Last week, our contributor Dora Naya Pons presented her final bachelor’s thesis on the conceptual design of a gyrodyne for rescue missions, achieving a grade of “Excelent”.

Below is a summary of the work he has carried out over the last few months:

CONCEPTUAL DESIGN OF A HYBRID-ELECTRIC GYRODYNE FOR RESCUE MISSIONS

With the aim of proposing a viable cleaner rescue aircraft design to reduce carbon emissions, this bachelor’s thesis, developed within the scope of the AURORA Project, explores the design, limits and capabilities of a gyrodyne configuration rotorcraft aimed at rescue mission operations. The gyrodyne configuration is a compound rotorcraft design that mounts horizontal propellers for forward thrust, allowing nominal cruise speeds of 300 km/h (83 m/s), unheard of for commercial rotorcraft performance.

EXTERNAL DESIGN

Figure 1: External gyrodyne design

The final external design (defined in OpenVSP) after rotor and fuselage iteration can be seen in Figure 1. The Airbus Racer and Eurocopter X3 have been taken as references for this design. This gyrodyne design counts with wing-like lifting surfaces such as the canard and engine support struts, that generate lift as the speed increases, de-loading the main rotor. The horizontal propellers are installed facing rearwards to clear the way and ensure the safety of the patient and personnel during loading and de-loading operations. Furthermore, as there is no tail rotor these horizontal propellers are used to implement the anti-toque system by differential propeller thrust.

PROPULSION SYSTEM

The entire sizing of the propulsion system and operational limits calculation was computed in MATLAB using BEMT to calculate the power needed at each instant (seen in Figure 7). The gravimetric and volumetric densities of the propulsion system components were used to determine their correspondent mass and volume. Fuel cells offer high gravimetric energy density with zero direct emissions, while batteries provide the power peak coverage that fuel cells alone cannot sustain. The battery percentage must therefore ensure the coverage of power peaks along all the mission duration. Changing the battery fraction also implies changes in the sizing tendencies and energy storage of the propulsion system. For the same MTOW of 4670 kg (after some iterations it was determined to approximately be the maximum weight for all the propulsion systems to fit inside the aircraft) the energy system mass (Figure 3) and volume (Figure 2) were computed for different battery fractions. It can be seen in Figure 4, that as battery percentage diminishes the maximum range and endurance increases. This is caused by the higher gravimetric density of fuel cells compared to batteries; for the same weight, fuel cells store more energy than batteries. One of the requirements of the proposed mission is a no-return range capability of 300 km. All battery fractions comply with this requirement, as seen in Figure 4, therefore to ensure that the batteries will be able to cover the power peaks the maximum fraction is taken, even though it is the most range penalizing one.

$Figure 2: Energy system volume as a function of battery fraction$

$Figure 3: Energy system mass as a function of battery fraction$

$Figure 4: Maximum range and endurance for different battery fractions$

FINAL INTERNAL LAYOUT

Figure 5: Interior configuration

In Figure 5 the final layout of the elements of the propulsion system can be observed. The center of gravity is located underneath the center of the main rotor, avoiding the creation of any significant longitudinal moments.

FINAL ASSESSMENT

As mentioned before, the sizing of the propulsion system was done using the computation of the power required for the entire trajectory at each instant, seen in Figure 7. This trajectory was defined at each instant by a collection of time, speed, altitude and distance vectors. To be able to define a trajectory within power and aerodynamic limitations, the flight envelope of the aircraft was computed for a MTOW of 4670 kg, seen in Figure 6.

Figure 6: Flight envelope for MTOW of 4670 kg

Figure 7: Power demand along the rescue trajectory

CONCLUDING REMARKS

As has been seen, an alternative to fossil fuel in aviation is viable, but a lot more research needs to be done into the components of the propulsion system to guarantee operational safety. The next steps to continue developing this design would be a CFD campaign, a structural analysis and a detail rotor design.

On behalf of everyone in the AURORA team, we would like to thank and congratulate Dora on her fantastic work.