Radial inflow gas turbine engine with advanced transition duct

A gas turbine engine comprising:

• a turbine comprising a plurality of radial inflow impellor blades;
• and
• a 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.

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, and
• wherein 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.