An airfoil for an aircraft engine is provided. The airfoil may include a plurality of composite plies extending in a chordwise direction from a leading edge to a trailing edge and in a spanwise direction between an airfoil tip and an airfoil base. The airfoil may include at least one discontinuous ply having a first ply segment, a second ply segment, and a butt joint disposed between the first ply segment and the second ply segment. The butt joint may be disposed along a first fragment profile to frangibly attach the first and second ply segments.

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. The interdigitated turbine assembly, the gear assembly, and the fan assembly together have a maximum AN.sup.2 at the second turbine rotor between 30 and 90.

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. The interdigitated turbine assembly, the gear assembly, and the fan assembly together have a maximum AN.sup.2 at the second turbine rotor between 30 and 90.

The gas turbine power generation system includes a core engine and a low pressure compressor. The core engine includes a high pressure compressor, a combustor, and a high pressure turbine configured in a serial flow arrangement. The high pressure compressor and the high pressure turbine are coupled together by a first shaft. The low pressure compressor is positioned axially forward of the core engine and is coupled to the first shaft.

The gas turbine power generation system includes a core engine and a low pressure compressor. The core engine includes a high pressure compressor, a combustor, and a high pressure turbine configured in a serial flow arrangement. The high pressure compressor and the high pressure turbine are coupled together by a first shaft. The low pressure compressor is positioned axially forward of the core engine and is coupled to the first shaft.

The present disclosure is directed to a method of operating a combustion system to attenuate combustion dynamics. The method includes flowing, via a compressor section, an overall supply of air to the combustion system; flowing, via a fuel supply system, an overall flow of fuel to the combustion system; flowing, to a first fuel nozzle of the combustion system, a first supply of fuel defining a richer burning fuel-air mixture at the first fuel nozzle; flowing, to a second fuel nozzle of the combustion system, a second supply of fuel defining a leaner burning fuel-air mixture at the second fuel nozzle; and igniting the richer burning fuel-air mixture and the leaner burning fuel-air mixture to produce an overall fuel-air ratio at a combustion chamber of the combustion system.

The present disclosure is directed to a method of operating a combustion system to attenuate combustion dynamics. The method includes flowing, via a compressor section, an overall supply of air to the combustion system; flowing, via a fuel supply system, an overall flow of fuel to the combustion system; flowing, to a first fuel nozzle of the combustion system, a first supply of fuel defining a richer burning fuel-air mixture at the first fuel nozzle; flowing, to a second fuel nozzle of the combustion system, a second supply of fuel defining a leaner burning fuel-air mixture at the second fuel nozzle; and igniting the richer burning fuel-air mixture and the leaner burning fuel-air mixture to produce an overall fuel-air ratio at a combustion chamber of the combustion system.

A compound engine assembly with at least one rotary internal combustion engine, an impulse turbine, and an exhaust pipe for each internal combustion engine providing fluid communication between the exhaust port of the respective internal combustion engine and the flow path of the turbine. Each exhaust pipe terminates in a nozzle. For each exhaust pipe, a ratio Vp/Vd between the pipe volume Vp and the displacement volume Vd of the respective internal combustion engine is at most 1.5. A minimum value of a cross-sectional area of each exhaust pipe is defined at the nozzle. In one embodiment, a ratio An/Ae between the minimum cross-sectional area An and the cross-sectional area Ae of the exhaust port of the respective internal combustion engine is at least 0.2. A method of compounding at least one rotary engine is also discussed.

A compound engine assembly with at least one rotary internal combustion engine, an impulse turbine, and an exhaust pipe for each internal combustion engine providing fluid communication between the exhaust port of the respective internal combustion engine and the flow path of the turbine. Each exhaust pipe terminates in a nozzle. For each exhaust pipe, a ratio Vp/Vd between the pipe volume Vp and the displacement volume Vd of the respective internal combustion engine is at most 1.5. A minimum value of a cross-sectional area of each exhaust pipe is defined at the nozzle. In one embodiment, a ratio An/Ae between the minimum cross-sectional area An and the cross-sectional area Ae of the exhaust port of the respective internal combustion engine is at least 0.2. A method of compounding at least one rotary engine is also discussed.

A particle separator includes a separator body in a primary fluid passageway of a machine. The primary fluid passageway includes one or more bleed holes through which a diverted portion of the fluid flowing in the primary fluid passageway toward a volume of the machine is diverted into an auxiliary flow passageway that bypasses the volume and directs the diverted portion of the fluid toward one or more other components of the machine. The separator body is coupled with the inner wall and/or outer wall of the primary fluid passageway. The separator body includes an upstream edge positioned to separate at least some particles carried by the fluid from the fluid as the diverted portion of the fluid bends around and flows over the at least one upstream edge of the separator body and into the auxiliary flow passageway.