*3.2. Trade-Off Wingspan*

Several trade-off studies are performed with the design iteration code. More wingspan usually leads to better aerodynamic performance; therefore, the calculations were performed for wingspans starting from the ATR 42 at 24.6 m and reaching to the end of Airport Category C at 36 m. The grading criteria are the energy requirements for the 400 km mission since this is a major part of the grading criteria for the whole aircraft, i.e., the DOC. Even though the best aircraft performance is at a span of 36 m, the overhanging WTP tips reduce the available wingspan to about 32 m. To account for structural integrity, the maximum taper ratio was set to 0.25, which limits the wingspan to 28.5 m as can be seen in Figure 5. Therefore, the wingspan of HAIQU was selected to be 28.5 m.

**Figure 5.** Energy requirements for a given wingspan.

#### *3.3. Result of the Propulsion System Sizing*

During the aircraft design process,we observed that using only wingtip propellers would lead to immense yawing moments in the case of a single engine failure regarding the electric motor, thus, leading to an oversizing of the VTP. To mitigate this problem, two further propellers were placed in a conventional location on the wing close to the fuselage. To optimize the power split between the outer and inner propellers, a tradeoff study was conducted with the result being that a slightly more powerful WTP would be the optimum solution regarding the boundary conditions.

However, the design iteration code does not include effects regarding disc loading, blown wing area and similar; hence, the results of Nathaniel [27] were used to define the

power split to achieve a higher level of fidelity. They investigated the power distribution of an ATR 42–500 with WTP and conventional propellers. A power split of 50–50 between the WTP and the inbound propellers was identified to be the optimum as the split reduces the disc loading and, therefore, reduces the total required power by 5% compared to a configuration with one propulsor per wing. While the induced drag reduction decreases when using more than one engine per wing, the effects of the disc loading dominate in comparison, thereby, leading to better performance with four propellers.

#### *3.4. Power Split between Battery and Fuel Cell*

The battery concept of HAIQU is to keep the mass as low as possible and to use it as emergency backup while also increasing the lifetime of the battery. With the calculated *H*<sup>2</sup> mass flow for the descent, the minimum fuel cell power setting is 1050 kW, of which 420 kW are not used by the propulsion system or climate control. This overproduction is directly fed into the battery to recharge it at 2C to further reduce the turnaround time. This maximum charging power during descent, in turn, limits the fuel cell takeoff power to 3400 kW. The remaining 440 kW for the takeoff power are provided by the battery system.

With the given power density, the resulting battery mass would be 300 kg; however, the self-set requirement to have a large enough battery to fly at a constant altitude for 10 min (a single traffic pattern in case of fuel cell failure for example) demands a storage capacity of 206 kWh resulting in a battery mass of 1950 kg. The fuel cell has a resulting mass of 1020 kg with the compressor and humidifier weighing 410 kg. The cable, propeller, inverter and gearbox masses are listed as sums in chapter Section 3.5.2. The results are listed in Table 11 for a better overview.

**Table 11.** Result of powertrain sizing regarding the power and mass.


To increase safety through redundancy, two identical fuel cells in the back of the cabin and two identical batteries in the front of the cabin are used. The energy from both systems is transformed using a DC/DC converter each to achieve the best transport voltage for the superconducting cables. Those cables run from the DC/DC converters to the DC/AC converters located at the motors in the wing as can be seen in Figure 6.

The motors need to provide a power of 900 kW each, resulting in a mass of 60 kg with the gearbox adding another 35 kg each. They are driving a six-bladed propeller with a diameter of 4 m running at low speed for lower noise emissions. Regarding the TMS, the hydrogen is pumped from the tanks through the HTS cables and in a separate line to the motors to supply the right temperature to both components. Both flows recombine at the DC/AC converters and merge back in the fuselage to cool the DC/DC converters before the gaseous hydrogen is heated to 85 °C using the fuel-cell system. This order is derived from Hartmann et al. [47] and shown in Figure 7 but with a few modifications to reduce the fuel line length.

For the cruise flight, a fuel consumption of 2.2 kg/min is achieved at a fuel cell efficiency of 55%. This results in fuel consumption for the whole 400 km flight mission of 172 kg. The electric green taxiing system complies with the requirements, and HAIQU fulfills the Flightpath 2050 goal of emission-free aircraft movements during the taxiing phase [73], as the electric motors produce relatively low noise, and they emit no emissions locally [34].

**Figure 6.** Powertrain component arrangement within HAIQU.

**Figure 7.** Hydrogen flow from the fuel tanks to the fuel-cell system; full line representing liquid hydrogen and dotted line representing gaseous hydrogen.

#### *3.5. Resulting Aircraft Design—HAIQU*

All the components mentioned in the previous chapter are part of the final design of HAIQU. In this section, the full aircraft will be shown and its performance characteristics described. In Figure 8, the three-side view of the final aircraft design is shown, and Figure 9 shows a rendering of the operating plane [74].
