**4.2. Aircraft retrofit with hybrid-electric propulsion system**

Retrofitting an existing aircraft with a hybrid-electric propulsion system enables having first insights about design parameters and constraints, the behavior of the propulsion system as well as in establishing the right interfaces between the propulsion system and any other aircraft systems. An under-wing podded twin engine narrow-body transport aircraft retrofitted with a hybrid-electric propulsion was investigated by Pornet et al. [21]. The retroï ˇn ˛Atting of the propulsion system consists of mounting an electric motor powered bybatteries in parallel to the low-pressure shaft of the gas-turbine to supportits operation during some segments of the mission. Due to the intrinsic nature of an aircraft retrofit, the certified Maximum Take-Off Weight (MTOW) of the baseline aircraft represents a design constraint not to be exceeded. The electrical system components including the batteries represent additional weight items. To provide the installation of the electrical system at aircraft level and still respect the MTOW limitation, the hybrid-electric propulsion system is utilized in off-design operations. As the baseline aircraft takes-off at a Take-Off Weight (TOW) lower than MTOW since less fuel is required to perform the off-design stage length, the delta weight available between TOW and MTOW makes the installation the electrical system possible. Sizing of the hybrid-electric propulsion system is the result of an interplay between the maximum power of the installed electric motor, which mainly determines the mass of the electrical system, the total battery mass required, the potential fuel burnreduction and the battery State-Of-Charge (SOC). The SOC characterizes the amount of energy available relative to the total energy of the battery. In order 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 a certain SOC limit typically set at 20%. The design parameters and constraints are illustrated in the design chart in Figure 4.

It represents the sizing characteristics of the hybrid-electric propulsion system for the case of utilizing the electric motor during cruise only assuming a battery speciï ˇn ˛Ac energy of 1500 Wh/kg at cell-level. The objective of this assessment was to achieve minimum fuel consumption at a stage length corresponding to peak in the utilization spectrum. For the baseline aircraft, the maximum utilization stage length was found to be 900 nm (1667 km). The total installed maximum power of the electric motor P*maxEM*,*totalinst*. is varied between 4 to 8 MW and the power setting of the electric motor during cruise (*P*/*Pmax*)*EM*,*cruise* between 0% to 100%. To maximize the number of installed batteries, the TOW of the hybrid-electric aircraft was set equal to MTOW. The total mass of the installed battery was consequently an outcome of the sizing process. When the electric motor is not used during cruise as indicated by the power setting of 0%, the hybrid-electric aircraft consumes more fuel than the baseline aircraft due to the higher aircraft weight. When increasing the power setting of theelectric motor, more electric energy is consumed increasing the potential block fuel reduction. According to the electric maximum power installed, the battery SOC can become

**Figure 4.** Design chart for hybrid-electric propulsion system in cruise [21]

Air Range (COSAR) was published by Pornet et al. [66]. The cost of the energy is not the only factor contributing to the total operating cost of an aircraft. Fixed costs and time dependent costs need to be also taken into account. Interested in minimizing the overall cost, airlines base their aircraft fleet operation on so-called Cost-Index which relates basically the cost of time to the cost of energy. A review of the Cost-Index traditionally used for fuel-based aircraft and the establishment of Cost-Index metric for hybrid-energy aircraft are found in

Retrofitting an existing aircraft with a hybrid-electric propulsion system enables having first insights about design parameters and constraints, the behavior of the propulsion system as well as in establishing the right interfaces between the propulsion system and any other aircraft systems. An under-wing podded twin engine narrow-body transport aircraft retrofitted with a hybrid-electric propulsion was investigated by Pornet et al. [21]. The retroï ˇn ˛Atting of the propulsion system consists of mounting an electric motor powered bybatteries in parallel to the low-pressure shaft of the gas-turbine to supportits operation during some segments of the mission. Due to the intrinsic nature of an aircraft retrofit, the certified Maximum Take-Off Weight (MTOW) of the baseline aircraft represents a design constraint not to be exceeded. The electrical system components including the batteries represent additional weight items. To provide the installation of the electrical system at aircraft level and still respect the MTOW limitation, the hybrid-electric propulsion system is utilized in off-design operations. As the baseline aircraft takes-off at a Take-Off Weight (TOW) lower than MTOW since less fuel is required to perform the off-design stage length, the delta weight available between TOW and MTOW makes the installation the electrical system possible. Sizing of the hybrid-electric propulsion system is the result of an interplay between the maximum power of the installed electric motor, which mainly determines the mass of the electrical system, the total battery mass required, the potential fuel burnreduction and the battery State-Of-Charge (SOC). The SOC characterizes the amount of energy available relative to the total energy of the battery. In order 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 a certain SOC limit typically set at 20%. The design parameters and

It represents the sizing characteristics of the hybrid-electric propulsion system for the case of utilizing the electric motor during cruise only assuming a battery speciï ˇn ˛Ac energy of 1500 Wh/kg at cell-level. The objective of this assessment was to achieve minimum fuel consumption at a stage length corresponding to peak in the utilization spectrum. For the baseline aircraft, the maximum utilization stage length was found to be 900 nm (1667 km). The total installed maximum power of the electric motor P*maxEM*,*totalinst*. is varied between 4 to 8 MW and the power setting of the electric motor during cruise (*P*/*Pmax*)*EM*,*cruise* between 0% to 100%. To maximize the number of installed batteries, the TOW of the hybrid-electric aircraft was set equal to MTOW. The total mass of the installed battery was consequently an outcome of the sizing process. When the electric motor is not used during cruise as indicated by the power setting of 0%, the hybrid-electric aircraft consumes more fuel than the baseline aircraft due to the higher aircraft weight. When increasing the power setting of theelectric motor, more electric energy is consumed increasing the potential block fuel reduction. According to the electric maximum power installed, the battery SOC can become

**4.2. Aircraft retrofit with hybrid-electric propulsion system**

constraints are illustrated in the design chart in Figure 4.

[66].

126 New Applications of Electric Drives

a limiting factor. The optimum sizing of the hybrid-electric propulsion system in view of achieving minimum fuel burn, indicated in the chart by HYC*cruise*, results in selecting the maximum power of the electric motor which can be used at 100% under the constraint of the 20% SOC limit of the battery. In the context of this analysis, the optimum sizing results in a 13% block fuel reduction compared to the baseline aircraft by installing a total electric motor power of 6 MW and 8400 kg of batteries. Further investigation presented in [21] demonstrated a potential block fuel reduction of 16% when utilizing the electric motor during climb and cruise. The prospects in block fuel reduction depend strongly on the assumption made in terms of battery technology level. For a battery speciï ˇn ˛Ac energy of 1000 Wh/kg at cell-level,results shown that the potential reduction block fuel were reduced by almost 50% compared to results obtained for 1500 Wh/kg [21].

In view of examining the sizing effects resulting from the integration of hybrid-electric propulsion at aircraft level and of investigating the potential market range application of hybrid-electric aircraft, clean sheet design of hybrid-electric aircraft need to be considered.
