Abstract
ABSTRACT
The combustion chamber of gas turbine unit is one of the most critical components to be designed. Scanning through literature reveals that the design methodologies for combustion chamber are available in a discrete manner and there exist a need to compile this information and evolve a systematic design procedure for combustion chamber. The present paper is an attempt towards presenting such a complete design methodology of combustion chamber for small gas turbine applications. The combustion chamber for the 20 kW gas turbine engine has been designed and fabricated as per these summarized design guidelines then checked for the axial and radial temperature profiles as well as liner wall temperatures, experimentally. It is worth to mention that the liner wall temperatures achieved is in the vicinity of 350°C when centerline temperature is of the order of 1300°C. The pressure loss is in the range of 10% of the delivery pressure which suggest the aerodynamic design of the combustion chamber.