Perpustakaan Digital ITB

With the rise of commercial space launching companies and the commercialization of space, cost effective measure are in pursue to make it as affordable as possible. One of which are designing a rocket that is efficient throughout the flight path and could handle all conditions without being too complex. This could be achieved with the use of altitude compensating nozzle such as aerospike nozzle. Aerospike nozzle is a nozzle that utilizing the atmospheric pressure to develop the flow by utilizing flow that expands after the nozzle throat. These expansion fans are controlled by the nozzle with a ramp geometry by the use of Angelino’s theory to reach the desired exit velocity based on the throat conditions and would have the same exit velocity throughout the flight path in an under-expanded conditions. The biggest weakness of the nozzle is heat management as the spike tapered off along the length of the nozzle, this could be mitigated by truncating the nozzle at a various percentage of area in the design. By truncating the spike, the expansion fan would have a different exit velocity with the same throat properties. Thus, the design of this study will follow the fixed exit mach number of 2.4 and the throat condition would follow those criteria. The difference in throat condition would be mitigated by the use of specific impulse as a point of comparisons between the designs as it accounts for the mass flow rate on the inlet. The design and analytical thrust calculation would be aided by MATLAB and the result of those geometry would be simulated in ANSYS environment by using ICEM CFD, and ANSYS CFX. It can be concluded that by choosing an exit mach number as a design point, the thrust and throat conditions would vary depending on the design and it only can be compared if the comparison is using specific impulse as the comparative variable. From the truncation variations also concluded that, a large truncation decreases thrust and specific impulse significantly at on-design conditions and more capable to harness more thrust at higher pressure ratio and truncation is only viable in the smaller range (Below 50%) with the on-design conditions.