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A gas turbine engine (10), including: a turbine having radial inflow impellor blades (38); and an array of advanced transition combustor assemblies arranged circumferentially about the radial inflow impellor blades (38) and having inner surfaces (34) that are adjacent to combustion gases (40). The i

A gas turbine engine (10), including: a turbine having radial inflow impellor blades (38); and an array of advanced transition combustor assemblies arranged circumferentially about the radial inflow impellor blades (38) and having inner surfaces (34) that are adjacent to combustion gases (40). The inner surfaces (34) of the array are configured to accelerate and orient, for delivery directly onto the radial inflow impellor blades (38), a plurality of discrete flows of the combustion gases (40). The array inner surfaces (34) define respective combustion gas flow axes (20). Each combustion gas flow axis (20) is straight from a point of ignition until no longer bound by the array inner surfaces (34), and each combustion gas flow axis (20) intersects a unique location on a circumference defined by a sweep of the radial inflow impellor blades (38).

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1. A gas turbine engine, comprising: a turbine comprising radial inflow impellor blades; andan array of transition combustor assemblies arranged circumferentially about the radial inflow impellor blades and comprising inner surfaces that are adjacent to combustion gases, wherein the inner surfaces o

1. A gas turbine engine, comprising: a turbine comprising radial inflow impellor blades; andan array of transition combustor assemblies arranged circumferentially about the radial inflow impellor blades and comprising inner surfaces that are adjacent to combustion gases, wherein the inner surfaces of the array are configured to accelerate and orient, for delivery directly onto the radial inflow impellor blades, a plurality of discrete flows of the combustion gases;wherein the array inner surfaces define respective combustion gas flow axes, wherein each combustion gas flow axis is straight from a point of ignition until reaching the radial inflow impellor blades, and wherein each combustion gas flow axis intersects a unique location on a circumference defined by a sweep of the radial inflow impellor blades. 2. The gas turbine engine of claim 1, wherein the array inner surfaces define a collimating length in each transition combustor assembly effective to collimate the combustion gases downstream of an acceleration region with respect to the respective combustion gas flow axis. 3. The gas turbine engine of claim 1, wherein cross sections of an inner surface of a respective transition combustor assembly, taken perpendicular to a longitudinal axis of the respective transition combustor assembly, narrow in a downstream direction with respect to a gas turbine engine longitudinal axis, and this narrowing is at least in part effective to accelerate the combustion gases. 4. The gas turbine engine of claim 1, wherein array inner surfaces leading to adjacent sides of adjacent outlets converge to form an aerodynamically sharp edge between the adjacent outlets. 5. The gas turbine engine of claim 1, wherein each transition combustor assembly comprises a geometry at downstream end with respect to a combustion gas flow axis that interlocks with a geometry at a downstream end with respect to a combustion gas flow axis of circumferentially adjacent transition combustor assemblies, thereby interlocking the transition combustor assemblies into the array. 6. The gas turbine engine of claim 4, wherein inner surfaces of two adjacent transition combustor assemblies define each combustion gas flow axis. 7. The gas turbine engine of claim 5, wherein an inner surface of a selected transition combustor assembly and an inner surface of the interlocking downstream end geometry of an adjacent transition combustor assembly define a respective combustion gas flow axis. 8. The gas turbine engine of claim 7, wherein the downstream end geometry of the adjacent transition combustor assembly comprises a hook that secures the adjacent transition combustor assembly to the selected transition combustor assembly. 9. The gas turbine engine of claim 6, comprising an aerodynamically smooth joint where the inner surfaces of two circumferentially adjacent transition combustor assemblies meet. 10. The gas turbine engine of claim 9, wherein the aerodynamically smooth joint is disposed at a location along a respective combustion gas flow longitudinal axis where the array inner surfaces only partially surround the combustion gas flow axis. 11. The gas turbine engine of claim 1, wherein the array is secured at only a forward side or an aft side. 12. The gas turbine engine of claim 11, wherein the aft side of the array is secured to a gas turbine engine casing. 13. A gas turbine engine comprising: a turbine comprising a plurality of radial inflow impellor blades; anda combustor assembly comprising a combustor and a transition, the combustor assembly configured to accelerate and orient combustion gases directly onto the radial inflow impellor blades;wherein the combustion gases travel along an axially straight flow path within the combustor assembly from the combustor until reaching the radial inflow impellor blades. 14. The gas turbine engine of claim 13, wherein each flow path narrows along a respective transition longitudinal axis. 15. The gas turbine engine of claim 14, wherein each flow path comprises: a fully bounded length immediately downstream of the narrowing with respect to a flow of the combustion gases, wherein a respective flow path is bounded on all sides; anda partially bounded length downstream of the fully bounded length, wherein a respective flow path transitions from being bounded on all sides at an upstream end to being gradually less bound toward a downstream end. 16. The gas turbine engine of claim 15, wherein a downstream end of a first partially bounded length defines a beginning of an upstream end of a partially bounded length of an adjacent flow path disposed downstream with respect to a direction of rotation of the radial inflow impellor blades. 17. The gas turbine engine of claim 15, wherein each flow path within the partially bounded length comprises aerodynamically smooth walls where bounded. 18. The gas turbine engine of claim 13, wherein an aft side of the combustor assembly is secured to a turbine section casing and a forward side is free to thermally expand and contract with respect to engine components adjacent the forward side. 19. A gas turbine engine, comprising: a turbine comprising radial inflow impellor blades, wherein a sweep of rotating tips of the impellor blades defines an annular swept shape;a plurality of advanced transition combustor assemblies arranged circumferentially in an array about the annular swept shape, the array comprising inner surfaces adjacent to combustion gases that define a plurality of combustion gas flow paths, wherein the array inner surfaces are configured to narrow each gas flow path in a downstream direction with respect to a flow of combustion gases therein effective to accelerate the combustion gases for delivery directly to the radial inflow impellor blades, and to define a respective gas flow path longitudinal axis that is straight from a point of ignition of the combustion gases until the combustion gases are not bound by any array inner surface, wherein each gas flow path longitudinal axis is within a plane perpendicular to a gas turbine engine longitudinal axis, and wherein the inner surfaces direct the combustion gasses directly onto the annular swept shape;wherein each advanced transition combustor assembly comprises an interlocking hook feature at downstream end with respect to the respective gas flow path longitudinal axis that interlocks with an adjacent advanced transition combustor assembly effective to interlock the advanced transition combustor assemblies into the array, wherein an inner surface of the interlocking hook feature and an inner surface of the adjacent advanced transition combustor assembly define a respective combustion gas flow path, andwherein an aft side of the array is secured to a turbine section casing and a forward side is free to thermally expand and contract with respect to engine components adjacent the forward side.