This paper presents the development and aerothermal investigation of the integrated combustor vane concept for power generation gas turbines with individual can combustors. In this novel concept, first introduced in 2010, the conventional nozzle guide vanes (NGVs) are removed and flow turning is achieved by vanes that extend the combustor walls. The concept was developed using the in-house computational fluid dyanamics (CFD) code TBLOCK. Aerothermal experiments were conducted using a modular high-speed linear cascade, designed to model the flow at the combustor–vane interface. The facility is comprised of two can combustor transition ducts and either four conventional vanes (CVs) or two integrated vanes (IVs). The experimental study validates the linear CFD simulations of the IV development. Annular full-stage CFD simulations, used to evaluate aerodynamics, heat transfer, and stage efficiency, confirm the trends of the linear numerical and experimental results, and thus demonstrate the concept's potential for real gas turbine applications. Results show a reduction of the total pressure loss coefficient at the exit of the stator vanes by more than 25% due to a reduction in profile and endwall loss. Combined with an improved rotor performance demonstrated by unsteady stage simulations, these aerodynamic benefits result in a gain in stage efficiency of above 1%. A distinct reduction in heat transfer coefficient (HTC) levels on vane surfaces, on the order of 25–50%, and endwalls is observed and attributed to an altered state of boundary layer (BL) thickness. The development of IV's endwall- and leading edge (LE)-cooling geometry shows a superior surface coverage of cooling effectiveness, and the cooling requirements for the first vane are expected to be halved. Moreover, by halving the number of vanes, simplifying the design and eliminating the need for vane LE film cooling, manufacturing and development costs can be significantly reduced.