In the quest for achieving high performance, gas turbine engines demand efficient design of various engine components, mainly the compressor stages. The compressor stages consume most of the energy produced by the engine to provide the required pressure ratio. CSIR-NAL is involved in the development of a small gas turbine engine for UAV applications. In this regard, a high transonic single stage axial flow compressor is designed with a mass flow of 4.6 kg/s and pressure ratio of 1.6, for technology demonstration. In this paper, the aerodynamic and structural design of a high transonic axial compressor stage is discussed along with its performance characteristics. Preliminary mean-line design of the compressor stage is carried out, followed by detailed 3D blade design. Aerodynamic performance of the compressor stage is investigated numerically. Grid independency study is carried out, and the flow un-altering grid is used for steady simulations. Steady 3D RANS CFD simulations with SST turbulence model are carried out for estimating the compressor stage performance. At the design speed, the compressor is able to produce the desired pressure ratio and efficiency. Detailed flow investigations across the compressor stage are studied from choke to near stall flow conditions for different speeds. The compressor rotor blisk made of titanium alloy (Ti6AL4V) is subjected to stress analysis. The von-Mises stress and radial deformation are observed to be well within the safe limits of the chosen material. Modal analysis is carried out to study the structural dynamics of the rotor.