Submitted in fulfillment of the requirements for the degree of Master of Science in Mechanical & Aerospace Engineering

The thesis made a significant contribution to the knowledge of HEPS for medium-range aircraft, and the findings can be used to improve the initial performance of aircraft under certain constraints. Due to his broad knowledge of the subject, invaluable patience and positive attitude towards the entire cycle of the thesis development, this time was valuable for me to overcome the difficulties I encountered along the way. Specific fuel consumption constant for BSFC brakes ηICE, ηEM, ηgen efficiency of ICE/ EM or generator .. mass of battery, fuel, ICE, EM or generator .. e battery specific power.

46 FIGURE 3.14: COMPARISON OF MASS COMPONENTS OF ALL ELECTRIC, ALL ICE, SERIES AND PARALLEL SILENT TAKEOFF AND LANDING COMPONENTS HEPS (ΗBAT=0.7).

Introduction

Motivation for Electric Aircraft Investigation
Literature Review

Overview of EA Classification in terms of Propulsion System
Advantages of Hybrid-Electric Aircraft
Successful cases of HEPS
Feasibility of HEPS
Electrification of Aircraft Propulsion System
Challenges in Implementing HEPS
Research Gap Analysis

Research Aim and Objectives
Thesis Layout

In addition, the power supply constitutes one of the most challenging components hindering the evolution of electric aircraft. One of the first proposals for the electrification of conventional fuel-based aircraft was obtained by extending the conventional sizing methodology [36]. During the procedure, the weight of the aircraft could be minimized, which proved the advantage of the HEPS implementation.

The first research was carried out in the evaluation of the performance of the Cessna 337 propulsion system [41]. To investigate the payload overrun of the considered aircraft's propulsion system during simulation of five considered scenarios. The evaluation of HEPS for business jet will be developed on the MATLAB® framework.

Figure 1.2: The prediction of the carbon emission contribution of aviation depends on the technology advance by 2050 [11]

Methodology

Aircraft performance evaluation

Aircraft specification
Mission profile and requirements
Mission power evaluation

Cruise mission
Climbing and landing missions

The requirement to successfully complete the mission is to maintain the maximum range and endurance even after hybridizing the aircraft. 2.2) (2.3) where k is a variable dependent on the aspect ratio and induced drag characteristic, , L is a lift, D – drag, T is a thrust force, 𝑐𝐿 and 𝑐𝐷 are the coefficients of respective forces such as light and drag , W assumed equal to the difference of MTOW and half fuel mass, 𝑞∞ is a dynamic pressure, which is characterized as. Knowing the pitch speed, the lift coefficient can be estimated, but the drag coefficient remains unknown.

From the figure it can be seen that at the minimum power speed, at which the maximum range of flight can be obtained, the drag coefficient is equal to the multiplication of the minimum drag coefficient twice. At this point (at the point where the lift to drag ratio is maximum) the cruising speed can be calculated. In addition, the minimum drag coefficient can be calculated from the known economy cruising speed, where the maximum endurance (at 𝑐𝐿3/2/ 𝑐𝐷) can be estimated.

2.6) The power required to complete the cruise mission is calculated from the general formula below. During the climb, excess force is generated as the value of the thrust force is much higher than the drag. 2.8) where W is equal to MTOW, R/C is the rate of climb and 𝑈∞ is equal to rate of climb.

2.9) where W is equal to MTOW, Wf – is the weight of consumed fuel, and 𝑈∞ is equal to landing speed. The entire power to complete landing and climbing missions can be calculated by (2.10), where U corresponds to climb or landing speeds.

Figure 2.2: Cruise speed and power relationship graph [50]

Assessing the weight of the airframe

Initial sizing assessment
HEPS Assessment
Fuel and battery weight estimation
HEPS components' weight estimation

In this thesis, the hybridization ratio varies with the change in the value of electric power (𝑃𝑒𝑙𝑒𝑐) to investigate the optimal proportion that guarantees the successful completion of the full mission profile of the aircraft. This is necessary to overcome unpleasant situations such as heating of the battery and so on. The battery will be used for short flight missions such as climb and landing to save on the fuel burned.

In terms of power management of series HEPS, due to the fact that only one power source can be used at a time, the only battery will be used during short duration missions. On cruise flights, fuel will only be used to power the battery and complete the mission. When using the parallel HEPS, both electrical and mechanical energy can be consumed at the same time.

Thus, the excess power during climb and landing will be provided by the battery, while energy from the fuel goes to the remaining power. In addition, the battery during cruise will be charged with two options, either slow charging for the entire cruise time or fast charging several times during the cruise flight. The weight of the fuel is estimated by the brake specific fuel consumption constant which is characterized by the ICE, thus the engine efficiency is used in the formula below. 2.13) where BSFC = 225 g/kWh for gasoline [53], which is mostly used in aircraft, t is the corresponding mission time, 𝜂𝐼𝐶𝐸 – internal combustion engine efficiency.

The above formula (2.14 and 2.15) is presented as a general formula to represent the relationship between the power and mass of the energy sources and HEPS components such as the electric motor, generator and ICE. The specific power of the generator (𝑃𝑔𝑒𝑛∗ ) and the electric motor (𝑃𝐸𝑀∗ ) is comparable and equal to 5 kW/kg, while the power of ICE 𝑃𝐼𝐶𝐸∗ is 1 kW/kg.

Figure 2.4: Scheme of series (a) and parallel (b) HEPS

Case Study

Aircraft selection

Zunum Aero ZA10 specification
Performance representation of Zunum Aero ZA10

For the electrification of the aircraft and its comparison with the conventional one, the model aircraft has been chosen. Zunum Aero began designing the aircraft in late 2017 in partnership with Boeing [54]. The first testing of the aircraft was carried out three years later and its delivery to the world was planned for this year.

ZA10 was designed for the regional airline in Washington to reduce operating costs by more than 60 percent. This was done by increasing the number of passengers on board, where the six in VIP, nine in premium, and 12 in economy class, which contribute to the operating cost of the plane as 250 dollars/hour. The technology of ZA10 is designed as a modification of the conventional Rockwell Turbo Commander, whose specific parameters are applied as reference values such as range, MTOW and cruise speed.

For this reason, the aircraft drawing provided by the manufacturer was printed to estimate the scale of the figure. Then, it enables us to estimate the near-exact value of the wing chord to determine other necessary values such as AR and S. The ZA10 engine power consists of a turboshaft engine from the Safran company, which has almost 2000 horsepower.

This engine was actually developed for a helicopter and then modified for the ZA10 and became the Ardiden 3Z. The presentation of the performance of the aircraft under consideration is a key part of the thesis, where each mission of the flight is characterized by the power, time of flight and speed distribution necessary to achieve the primary objective of the project.

Figure 3.2: Illustration of reference aircraft for estimation of necessary dimensions

Simulation scenarios

Pure Electric and Conventional
Series Hybrid Electric
Parallel Hybrid Electric
Quiet Takeoff and Landing Scenario
Discussion

Furthermore, comparing the two configurations reveals the limitation of the all-electric drive system. The drive system of the series hybrid connects the components to obtain all the power from the fuel to run the electric car. The consumption of the battery is introduced in Figure 3.6c to illustrate when the battery is switched on and how quickly it is discharged.

The propulsion efficiency is evaluated according to key parameters such as the state of charge of the battery and the mass of the fuel consumption during the mission flight. This increase may allow for a decrease in battery mass as less current is drawn from the electrical source. From the numerical values, it is obvious that the efficiency of the battery significantly affects its weight and thus the weight of the entire propulsion system.

This option benefits the weight of the HEPS, which significantly reduced the weight of the battery and the overall system compared to the previous type (slow charging). From the cases described in previous sections, it became clear that the parallel case can satisfy the battery's recharging and its consumption for the landing. Since the battery under the parallel HEPS is charged during the cruise flight, the mass of the fuel is now affected from the efficiency factor of the battery.

From the obtained results, which can be seen in Figures 3.15 and 3.16, it is obvious that the impact of battery performance is negligible compared to HEPS. From the obtained results, it is obvious that increasing the HR reduces the noise and emissions of the aircraft. The numbers show that the efficiency of the battery has a big impact on how much it weighs, which affects the mass of the entire drive system.

As the combustion power value changed, the mass of the ICE and the generator also changed.

Figure 3.4: Conventional propulsion system`s a) components connection b) power distribution graph during full flight mission, kW vs h

Conclusion and Future Research Directions

Conclusion

Contribution to Knowledge

Achieving the best ratio of fuel to battery weights for such medium range aircraft enables us to meet the constant range and MTOW requirement.

Future Research Directions

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