Abstract: Disclosed is a tangential on-board injector (TOBI) system that includes an annulus and a plurality of cooling airflow passages disposed about the annulus. Each cooling airflow passage of the plurality of cooling airflow passages includes an inlet opening having a polygonal inlet cross-section, the inlet opening having an inlet cross-sectional area. Each cooling airflow passage of the plurality of cooling airflow passages further includes an outlet opening having an outlet cross-section and an outlet cross-sectional area. The inlet cross-sectional area is greater in magnitude than the outlet cross-sectional area. Also disclosed are additive manufacturing methods for manufacturing the tangential on-board injector system and gas turbine engines that incorporate the tangential on-board injector system.

Abstract: Disclosed is a tangential on-board injector (TOBI) system that includes an annulus and a plurality of cooling airflow passages disposed about the annulus. Each cooling airflow passage of the plurality of cooling airflow passages includes an inlet opening having a polygonal inlet cross-section, the inlet opening having an inlet cross-sectional area. Each cooling airflow passage of the plurality of cooling airflow passages further includes an outlet opening having an outlet cross-section and an outlet cross-sectional area. The inlet cross-sectional area is greater in magnitude than the outlet cross-sectional area. Also disclosed are additive manufacturing methods for manufacturing the tangential on-board injector system and gas turbine engines that incorporate the tangential on-board injector system.

Abstract: Embodiments of an impeller shroud support for disposition around an impeller are provided, as are embodiments of gas turbine engine including impeller shroud supports. In one embodiment, the impeller shroud support includes a shroud body, a support arm joined to and extending around the shroud body, and a plurality of Mid-Impeller Bleed (MIB) flow passages. Each MIB flow passage includes, in turn, an inlet formed in the shroud body and configured to receive bleed air extracted from the impeller, a throat portion, an outlet formed in the support arm and through which the bleed air is discharged, and a curved intermediate section between the inlet and the outlet. During usage of the impeller shroud support, the curved intermediate section turns the bleed air flowing through the MIB passage in a radially outward direction prior to discharge from the outlet of the MIB flow passage.

Abstract: A system and method is provided that combines fan bypass components and minimizes assembly interfaces in a turbofan engine. The system and method provide a front frame structure of reduced weight that slidably installs/removes from within the combined fan bypass components.

Abstract: A combustor is provided for a turbine engine. The combustor includes a first liner having a first side and a second side and a second liner having a first side and a second side. The second side of the second liner forms a combustion chamber with the second side of the first liner, and the combustion chamber is configured to receive an air-fuel mixture for combustion therein. The first liner defines a plurality of effusion cooling holes configured to form a film of cooling air on the second side of the first liner. The plurality of effusion cooling holes including a first effusion cooling hole extending from the first side to the second side with a non-linear line of sight.

Abstract: Embodiments of an impeller shroud support for disposition around an impeller are provided, as are embodiments of gas turbine engine including impeller shroud supports. In one embodiment, the impeller shroud support includes a shroud body, a support arm joined to and extending around the shroud body, and a plurality of Mid-Impeller Bleed (MIB) flow passages. Each MIB flow passage includes, in turn, an inlet formed in the shroud body and configured to receive bleed air extracted from the impeller, a throat portion, an outlet formed in the support arm and through which the bleed air is discharged, and a curved intermediate section between the inlet and the outlet. During usage of the impeller shroud support, the curved intermediate section turns the bleed air flowing through the MIB passage in a radially outward direction prior to discharge from the outlet of the MIB flow passage.

Abstract: A combustor for a gas turbine engine is provided. The combustor includes an annular inner liner; an annular outer liner circumscribing the annular inner liner; and a combustor dome having a first edge coupled to the annular inner liner and a second edge coupled to the annular outer liner, the combustor dome forming a combustion chamber with the annular inner liner and the annular outer liner. The combustion chamber accommodates fluid flow through the annular inner and annular outer liners. The combustion chamber converges in the direction of the air flow to reduce a diameter of the combustion chamber. The combustor dome is configured to bifurcate the air flow at the combustor dome into a first stream directed to the annular inner liner and a second stream directed to the annular outer liner.

Abstract: A system and method is provided that combines fan bypass components and minimizes assembly interfaces in a turbofan engine. The system and method provide a front frame structure of reduced weight that slidably installs/removes from within the combined fan bypass components.

Abstract: A system and method of controlling turbine blade tip-to-static structure clearance in a gas turbine engine includes processing aircraft data to determine when the aircraft is flying at altitude cruise conditions. When the aircraft is not flying at altitude cruise conditions, then compressor discharge air is directed to impinge upon at least a portion of the turbine shroud. Upon determining that the aircraft is flying at altitude cruise conditions, the compressor discharge air is directed to flow across at least a portion of the turbine case and to impinge upon at least a portion of the turbine shroud.

Abstract: A system and method of controlling turbine blade tip-to-static structure clearance in a gas turbine engine includes processing aircraft data to determine when the aircraft is flying at altitude cruise conditions. When the aircraft is not flying at altitude cruise conditions, then compressor discharge air is directed to impinge upon at least a portion of the turbine shroud. Upon determining that the aircraft is flying at altitude cruise conditions, the compressor discharge air is directed to flow across at least a portion of the turbine case and to impinge upon at least a portion of the turbine shroud.

Abstract: A combustor for a gas turbine engine is provided. The combustor includes an annular inner liner; an annular outer liner circumscribing the annular inner liner; and a combustor dome having a first edge coupled to the annular inner liner and a second edge coupled to the annular outer liner, the combustor dome forming a combustion chamber with the annular inner liner and the annular outer liner. The combustion chamber accommodates fluid flow through the annular inner and annular outer liners. The combustion chamber converges in the direction of the air flow to reduce a diameter of the combustion chamber. The combustor dome is configured to bifurcate the air flow at the combustor dome into a first stream directed to the annular inner liner and a second stream directed to the annular outer liner.

Abstract: An improved type of cooled component is disclosed. The cooled components, such as turbine airfoils and combustors, have a base metal structural member, including a metal sheet or blank, with a featured surface. The features, such as fins, pins, and other protrusions, as well as divots, indentions, slots, channels, and holes, are formed by lithographic processes. The protrusions can be formed from a different material than the metal sheet or blank, preferably one with superior heat-conductive or oxidation properties. A related method of manufacturing is also disclosed.

Abstract: Bonded assemblies and methods for improving the bond strength of a joint are provided. The joint may be formed between bonding surfaces of a first and a second component. The first component has a bonding surface adapted to be bonded to a bonding surface of a second component. A plurality of features may be formed in at least one of the bonding surfaces. The bonding surfaces are bonded together to form the joint. A bonding material layer between the bonding surfaces may form the joint or the first and second components may be co-cured to form the joint. A chemical and mechanical bond between the bonding surfaces is formed with the plurality of features forming the mechanical bond and increasing the bond area of the chemical bond. The features may be openings and/or protrusions such as holes, slots, and bosses.

Abstract: An improved type of cooled component is disclosed. The cooled components, such as turbine airfoils and combustors, have a base metal structural member, including a metal sheet or blank, with a featured surface. The features, such as fins, pins, and other protrusions, as well as divots, indentions, slots, channels, and holes, are formed by lithographic processes. The protrusions can be formed from a different material than the metal sheet or blank, preferably one with superior heat-conductive or oxidation properties. A related method of manufacturing is also disclosed.