to substantially follow a helical curve around an axis of the engine component at an angle inclined relative to a line normal to the axis of the engine component. 
     In a further embodiment of any of the above, the engine component includes a first helical wall and a second helical wall between the first wall and the second wall, the first helical wall spaced apart from the second helical wall. 
     In a further embodiment of any of the above, the component includes a leading edge cooling passageway provided between a leading edge of the engine component and a lateral divider adjacent the leading edge. 
     In a further embodiment of any of the above, the component includes a trailing edge passageway provided between a trailing edge of the engine component and a lateral divider adjacent the trailing edge. 
     Another exemplary embodiment of this disclosure relates to a method of forming an engine component. The method includes forming an engine component using an additive manufacturing technique. The engine component has a cooling passageway provided in part by a helical wall. 
     In a further embodiment of any of the above, the engine component includes an airfoil wall and an inner wall spaced inwardly from the airfoil wall, and wherein the airfoil wall, the inner wall, and the helical wall are integrally formed with one another. 
     In a further embodiment of any of the above, the additive manufacturing technique includes building-up layers of material to form the engine component. 
     The embodiments, examples and alternatives of the preceding paragraphs, the claims, or the following description and drawings, including any of their various aspects or respective individual features, may be taken independently or in any combination. Features described in connection with one embodiment are applicable to all embodiments, unless such features are incompatible. 
    
    
     
       BRIEF DESCRIPTION OF THE DRAWINGS 
       The drawings can be briefly described as follows: 
         FIG. 1  schematically illustrates a gas turbine engine embodiment. 
         FIG. 2  is perspective view of an example engine component. 
         FIG. 3  is a broken-away view of the airfoil section of the component of  FIG. 2 , and illustrates a cooling arrangement according to this disclosure. 
     
    
    
     DETAILED DESCRIPTION 
       FIG. 1  schematically illustrates an example gas turbine engine  20  that includes a fan section  22 , a compressor section  24 , a combustor section  26  and a turbine section  28 . Alternative engines might include an augmenter section (not shown) among other systems or features. The fan section  22  drives air along a bypass flow path B while the compressor section  24  draws a core air flow C along a core flow path where air is compressed and communicated to a combustor section  26 . In the combustor section  26 , air is mixed with fuel and ignited to generate a high pressure exhaust gas stream that expands through the turbine section  28  where energy is extracted and utilized to drive the fan section  22  and the compressor section  24 . 
     Although the disclosed non-limiting embodiment depicts a turbofan gas turbine engine, it should be understood that the concepts described herein are not limited to use with turbofans as the teachings may be applied to other types of turbine engines; for example a turbine engine including a three-spool architecture in which three spools concentrically rotate about a common axis and where a low spool enables a low pressure turbine to drive a fan via a gearbox, an intermediate spool that enables an intermediate pressure turbine to drive a first compressor of the compressor section, and a high spool that enables a high pressure turbine to drive a high pressure compressor of the compressor section. The concepts disclosed herein can further be applied outside of gas turbine engines, such as in the context of wind turbines. 
     The example engine  20  generally includes a low speed spool  30  and a high speed spool  32  mounted for rotation about an engine central longitudinal axis A relative to an engine static structure  36  via several bearing systems  38 . It should be understood that various bearing systems  38  at various locations may alternatively or additionally be provided. 
     The low speed spool  30  generally includes an inner shaft  40  that connects a fan  42  and a low pressure (or first) compressor section  44  to a low pressure (or first) turbine section  46 . The inner shaft  40  drives the fan  42  through a speed change device, such as a geared architecture  48 , to drive the fan  42  at a lower speed than the low speed spool  30 . The high-speed spool  32  includes an outer shaft  50  that interconnects a high pressure (or second) compressor section  52  and a high pressure (or second) turbine section  54 . The inner shaft  40  and the outer shaft  50  are concentric and rotate via the bearing systems  38  about the engine central longitudinal axis A. 
     A combustor  56  is arranged between the high pressure compressor  52  and the high pressure turbine  54 . In one example, the high pressure turbine  54  includes at least two stages to provide a double stage high pressure turbine  54 . In another example, the high pressure turbine  54  includes only a single stage. As used herein, a “high pressure” compressor or turbine experiences a higher pressure than a corresponding “low pressure” compressor or turbine. 
     The example low pressure turbine  46  has a pressure ratio that is greater than about five (5). The pressure ratio of the example low pressure turbine  46  is measured prior to an inlet of the low pressure turbine  46  as related to the pressure measured at the outlet of the low pressure turbine  46  prior to an exhaust nozzle. 
     A mid-turbine frame  57  of the engine static structure  36  is arranged generally between the high pressure turbine  54  and the low pressure turbine  46 . The mid-turbine frame  57  further supports bearing systems  38  in the turbine section  28  as well as setting airflow entering the low pressure turbine  46 . 
     The core airflow C within the core flow path is compressed by the low pressure compressor  44 , then by the high pressure compressor  52 , mixed with fuel and ignited in the combustor  56  to produce high speed exhaust gases that are then expanded through the high pressure turbine  54  and low pressure turbine  46 . The mid-turbine frame  57  includes vanes  59 , which are in the core airflow path and function as an inlet guide vane for the low pressure turbine  46 . Utilizing the vane  59  of the mid-turbine frame  57  as the inlet guide vane for low pressure turbine  46  decreases the length of the low pressure turbine  46  without increasing the axial length of the mid-turbine frame  57 . Reducing or eliminating the number of vanes in the low pressure turbine  46  shortens the axial length of the turbine section  28 . Thus, the compactness of the gas turbine engine  20  is increased and a higher power density may be achieved. 
     The disclosed gas turbine engine  20  in one example is a high-bypass geared aircraft engine. In a further example, the gas turbine engine  20  includes a bypass ratio greater than about six (6), with an example embodiment being greater than about ten (10). The example geared architecture  48  is 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. 
     In one disclosed embodiment, the gas turbine engine  20  includes a bypass ratio greater than about ten (10:1) and the fan diameter is significantly larger than an outer diameter of the low pressure compressor  44 . 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. 
     A significant amount of thrust is provided by the bypass flow B due to the high bypass ratio. The fan section  22  of the engine  20  is designed for a particular flight condition—typically cruise at about 0.8 Mach and about 35,000 feet. The flight condition of 0.8 Mach and 35,000 ft., with the engine at its best fuel consumption—also known as “bucket cruise Thrust Specific Fuel Consumption (‘TSFC’)”—is the industry standard parameter of pound-mass (lbm) of fuel per hour being burned divided by pound-force (lbf) of thrust the engine produces at that minimum point. 
     “Low fan pressure ratio” is the pressure ratio across the fan blade alone, without a Fan Exit Guide Vane (“FEGV”) system. The low fan pressure ratio as disclosed herein according to one non-limiting embodiment is less than about 1.50. In another non-limiting embodiment the low fan pressure ratio is less than about 1.45. 
     “Low corrected fan tip speed” is the actual fan tip speed in ft/sec divided by an industry standard temperature correction of [(Tram ° R)/(518.7 ° R)] 0.5 . The “Low corrected fan tip speed,” as disclosed herein according to one non-limiting embodiment, is less than about 1150 ft/second. 
       FIG. 2  illustrates an example gas turbine engine component  60 . In the example, the component  60  is a stator vane. However, this disclosure extends to rotor blades, inlet guide vanes (e.g., the vanes  59  of the mid-turbine frame  57 ), and other engine components such as blade outer air seals (BOAS). This disclosure may be particularly beneficial when used in the turbine section  28 , as components in this portion of the engine  20  are exposed to relatively high temperatures during engine operation. 
     The component  60  includes an inner platform  62 , an outer platform  64  radially outward of the inner platform  62 , and an airfoil section  66  extending radially between the inner and outer platforms  62 ,  64  along a component axis  68 . In this example, the component  60  is mounted such that the component axis  68  extends in a radial direction Z, which is normal to the engine central longitudinal axis A. 
     The airfoil section  66  is primarily defined by an airfoil wall  70 , which provides a leading edge  72 , a trailing edge  74 , and opposed pressure and suction sides  76 ,  78  extending between the leading edge  72  and the trailing edge  74 . In this example the airfoil wall  70  is a gas path wall, in that an exterior surface of the airfoil wall  70  will be exposed to a core air flow C during engine operation. 
       FIG. 3  illustrates the airfoil section  66  of the engine component  60 . For clarity, a portion of the airfoil wall  70  is broken-away to show the detail of the interior of the airfoil section  66 . Further, the inner and outer platforms  62 ,  64  are not illustrated in  FIG. 3 . The airfoil section  66  includes a radially inner surface  80 , and a radially outer surface  82 , which would be positioned adjacent the inner and outer platform  62 ,  64 , respectively. 
     In this example, the engine component  60  includes an inner wall  84 , which is substantially hollow, and circumscribes the component axis  68  to provide a core passageway  86 . The inner wall  84  includes an interior surface  88  facing the core passageway  86 , and an exterior surface  90  facing away from the component axis  68 . The inner wall  84  in this example is a non-gas-path wall, in that the inner wall  84  will not be directly exposed to a core air flow C during engine operation. 
     At least one cooling passageway is provided between the exterior surface  90  of the inner wall  84  and an interior surface  92  of the airfoil section  66 . In this example, the interior surface  92  is provided by an inner surface of the airfoil wall  70  (in particular, the pressure and suction sides  76 ,  78  thereof), a first lateral divider  70 A, and a second lateral divider  70 B. As will be explained below, the first and second lateral dividers  70 A,  70 B are not required in all examples. In such examples, an interior surface of the airfoil wall  70  alone may provide the interior surface  92 . 
     The component  60  includes first and second helical passageways  94 ,  96  provided between the exterior surface  90  and the interior surface  92 . The component further includes first and second helical walls  98 ,  100  extending in a direction normal to the component axis  68  to connect the exterior surface  90  to the interior surface  92 . In another example, the first and second helical walls  98 ,  100  are sloped relative to the component axis. Together with the exterior surface  90  and the interior surface  92 , the first and second helical walls  98 ,  100  bound the first and second helical passageways  94 ,  96 . The first and second helical walls  98 ,  100  are arranged to substantially follow a helical curve around the central component axis  68  at an angle A 1  inclined relative to a line L 1  normal to the component axis  68 . The term “substantially” is used herein to account for acceptable tolerances in this industry. 
     While the illustrated component  60  includes two helical walls (e.g., the first and second helical walls  98 ,  100 ), it should be understood that components having a different number of helical walls come within the scope of this disclosure. For example, the component  60  could include a single helical wall, or could include additional helical walls (e.g., three or four), as desired. 
     In this example, the airfoil wall  70  includes a first lateral divider  70 A adjacent the leading edge  72  of the airfoil section  66 , and a second lateral divider  70 B adjacent the trailing edge  74  of the airfoil section  66 . The first and second lateral dividers  70 A,  70 B extend between the pressure and suction sides  76 ,  78 . As mentioned above, the surfaces of the first and second lateral dividers  70 A,  70 B facing the component axis  68  provide a portion of the interior surface  92 . The lateral dividers  70 A,  70 B in this example further bound leading and trailing edge cooling passageways  102 ,  104 , respectively. 
     A leading edge cooling passageway  102  is provided between the first lateral divider  70 A and the airfoil wall  70  adjacent the leading edge  72  of the airfoil section  66  Likewise, a trailing edge cooling passageway  104  is provided between the second lateral divider  70 B and airfoil wall  70  adjacent the trailing edge  74 . While illustrated in  FIG. 3 , the first and second lateral dividers  70 A,  70 B are not required. In such an example, the first and second helical passageways  94 ,  96  alone would cool the leading edge  72  and trailing edge  74  of the component  60 . 
     During operation of the engine  20 , a cooling flow is directed toward the component  60  from a radially outer source, in one example. A first portion S 1  of the cooling flow is directed into the core passageway  84 , and flows downstream toward a rotor hub, for example. 
     Another portion of a cooling flow of fluid is provided between the exterior surface  90  and the interior surface  92 . In this example, because there are first and second helical walls  98 ,  100 , the cooling flow is illustrated as two flows S 2-1  and S 2-2 . As the fluid is directed toward the component  60 , the fluid will divide into one of the two flows S 2-1 , S 2-2 , and enter one of the two helical passageways  94 ,  96 . Alternatively, the passageways  94 ,  96  could be provided with separate, dedicated flows. The flows S 2-1  and S 2-2  are circulated helically around the inner wall  84 , about the component axis  68 , and travel axially from the radially outer surface  82  toward the radially inner surface  80  of the airfoil section  66  within a respective one of the passageways  94 ,  96 . The flows S 2-1  and S 2-2  provide effective cooling of the airfoil wall  70 , which is exposed to relatively high temperatures from the core airflow C during operation of the engine  20 . 
     Third and fourth portions S 3 , S 4  of the secondary cooling flow are directed to the leading edge and trailing edge passageways  102 ,  104  in the illustrated example. These dedicated passageways  102 ,  104  for the leading edge  72  and the trailing edge  74  provide increased cooling at these locations. Again, however, the leading edge and trailing edge passageways  102 ,  104  are not required, and the level of cooling provided by the helical passageways  94 ,  96  may be adequate in some examples. 
     In one example, the first and second lateral dividers  70 A,  70 B completely separate the third and fourth flows S 3 , S 4  from the flows S 2-1  and S 2-2 . In other examples, however, the first and second lateral dividers  70 A,  70 B allow fluid to communicate between the passageways  102 ,  104  and the helical passageways  94 ,  96 . 
     The cross-sectional area of the first and second helical passageways  94 ,  96  affects the level of heat transfer between the flows S 24 , S 2-2  and the airfoil wall  70 . Consistent with this disclosure, one can tailor the cross-sectional area of the helical passageways  94 ,  96  to selectively increase and decrease the level of cooling along the airfoil section  66 . 
     For instance, the first helical wall  98  and the second helical wall  100  are spaced apart from one another, relative to the component axis  68 , by a first distance D 1 . This distance D 1  may be selected to provide an appropriate cross-section for the first and second helical passageways  94 ,  96 . In the illustrated embodiment, the first and second helical walls  98 ,  100  maintain a constant angle A 1  as the first and second helical walls extend along the component axis  68 . In this example, the first distance D 1  is held constant along the component axis  68 . However, it should be understood that the first distance D 1  could vary to change the cross-sectional area of the helical passageways  94 ,  96  and to selectively provide an appropriate level of cooling at desired locations. 
     Similarly, the distance between the exterior surface  90  and the interior surface  92  can be varied. In the disclosed embodiment, the exterior surface  90  is spaced a distance D 2  from the interior surface  92  adjacent the leading edge  72 , while the exterior surface  90  is spaced a distance D 3  from the interior surface  92  adjacent the trailing edge. The distance D 3  is larger than the distance D 2 , which provides the first and second helical passageways  94 ,  96  with a smaller cross-sectional area adjacent the leading edge  72 , and a greater cross-sectional area adjacent the trailing edge  74 . This, in turn, provides for increased heat transfer (e.g., cooling) of the airfoil wall  70  adjacent the leading edge  72 . 
     It should be understood that this disclosure may be built using any known technique. In one example, the component  60  is formed using an additive manufacturing technique, in which layers of material are successively built-up onto one another to form the component  60 . That is, in one example, the entirety of the component  60  is formed using additive manufacturing, such that the various structures described above are integrally formed with one another. For instance, the inner wall  84 , the airfoil wall  70 , and the first and second helical walls  98 ,  100  are each integrally formed together with one another as one structural piece. 
     In one known additive manufacturing technique, a layer of powdered material (e.g., metal) is provided onto a bed and selectively heated (e.g., melted) by a laser, using electron beam melting, or by high-speed accumulation combined with laser sintering, as examples. The technique may use materials capable of equiax grain boundaries. Further, the technique may use ceramic-to-metal mixing of powders to provide a functionally graded material (FGM). 
     The heated material is allowed to cool to form a first layer of the component. Additional powdered material is provided onto the bed, and another layer of material is melted onto the first layer, and allowed to cool. The process repeats, incrementally building-up the component with each layer. Additive manufacturing techniques are particularly beneficial for forming components with intricate internal passageways, such as the first and second helical walls  98 ,  100 , which are not easily formed using other, conventional machining techniques. 
     The component  60  may be manufactured of separate cores. In one example, at least an interior core (e.g., an interior core providing at least the inner wall  84  and the first and second helical passageways  98 ,  100 ) is additively manufactured and is then inserted into an outer core. The outer core may or may not be additively manufactured. The inner and outer cores are then joined together. 
     While the flows S 1 , S 2-1 , S 2-2 , S 3  and S 4  have been described above as being portions of a secondary cooling airflow, these flows could be provided by separate, dedicated flows of fluid. 
     Although the different examples have the specific components shown in the illustrations, embodiments of this disclosure 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. 
     One of ordinary skill in this art would understand that the above-described embodiments are exemplary and non-limiting. That is, modifications of this disclosure would come within the scope of the claims. Accordingly, the following claims should be studied to determine their true scope and content.