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3.2 Methods for Sizing and Performance of Hybrid-Electric Aircraft 39 3.2.4 Methods for Propulsion System Integration In the previous section, the component models constituting the hybrid-electric propulsion system were considered independent of each other. According to a given hybrid-electric topo- logy, the component characteristics need to be linked together to interface the performance of the propulsion system to the aircraft design environment. The calculation of the max- imum thrust available and the determination of the energy consumed during the mission are the main aspects that the methods have to address as detailed in Section 3.2.4.1 and Section 3.2.4.3 respectively. The methods are described for the serial, parallel and and partial parallel hybrid topologies. 3.2.4.1 Thrust Characteristics The determination of the maximum available thrust is the focus of the mathematical relations developed in the following paragraphs. According to the characteristics of the hybrid-electric propulsion system, the maximum thrust is computed at each flight state (π₯πΌππ΄, π΄ππ‘ and π) for the different ratings MTO, MCL, MCR and MCO. The commonality with a tradi- tional sizing and performance program is provided by the utilization of an identical thrust table format with π₯πΌππ΄, π΄ππ‘ and π as inputs and ππππ₯ as output. This table enables the interfacing of the propulsion system thrust characteristics with the flight performance mod- ule as illustrated in Figure 3.2. The calculation of the maximum thrust for parallel, serial and partial parallel hybrid-electric propulsion system is detailed in Section 3.2.4.1.1, Section 3.2.4.1.2 and Section 3.2.4.1.3 respectively. 3.2.4.1.1 Parallel Hybrid System The equations provided for the parallel system were published by Pornet et al. [27]. The formulation of the mathematical relations is based upon the generic parallel propulsion system illustrated in Figure 3.4. For a hybrid-electric parallel system, the maximum thrust is determined by the sum of the maximum shaft power delivered by the conventional system (ππhπππ‘,πΆπππ£,πππ₯) and the maximum shaft power delivered by the electrical system (ππhπππ‘,πΈπππ,πππ₯) available at the shaft of the propulsive device at each flight state as established in Equation 3.8, ππππ₯ = πππ· Β· ππhπππ‘,πΆπππ£,πππ₯ + ππhπππ‘,πΈπππ,πππ₯ (3.8) πΒ·π where πππ· is computed according to the propulsive device performance map described in Section 3.2.3.2. ππhπππ‘,πΆπππ£,πππ₯ is determined by Equation 3.9. The power lever of the combustion engine (ππΏπΆπΈ) determines the power setting of the combustion engine according to the maximum power available, ππΆπΈ,πππ₯ as expressed in Equation 3.10. ππΏπΆπΈ is introduced in the equation in view of conducting energy management strategy to adapt the utilization of the combustion engine according to the segment of the mission. In addition, the power off-take at the combustion engine (ππΆπΈ,ππππ‘πππ) is taken into account in the determination ofPDF Image | Conceptual Design Methods Hybrid-Electric Transport Aircraft
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