Patent Application
20170101870
This disclosure relates to a gas turbine engine, and more particularly to cooling features of a gas turbine engine.
Gas turbine engines, such as those used to power modern commercial and military aircrafts, generally include a compressor section to pressurize an airflow, a combustor section for burning hydrocarbon fuel in the presence of the pressurized air, and a turbine section to extract energy from the resultant combustion gases.
Gas turbine engine components, such as airfoils, combustor sections, augmentor sections, and exhaust duct liner for example, are subject to high thermal loads for prolonged periods of time. Conventionally, various cooling arrangements are provided to film cool these components. Among these are impingement cooling on a backside of the component and film cooling on a gas side of the component to maintain temperature within material limits.
In some aspects of the disclosure, a component of a gas turbine engine is provided including at least one cooling hole formed in the component. The cooling hole has an interior surface that defines a flow path for air configured to cool a portion of the component. A feature is arranged within at least a portion of the cooling hole. The feature is configured to generate non-linear movement of the air as it flows there through.
In addition to one or more of the features described above, or as an alternative, further embodiments may include that the feature extends over a portion of the cooling hole.
In addition to one or more of the features described above, or as an alternative, further embodiments may include that the feature extends over substantially all of a length the cooling hole.
In addition to one or more of the features described above, or as an alternative, further embodiments may include that the cooling hole has a substantially uniform hydraulic diameter over its length.
In addition to one or more of the features described above, or as an alternative, further embodiments may include that the cooling hole includes a metering section and a diffusion section. The diffusion section is fluidly coupled to and arranged downstream from the metering section.
In addition to one or more of the features described above, or as an alternative, further embodiments the hydraulic diameter of the diffusion section increases in a direction away from the metering section.
In addition to one or more of the features described above, or as an alternative, further embodiments may include that the cooling hole has a tapered configuration.
In addition to one or more of the features described above, or as an alternative, further embodiments may include that the cooling hole has a conical configuration.
In addition to one or more of the features described above, or as an alternative, further embodiments may include that the feature extends about at least a portion of a periphery of an inner surface of the cooling hole.
In addition to one or more of the features described above, or as an alternative, further embodiments may include that the feature is integrally formed with an interior surface of the cooling hole.
In addition to one or more of the features described above, or as an alternative, further embodiments may include that the feature includes rifling having interleaved lands and grooves, the rifling being arranged spirally about an axis of defined by the cooling hole.
In addition to one or more of the features described above, or as an alternative, further embodiments may include that a height of the lands is between about 3% and about 30% of a hydraulic diameter of the metering section.
In addition to one or more of the features described above, or as an alternative, further embodiments may include that at least one of the lands is configured to rotate 360° occur over a distance between about 2 and about 15 times a hydraulic diameter of the cooling hole.
In addition to one or more of the features described above, or as an alternative, further embodiments may include that the component is an airfoil.
In some aspects of the disclosure, a turbine engine is provided including a component exposed to hot gas flow. At least one cooling hole is formed in the interior of the component. The cooling hole defines a flow path for air configured to cool a portion of the component and includes a feature configured to generate non-linear movement of air as it flows there through.
In addition to one or more of the features described above, or as an alternative, further embodiments may include that the feature is configured to cause the air to rotate about the central axis of the cooling hole.
In addition to one or more of the features described above, or as an alternative, further embodiments may include that the feature extends over at least a portion of a length of the cooling hole.
In addition to one or more of the features described above, or as an alternative, further embodiments may include that the feature includes rifling.
In some aspects of the disclosure, a method of cooling a component of a turbine engine is provided including providing air to a flow path defined by a cooling hole formed in the component. The air within the flow path is rotated about a central axis of the cooling hole.
In addition to one or more of the features described above, or as an alternative, further embodiments may include that the feature is integrally formed with the interior surface of the cooling hole via an additive manufacturing process.
The subject matter which is regarded as the present disclosure is particularly pointed out and distinctly claimed in the claims at the conclusion of the specification. 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:
Referring now to the FIGS., a cross-sectional view of an example of a gas turbine engine 10 is illustrated in
In the turbofan configuration of
In the two-spool, high bypass configuration of
Flow F at inlet 18 divides into primary (core) flow FP and secondary (bypass) flow FS downstream of fan rotor 26. Fan rotor 26 accelerates secondary flow FS through bypass duct 28, with fan exit guide vanes (FEGVs) 42 to reduce swirl and improve thrust performance. In some designs, structural guide vanes (SGVs) 42 are used, providing combined flow turning and load bearing capabilities.
Primary flow FP is compressed in low pressure compressor 30 and high pressure compressor 32, then mixed with fuel in combustor 14 and ignited to generate hot combustion gas. The combustion gas expands to provide rotational energy in high pressure turbine 34 and low pressure turbine 36, driving high pressure compressor 32 and low pressure compressor 30, respectively. Expanded combustion gases exit through exhaust section (or exhaust nozzle) 20, which can be shaped or actuated to regulate the exhaust flow and improve thrust performance.
Low pressure shaft 38 and high pressure shaft 40 are mounted coaxially about centerline CL, and rotate at different speeds. Fan rotor (or other propulsion stage) 26 is rotationally coupled to low pressure shaft 38. In advanced designs, fan drive gear system 44 is provided for additional fan speed control, improving thrust performance and efficiency with reduced noise output.
Fan rotor 26 may also function as a first-stage compressor for gas turbine engine 10, and LPC 30 may be configured as an intermediate compressor or booster. Alternatively, propulsion stage 26 has an open rotor design, or is absent, as described above. Gas turbine engine 10 thus encompasses a wide range of different shaft, spool and turbine engine configurations, including one, two and three-spool turboprop and (high or low bypass) turbofan engines, turboshaft engines, turbojet engines, and multi-spool industrial gas turbines.
In each of these applications, turbine efficiency and performance depend on the overall pressure ratio, defined by the total pressure at inlet 18 as compared to the exit pressure of compressor section 12, for example at the outlet of high pressure compressor 32, entering combustor 14. Higher pressure ratios, however, also result in greater gas path temperatures, increasing the cooling loads on rotor airfoils 22, stator airfoils 24 and other components of gas turbine engine 10. To reduce operating temperatures, increase service life and maintain engine efficiency, these components are provided with improved cooling configurations, as described below. Suitable components include, but are not limited to, cooled gas turbine engine components in compressor sections 30 and 32, combustor 14, turbine sections 34 and 36, and exhaust section 20 of gas turbine engine 10.
For improved service life and reliability, components of gas turbine engine 10 are provided with an improved cooling configuration, as described below. Suitable components for the cooling configuration include rotor airfoils 22, stator airfoils 24 and other gas turbine engine components exposed to hot gas flow, including, but not limited to, platforms, shrouds, casings and other endwall surfaces in hot sections of compressor 12 and turbine 16, and liners, nozzles, afterburners, augmentors and other gas wall components in combustor 14 and exhaust section 20.
Pressure and suction surfaces 53 and 54 form the major opposing surfaces or walls of airfoil 22, extending axially between leading edge 51 and trailing edge 52, and radially from root section 55, adjacent inner diameter (ID) platform 56, to tip section 57, opposite ID platform 56. In some designs, tip section 57 is shrouded.
Cooling holes or outlets 60 are provided on one or more surfaces of airfoil 22, for example along leading edge 51, trailing edge 52, pressure (or concave) surface 53, or suction (or convex) surface 54, or a combination thereof Cooling holes or passages 60 may also be provided on the endwall surfaces of airfoil 22, for example along ID platform 56, or on a shroud or engine casing adjacent tip section 57.
Cooling holes or outlets 60 are provided along one or more surfaces of airfoil 24, for example leading or trailing edge 61 or 62, pressure (concave) or suction (convex) surface 63 or 64, or a combination thereof Cooling holes or passages 60 may also be provided on the endwall surfaces of airfoil 24, for example along ID platform 66 and OD platform 68.
Rotor airfoils 22 (
While surface cooling extends service life and increases reliability, injecting cooling fluid into the gas path also reduces engine efficiency, and the cost in efficiency increases with the required cooling flow. Cooling holes 60 are thus provided with improved metering and inlet geometry to reduce jets and blow off, and improved diffusion and exit geometry to reduce flow separation and corner effects. Cooling holes 60 reduce flow requirements and improve the spread of cooling fluid across the hot outer surfaces of airfoils 22 and 24, and other gas turbine engine components, so that less flow is needed for cooling and efficiency is maintained or increased.
With reference now to
As shown in
In the illustrated, non-limiting embodiment, the feature 84 includes rifling formed about the interior surface 86 of a cooling hole 60. The rifling may be formed by adding material to the interior surface 86 of the cooling hole 60, such as via an additive manufacturing process for example. The rifling 84 may extend over only a portion of the cooling hole 60, such as only the metering section 80 as shown in
Rifling 84 usually includes a plurality of helical or spiral inward facing lands 88 with interleaved grooves 90. In the illustrated, non-limiting embodiment, the plurality of lands 88 extend inwardly from the surface 86 of the cooling hole 60 towards the central axis, and the plurality of grooves 90 are formed by the interior surface 86. By adding the thickness of the plurality of lands 88 to the interior surface 86 such that the lands 88 extend into the cooling hole 60, the lands 88 more effectively interrupt and swirl the cooling flow passing through the cooling hole 60 when compared to traditional rifling that is formed by removing material from the interior surface 86.
The surfaces of these lands 88 and grooves 90 commonly include a curvature complementary to the respective radius of the cooling hole 60. With reference now to
The cooling holes 60 described herein including a feature 84, such as rifling formed about the interior surface 86, provide a cooling solution that offers improved film cooling coverage and eliminates or reduces the problems associated with conventional cooling holes by increasing the movement of the air along the flow path defined by the cooling holes 60. As a result of this movement, the air provided at the outlet of the diffusion section 82 is better able to overcome the vortices around the cooling jets which are typically detrimental to the cooling air flow. Because a reduced portion of the cooling air flow is diverted away from the hot surface, the air is more effective at cooling an adjacent component of the turbine engine 10.
While the present disclosure has been described in detail in connection with only a limited number of embodiments, it should be readily understood that the present disclosure is not limited to such disclosed embodiments. Rather, the present disclosure can be modified to incorporate any number of variations, alterations, substitutions or equivalent arrangements not heretofore described, but which are commensurate with the spirit and scope of the present disclosure. Additionally, while various embodiments of the present disclosure have been described, it is to be understood that aspects of the present disclosure may include only some of the described embodiments. Accordingly, the present disclosure is not to be seen as limited by the foregoing description, but is only limited by the scope of the appended claims.
