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This process needs constant feedback between the different disciplines, and changes in one may lead to changes in another. Figure 1 shows a schematic diagram representing a general design procedure. This thesis will cover the thermodynamic, aerodynamic and a small part of the mechanical design. The full specification of the engine requirements can be found in the RFP [1]. 3.2 Thermodynamic analysis 3.2.1 Design point and off design performance calculations At the design point, the thermodynamic performance of the engine is evaluated at a fixed operating condition. In the case of the MJ – Haran S14 the operating condition chosen is supersonic cruise at M 1.5 at 51000ft, where the engine is expected to spend most of its time. Changing the operating condition in the design point stage would result in a different engine with a different geometry. This conditions used as a reference point. The next step is to run off design performance calculations. The engine geometry is now fixed and the performance of the engine at different operating conditions is evaluated. Many off design cases may be run to ensure the engine performance over the entire operating range. Both design point and off design performance calculations are highly iterative processes [11]. 3.2.2 Software description For the thermodynamic analysis of the engine, the Numerical Propulsion System Simulation (NPSS) code has been used. NPSS was developed by NASA in cooperation with U.S aerospace industries with the aim to develop a state of the art simulation tool for complex systems with an open architecture for user flexibility. The code is written in C++ and allows users to add their own unique objects and calculations. The program has a built in solver with the ability to make one parameter dependent on another, and hence varying the user input of these parameters to satisfy a certain condition specified by the user [12]. 3.2.3 Cycle modeling and optimization A replica of the baseline engine was modelled as a starting point for the cycle optimization. The cycle was then modified to achieve the required thrust levels at design point. The key cycle performance data can be seen in chapter 4. All comparisons herein are to the modified baseline engine. For the initial cycle optimization, turbomachinery efficiencies and pressure losses were assumed to be the same as for the baseline engine. 3.2.3.1 Cycle efficiency A number of different types of efficiencies can be used to characterize the performance of an aero- engine: Propulsive efficiency – the ratio of useful propulsive energy. This is the product of thrust and flight velocity divided by the wasted kinetic energy of the jet. Thermal efficiency – the increase of the kinetic energy of the gas stream divided by the energy of the fuel, which is the product of the fuel mass flow and the fuel LHV. For a turbofan engine, because of 17PDF Image | Conceptual Design of a Supersonic Jet Engine
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