Patent Publication Number: US-2016245094-A1

Title: Engine component

Description:
BACKGROUND OF THE INVENTION 
     Turbine engines, and particularly gas or combustion turbine engines, are rotary engines that extract energy from a flow of combusted gases passing through the engine onto a multitude of turbine blades. Gas turbine engines have been used for land and nautical locomotion and power generation, but are most commonly used for aeronautical applications such as for aircraft, including helicopters. In aircraft, gas turbine engines are used for propulsion of the aircraft. In terrestrial applications, turbine engines are often used for power generation. 
     Gas turbine engines for aircraft are designed to operate at high temperatures to maximize engine efficiency, so cooling of certain engine components, such as the high pressure turbine and the low pressure turbine, may be necessary. Some engine components include film holes that supply a thin layer or film of cooling fluid on a hot surface of the engine component to protect the engine component from hot combustion gas. Typically, cooling is accomplished by ducting cooler air from the high and/or low pressure compressors to the engine components which require film cooling. The cooling air from the compressor is about 500° C. to 700° C. While the compressor air is a high temperature, it is cooler relative to the air that passes through the combustion chamber, which may be around 1000° C. to 2000° C. 
     A prior art film hole  200  in an engine component  202  is shown in cross-section in  FIG. 1 . The engine component  202  includes a hot surface  204  facing a hot combustion gas flow H and a cooling surface  206  facing a cooling fluid flow C. During operation, the cooling fluid flow C is supplied out of the film hole  200  to create a thin layer or film of cool air on the hot surface  204 , protecting it from the hot combustion gas flow H. The film hole  200  includes an inlet  208  provided on a cooling surface  206 , an outlet  210  provided on the hot surface  204 , and a passage  212  connecting the inlet  208  and the outlet  210 . The passage  212  can include sections  214 ,  216  oriented at an angle to each other when viewed from a plane orthogonal to the hot and cooling surfaces  204 ,  206 , i.e. in the cross-sectional view shown in  FIG. 1 . In the illustrated embodiment, the sections include a metering section  214  for metering of the mass flow rate of the cooling fluid flow C, and a diffusing section  216  in which the cooling fluid C may expand to form a wider cooling film. 
     The film hole  200  lies along a longitudinal axis of the passage  212 , also referred to herein as the centerline  218 , which passes through the geometric center of the cross-sectional area of the metering section  214 . The diffusing section  216  can define its own centerline  220 , which passes through the geometric center of the cross-sectional area of the diffusing section  216 . The two centerlines  218 ,  220  intersect at an angle X when viewed from a plane orthogonal to the hot and cooling surfaces  204 ,  206 , i.e. in the cross-sectional view shown in  FIG. 1 . When viewed from the hot surface  204 , as shown in  FIG. 2 , the centerlines  218 ,  220  are collinear, such that there is no lateral change in the direction of cooling air flow C through the passage  212 . 
     BRIEF DESCRIPTION OF THE INVENTION 
     In one aspect, the invention relates to an engine component for a gas turbine engine, the gas turbine engine generating hot combustion gas flow, having a substrate having a hot surface facing the hot combustion gas flow and a cooling surface facing a cooling fluid flow, and a film hole extending through the substrate and having an inlet provided on the cooling surface, an outlet provided on the hot surface, and a passage connecting the inlet and the outlet, wherein the passage comprises an inlet portion defining an inlet portion centerline and an outlet portion defining an outlet portion centerline, which forms a first angle relative to the inlet portion centerline such that the outlet portion centerline is non-collinear with the inlet portion centerline when viewed from the hot surface. 
     In another aspect, the invention relates to an engine component for a gas turbine engine, the gas turbine engine generating hot combustion gas flow, having a substrate having a hot surface facing the hot combustion gas flow and a cooling surface facing a cooling fluid flow, and a film hole extending through the substrate and having an inlet provided on the cooling surface, an outlet provided on the hot surface, and a passage connecting the inlet and the outlet, wherein the passage comprises a first portion defining a first centerline and a second portion, located downstream of the first portion relative to the direction of the cooling fluid flow through the passage, defining a second centerline, wherein the second centerline forms a first angle relative to the first centerline about an axis perpendicular to a plane defined by the hot surface. 
     In yet another aspect, the invention relates to an engine component for a gas turbine engine, the gas turbine engine generating hot combustion gas flow, having a substrate having a hot surface facing the hot combustion gas and a cooling surface facing a cooling fluid flow, and a film hole extending through the substrate and having an inlet provided on the cooling surface, an outlet provided on the hot surface, and a passage connecting the inlet and the outlet, wherein the passage comprises an inlet portion including the inlet and defining an inlet portion centerline, and an outlet portion including the outlet and defining an outlet portion centerline, and wherein the outlet portion centerline intersects the inlet portion centerline at a point and forms a first angle relative to the inlet portion centerline about an axis passing through the point and perpendicular to a plane defined by the hot surface. 
    
    
     
       BRIEF DESCRIPTION OF THE DRAWINGS 
       In the drawings: 
         FIG. 1  is a schematic, sectional view through a film hole of an engine component according to the prior art. 
         FIG. 2  is a top view of the hot surface of the prior art engine component from  FIG. 1 . 
         FIG. 3  is a schematic cross-sectional diagram of a gas turbine engine for an aircraft. 
         FIG. 4  is a side section view of a combustor and a high pressure turbine of the engine from  FIG. 3 . 
         FIG. 5  is a schematic, perspective view of an engine component having a film hole according to a first embodiment of the invention. 
         FIG. 6  is a top view of a hot surface of the engine component from  FIG. 5 . 
         FIG. 7  is a side view of the engine component from  FIG. 5 . 
     
    
    
     DESCRIPTION OF EMBODIMENTS OF THE INVENTION 
     The described embodiments of the present invention are directed to a film-cooled engine component, particularly in a gas turbine engine. For purposes of illustration, aspects of the present invention will be described with respect to an aircraft gas turbine engine. It will be understood, however, that the invention is not so limited and may have general applicability in non-aircraft applications, such as other mobile applications and non-mobile industrial, commercial, and residential applications. 
       FIG. 3  is a schematic cross-sectional diagram of a gas turbine engine  10  for an aircraft. The engine  10  has a generally longitudinally extending axis or centerline  12  extending forward  14  to aft  16 . The engine  10  includes, in downstream serial flow relationship, a fan section  18  including a fan  20 , a compressor section  22  including a booster or low pressure (LP) compressor  24  and a high pressure (HP) compressor  26 , a combustion section  28  including a combustor  30 , a turbine section  32  including a HP turbine  34 , and a LP turbine  36 , and an exhaust section  38 . 
     The fan section  18  includes a fan casing  40  surrounding the fan  20 . The fan  20  includes a plurality of fan blades  42  disposed radially about the centerline  12 . 
     The HP compressor  26 , the combustor  30 , and the HP turbine  34  form a core  44  of the engine  10  which generates combustion gases. The core  44  is surrounded by a core casing  46  which can be coupled with the fan casing  40 . 
     A HP shaft or spool  48  disposed coaxially about the centerline  12  of the engine  10  drivingly connects the HP turbine  34  to the HP compressor  26 . A LP shaft or spool  50 , which is disposed coaxially about the centerline  12  of the engine  10  within the larger diameter annular HP spool  48 , drivingly connects the LP turbine  36  to the LP compressor  24  and fan  20 . 
     The LP compressor  24  and the HP compressor  26  respectively include a plurality of compressor stages  52 ,  54 , in which a set of compressor blades  56 ,  58  rotate relative to a corresponding set of static compressor vanes  60 ,  62  (also called a nozzle) to compress or pressurize the stream of fluid passing through the stage. In a single compressor stage  52 ,  54 , multiple compressor blades  56 ,  58  may be provided in a ring and may extend radially outwardly relative to the centerline  12 , from a blade platform to a blade tip, while the corresponding static compressor vanes  60 ,  62  are positioned downstream of and adjacent to the rotating blades  56 ,  58 . It is noted that the number of blades, vanes, and compressor stages shown in  FIG. 3  were selected for illustrative purposes only, and that other numbers are possible. 
     The HP turbine  34  and the LP turbine  36  respectively include a plurality of turbine stages  64 ,  66 , in which a set of turbine blades  68 ,  70  are rotated relative to a corresponding set of static turbine vanes  72 ,  74  (also called a nozzle) to extract energy from the stream of fluid passing through the stage. In a single turbine stage  64 ,  66 , multiple turbine blades  68 ,  70  may be provided in a ring and may extend radially outwardly relative to the centerline  12 , from a blade platform to a blade tip, while the corresponding static turbine vanes  72 ,  74  are positioned upstream of and adjacent to the rotating blades  68 ,  70 . It is noted that the number of blades, vanes, and turbine stages shown in  FIG. 3  were selected for illustrative purposes only, and that other numbers are possible. 
     In operation, the rotating fan  20  supplies ambient air to the LP compressor  24 , which then supplies pressurized ambient air to the HP compressor  26 , which further pressurizes the ambient air. The pressurized air from the HP compressor  26  is mixed with fuel in combustor  30  and ignited, thereby generating combustion gases. Some work is extracted from these gases by the HP turbine  34 , which drives the HP compressor  26 . The combustion gases are discharged into the LP turbine  36 , which extracts additional work to drive the LP compressor  24 , and the exhaust gas is ultimately discharged from the engine  10  via the exhaust section  38 . The driving of the LP turbine  36  drives the LP spool  50  to rotate the fan  20  and the LP compressor  24 . 
     Some of the ambient air supplied by the fan  20  may bypass the engine core  44  and be used for cooling of portions, especially hot portions, of the engine  10 , and/or used to cool or power other aspects of the aircraft. In the context of a turbine engine, the hot portions of the engine are normally downstream of the combustor  30 , especially the turbine section  32 , with the HP turbine  34  being the hottest portion as it is directly downstream of the combustion section  28 . Other sources of cooling fluid may be, but is not limited to, fluid discharged from the LP compressor  24  or the HP compressor  26 . 
       FIG. 4  is a side section view of the combustor  30  and HP turbine  34  of the engine  10  from  FIG. 3 . The combustor  30  includes a deflector  76  and a combustor liner  77 . Adjacent to the turbine blade  68  of the turbine  34  in the axial direction are sets of radially-spaced, static turbine vanes  72 , with adjacent vanes  72  forming nozzles therebetween. The nozzles turn combustion gas to better flow into the rotating blades so that the maximum energy may be extracted by the turbine  34 . A cooling fluid flow C passes through the vanes  72  to cool the vanes  72  as hot combustion gas flow H passes along the exterior of the vanes  72 . A shroud assembly  78  is adjacent to the rotating blade  68  to minimize flow loss in the turbine  34 . Similar shroud assemblies can also be associated with the LP turbine  36 , the LP compressor  24 , or the HP compressor  26 . 
     One or more of the engine components of the engine  10  includes a film-cooled substrate in which a film hole of an embodiment disclosed further herein may be provided. Some non-limiting examples of the engine component having a film-cooled substrate can include the blades  68 ,  70 , vanes or nozzles  72 ,  74 , combustor deflector  76 , combustor liner  77 , or shroud assembly  78 , described in  FIGS. 3-4 . Other non-limiting examples where film cooling is used include turbine transition ducts and exhaust nozzles. 
       FIG. 5  is a schematic, perspective view showing a portion of an engine component  80  according to a first embodiment of the invention. The engine component  80  may be an engine component of the engine  10  from  FIG. 3 , and can be disposed in a flow of hot gas represented by arrow H. A cooling fluid flow, represented by arrow C may be supplied to cool the engine component. As discussed above with respect to  FIGS. 3-4 , in the context of a turbine engine, the cooling air can be ambient air supplied by the fan  20  which bypasses the engine core  44 , fluid from the LP compressor  24 , or fluid from the HP compressor  26 . 
     The engine component  80  includes a substrate  82  having a hot surface  84  facing the hot combustion gas flow H and a cooling surface  86  facing the cooling fluid C. The substrate  82  may form a wall of the engine component  80 ; the wall may be an exterior or interior wall of the engine component  80 . The first engine component  80  can define at least one interior cavity  88  comprising the cooling surface  86 . The hot surface  84  may be an exterior surface of the engine component  80 . In the case of a gas turbine engine, the hot surface  84  may be exposed to gases having temperatures in the range of 1000° C. to 2000° C. Suitable materials for the substrate  82  include, but are not limited to, steel, refractory metals such as titanium, or superalloys based on nickel, cobalt, or iron, and ceramic matrix composites. The superalloys can include those in equi-axed, directionally solidified, and single crystal structures. 
     The engine component  80  further includes one or more film hole(s)  90  extending through the substrate  82  that provide fluid communication between the interior cavity  88  and the hot surface  84  of the engine component  80 . During operation, the cooling fluid flow C is supplied to the interior cavity  88  and out of the film hole  90  to create a thin layer or film of cool air on the hot surface  84 , protecting it from the hot combustion gas flow H. While only one film hole  90  is shown in  FIG. 5 , it is understood that the engine component  80  may be provided with multiple film holes  90 , which may be arranged in any desired configuration on the engine component  80 . 
     The film hole  90  can have an inlet  92  provided on the cooling surface  86  of the substrate  82 , an outlet  94  provided on the hot surface  84 , and a passage  96  connecting the inlet  92  and the outlet  94 . Cooling fluid flow C enters the film hole  90  at the inlet  92  and passes through the passage  96  before exiting the film hole  90  at the outlet  94 . 
     The passage  96  can include a first portion  98  and a second portion  100  that is downstream of the first portion  98  with respect to the direction of cooling fluid flow C through the passage  96 . Each portion  98 ,  100  of the passage can define a distinct centerline  102 ,  104 , which is the longitudinal axis which passes through the geometric center of the cross-sectional area of the portion  98 ,  100 . As used herein with respect to the film hole  90 , the term “axial direction” and variants thereof refer to the direction of cooling fluid flow C along the centerlines  102 ,  104  from the cooling surface  86  to the hot surface  84 , and the term “radial direction” and variants thereof refer to the direction orthogonal to the centerlines  102 ,  104 . As illustrated herein, the centerlines  102 ,  104  are linear; in other embodiments the centerlines  102 ,  104  may be non-linear or curved, depending on the shape of the film hole  90 . 
     With additional reference to  FIG. 6 , which is a top view of the hot surface  84  of the engine component  80  from  FIG. 5 , the second portion  100  can be rotated relative to the first portion  98 , such that the second centerline  104  is non-collinear with the first centerline  102  when viewed from the hot surface  84 . The second centerline  104  can form a first angle A relative to the first centerline  102 . In the illustrated embodiment, the first angle A can be determined about an axis  106  perpendicular to a plane defined by the hot surface  84 . When viewed from the hot surface  84  as in  FIG. 6 , the axis  106  comes out of the page. 
     Many prior art film holes, including the film hole  200  shown in  FIGS. 1-2 , lie along a single centerline or collinear centerlines when viewed from above. Here, because the second portion  100  is rotated out of the plane in which the first centerline  104  lies, the centerlines  102 ,  104  are non-collinear when viewed from the hot surface  84 . The rotated second portion  100  provides a lateral change in direction for the cooling fluid flow C through the passage  96 . One general benefit to this is the added degree of design freedom and/or flexibility. The inlet  92  and outlet  94  are no longer constrained to be on the same collinear axes. The inlet  92  can be placed as needed with respect to internal features or walls of the engine component  80 , while still maintaining a desired beneficial location for the outlet  94 . 
     The two centerlines  102 ,  104  can intersect at an intersection point  108 . The intersection point  108  may lie within the passage  96  as illustrated herein. The axis  106  may pass through the intersection point  108 , such that the vertex of the first angle A lies at the intersection point  108 . In addition to passing through the intersection point  108 , in some embodiments of the invention the axis  106  may further be perpendicular to both the hot and cooling surfaces  84 ,  86 . 
     It is noted that, in any of the embodiments discussed herein, although the substrate  82  is schematically shown as being generally planar, it is understood that the substrate  82  may be curved for many engine components  80 . However, the curvature of the substrate  82  may be slight in comparison to the size of the film hole  90 , and so for the purposes of discussion and illustration, the substrate  82  is shown as planar. Whether the substrate  82  is planar or curved, the axis  106  may be perpendicular to a plane defined by the hot surface  84  in the localized area of the substrate  82  through which the axis  106  passes. Furthermore, whether the substrate  82  is planar or curved local to the film hole  90 , the hot and cooling surfaces  84 ,  86  may be parallel to each other as shown herein, or may lie in non-parallel planes. 
     The first portion  98  can be inclined in a downstream direction of cooling fluid flow C through the passage  96  such that the first centerline  102  is non-orthogonal to the hot and cooling surfaces  84 ,  86 . The second portion  100  can also be inclined in a downstream direction of cooling fluid flow C through the passage  96  such that the second centerline  104  is non-orthogonal to the hot and cooling surfaces  84 ,  86 . Alternatively, either centerline  102 ,  104  may be orthogonal to one or both of the hot or cooling surfaces  84 ,  86 . 
     The first and second portions  98 ,  100  may have a circular or non-circular cross-sectional shape, where the cross-sectional shape is defined radially relative to the centerline  102 ,  104 , respectively. Non-circular cross-sections may include, but not are limited to, rectangular, elliptical, trapezoidal, or other irregular shapes. Furthermore, the cross-sectional shape of the first and second portions  98 ,  100  may remain substantially constant long the centerline  102 ,  104 , respectively, or may vary. For example, the first and second portions  98 ,  100  may converge or diverge from the centerline  102 ,  104 , respectively, along the axial direction. 
     In the embodiment illustrated, the first portion  98  can be an inlet portion of the passage  96 , such that the first portion  98  includes the inlet  92 . The second portion  100  can be an outlet portion of the passage  96 , such that the second portion  100  includes the outlet  94  of the passage  96 . The inlet portion  98  can extend from the inlet  92  to the intersection point  108 , and the outlet portion  100  can extend from the intersection point  108  to the outlet  94 . 
     More specifically, in the illustrated embodiment, the first portion  98  is defined by a metering section  110  of the passage  96  for metering of the mass flow rate of the cooling fluid flow C, and the second portion  100  is defined by a diffusing section  112  in which the cooling fluid C may expand to form a wider cooling film. The diffusing section  112  may be in serial flow communication with the metering section  110 . The metering section  110  can be provided at or near the inlet  92 , while the diffusing section  112  can be defined at or near the outlet  94 . 
     The metering section  110  is a portion of the passage  96  with the smallest cross-sectional area perpendicular to the direction of cooling fluid flow C through the passage  96 . The metering section  110  may be a discrete location at which the passage  96  has the smallest cross-sectional area, or an elongated section of the passage  96 . 
     An inlet to the metering section  110  communicates with the inlet  92  to the passage  96  and receives the cooling fluid flow C therefrom; in some embodiments of the invention, including the embodiment of  FIG. 5 , the inlet to the metering section  110  may further be coincident with the inlet  92  to the passage  96 . An outlet of the diffusing section  112  is coincident with the outlet  94  of the passage  96 . An outlet of the metering section  110  is coincident with an inlet to the diffusing section  112 , defines a transition where the cooling fluid flow C may begin to expand. 
     The intersection point  108  of the two centerlines  102 ,  104  can lie at the transition between the metering section  110  and the diffusing section  112 . In the illustrated embodiment, the metering section  110  is an elongated section of the passage  96 , and the intersection point  108  lies at a distal or downstream end of the metering section  110 . 
     The overall shape of the second portion  100  shown in  FIG. 5  is substantially similar conical in shape, such that, in the axial direction, the diffusing section  100  generally diverges from the second centerline  104  but has a substantially circular cross-section perpendicular to the second centerline  104 . Alternatively, the second portion  100  may have a substantially elliptical or rectilinear cross-section. 
     In operation, cooling fluid flow C enters the film hole  90  through the inlet  92  and passes through the metering section  98 , turns at the intersection point  108 , and passes through the diffusing section  100  before exiting the film hole  90  at the outlet  94  along the hot surface  84 . The first angle A can represent a lateral change in the general direction of cooling fluid flow C though the passage  96 . The first angle A can be the minimum angle between the first centerline  102  and the second centerline  104 , such that it represents the magnitude or absolute value of the lateral change in direction, regardless of the direction in which the second portion  100  is rotated relative to the first portion  98 , i.e. whether the second portion  100  is rotated up or down, relative to the view shown in  FIG. 6 , about the axis  106 . 
     In one example, the first angle A can be acute. More specifically, the first angle A can be greater than 0 and less than 90 degrees. Still more specifically, the first angle A can range from 0 to 45 degrees. An acute angle A may lower the potential for material damage when manufacturing the film hole  90  and also lower the effects of manufacturing tolerances on cooling performance. Higher angles may also decrease the hole discharge coefficients such that the cooling fluid flow rate through the film hole decreases. In other embodiments of the invention, the first angle A may be obtuse, i.e. greater than 90 degrees. 
       FIG. 7  is a side view of the engine component  80 , in which the film hole  90  is shown in dotted line. When viewed from a plane orthogonal to the hot surface  84 , the second centerline  104  can form a second angle B relative to the first centerline  102 . In the illustrated embodiment, the second angle B can be determined about an axis  114  perpendicular to the axis  106 . The axis  114  may thus be parallel to the plane defined by the hot surface  84 . The axis  114  may pass through the intersection point  108 , such that the vertex of the second angle B lies at the intersection point  108 . 
     The second angle B can represent a longitudinal change in the general direction of cooling fluid flow C through the passage  96 . The second angle B can be the minimum angle between the first centerline  102  and the second centerline  104 , such that it represents the magnitude or absolute value of the longitudinal change in direction, regardless of the direction in which the second portion  100  is rotated relative to the first portion  98 , i.e. whether the second portion  100  is rotated up or down, relative to the view shown in  FIG. 7 , about the axis  114 . 
     In one example, the second angle B can be acute. More specifically, the second angle B can be greater than 0 and less than 90 degrees. Still more specifically, the second angle B can range from 0 to 25 degrees. In other embodiments of the invention, the second angle B may be obtuse, i.e. greater than 90 degrees. 
     It is noted that for many prior art film holes, including the film hole shown in  FIGS. 1-2 , the film hole lies along one or more centerlines which can be viewed in a cross-sectional plane. Here, because the second portion  100  is rotated out of plane with the first portion  98 , the centerlines  102 ,  104  of the film hole  90  cannot be viewed in a single cross-sectional view orthogonal to the hot surface  84 . However, both centerlines  102 ,  104  can be viewed from a plane orthogonal to the hot surface  84  from outside the film hole  90 . 
     Embodiments of the present invention may be combined with shaping or contouring of the metering section and/or diffusing section of the film hole  90 . Embodiments of the present invention may also be applied to film holes without a diffusing section. Embodiments of the present invention may also apply to slot-type film cooling, in which case the outlet  94  is provided within a slot on the hot surface  84 . Further, in any of the above embodiments, a coating can be applied to the hot surface  84  of the substrate  82 . Some non-limiting examples of coatings include a thermal barrier coating, an oxidation protection coating, or combinations thereof. 
     The various embodiments of devices and methods related to the invention disclosed herein provide improved cooling for engine structures, particularly in a turbine component having film holes. One advantage that may be realized in the practice of some embodiments of the described systems is that the film hole has includes portions that are angled relative to each other when viewed from the hot surface. Previous shaped diffuser film holes have significant limitations when applied in compound angle orientations. By allowing the axes of differ portion of the film hole to be different, particularly the axes of the inlet and outlet portions, the outlet portion can be rotated from the inlet portion. The multi-axis film hole provides the ability to rotate the outlet portion as needed for maximum benefit, such as increasing film cooling effectiveness in highly curved airfoil regions, allowing film holes to be preferentially positioned within cooling passages without sacrificing the benefits of full shaping and coverage, and/or allowing decoupling of the flow entry and exit effects due to the relative orientation of the inlet or outlet of the film hole to the cooling fluid or hot combustion gas flow direction local to the film hole, respectively. 
     This written description uses examples to disclose the invention, including the best mode, and also to enable any person skilled in the art to practice the invention, including making and using any devices or systems and performing any incorporated methods. The patentable scope of the invention is defined by the claims, and may include other examples that occur to those skilled in the art. Such other examples are intended to be within the scope of the claims if they have structural elements that do not differ from the literal language of the claims, or if they include equivalent structural elements with insubstantial differences from the literal languages of the claims.