A transition duct system (100) for routing a gas flow from a combustor (102) to the first stage (104) of a turbine section (106) in a combustion turbine engine (108), wherein the transition duct system (100) includes one or more converging flow joint inserts (120) forming a trailing edge (122) at an intersection (124) between adjacent transition ducts (126, 128) is disclosed. The transition duct system (100) may include a transition duct (126, 128) having an internal passage (130) extending between an inlet (132, 184) to an outlet (134, 186) and may expel gases into the first stage turbine (104) with a tangential component. The converging flow joint insert (120) may be contained within a converging flow joint insert receiver (136) and disconnected from the transition duct bodies (126, 128) by which the converging flow joint insert (120) is positioned. Being disconnected eliminates stress formation within the converging flow joint insert (120), thereby enhancing the life of the insert. The converging flow joint insert (120) may be removable such that the insert (120) can be replaced once worn beyond design limits.

A turbine stator vane assembly comprises a plurality of circumferentially arranged struts and a plurality of circumferentially arranged stator vanes. Each stator vane has a leading edge, a trailing edge and a chord length. Each strut has a leading edge, a trailing edge, a chord length, a pressure surface and a suction surface. The chord length of the struts is greater than the chord length of the stator vanes. The stator vanes are arranged in groups between adjacent pairs of struts and each group of stator vanes comprises a plurality of stator vanes. The circumferential spacing between adjacent struts is substantially the same and each stator vane in a group of stator vanes has a different chord length to the other stator vanes in that group of stator vanes.

An injector for a liquid rocket engine includes an array of injector elements. Each injector element includes a central passage and a plurality of peripheral transverse passages. The central passages are configured to provide axial injection flow and the peripheral transverse passages are configured to provide swirl injection flow. A portion of the injector elements provide the swirl injection flow in a clockwise direction and another portion of the injector elements provide the swirl injection flow in a counter-clockwise direction. The injector elements are arranged with a single, center injector element and a plurality of circumferential rows disposed around the single, center injector element. The injector elements of each one of the circumferential rows are either all of the clockwise direction or all of the counter-clockwise direction, and from the single, center injector element moving radially outwardly the circumferential rows alternate between the clockwise direction and the counter-clockwise direction.

A circumferential row of vanes for a gas turbine engine, the circumferential row of vanes has non-uniform spacing. The circumferential row of vanes includes a plurality of stator vanes arranged circumferentially about an inner ring. The plurality of stator vanes include a first group of stator vanes having a first spacing between adjacent stator vanes of the first group of stator vanes and a second group of stator vanes having a second spacing between adjacent stator vanes of the second group of stator vanes. The second spacing is from two percent to eleven percent greater than a nominal uniform vane spacing or from two percent to eleven percent lesser than the nominal uniform vane spacing, the nominal uniform vane spacing being defined by a total number of the plurality of stator vanes. An engine includes the circumferential row of vanes.

An atmospheric re-entry vehicle includes a main body defining a forward end of the vehicle, a base defining an aft end of the vehicle, and a heat shield at the base. The heat shield includes a heat shield outer surface. The main body and the heat shield are configured such that a centerline of the heat shield is offset relative to a centerline of the main body. At least a portion of the heat shield outer surface is at least substantially axisymmetric relative to the centerline of the heat shield.

A turbofan engine defining a circumferential direction is provided. The turbofan engine includes: a turbomachine, the turbomachine defining a pylon attachment location along the circumferential direction; an unducted rotor assembly drivingly coupled to the turbomachine, the unducted rotor assembly including a plurality of unducted rotor blades; and a plurality of outlet guide vanes positioned downstream of the plurality of unducted rotor blades, the plurality of outlet guide vanes including N.sub.OGV number of outlet guide vanes, the plurality of outlet guide vanes including a first outlet guide vane and a second outlet guide vane adjacent the first outlet guide vane, the first and second outlet guide vanes defining a gap spacing greater than 360 degrees divided by N.sub.OGV, and the pylon attachment location positioned outside of between the first and second outlet guide vanes along the circumferential direction.