Abstract: A gas turbine engine includes a fan located at a forward portion of the gas turbine engine, and a compressor section and a turbine section arranged in serial flow order. The compressor section and the turbine section together define a core airflow path. A rotary member is rotatable with the fan and with a low pressure turbine of the turbine section. The low pressure turbine includes a rotating drum to which a first airfoil structure is connected and extends radially inward toward the rotary member. A torque frame connects the rotating drum to the rotary member and transfers torque from the first airfoil structure mounted to the rotating drum to the rotary member. The torque frame includes an inner disk mounted to the rotary member, an outer ring and a second airfoil structure formed separately from the outer ring and connected thereto by a releasable connecting structure. The second airfoil structure extends radially inward from the outer ring toward the inner disk.

Abstract: A turbomachine engine including a high-pressure compressor, a high-pressure turbine, a combustion chamber in flow communication with the high-pressure compressor and the high-pressure turbine, and a power turbine in flow communication with the high-pressure turbine. At least one of the high-pressure compressor, the high-pressure turbine, and the power turbine comprises a ceramic matrix composite (CMC) material. The turbomachine engine includes a low-pressure shaft coupled to the power turbine and characterized by a midshaft rating (MSR) between two hundred (ft/sec)1/2 and three hundred (ft/sec)1/2. The low-pressure shaft has a redline speed between fifty and two hundred fifty feet per second (ft/sec). The turbomachine engine is configured to operate up to the redline speed without passing through a critical speed associated with a first-order bending mode of the low-pressure shaft.

Abstract: A rotor blade assembly for a turbine engine, including an airfoil blade including an inner diameter end and an outer diameter end, a lower blade carrier coupled to the inner diameter end of the airfoil blade and rigidly coupled to a disk via a pin, an upper blade carrier coupled to the outer diameter end of the airfoil blade, and an outer drum coupled to the upper blade carrier via a radial joint. The radial joint supports radial motion of the upper blade carrier relative to an axis extending through a center of the rotor blade assembly.

Abstract: A turbine engine having a drive shaft rotatable about an axis, a multi-stage compressor, a turbine section, a thrust bearing, and a balance cavity. The thrust bearing being provided between the drive shaft and at least a portion of the multi-stage compressor section and rotationally supporting the drive shaft. During operation of the turbine engine, a first axial force is applied to the thrust bearing by the drive shaft and a second axial force is applied to the thrust bearing in an opposite direction of the first axial force by the balance cavity.

Abstract: A turbomachine engine includes an engine core including a high-pressure compressor, a high-pressure turbine, and a combustion chamber. The engine core has a length (LCORE), and the high-pressure compressor has an exit stage diameter (DCORE). A power turbine is in flow communication with the high-pressure turbine. A low-pressure shaft is coupled to the power turbine and characterized by a midshaft rating from one hundred fifty (ft/sec)1/2 to three hundred thirty (ft/sec)1/2. The low-pressure shaft has a redline speed between fifty and two hundred fifty feet per second. The turbomachine engine is configured to operate up to the redline speed without passing through a critical speed associated with a first-order bending mode of the low-pressure shaft. A high-pressure shaft is coupled to the high-pressure compressor and the high-pressure turbine and is characterized by a high-speed shaft rating from 1.5 to 6.2, and a ratio of LCORE/DCORE is from 2.1 to 4.3.

Abstract: A turbine containment system is provided. The turbine containment system includes a first containment member surrounding a portion of a turbine including a plurality of rotor disks and rotor blades; and a second containment member in communication with the first containment member, wherein the first containment member and the second containment member together contain each of the rotor disks and the rotor blades therein.

Abstract: An unducted single fan engine includes a housing having one or more fan blades coupled to the housing and configured to rotate circumferentially. The engine has one or more outlet guide vanes coupled to the housing. Each of the one or more guide vanes has a leading edge portion having a variable leading edge. The engine has one or more actuation devices coupled to each of the one or more outlet guide vanes. The one or more actuation devices are configured to control the variable leading edge of the respective outlet guide vane. The variable leading edge is controllable to vary the pitch, camber, lean angle, or sweep of the respective outlet guide vane.

Abstract: A turbine engine having a drive shaft rotatable about an axis, a multi-stage compressor, a turbine section, a thrust bearing, and a balance cavity. The thrust bearing being provided between the drive shaft and at least a portion of the multi-stage compressor section and rotationally supporting the drive shaft. During operation of the turbine engine, a first axial force is applied to the thrust bearing by the drive shaft and a second axial force is applied to the thrust bearing in an opposite direction of the first axial force by the balance cavity.

Abstract: A turbomachine engine including a high-pressure compressor, a high-pressure turbine, a combustion chamber in flow communication with the high-pressure compressor and the high-pressure turbine, and a power turbine in flow communication with the high-pressure turbine. At least one of the high-pressure compressor, the high-pressure turbine, and the power turbine comprises a ceramic matrix composite (CMC) material. The turbomachine engine includes a low-pressure shaft coupled to the power turbine and characterized by a midshaft rating (MSR) between two hundred (ft/sec)1/2 and three hundred (ft/sec)1/2. The low-pressure shaft has a redline speed between fifty and two hundred fifty feet per second (ft/sec). The turbomachine engine is configured to operate up to the redline speed without passing through a critical speed associated with a first-order bending mode of the low-pressure shaft.

Abstract: An unducted single fan engine includes a housing having one or more fan blades coupled to the housing and configured to rotate circumferentially. The engine has one or more outlet guide vanes coupled to the housing. Each of the one or more guide vanes has a leading edge portion having a variable leading edge. The engine has one or more actuation devices coupled to each of the one or more outlet guide vanes. The one or more actuation devices are configured to control the variable leading edge of the respective outlet guide vane. The variable leading edge is controllable to vary the pitch, camber, lean angle, or sweep of the respective outlet guide vane.

Abstract: A turbomachine engine including a high-pressure compressor, a high-pressure turbine, a combustion chamber in flow communication with the high-pressure compressor and the high-pressure turbine, and a power turbine in flow communication with the high-pressure turbine. At least one of the high-pressure compressor, the high-pressure turbine, and the power turbine comprises a ceramic matrix composite (CMC) material. The turbomachine engine includes a low-pressure shaft coupled to the power turbine and characterized by a midshaft rating (MSR) between two hundred (ft/sec)1/2 and three hundred (ft/sec)1/2. The low-pressure shaft has a redline speed between fifty and two hundred fifty feet per second (ft/sec). The turbomachine engine is configured to operate up to the redline speed without passing through a critical speed associated with a first-order bending mode of the low-pressure shaft.

Abstract: Structures, such as compressor casings, for reducing a thermal gradient are provided. For example, a compressor case is split such that it includes first and second case segments. The first case segment extends over a first portion of the compressor case circumference and comprises a first inner structure, a first outer structure, and a first porous structure integrally formed as a monolithic component. The first porous structure is defined between the first inner structure and the first outer structure. The second case segment extends over a second portion of the compressor case circumference and comprises a second inner structure, a second outer structure, and a second porous structure integrally formed as a monolithic component. The second porous structure is defined between the second inner structure and the second outer structure. Methods of cooling structures such as compressor casings also are provided.

Abstract: A turbomachine engine including an engine core including a high-pressure compressor, which has an exit stage having an exit stage diameter (DCORE), a high-pressure turbine, and a combustion chamber in flow communication with the high-pressure compressor and the high-pressure turbine, a power turbine in flow communication with the high-pressure turbine, and a low-pressure shaft coupled to the power turbine and characterized by a midshaft rating (MSR) between two hundred (ft/sec)1/2 and three hundred (ft/sec)1/2. The low-pressure shaft has a redline speed between fifty and two hundred fifty feet per second (ft/sec). The turbomachine engine is configured to operate up to the redline speed without passing through a critical speed associated with a first-order bending mode of the low-pressure shaft. The low-pressure shaft has a length (LMSR) defined by an engine core length (LCORE) given by: LCORE=[m(20+m)*n(10+n)](1/100)*DCORE+CIS.

Abstract: Methods, apparatus, systems and articles of manufacture for compact compressors are disclosed including a gas turbine engine defining an axial direction and a radial direction, the gas turbine engine including an axial flow compressor and a radial flow compressor, wherein the axial flow compressor is located axially forward of the radial flow compressor, a blade assembly including a splitter shroud to divide incoming air into axial air flow for the axial flow compressor and radial air flow for the radial flow compressor, the blade assembly rotating relative to the axial flow compressor and counter-rotating relative to the radial flow compressor, and wherein the blade assembly is located axially aft of the radial flow compressor.

Abstract: A passive clearance control limits thermal expansion between stator components relative to rotor components. A control ring controls clearance in a passive manner and is located on or adjacent to stationary components which thermally expand during engine operation. The control ring is formed of material having low coefficient of thermal expansion such as CMCs (Ceramic Matrix Composites) and therefore limits, inhibits or restrains expansion of the adjacent stator components as temperatures increase. Limiting expansion of the stator component reduces rotor/stator clearances and limits parasitic leakage of fluid along the flow path through the engine core.

Abstract: A turbine containment system is provided. The turbine containment system includes a first containment member surrounding a portion of a turbine including a plurality of rotor disks and rotor blades; and a second containment member in communication with the first containment member, wherein the first containment member and the second containment member together contain each of the rotor disks and the rotor blades therein.

Abstract: Structures, such as compressor casings, for reducing a thermal gradient are provided. For example, a compressor case is split such that it includes first and second case segments. The first case segment extends over a first portion of the compressor case circumference and comprises a first inner structure, a first outer structure, and a first porous structure integrally formed as a monolithic component. The first porous structure is defined between the first inner structure and the first outer structure. The second case segment extends over a second portion of the compressor case circumference and comprises a second inner structure, a second outer structure, and a second porous structure integrally formed as a monolithic component. The second porous structure is defined between the second inner structure and the second outer structure. Methods of cooling structures such as compressor casings also are provided.

Abstract: A gas turbine engine having an interdigitated turbine assembly including a first turbine rotor and a second turbine rotor, wherein a total number of stages at the interdigitated turbine assembly is between 3 and 8, and an average stage pressure ratio at the interdigitated turbine assembly is between 1.3 and 1.9. A gear assembly is configured to receive power from the interdigitated turbine assembly, and a fan assembly is configured to receive power from the gear assembly. The interdigitated turbine assembly and the gear assembly are together configured to allow the second turbine rotor to rotate at a second rotational speed greater than a first rotational speed at the first turbine rotor. The fan assembly and the gear assembly are together configured to allow the fan assembly to rotate at a third rotational speed less than the first rotational speed and the second rotational speed.

Abstract: Methods, apparatus, systems and articles of manufacture for compact compressors are disclosed including a gas turbine engine defining an axial direction and a radial direction, the gas turbine engine including an axial flow compressor and a radial flow compressor, wherein the axial flow compressor is located axially forward of the radial flow compressor, a blade assembly including a splitter shroud to divide incoming air into axial air flow for the axial flow compressor and radial air flow for the radial flow compressor, the blade assembly rotating relative to the axial flow compressor and counter-rotating relative to the radial flow compressor, and wherein the blade assembly is located axially aft of the radial flow compressor.

Abstract: A passive clearance control limits thermal expansion between stator components relative to rotor components. A control ring controls clearance in a passive manner and is located on or adjacent to stationary components which thermally expand during engine operation. The control ring is formed of material having low coefficient of thermal expansion such as CMCs (Ceramic Matrix Composites) and therefore limits, inhibits or restrains expansion of the adjacent stator components as temperatures increase. Limiting expansion of the stator component reduces rotor/stator clearances and limits parasitic leakage of fluid along the flowpath through the engine core.