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3.2 Methods for Sizing and Performance of Hybrid-Electric Aircraft 47 The flight performance analysis as represented in Figure 3.2 is composed of two modules, the integrated mission performance module and the point performance module. The integrated mission performance analysis consists of determining the amount of energy an aircraft must carry on-board to accomplish its design mission. The calculation includes the energy require- ment for the mission as well as for the reserves to comply with operational regulations. The required thrust is computed by resolving the flight mechanics equations at each discretized point of the mission segments according to the flight state π₯πΌππ΄, π΄ππ‘ and π, the aircraft gross-weight and the aerodynamic efficiency delivered by the drag polar. With respect to the flight state and the computed thrust, the value of the πΉπΉ is calculated by interpolating in the energy table according to the rating corresponding to the segment of the mission. The computation of the πΉ πΉ is performed at each discretized point of the segment. The integration of the πΉ πΉ over time enables the computation of the fuel mass consumed and correspondingly the change in aircraft gross-weight over time. Concurrently, the electric power required to produce the thrust, ππΈπππ, is interpolated in the energy table at each discretized point of the segment according to the flight state. The required electrical power ππΈπππ is the input for the electric energy and power device model. ππΈπππ defines the required electric power profile. In the case, of an electrical system powered by fuel cells, the required πΉπΉπΉπΆ is determined according to ππΈπππ and the computed fuel cell efficiency ππΉ πΆ as formulated in Equation 3.6. The πΉ πΉπΉ πΆ is integrated over time to determine the fuel mass consumed and the change in aircraft gross-weight along the mission. If the EEPD is a battery system, the available output voltage at the required discharge rate is calculated using the discharge curve of the battery model according to ππΈπππ and the SOC of the battery as highlighted in Section 3.2.3.6.1. According to the discharge rate fulfilling the power requirements, the SOC is calculated at each discretized point of the mission segment. As a result, the SOC of the battery at the end of the mission can be determined. The know- ledge of the battery SOC at the end of the mission enables the sizing of the battery mass according to the energy requirement as described in Section 3.2.7. In addition, the maximal electrical power occurring in the profile, πππ₯(ππΈπππ), is assessed in order to determine the sizing criterion of the battery mass. 3.2.6 Methods for Hybrid-Electric Propulsion System Sizing Freeman et al. [18] highlighted that the sizing of hybrid-electric propulsion system should pro- ceed from the propulsive device back to the electric energy and power device. It is explained by the fact that the sizing of the upstream components (close to the production of thrust) influences the power required of all-downstream elements according to the efficiency value chain. The same strategy is followed in this thesis for the sizing of the hybrid-electric propul- sion systems. The sizing characteristics of the propulsive device and of the electric motor vary according to the investigated hybrid-electric propulsion system topology. Consequently, the sizing specificities are discussed with respect to the topologies investigated.PDF Image | Conceptual Design Methods Hybrid-Electric Transport Aircraft
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