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52 3 Methods for Sizing and Performance of Hybrid-Electric Aircraft design mission is determined by the two main following parameters: the design payload and the design range. According to an initial estimated MTOW and with respect to initial val- ues for the main sizing parameters of the aircraft namely ππππ/ππππ and ππ‘ππ‘ππ/ππππ, the total power required ππ‘ππ‘ππ and the wing reference area ππππ are determined. With the specification of a degree of hybridization for power π»π , PSLS and ππΈπ,πππ₯ are computed so that the total aircraft power requirement ππ‘ππ‘ππ is fulfilled as detailed in Section 3.2.6.2. As illustrated in Figure A.1, in case of a partial parallel hybrid-electric propulsion system, instead of expressing the power-to-weight ratio ππ‘ππ‘ππ/ππππ, a more practical approach is to specify the ratio of π/ππππ. Similarly, the total design thrust requirement ππππ is calculated according to an initial value of MTOW. With the formulation of a degree of hy- bridization for useful power π»ππ’π π, the design thrust of the conventional system ππππ ,ππΉ and of the electrical system ππππ ,πΈπΉ is computed as demonstrated in Section 3.2.6.1. The sizing characteristics of the propulsive device were discussed for a serial, a parallel and a partial parallel hybrid system in Section 3.2.6.1. The particularities of the electric motor sizing were formulated in Section 3.2.6.2. The geometrical properties of the propulsion sys- tem are linked to the drag component build-up functions to determine the complete drag polar of the aircraft. The overall weight of the electrical propulsion system is determined by summing up the weight of the electrical components. The electrical propulsion system weight is mapped as an additional item to the propulsion weight of the aircraft. According to the hybrid-electric propulsion system topology selected and with respect to the design and off-design performance characteristics of the components, the thrust tables and the energy tables are computed according to the different mission segment as detailed in Section 3.2.4.1 and Section 3.2.4.3 respectively. By interpolating within the energy table according to the flight state and the thrust, the πΉπΉ is calculated. The fuel mass required is calculated by integrating the πΉ πΉ along the integrated mission. Concurrently, the electric power required at the electric energy and power device, ππΈπππ, is interpolated within the energy table. ππΈπππ serves as input for the sizing of the electric energy and power device. If a fuel cell system is implemented as electric energy and power device, the number of fuel cells required is sized according to the maximum electric power occurring within the flight envelope as highlighted in Section 3.2.3.6.2. According to ππΈπππ, the efficiency of the fuel cell system is evaluated based for instance on the physics-based model of a fuel cell as discussed in Section 3.2.3.6.2. The determination of the fuel cell efficiency enables the computation of the πΉ πΉπΉ πΆ according to Equation 3.6. By integrating the πΉ πΉπΉ πΆ of the fuel cell over time, the total fuel mass supply of the fuel cell system to provide the electric energy and power over the mission is determined. Utilizing batteries as an energy and power device, the total battery system mass required to provide the electric energy and power has to be determined. According to the battery model, the SOC is computed along the mission as detailed in Section 3.2.5. To protect the battery from any damage and to prolong design service goal suitable for use in aerospace, the battery must not be discharged below πππΆπππππ‘ typically set at 20%. This is represented generically in Figure 3.19 for the mission power profile discussed in Figure 3.18.PDF Image | Conceptual Design Methods Hybrid-Electric Transport Aircraft
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