An engine system that produces all required thrust for an aerospace vehicle from takeoff through space operation, consisting of airbreathing and non-airbreathing propulsion apparatuses. The airbreathing system consists of a turbine engine, a ram and or scram jet, and the non-airbreathing system is a unique liquid rocket motor. The turbine engine may consist of a turbojet or turbofan configuration. The air breathing turbine, ramjet and scram jet feature a single air inlet system, and combustion fuel. The non-airbreathing rocket system includes separate oxidizer system, and either a separate or same source of combustion fuel as the turbine. Airflow velocities in the turbine bypass duct, and burner system, include subsonic and supersonic velocities for ramjet or scramjet operation. The rocket engine may utilize either cryogenic or a non-cryogenic fuel and oxidizer system.

A film cooled component may comprise a cooling chamber and a first ligament centered about a first axis. The first ligament may be in fluid communication with the cooling chamber. A first meter may be disposed at an end of the first ligament. A first diffuser may extend from the first meter to a surface of the film cooled component. The first diffuser may comprise a first tapered sidewall oriented at an angle of between 5 degrees to 15 degrees relative to the first axis. The first diffuser may further comprise a first non-tapered sidewall oriented at an angle less than 5 degrees relative to the first axis.

An internally cooled component of a gas turbine engine is provided. The component may include a cooling passage at least partially defined by a first wall and a second wall with a first pedestal extending from the first wall to the second wall. The first pedestal may have a chevron geometry. A second pedestal may extend from the first wall to the second wall and also have a chevron geometry. A gap may be defined by the first pedestal and the second pedestal with the gap oriented between the first pedestal and the second pedestal.

A shroud assembly for a turbine engine comprising a plurality of circumferentially arranged shroud segments having confronting end faces defining first and second radially spaced surfaces. The shroud assembly includes a forward edge spanning to an aft edge to define an axial direction and a set of confronting seal channels formed in each of the confronting end faces with a spline seal located within the confronting seal channels.

A shroud assembly for a turbine engine comprising a plurality of circumferentially arranged shroud segments having confronting end faces defining first and second radially spaced surfaces. The shroud assembly includes a forward edge spanning to an aft edge to define an axial direction and a set of confronting seal channels formed in each of the confronting end faces with a spline seal located within the confronting seal channels.

A shroud assembly for a turbine engine comprising a plurality of circumferentially arranged shroud segments having confronting end faces defining first and second radially spaced surfaces. The shroud assembly includes a forward edge spanning to an aft edge to define an axial direction and a set of confronting seal channels formed in each of the confronting end faces with a spline seal located within the confronting seal channels.

A shroud assembly for a turbine engine comprising a plurality of circumferentially arranged shroud segments having confronting end faces defining first and second radially spaced surfaces. The shroud assembly includes a forward edge spanning to an aft edge to define an axial direction and a set of confronting seal channels formed in each of the confronting end faces with a spline seal located within the confronting seal channels.

A flight vehicle has a propulsion system that includes an air inlet, an isolator (or diffuser) downstream of the air inlet, and a combustor downstream of the isolator. The isolator includes an obstruction that protrudes inwardly from an inner wall of the isolator, into the flow channel in which air flows through the isolator. The obstruction diverts the flow to either side of it. Downstream of the obstruction the flow on either side of the obstruction comes together again, leading to mixing of the flow, for example including mixing of low energy and boundary layer flow with high energy flow. This mixing of flow may make for a more uniform flow at the exit of the isolator. In addition the obstruction may help fix the location of shocks within the isolator, providing longer flow mixing length in the isolator.

In some embodiments, a plenum of an apparatus for transferring energy between a rotating element and a fluid may include a through hole disposed through the plenum; a plurality of inlet guide vanes disposed proximate a peripheral edge of the through hole, the plurality of inlet guide vanes comprising a first group of inlet guide vanes having a symmetrical profile, a second group of inlet guide vanes, and a third group of inlet guide vanes, wherein each inlet guide vane of the second group and third group have a cambered profile, wherein each inlet guide vane of the second group has same cambered profile, and further wherein each inlet guide vane of the third group has a different cambered profile from each other inlet guide vane of the third group.

A stator ring for an aircraft turbine engine includes an inner shroud coaxial with an outer shroud. The shrouds are connected to each other by vanes that are each fully integral with the shrouds. The outer shroud has an outer annular surface connected to at least one catch for attaching the stator ring. The inner shroud has an inner annular surface connected to a support member with an abradable coating. At least one of the inner and outer surfaces includes recesses that are situated in line with the vanes and are configured such that the vanes are connected to the corresponding shroud.