Patent Application
20240044288
This disclosure relates generally to a gas turbine engine and, more particularly, to a fluid system for the gas turbine engine with a takeoff port.
A fluid system for a gas turbine engine may bleed air from a flowpath within the gas turbine engine for use with one or more components. Various types and configurations of engine fluid systems are known in the art. While these known engine fluid systems have various advantages, there is still room in the art for improvement.
According to an aspect of the present disclosure, a system is provided for a gas turbine engine. This engine system includes a flowpath wall, a takeoff conduit and a porous cover. The flowpath wall forms a peripheral boundary of an internal engine flowpath. The flowpath wall includes a takeoff port. The takeoff conduit includes an internal conduit passage fluidly coupled with the internal engine flowpath through the takeoff port. The takeoff conduit projects out from the flowpath wall. The porous cover for the internal conduit passage is disposed at the takeoff port.
According to another aspect of the present disclosure, another system is provided for a gas turbine engine. This engine system includes a flowpath wall, a takeoff conduit and a porous cover. The flowpath wall forms a peripheral boundary of an internal engine flowpath. The flowpath wall includes a takeoff port. The takeoff conduit includes an internal conduit passage fluidly coupled with the internal engine flowpath through the takeoff port. The porous cover extends across the takeoff port. The porous cover is configured from or otherwise includes mesh.
According to still another aspect of the present disclosure, another system is provided for a gas turbine engine. This engine system includes a flowpath wall, a takeoff conduit and a porous cover. The flowpath wall forms a peripheral boundary of an internal engine flowpath. The flowpath wall includes a takeoff port with a cross-sectional geometry having one of a circular shape or a non-circular shape with a minor axis dimension and a major axis dimension that is less than five times the minor axis dimension. The takeoff conduit includes an internal conduit passage fluidly coupled with the internal engine flowpath through the takeoff port. The porous cover for the internal conduit passage is disposed at the takeoff port.
The mesh may include wire with a diameter of between 0.030 inches and 0.040 inches. The mesh may also or alternatively include a percentage of open area between sixty percent and sixty-five percent.
The porous cover may be configured to alter a shear layer region of air flowing through the internal engine flowpath at the takeoff port.
The porous cover may be disposed within the takeoff port.
The porous cover may be disposed within the internal engine flowpath adjacent the takeoff port.
The porous cover may be disposed within the internal conduit passage adjacent the takeoff port.
The porous cover may extend across the takeoff port.
The porous cover may be configured as a single layer of porous material.
The porous cover may be configured as or otherwise include mesh.
The mesh may include a mesh element with a diameter of between 0.025 inches and 0.045 inches.
The porous cover may be configured as or otherwise include a perforated plate.
The porous cover may have a percentage of open area between thirty percent and forty-five percent.
The porous cover may have a percentage of open area between forty-five percent and sixty percent.
The porous cover may have a percentage of open area between sixty percent and seventy-five percent.
The takeoff port may have a cross-sectional geometry with a circular shape. The takeoff port may alternatively have a cross-sectional geometry with a non-circular shape with a first dimension and a second dimension angularly offset from the first dimension. The second dimension may be less than five times the first dimension.
The engine system may also include an engine component and a flow regulator.
The flow regulator may be fluidly coupled between the internal conduit passage and the engine component. The flow regulator may be configured to regulate a flow of gas bled from the internal engine flowpath through the takeoff port and directed to the engine component.
The engine system may also include a gas turbine engine core. The internal engine flowpath may be configured as or otherwise include a bypass flowpath that bypasses the gas turbine engine core.
The engine system may also include a gas turbine engine core. The internal engine flowpath may be configured as or otherwise include a core flowpath that extends within the gas turbine engine core.
The present disclosure may include any one or more of the individual features disclosed above and/or below alone or in any combination thereof.
The foregoing features and the operation of the invention will become more apparent in light of the following description and the accompanying drawings.
The engine structure 22 partially or completely forms an internal engine flowpath 26 within the gas turbine engine. This engine flowpath 26 extends longitudinally along a longitudinal centerline 28 within (e.g., through) the gas turbine engine and its engine structure 22, which longitudinal centerline 28 may also be a rotational axis for one or more rotating structures (e.g., spools) within the gas turbine engine. The engine flowpath 26 may extend circumferentially about (e.g., completely around) the longitudinal centerline 28 providing the engine flowpath 26 with, for example, an annular cross-sectional geometry. Examples of the engine flowpath 26 include, but are not limited to, a bypass flowpath which bypasses a core of the gas turbine engine, and a core flowpath which extends within (e.g., through) the gas turbine engine core.
The engine structure 22 includes a flowpath wall 30 such as, but not limited to, a duct wall, an engine casing, a shroud, a platform or a liner. The flowpath wall 30 forms a peripheral boundary of the engine flowpath 26 within the gas turbine engine and its engine structure 22. The peripheral boundary of
The flowpath wall 30 includes a takeoff port 36 such as an air bleed port or any other through-aperture. This takeoff port 36 is disposed (e.g., intermediately) along a longitudinal length of the engine flowpath 26. The takeoff port 36 extends vertically through the flowpath wall 30 (e.g., along a centerline 38 of the takeoff port 36) between and to the wall interior side 32 and the wall exterior side 34. The takeoff port 36 is thereby vertically adjacent and fluidly coupled with the engine flowpath 26.
Referring to
The fluid system 24 of
The takeoff conduit 44 has a tubular sidewall 50 that forms an internal conduit passage 52 (e.g., an inner bore) within the takeoff conduit 44. This conduit passage 52 is fluidly coupled with the takeoff port 36. The conduit passage 52 is thereby fluidly coupled with the engine flowpath 26 through the takeoff port 36. The takeoff conduit 44 and its sidewall 50 project (e.g., vertically) out from the flowpath wall 30. The takeoff conduit 44 and its sidewall 50 may also be formed integral with or otherwise connected (e.g., mechanically fastened, bonded, etc.) to the flowpath wall 30. The conduit passage 52 may thereby be disposed adjacent the takeoff port 36, for example, without any other volumes (e.g., a plenum, etc.) between the conduit passage 52 and the takeoff port 36. The takeoff port 36, for example, may form an inlet to/of the conduit passage 52. The present disclosure, however, is not limited to such an exemplary relationship/fluid coupling between the takeoff port 36 and the conduit passage 52.
The flow regulator 46 is fluidly coupled with and (e.g., inline) between the takeoff conduit 44 and its conduit passage 52 and the engine component 48. The flow regulator 46 is configured to regulate a flow of the takeoff gas directed (e.g., bled) out of the engine flowpath 26 through the takeoff port 36 and directed to the engine component 48. During a first mode of operation, the flow regulator 46 may close and thereby fluidly decouple the conduit passage 52 from the engine component 48. During a second mode of operation, the flow regulator 46 may open and thereby fluidly couple the conduit passage 52 with the engine component 48.
During the first mode of operation when the flow regulator 46 closes and fluidly decouples the conduit passage 52 from the engine component 48, the conduit passage 52 may become a closed-ended passage; e.g., a blind passage, a deadheaded passage, etc. Under certain conditions, this closed-ended passage may operate as an acoustic resonance chamber. For example, while the takeoff gas may no longer flow through the conduit passage 52 to the engine component 48, the gas flowing within the engine flowpath 26 still flows across the open takeoff port 36. The gas flow across the takeoff port 36 may be a source for flow instabilities. These flow instabilities may excite an intake structure (e.g., an adjacent portion of the flowpath wall 30, the takeoff conduit 44, etc.) causing vibrations and/or sound; e.g., noise. The vibrations and/or the sound may result from a self-sustained flow oscillation at the takeoff port 36 (e.g., the inlet to the conduit passage 52) where a dominant vortex mode may lock onto intake structure modes and/or acoustic mode—resonance. As an analogy, consider a whistling sound generated by directing an airflow across an opening of an empty bottle.
Referring to
The porous cover 54 of
In some embodiments, referring to
Referring to
The porous cover 54 includes a plurality of pores 60 such as perforations and/or any other apertures. Each of the pores 60 may extend vertically through the porous cover 54 and its layer of material from the cover interior side 56 to the cover exterior side 58. For example, referring to
A quantity of the pores 60, dimensions of the pores 60 and/or a density of the pores are selected to provide the porous cover 54 with a percentage of open area (POA). This percentage of open area may describe a ratio between a total cross-sectional area of all of the pores to a total cross-sectional area of the porous cover 54. The percentage of open area of the porous cover 54 (e.g., the mesh cover or the perforated plate cover) may be between thirty percent (30%) and forty-five percent (45%), between forty-five percent (45%) and sixty percent (60%), or between sixty percent (60%) and seventy-five percent (75%). In particular, where the porous cover 54 comprises the mesh 62 (e.g., with a mesh element diameter of between 0.030 inches (0.0762 centimeters) and 0.040 inches (0.1016 centimeters)), the percentage of open area may be between sixty percent (60%) and sixty-five percent (65%); e.g., exactly or about sixty-three percent (63%). The present disclosure, however, is not limited to such exemplary arrangements.
The fan section 80 includes a fan rotor 84. The compressor section 81 includes a compressor rotor 85. The turbine section 83 includes a high pressure turbine (HPT) rotor 86 and a low pressure turbine (LPT) rotor 87, where the LPT rotor 87 is configured as a power turbine rotor. Each of these rotors 84-87 includes a plurality of rotor blades arranged circumferentially around and connected to one or more respective rotor disks.
The fan rotor 84 is connected to the LPT rotor 87 through a low speed shaft 90. The compressor rotor 85 is connected to the HPT rotor 86 through a high speed shaft 92. The low speed shaft 90 extends through a bore of the high speed shaft 92 between the fan rotor 84 and the LPT rotor 87.
During operation, air enters the gas turbine engine 72 through the airflow inlet 76. This air is directed through the fan section 80 and into a core flowpath 94 and a bypass flowpath 96, where either the core flowpath 94 or the bypass flowpath 96 may be or otherwise include the engine flowpath 26. The core flowpath 94 extends sequentially through the engine sections 81-83; e.g., a core of the gas turbine engine 72. The air within the core flowpath 94 may be referred to as “core air”. The bypass flowpath 96 extends through a bypass duct, which bypasses the engine core. The air within the bypass flowpath 96 may be referred to as “bypass air”.
The core air is compressed by the compressor rotor 85 and directed into a (e.g., annular) combustion chamber 98 of a (e.g., annular) combustor 100 in the combustor section 82. Fuel is injected into the combustion chamber 98 via one or more of the fuel injectors 102 and mixed with the compressed core air to provide a fuel-air mixture. This fuel-air mixture is ignited and combustion products thereof flow through and sequentially cause the HPT rotor 86 and the LPT rotor 87 to rotate. The rotation of the HPT rotor 86 drives rotation of the compressor rotor 85 and, thus, compression of air received from an inlet into the core flowpath 94. The rotation of the LPT rotor 87 drives rotation of the fan rotor 84, which propels bypass air through and out of the bypass flowpath 96. The propulsion of the bypass air may account for a significant portion (e.g., a majority) of thrust generated by the turbine engine.
The engine system 20 may be configured with various gas turbine engines other than the one described above. The engine system 20, for example, may be configured with a geared gas turbine engine where a geartrain connects one or more shafts to one or more rotors in a fan section, a compressor section and/or any other engine section. Alternatively, the engine system 20 may be configured with a gas turbine engine configured without a geartrain. The engine system 20 may be configured with a geared or non-geared gas turbine engine configured with a single spool, with two spools (e.g., see
While various embodiments of the present disclosure have been described, it will be apparent to those of ordinary skill in the art that many more embodiments and implementations are possible within the scope of the disclosure. For example, the present disclosure as described herein includes several aspects and embodiments that include particular features. Although these features may be described individually, it is within the scope of the present disclosure that some or all of these features may be combined with any one of the aspects and remain within the scope of the disclosure. Accordingly, the present disclosure is not to be restricted except in light of the attached claims and their equivalents.
