Abstract: Embodiments of a combustor for a gas turbine engine are provided herein. In some embodiments, a combustion chamber for a gas turbine engine comprising may include a combustor having an inner volume defined at least partially by a front wall, wherein the wall comprises a plurality of facets each having a through hole fluidly coupled to the inner volume, and wherein the plurality of facets are oriented such that an axis of each of the plurality of facets is offset from a central axis of the combustor by an angle.

Abstract: The present disclosure is directed to a rotor blade having a passive airflow modifying assembly to create an airflow feature along the blade, based on the instant pressure gradient around the blade during operation. The present disclosure also is directed to a rotor blade that passively channels airflow through the passive airflow modifying assembly to create an air feature that decreases the aerodynamic load, at times when the aerodynamic load experienced by the blade is bearing on the rotatable hub, and one the passively channels airflow through the passive airflow modifying assembly to create an air feature that increases the aerodynamic load, at times when the aerodynamic load is not bearing on the rotatable hub, and one that passively operates to not create an air feature, at times when the requisite pressure gradient is not met and/or when the load conditions are not an issue.

Abstract: A gas turbine engine includes: a compressor section including a compressor mean radius; a combustor section fluidly coupled downstream of the compressor section and include a combustor mean radius; and a turbine section fluidly coupled downstream of the combustor section and a turbine mid-span radius. The combustor mean radius is greater than each of the compressor mean radius and the turbine mid-span radius.

Abstract: A gas turbine engine includes: a compressor section including a compressor mean radius; a combustor section fluidly coupled downstream of the compressor section and include a combustor mean radius; and a turbine section fluidly coupled downstream of the combustor section and a turbine mid-span radius. The combustor mean radius is greater than each of the compressor mean radius and the turbine mid-span radius.

Abstract: A gas turbine engine includes: a compressor section including a compressor mean radius; a combustor section fluidly coupled downstream of the compressor section and include a combustor mean radius; and a turbine section fluidly coupled downstream of the combustor section and a turbine mid-span radius. The combustor mean radius is greater than each of the compressor mean radius and the turbine mid-span radius.

Abstract: Embodiments of a combustor for a gas turbine engine are provided herein. In some embodiments, a combustion chamber for a gas turbine engine comprising may include a combustor having an inner volume defined at least partially by a front wall, wherein the wall comprises a plurality of facets each having a through hole fluidly coupled to the inner volume, and wherein the plurality of facets are oriented such that an axis of each of the plurality of facets is offset from a central axis of the combustor by an angle.

Abstract: Embodiments of a combustor for a gas turbine engine are provided herein. In some embodiments, a combustion chamber for a gas turbine engine comprising may include a combustor having an inner volume defined at least partially by a front wall, wherein the wall comprises a plurality of facets each having a through hole fluidly coupled to the inner volume, and wherein the plurality of facets are oriented such that an axis of each of the plurality of facets is offset from a central axis of the combustor by an angle.

Abstract: A gas turbine engine includes: a compressor section including a compressor mean radius; a combustor section fluidly coupled downstream of the compressor section and include a combustor mean radius; and a turbine section fluidly coupled downstream of the combustor section and a turbine mid-span radius. The combustor mean radius is greater than each of the compressor mean radius and the turbine mid-span radius.

Abstract: A transverse fan propulsion system includes a first transverse fan configured to rotate in a first direction, and a second transverse fan configured to rotate in a second direction. The second direction is opposite to the first direction. The second transverse fan is located coaxially with and radially inward of the first transverse fan.

Abstract: Embodiments of a combustor for a gas turbine engine are provided herein. In some embodiments, a combustion chamber for a gas turbine engine comprising may include a combustor having an inner volume defined at least partially by a front wall, wherein the wall comprises a plurality of facets each having a through hole fluidly coupled to the inner volume, and wherein the plurality of facets are oriented such that an axis of each of the plurality of facets is offset from a central axis of the combustor by an angle.

Abstract: A turbine engine includes a rotating detonation combustor including a housing defining at least one combustion chamber. The rotating detonation combustor is configured for a rotating detonation process to occur within the at least one combustion chamber to generate a combustion flow including a first portion and a second portion. The turbine engine also includes a turbine coupled in flow communication with the rotating detonation combustor. The turbine is configured to receive the combustion flow from the rotating detonation combustor. The turbine includes a first blade and a second blade that rotate about an axis at a rotational frequency. The rotating detonation combustor and the turbine are configured for the combustion flow first portion to contact the first blade substantially continuously as the first blade rotates.

Abstract: A turbine engine assembly including a plurality of rotating detonation combustors configured for a rotating detonation process to occur to produce a flow of combustion gas. The plurality of rotating detonation combustors are oriented such that the flow of combustion gas discharged therefrom flows helically relative to a centerline of the turbine engine assembly. The assembly also includes a turbine coupled downstream from the plurality of rotating detonation combustors. The turbine is configured to receive the flow of combustion gas.

Abstract: A turbine engine is provided. The turbine engine includes a compressor configured to discharge a flow of air at a first flow angle and a combustor coupled downstream from the compressor. The combustor is configured to combust the flow of air with fuel to form a flow of combustion gas discharged from the combustor at a second flow angle. The turbine engine also includes a turbine coupled downstream from the combustor. The turbine includes an inlet configured to receive the flow of combustion gas having a flow angle within a predetermined range, wherein the combustor is oriented such that the second flow angle is within the predetermined range.