Estimating Aircraft Power Requirements: A Study of Electrical Power Demand Across Various Aircraft Models and Flight Phases

Abstract

This research addresses the increasing electrification of aircraft systems, driven by the need to improve energy efficiency and reduce CO₂ emissions in global aviation. The transition to more-electric aircraft (MEA) is advocated as a promising strategy, as it is expected to improve environmental performance and economic viability. However, this shift significantly increases the demand for on-board electrical power. One alternative to traditional engine generators is novel power supply systems such as fuel cell systems. In order to design these systems effectively, it is essential to determine the electrical power requirements that the fuel cells must supply. Estimating the electrical power consumption of individual aircraft systems is critical given the proprietary nature of manufacturer data. Using existing literature methods, this study aims to identify the essential variables for estimating the magnitude of power consumption. The research focuses on different aircraft models, taking into account their system architectures and electrification trends, in particular for Airbus and Boeing models. The study includes a detailed description of the aircraft systems, calculation methods, and presentation and analysis of the estimated electrical power requirements. Despite a lack of available data for comparison, the calculated results appear to be reasonably consistent with existing literature and provide valuable insights into the electrical power requirements of aircraft systems.

1. Introduction

Current research indicates that global aviation is responsible for approximately 2.5% of anthropogenic CO₂ emissions. If the additional effects of non-CO₂ emissions such as water vapor and nitrogen oxides are taken into account, the contribution to anthropogenic climate change rises to around 5% [1]. The situation is further exacerbated by the increasing rate of growth in air traffic, around 5% per year [2]. To significantly reduce aircraft emissions and meet the climate targets set out in Flightpath 2050, it is imperative to improve the energy efficiency of future commercial aircraft [3]. One promising approach is to replace hydraulically and pneumatically powered aircraft systems with more energy-efficient electrical systems, collectively known as “more-electric aircraft” (MEA). Complete electrification is referred to as an “all-electric aircraft” (AEA). This transition is expected to improve environmental performance and increase economic viability due to higher energy efficiency compared to conventional systems [4,5,6]. In addition, researchers such as [7,8] anticipate weight savings and improved maintainability of components. The trend towards electrification of aircraft systems is already evident in current aircraft models from Airbus (A350, A380) and Boeing (B787) [9,10].

However, the increasing electrification of aircraft systems is also leading to an increase in the demand for electrical power on board. Until now, electrical power has been generated by engine generators. This means that the generation of electrical power requires kerosene, which is also responsible for aircraft emissions from combustion in the engine combustor. To mitigate this reliance on engine generators for electrical power, one discussed solution is the implementation of fuel cell systems to provide electrical power [11]. These electrochemical energy converters efficiently convert chemical energy in the form of hydrogen and oxygen into electrical energy without emitting pollutants or noise [10]. The first step in the design and sizing of fuel cell systems is to determine the electrical power required. The power consumption of individual systems is calculated and measured by aircraft manufacturers and recorded in electrical load analysis (ELA) documents. This information is proprietary and not publicly available. This paper aims to estimate the electrical power consumption of individual aircraft systems. To achieve this, existing calculation methods from the literature are used. As the results of these methods are typically normalized, this work seeks to identify and estimate all necessary variables to provide an estimate of the order of magnitude of the electrical power consumption.

This paper is structured as follows: In Section 2 conventional aircraft systems, MEA, and AEA are explained. Section 3 defines the scope of this paper. The aircraft systems under consideration are identified, different aircraft models are introduced, the output parameters are described, and the assumed flight mission is explained, along with the assumptions made regarding altitude, speed, and environmental conditions. Section 4 describes the operation and functionality of considered aircraft systems. The various calculation methods are then explained in Section 5. The required variables are identified and, if necessary, estimated. The results of the electrical power estimation can be found in Section 6. Various graphs and their descriptions are presented. The validation of the calculated power requirements is also described in the same chapter, including a comparison with existing data from the literature. Based on those results, the total power requirements per flight phase and aircraft type are presented. We conclude those results in Section 7 by evaluating and discussing their general validity. Section 8 presents the summary and outlook.

2. State of the Art

In contemporary civil commercial aircraft, jet engines are responsible for both providing thrust and powering the aircraft’s systems. In this configuration, ambient air is drawn into the engine from the surrounding area. It is compressed by a fan and low-pressure and high-pressure compressors, simultaneously raising its temperature. The air then flows into the downstream combustion chamber where kerosene is injected. The resulting air–kerosene mixture is ignited. This expansion of the air drives a rotary motion of the shaft with the turbine blades mounted on top of it. This rotational energy also drives the upstream compressor, which is connected to the turbine by one, two, or three shafts, depending on the engine [12,13].

To provide power to the different aircraft systems, three forms of power—hydraulic, pneumatic, and electrical—are taken from the engine at various points as secondary power, as shown in Figure 1.

Pneumatic power is extracted from the engine in the form of compressed air from the compressor and is referred to as bleed air or secondary air. The high temperature and high pressure of the bleed air are used for a variety of tasks related to sealing, cooling, heating, or ventilation [14,15,16].

Hydraulic power is used for mechanical tasks on board. These usually primary and secondary flight control, landing gear operation, cargo door and hatch operation, brake operation, and thrust reversal. Therefore, a hydro-mechanical constant-pressure axial piston pump draws hydraulic fluid from a reservoir and compresses it to a pressure typically in the range of 196 to 350 bar in commercial aviation [17]. The piston pump is mechanically connected to the auxiliary gearbox, which in turn is coupled to the high-pressure shaft of the engine via a radial shaft and another angular gear [12].

Engine-driven generators are used to generate electrical power. These generators, like the hydraulic pumps, are driven by the gearbox and thus by the high-pressure shaft of the engine [12,16]. The electrical power is used to supply for example the cabin systems, the lighting system, the avionic, the fuel system, and the de-icing system of the windows and windshield.

The MEA focuses on the gradual electrification of aircraft systems to reduce dependence on mechanical and hydraulic systems. This involves the use of electrified supply systems for specific aircraft components such as hydraulics and pneumatics systems. A key component of MEA is the deployment of potent electric generators driven by the engines, which grant the necessary electrical energy to power the electrified systems of the aircraft. MEA additionally enables the integration of energy storage systems, such as batteries or supercapacitors, to boost the efficiency and redundancy of electrical systems [18].

In contrast to MEA, AEA architecture represents a more radical transformation, rethinking the entire aircraft design. AEA concepts encompass not only electrifying specific systems but also transitioning the entire aircraft to an electric platform. As a result, both the primary engines and all other systems, including flight control and cabin infrastructure, operate using electrical power. AEA may offer several advantages, such as a considerable weight reduction since heavy hydraulic and mechanical systems can be eliminated. AEA also allows for more flexible placement of engines and propulsion systems due to the simplicity of laying electrical wiring compared to mechanical connections [19]. The incorporation of MEA and AEA in aviation presents abundant prospects to create aircraft that are both more eco-friendly and efficient. Despite these opportunities, there are also challenges to surmount, comprising the creation of dependable electrical propulsion systems, guaranteeing the safety and certification of electrified aircraft, and developing suitable infrastructure for electric flight.

3. Scope

After acknowledging the challenge of achieving high energy efficiency within the aviation sector and outlining conventional aircraft system architectures, this chapter will explain the basic assumptions for electrical power estimation in this paper. These include an explanation of the assumed aircraft system architecture, the considered commercial aircraft, and the assumed flight missions.

3.1. Assumed Aircraft Architecture

The aircraft architecture assumed and examined in this paper is based on the MEA concept, where hydraulically and pneumatically powered aircraft systems are replaced by electric ones [9].

As shown in Figure 2, the electrified aircraft systems are intended to be powered by one or more fuel cell systems (FuCSs), in contrast to Figure 1. This results in an independent power supply for aircraft systems, separate from the propulsion system.

As the electrical power for the aircraft systems in this concept is provided by the fuel cell system, there is no need to draw secondary power from the engine, which offers several advantages:

• Improvement of the engine high-pressure compressor efficiency: As no bleed air is drawn from any of the high-pressure compressor stages for air conditioning and wing de-icing systems [20].
• Improved turbine efficiency of the engine: The supply of electrical systems via the fuel cell system and the electrification of hydraulic systems eliminates the need for mechanical coupling of the accessory gearbox to the high-pressure shaft via angular gears.
• Weight saving: By providing power to aircraft systems through alternative means, the elimination of the gearbox results in weight savings.
• Reduced maintenance costs and increased engine reliability: Eliminating the maintenance-intensive bleed air system reduces expected maintenance costs. In addition, engine complexity is reduced by eliminating the pneumatic system consisting of a pre-cooler, control valves, and piping [20].

3.2. Studied Aircraft Systems

The aircraft systems integrated into current commercial aircraft are listed in

Table 1. Relevant ATA chapters with the aircraft systems to be examined.

ATA-Chapter	Description	El. The Power Demand
ATA 21	Environmental Control System	Yes
ATA 22	Auto flight System	Yes
ATA 23	Communication System	Yes
ATA 24	Electrical Power System	No
ATA 25	Equipment	Yes
ATA 26	Fire Protection	No
ATA 27	Flight Controls	Yes
ATA 28	Fuel System	Yes *
ATA 29	Hydraulic Power System	Yes
ATA 30	Ice and Rain Protection System	Yes
ATA 31	Indicating and Recording Systems	Yes
ATA 32	Landing Gear	Yes *
ATA 33	Lights	Yes
ATA 34	Navigation	Yes
ATA 35	Oxygen System	No
ATA 36	Pneumatic System	No
ATA 38	Water and Waste System	Yes
ATA 44	Cabin Systems	Yes

* Considered in this paper by the electrified hydraulic system.

, sorted by their ATA chapters. The ATA chapters assign a designated system identifier, consisting of a numerical sequence, to various systems [16]. This system was established by the Air Transport Association of America (ATA) and applies to all aircraft certified to FAR-25 and CS-25. It enables standardized communication and organization within the industry, which is particularly important for efficient maintenance, repair, and documentation. Although the ATA chapters are considered somewhat outdated, they are still widely used to structure and understand the complexity of aircraft technology. ATA Chapters 21 to 49 are the chapters associated with aircraft systems.

Whether a system requires electrical power can be seen in the last column of Table 1. The ATA 26, 35, and 36 systems are omitted for the following reasons:

• ATA 26—Fire protection system: This system is an emergency system consisting of sensors for fire and smoke detection, as well as fire extinguishing systems [16]. Since it is typically an inactive system and most sensors do not require electrical energy, this system is not further investigated.
• ATA 35—Oxygen system: This system is also an emergency system that provides oxygen either through pressure bottles or oxygen generators through oxygen masks in case of a sudden cabin pressure loss. This system is self-activating and does not require external electrical energy [16].
• ATA 36—Pneumatic system: This system, according to [16], consists of channels, sensors, and valves that deliver compressed air from the power source, such as the engine and an electrically operated pump, to a consumer. No electrical power is required for this system.

• ATA 24, the electrical power supply, is also not considered in this paper as this system provides the distribution of electrical power to aircraft systems via buses. No electrical power is required for distribution. Additional assumptions have been made for some of the aircraft systems:

• ATA 21—Environmental control system (ECS): Within this paper, we calculate the electrical power requirements for the bleed air-free and electrified ECS currently employed in the Boeing 787–800, excluding the conventional bleed air method.
• ATA 29—Hydraulic System A complete substitution of the hydraulic system is currently not feasible due to safety considerations. This paper assumes an electro-hydraulic system, resembling the Boeing 787–800, where hydraulic power is supplied by electrically driven pumps. All hydraulic pumps are assumed to be electrically powered. This implies that this work will not delve further into ATA 27 (flight control system) and ATA 32 (landing gear system), as both systems are believed to be energized via an electrified hydraulic system.
• ATA 30—Ice and rain protection system: This system consists of the engine cowl anti-ice system, the wing anti-ice system (WIPS), the window, the windshield heater, as well as the probe heater. The engine cowl anti-ice system and probe heater are not considered in this paper. For the WIPS it is assumed that it is an electrical resistance system, as is the case with the Boeing 787–800. Here, the leading edge of the aircraft’s wing can undergo thermal de-icing through the use of an electrical resistance system. For the estimation of the electrical power consumption of the windows and windshield heater, no additional assumptions are required.

3.3. Calculated Parameters

Based on the assumptions outlined in Section 3.1 and Section 3.2 of the complete electrification of aircraft systems, the required power is calculated as power per flight phase. This is essential to useful estimations of electrical power requirements. The given power is calculated as the maximum necessary power required by the aircraft system for the given flight phase. It does not consider variable load profiles or multiple activity segments over any single flight phase.

The results can be used as a reference for determining the necessary power required by a supplying system. Such systems are usually sized by necessary power and required energy. For novel supply systems and fuel cell systems in particular, required power becomes increasingly important, as implied cooling demands influence system design.

3.4. Analyzed Aircraft Types

This paper solely concentrates on commercial aircraft. The aircraft models investigated can be found in

Table 2. Aircraft types under investigation.

Aircraft Types	First Flight	Engine Number	Assumed Number of Passengers ${\mathbf{n}}_{\mathbf{occ}}$
A320-200	1987	2	168
A330-200	1997	2	293
A380-800	2005	4	540
B737-800	1997	2	189
B757-300	1998	2	243
B777-300	1997	2	383

. The selection of aircraft is based on several factors. The selection of aircraft with different numbers of engines and passenger capacities allowed for a comprehensive study of the electrical power requirements that are necessary across a broad range of commercial planes. Another important factor is the availability of data. The required data for estimation could be ascertained or reasonably estimated for all the aircraft types listed in this study.

3.5. Assumed Flight Mission

The power consumption of aircraft systems is affected by the flight phase. A reference mission is established with distinct phases, as illustrated in Figure 3. Altitude and airspeed define each flight phase and also act as input parameters for the power models. Additionally, environmental factors, specifically pressure, temperature, and density, are determined by the altitude.

3.5.1. ICAO Standard Atmosphere

The electrical power requirement is affected by atmospheric conditions. All of the following data are taken from the International Civil Aviation Organization (ICAO) [21]. The ICAO has established a standardized model for reference, known as the International Standard Atmosphere (ISA) [22]. This model defines specific values for temperature, air pressure, and density at different altitudes up to 80 km. For this paper, however, only the values up to an altitude of 11 km are taken into account. According to ISA, the mean sea level temperature is 15 °C, and the air pressure is 1013.25 hPa. As altitude increases, temperature decreases at a uniform rate of around 6.5 °C per kilometer until the tropopause is reached at an altitude of 11 km [21]. The formula for calculating the temperature $T_{0-11}$ at a given height is given as follows:

$$T_{0-11}=T_{ambient}=T_0+{\displaystyle \frac{dT}{dH}}·H=T_0-0.0065\left[{\displaystyle \frac{K}{m}}\right]·H$$

(1)

$T_0$ refers to the temperature at mean sea level, which is given as 15 °C or 288.15 K according to the ISA. Above the tropopause (above 11 km), there is a constant temperature of −56.5 °C. H represents the altitude in meters above mean sea level. Air pressure $p_{0-11}$ also varies with altitude and can be calculated using the barometric altitude formula. Within the troposphere, the following applies:

$$p_{0-11}=p_{ambient}=p_0·{\left(1-{\displaystyle \frac{0.0065·H}{T_0}}\right)}^{5.255}$$

(2)

The assumed air pressure at mean sea level is described by $p_0=1013.25\mathrm{hPA}$ in accordance with ISA.

3.5.2. Input Parameters for Power Calculation

The input parameters for the electrical power calculation, as determined and set for the mission, can be found in

Table 3. Assumed mission data.

	Height Above Ground H [ft]	Height Above Ground H [m]	Ground Speed ${\mathbf{v}}_{\mathbf{flight}}$ [kt]	Ground Speed ${\mathbf{v}}_{\mathbf{flight}}$ [m/s]	Speed of Sound [m/s]	Ma	${\mathbf{T}}_{\mathbf{ambient}}$ [°C]	${\mathbf{T}}_{\mathbf{ambient}}$ [K]	${\mathbf{p}}_{\mathbf{ambient}}$ [bar]
Taxi-Out
Taxi	0	0	15	7.71	340.5	0.02	15	288.15	1.01325
Take-Off
Acceleration	35	10.67	150	77.1	340.46	0.23	14.93	288.08	1.012
Climb
Initial Climb	4000	1219	190	97.66	335.75	0.29	7.08	280.23	0.88
Climb 1	15,000	4572	290	149.06	322.67	0.46	−14.72	258.43	0.572
Climb 2	24,000	7315.2	350	179.9	311.97	0.58	−32.55	240.60	0.393
Climb 3	40,000	12,192	410	210.74	292.95	0.72	−56.5	208.90	0.187
Cruise
Cruise	40,000	12,192	468.98	241.07	297.6	0.81	−56.5	216.65	0.187
Descent
Descent 1	30,000	9144	480.39	246.92	304.84	0.81	−44.44	228.72	0.3009
Descent 2	10,000	3048	250	128.5	328.62	0.391	−4.81	268.3	0.697
Descent 3	4000	1219.2	240	123.36	335.75	0.367	7.08	280.23	0.875
Approach
Approach Initial	2500	762	240	123.36	337.53	0.365	10.047	283.197	0.925
Approach Final	1000	304.8	160	82.24	339.31	0.242	13.018	286.169	0.977
Landing
Thresh. Cros.	100	30.48	140	71.96	340.38	0.212	14.802	287.952	1.0095
Touchdown	0	0	137	70.418	340.5	0.207	15	288.15	1.01325
Deceleration	0	0	15	7.71	340.5	0.023	15	288.15	1.01325
Taxi-In
Taxi	0	0	15	7.71	340.5	0.023	15	288.15	1.01325

. Various estimates exist in the literature, and the following estimations are an assumption. The subsequent sections provide a concise explanation of the different stages of a flight, which are considered in this paper:

• Taxi-out: This phase encompasses the period before takeoff, during which the aircraft taxis from the gate to the runway. It involves various operational processes, including the pushback maneuver, engine start, communication with air traffic control, taxiway, and runway navigation, and internal checks to ensure flight readiness. According to [23], the aircraft typically taxis at an average speed of 15 to 20 knots.
• Take-Off: This phase involves the acceleration on the runway, the aircraft’s ascent, and the transition from the ground to a controlled flight state, ending when the aircraft reaches obstacle height (35 feet above ground).
• Climb: This phase encompasses the entire climb process, divided into several sub-phases [24]. The initial climb begins immediately after takeoff and involves ascending to a specific altitude, with speed gradually increasing by 10 knots, 30 knots, and 60 knots. In the subsequent climb phases (Climb 1–3), speed and altitude are continuously increased, with the condition that the airspeed should be greater than 250 knots only at an altitude above 10,000 feet.
• Cruise: This phase refers to the portion of a flight mission during which the aircraft maintains a constant altitude. Commercial aircraft typically fly at altitudes of up to 40,000 feet and speeds of typically 0.81 Mach. This paper assumes a cruising altitude of 40,000 feet and a speed of 0.81 Mach. During this phase, efficiency and cost-effectiveness are prioritized by considering optimal cruise profiles and fuel consumption.
• Descent: This phase involves the aircraft descending from cruise altitude to the destination airport or runway. During this phase, altitude is gradually reduced to facilitate a safe and controlled landing. In the first phase (Descent 1), the cruising speed is maintained. Starting from the “crossover altitude” at 30,000 feet, speed is reduced and limited to 300 knots (CAS). At altitudes below 10,000 feet, the airspeed must be less than 250 knots, as in the Climb phase.
• Approach: In this phase, the aircraft is in the immediate lead-up to landing [10]. It involves several sub-phases. During the “Threshold Crossing” phase, the aircraft flies at a low altitude over the runway approach area to reach a suitable landing position. During the “Touchdown”, the aircraft’s wheels make contact with the runway, and the landing roll begins. “Deceleration” refers to the process of slowing down the aircraft, either through the use of brakes or thrust reversal, to bring the aircraft to a stop or significantly reduce its speed.
• Taxi-in: In this phase, the aircraft taxis from the runway to the gate. Similar to the Taxi-out phase, the aircraft taxis at an average speed of 15 to 20 knots.

4. Functionality of the Aircraft Systems

This chapter describes the operation of the relevant aircraft systems sorted by their ATA classification.

4.1. ATA 21: Environmental Control System

Commercial aircraft typically cruise at altitudes between 11 and 15 km in the stratosphere. At these heights, the temperature of the air remains steady at around −56.5 °C while air pressure fluctuates along with altitude. At the same time, the partial pressure of oxygen declines in proportion to the decline in air pressure. Due to the decline in oxygen partial pressure with decreasing air pressure, continuous pressurization is crucial for the aircraft cabin. In conventional aircraft architectures, the air supply is taken from the engines, as shown in Figure 4 as a pneumatic air conditioning system (pACS). This involves extracting between 0.05% and 2% of the high-pressure compressor airflow for cabin air conditioning [25]. The air, maintained at a constant pressure level, is then directed to an air-to-air cross-flow heat exchanger. This pre-cooler, as indicated in [12,26], cools the hot bleed air to approximately 180 °C using cold air from, for example, one of the fan stages. A portion of the bleed air mass flow is diverted and reintroduced into the process at a later stage. The remaining mass flow is directed to the aircraft’s air conditioning packs (ACPs). Modern commercial aircraft usually employ two ACPs for air conditioning. In these packs, the air temperature is reduced and the pressure is lowered compared to the extracted air from the high-pressure stage compressor to match the cabin pressure. Further details about this process can be found in [15,16,27]. Following the air conditioning packs, depending on the literature source, the bleed air has a temperature ranging from −12 °C to 9 °C and a pressure of approximately 0.8 bar before entering the mixing unit [26,28,29]. A minimum fresh air supply of $0.25$ kg/min per passenger is required [30]. Afterward, the conditioned bleed air enters the mixing unit where it mixes with hot trim air and recirculated cabin air.

An alternative system is the bleed air-free electric system used in the Boeing 787, as illustrated in Figure 4. In this system, ambient air is drawn through special inlets on the aircraft’s fuselage and directed to two redundantly installed electric compressors, known as cabin air compressors (CACs). These compressors are responsible for pressurizing the air to a level of 1 bar, leading to a simultaneous increase in temperature [26]. Subsequently, the air is directed to air conditioning packs, which are equivalent to those in the bleed air system described above. From there, the air flows to the mixing unit and is maintained at a pressure level of approximately 0.8 bar. It is mixed with trim air and recirculated cabin air and is distributed into the cabin.

4.2. ATA 22, 23, 27, 31, 34: Avionic Systems

The term avionics, derived from “aviation” and “electronics”, encompasses electronic aircraft systems that receive, process, and exchange information [16,31]. Associated with specific ATA chapters, key functionalities are outlined:

• ATA 22—Auto Flight: The auto flight system (AFS) performs functions such as attitude control, automatic speed control, altitude guidance, approach, and landing. The system relieves pilots during cruise flights, enhances flight guidance accuracy, and contributes to the efficiency and safety of flight operations.
• ATA 23—Communications: Communication system enables transmission of information between the aircraft, air traffic control, crew, and passengers.
• ATA 27—Flight Control System: The flight control system (FCS) is responsible for the movement of the aircraft, including roll, pitch, and yaw, to maintain stable flight and perform all necessary maneuvers.
• ATA 31—Indicating Recording System: Flight monitoring systems continuously monitor various flight parameters and conditions using sensors and systems to inform the crew about system flight and system status as well as operating parameters.
• ATA 34—Navigation: The navigation system ensures precise navigation. For this purpose, it provides precise positioning determination, environmental data collection, attitude, and direction indication. It not only enables precise flight positioning but can also integrate weather data, terrain information, and air traffic control transponder data, which are essential for flight safety, efficiency, and compliance with flight plans. Additionally, it provides the necessary sensor data for ATA 27 and ATA 31.

Due to the critical nature of some avionics systems and their potential to lead to an aircraft crash if they fail, the highest safety standards are imposed on these systems [9]. This requires ensuring electrical power to the systems in all scenarios.

4.3. ATA 25: Equipment

According to [16,32], equipment includes all removable components of the cockpit and cabin, such as pilot and passenger seats, wall paneling, floor coverings, overhead bins, and insulation. It also encompasses toilets, emergency equipment (evacuation equipment, life rats, etc.), and galleys (cabinets, ovens, etc.). In terms of electrical power demand, as stated by [32], galleys are the primary consumers among the equipment systems, accounting for over 95% of the power consumption. Due to the minimal electrical energy needs of most equipment items, only galleys are considered for electrical power consumption estimates. Required power depends on factors like passenger count, flight duration, passenger behavior, and airline specifications [32,33].

4.4. ATA 28: Fuel System

The primary function of the fuel system is to reliably store and supply the engines with fuel throughout the entire flight. Typical components and functionalities include:

• Tanks: The fuel system comprises multiple fuel tanks that store fuel for the engines and the auxiliary power unit (APU).
• Fuel Delivery: It includes pumps and pipelines that transport fuel from the tanks to the engines or transfer fuel between the tanks, maintaining the necessary pressure and flow rate.
• Monitoring and Indicating System: The system monitors fuel levels, consumption, and fuel temperatures to ensure an adequate fuel supply for the flight and to detect potential issues early.
• Venting and Drainage: The system contains venting and drainage devices to remove excess air and moisture from the fuel system.
• Safety: The fuel system is equipped with safety measures to minimize the risk of leaks and fires, such as pressure relief valves and fire protection devices.

• The fuel system plays a critical role in flight operations by ensuring a continuous fuel supply to the engines. It must adhere to the highest safety standards to minimize the risk of fuel leaks and fires and ensure flight safety. According to [32], electrically driven fuel pumps, particularly transfer pumps and booster pumps, significantly contribute to electrical power consumption. Transfer pumps are used to transfer fuel between aircraft tanks. Fuel transfer, as described in [9], is employed to shift the aircraft’s center of gravity, thereby reducing structural loads on the wings and fuselage. If there is a fuel imbalance due to increased fuel consumption by one engine, a fuel transfer is also performed. Booster pumps, on the other hand, are used to increase fuel pressure and support delivery to the engine’s combustion chamber. Depending on the aircraft type, however, they can also be used purely as backup pumps in case the main ejector pumps fail.

4.5. ATA 29: Hydraulic System

Hydraulic power is utilized in commercial aircraft to accomplish numerous mechanical tasks. The flight control systems, landing gear, cargo and passenger doors, brakes, and thrust reversers are among the primary consumers of hydraulic power [16]. It has been stated by [17] that the framework of the hydraulic system has remained fundamentally unchanged for several decades. A constant-pressure-regulated axial piston pump draws hydraulic fluid from a reservoir via a suction line and compresses it to the required aircraft pressure. The regulated pressure is usually set to either 3000 psi (≈206 bar) or 5000 psi (≈345 bar) in aircraft applications [17]. The compressed hydraulic fluid is conveyed through high-pressure lines to either a hydraulic consumer within the aircraft or a pressure accumulator.

The hydraulic system is classified as a safety-critical system because a failure in the hydraulic supply could potentially lead to the failure of affected components. To mitigate such failures, the hydraulic system usually consists of two to three independent circuits. In the case of the Airbus A320, the hydraulic subsystem is divided into blue, green, and yellow hydraulic subsystems. In the green and yellow systems, axial piston pumps are mechanically connected to an auxiliary gear support, which, in turn, is linked to the high-pressure shaft of the engine through a radial shaft and angle gear, as stated by [12]. These engine-driven axial piston pumps are referred to as EDPs (engine-driven pumps). The blue subsystem is powered by an electrically driven hydraulic pump [16].

4.6. ATA 30: Ice and Rain Protection System

The aircraft surface can accumulate ice under certain conditions (icing conditions). Critical factors for this include elevated humidity, an air temperature between −30 °C and 0 °C, and an aircraft surface temperature below 0 °C, as documented by [34]. When supercooled water droplets encounter the aircraft surface under these conditions, they freeze and form an ice layer. Initial icing particularly affects the leading edges of the wings, air inlets, sensors, and cockpit windows. The negative consequences of icing are diverse. According to [16,23], potential effects include:

• Alteration of the leading-edge profile shape, leading to changes in the angle of attack and an increase in the stalling speed, potentially resulting in a stall.
• Reduction in lift and an increase in drag.
• Loss of stability.
• Increased weight.
• Aircraft antenna breakage.
• Blockage of the pitot tube opening, which prevents the measurement of static pressure.
• Reduced visibility.
• Blockage of control surfaces and control mechanisms.

To protect the aircraft from these dangers, de-icing critical areas under icing conditions is essential. The ice protection system is designed based on certification requirements and can be divided into the following categories according to ATA chapters, as described by [16]:

• Ice protection for aircraft surfaces such as the leading edge and air inlets.
• Ice protection for external components such as pitot probes and antennas.
• Ice protection for internal components like water lines.
• Ice protection for the windshield.

We assume that the majority of power is required for protecting the wing’s leading edge and the windshield. Therefore, the focus for estimating the total electrical consumption is placed on these two subsystems. In addition, the engine inlets are also de-iced if necessary. However, the power estimation for de-icing the inlets is not part of this paper.

4.6.1. Ice Protection for Wing Leading Edge

For protecting the wing leading edge, two fundamental principles are employed: de-icing and anti-icing. According to AIR 1168/4 [35], de-icing involves the periodic removal of ice particles. Anti-icing prevents ice formation by continuously evaporating super-cooled water droplets, ensuring an ice-free wing surface throughout flight [16].

Various methods can be used for de-icing or anti-icing. In the case of pneumatic-mechanical de-icing, pneumatic de-icing boots periodically inflate to remove ice accumulations. This ice protection method is typically used in propeller-driven aircraft, as it can only provide less hot air compared to turbine-driven aircraft. The second method is the bleed air method. Compressed high-temperature air is extracted from the high-pressure compressor of the engines, cooled to 232 °C using a pre-cooler, and then directed toward the aircraft’s leading edges [12]. The bleed air warms the wing surface, causing the ice to melt. This method is employed in all turbine-driven commercial aircraft for de-icing and anti-icing of wing leading edges, except for the B787-800. In MEA, wing leading edges are for example thermally de-iced using an electrical resistance system, as seen in the B787-800 [20]. The protected wing edge surface is divided into smaller areas using parting strips, as shown in Figure 5. The parting strips are permanently heated. The sequentially heated areas are active in fixed time intervals, partially melting the ice. This allows aerodynamic forces to remove the ice.

4.6.2. Ice and Fog Protection for Cockpit Windows

Commercial aircraft are equipped with an anti-icing system for cockpit windows, which protects the cockpit window from fogging and icing throughout the entire flight and enhances the window’s resilience against impact forces, such as hail or bird strikes [9]. In most cases, an electrical resistance system is used as the anti-icing system. A transparent, electrically conductive film is applied to the inner surface of the outermost layer of the window. To maintain a fog-free view and provide de-icing, an external window surface temperature of 1.6 °C is required, necessitating an electrical power consumption of approximately 3 to 4.5 kW per square inch of window area [35]. As an alternative to the electrical resistance system, some aircraft types employ a hot air system. In this approach, hot air is directed onto the cockpit window from the exterior, which can also serve as a rain protection system, removing droplets from the cockpit window [16].

4.7. ATA 33: Lighting System

Every commercial aircraft is equipped with a multitude of lights, all integrated into the aircraft’s lighting system. In this paper, they are grouped into internal and external lighting.

• Internal lighting encompasses general cabin illumination, reading lights, and various informational and warning lights within the cabin. General cabin lighting employs LED strips along both the cabin ceiling and the lateral cabin walls. The number of LED strips used varies depending on the seating configuration.
• External lighting can be categorized into two groups: the first consists of lights used throughout the entire flight, while the second includes lights specifically required during takeoff and landing.

The first category includes various navigation positions and anti-collision and beacon lights. Lights exclusively used during the takeoff and landing phases comprise landing lights, taxi lights, and logo lights that illuminate the airline’s logo.

4.8. ATA 38: Water and Waste System

According to [16], this system can be divided into three subsystems: the drinking water, the wastewater, and the toilet system. The drinking water system supplies the aircraft cabin with potable water stored in a tank located in the aircraft’s nose. The potable water is distributed to the galley and lavatory sinks through a piping system [16]. The water used for hand-washing in the lavatories is heated by a boiler. Assuming that each toilet is equipped with a faucet, the number of boilers depends on the number of toilets.

The water used for hand washing in the toilets is heated by a boiler. Assuming that each toilet is equipped with a faucet, the number of boilers depends on the number of toilets.

The wastewater system, according to [16], mainly consists of drainage pipes and drainage openings through which wastewater from the galleys and sinks is directed into the environment. To prevent these drainage openings from freezing, they are equipped with heating elements.

Most aircraft are equipped with a vacuum toilet system. In this system, fecal matter is conveyed to the toilet tank through a vacuum created at altitudes below 16,000 feet using a device known as a vacuum generator. This vacuum generator creates a reduced pressure within the toilet tank. However, at altitudes above 16,000 feet, the differential pressure is exclusively maintained using ambient pressure [16].

4.9. ATA 44: Cabin System

Aircraft cabin systems, crucial for passenger and crew comfort and safety, usually encompass:

• Cabin intercommunication data system (CIDS).
• In-flight entertainment system (IFE).
• Cabin monitoring system.

• CIDS integrates communication in aircraft cabins, facilitating passenger-crew interaction via in-flight telephones, emergency systems, and entertainment. It serves as a critical interface for monitoring and controlling the cabin environment, covering lighting, air conditioning, and security [37].The IFE enhances the passenger experience with movies, music, and games, addressing stress during air travel. Installed in most commercial aircraft, IFE allows individual consumption of entertainment and real-time flight data, significantly boosting comfort [38]. Integrated into seats, visual display units (VDUs) are commonly placed on the backs of front seats. Airlines may permit wireless Wi-Fi connectivity for passengers’ own devices, enhancing Internet access [39]. The cabin monitoring system is responsible for monitoring the cabin environment, including collecting data on air quality, temperature, and other environmental conditions to ensure that passengers and crew travel in a comfortable and safe environment [40].

5. Calculation Methods

This chapter presents the calculation methods used in this paper for estimating electrical power, arranged according to their ATA chapters.

5.1. ATA 21: Environmental Control System

The ECS can be grouped into two systems. One is the power required for the electric air conditioning system (ACS) and the other is the power required for the recirculation fan. The total power required for the A/C system is calculated as follows:

$$P_{el}=P_{ACS}+P_{Fan}$$

(3)

The calculation methods used are described separately below.

Estimating the Electrical Power Requirement for the Electrical ACS

As described in Section 4.1, there are two different ways of air conditioning. In conventional aircraft architectures, engine bleed air is used to supply fresh air to the aircraft cabin. An alternative to this system is the electrical system installed on the B787-800. In this system, external electrical power is required to condition the ambient air. The required power depends on the specific heat capacity $c_p$ of the air, the mass flow of air to be transported ${\dot{m}}_{ram}$, the temperature in the ram air duct $T_{ram}$, the pressure ratio $\pi =p_{cabin}/p_{ram}$ and the heat capacity ratio $\kappa$. Assuming reversible adiabatic compression, the minimum ideal power requirement is calculated using the following equation:

$$P_{ACS}=c_p·{\dot{m}}_{ram}·T_{ram}·\left\{{\left({\displaystyle \frac{p_{Cabin}}{p_{ram}}}\right)}^{{\displaystyle \frac{\kappa -1}{\kappa }}}-1\right\}$$

(4)

The ram air temperature $T_{ram}$ is higher than the ambient temperature due to compression in the ram air inlet. The ram air temperature is calculated based on the flight speed and the ambient temperature $T_{amb}$ as follows:

$$T_{ram}=T_{amb}·\left(1+{\displaystyle \frac{\kappa -1}{2}}·M{a}^2\right)$$

(5)

For cruise flight ($T_{amb}=-56$ °C $=217$ K, $Ma=0.81$, $\kappa =1.4$) this results in a ram air temperature of $245.5$ K = $-27$ °C. After entering the ram air duct, the air is compressed to a certain pressure level depending on the altitude. The resulting pressure ratio is calculated using the following equation:

$$\pi ={\displaystyle \frac{p_{Cabin}}{p_{ram}}}$$

(6)

where $p_{cabin}$ is the pressure in the aircraft cabin. For the purpose of this analysis, it is assumed that the cabin pressure is equal to the pressure in the ram air duct + 0.2 bar up to an altitude of 8000 feet (approx. 2440 m). Above 8000 ft a constant cabin pressure of 0.78 bar is assumed [41]. The pressure in the ram air duct $p_{ram}$, as well as the ram air temperature, is calculated as a function of airspeed and ambient pressure $p_{amb}$ as follows:

$$p_{ram}=p_{amb}·\left(1+{\displaystyle \frac{\kappa -1}{2}}·M{a}^2\right)$$

(7)

According to Equation (4), the electrical power required for compression depends directly on the mass flow rate of the air to be conveyed. This flow rate is first determined by the following equation:

$${\dot{m}}_{nom}={\dot{m}}_{min,perocc}·n_{occ}$$

(8)

According to CS 25.831 [41], the minimum fresh air supply per person ${\dot{m}}_{min,perocc}$ must be at least 0.25 kg/min. Multiplying this by the total number of aircraft occupants $n_{occ}$,

$$n_{occ}=n_{pax}+n_{cabincrew}+n_{flightcrew}$$

(9)

gives a required fresh air mass flow of 0.725 kg/s for the A320-200 example ($n_{occ}=174$). In addition to providing fresh air to aircraft passengers, it is also important to ensure that the heat loads generated in the cabin are removed by the cabin airflow to maintain acceptable cabin temperatures. To prevent the cabin temperature from rising due to the continuously generated heat loads, the air must be introduced into the cabin at a specific inlet temperature $T_{in}$. This inlet temperature is calculated based on the total heat load ${\dot{Q}}_{total}$, the desired cabin temperature $T_{cabin}$ (assumed to be 24 °C in this paper), and the minimum air mass flow rate ${\dot{m}}_{nom}$, according to the CS 25.831 guidelines, assuming a steady-state heat balance as defined by the following [23]:

$${\dot{Q}}_{total}={\dot{m}}_{nom}·c_p·(T_{cabin}-T_{in})⟹T_{in}=T_{cabin}-{\displaystyle \frac{{\dot{Q}}_{total}}{c_p·{\dot{m}}_{nom}}}$$

(10)

According to the SAE ARP85G guidelines [42], the calculated cabin inlet temperature should be within the range of $T_{in}^{min}=18$ °C and $T_{in}^{max}=29$ °C [15]. If temperature control requires inlet temperatures outside the allowable range, the mass flow must be adjusted. The corrected mass flow ${\dot{m}}_{cabin}$ can be estimated as follows, based on [23]:

$$\begin{array}{c}\hfill {\dot{m}}_{cabin}=\left\{\begin{array}{cc}{\dot{m}}_{nom},\hfill & \mathrm{wenn}{T}_{in}^{min}<T_{in}>T_{in}^{max}\hfill \\ {\displaystyle \frac{{\dot{Q}}_{total}}{c_p·(T_{cabin}-T_{in}^{min})}},\hfill & \mathrm{wenn}{T}_{in}<T_{in}^{min}\hfill \\ {\displaystyle \frac{{\dot{Q}}_{total}}{c_p·(T_{in}^{max}-T_{cabin})}},\hfill & \mathrm{wenn}{T}_{in}>T_{in}^{max}\hfill \end{array}\right.\end{array}$$

(11)

The total heat load to be removed from the cabin, ${\dot{Q}}_{total}$, consists of several individual heat loads:

$${\dot{Q}}_{total}={\dot{Q}}_{met}+{\dot{Q}}_{system}+{\dot{Q}}_{solar}+{\dot{Q}}_{conduction}$$

(12)

${\dot{Q}}_{met}$ describes the metabolic heat load given off by the human body. This load varies and can be estimated as follows:

$${\dot{Q}}_{met}={\dot{Q}}_{pax}·n_{pax}+{\dot{Q}}_{flightcrew}·n_{flightcrew}+{\dot{Q}}_{cabincrew}·n_{cabincrew}$$

(13)

Standard values for human metabolic heat emissions can be used according to [24] and can be found in

Table 4. Assumed metabolic heat loads from the literature [23,24].

Heat Load	Formula Symbol	Value
Heat load from one passenger	${\dot{Q}}_{pax}$	70 W/Pax
Heat load from one flight crew member	${\dot{Q}}_{flightcrew}$	100 W/Pax
Heat load from one cabin crew member	${\dot{Q}}_{cabincrew}$	200 W/Pax

.

In addition to the passengers, the systems in the cabin also emit heat. These loads mainly include the electrical equipment of the IFE, the galley, and the passenger lighting.

$${\dot{Q}}_{system}=({\dot{Q}}_{ife}+{\dot{Q}}_{galley}+{\dot{Q}}_{cabinlights})·n_{pax}$$

(14)

As precise values for the individual system heat loads are not given in the literature, a value of 40 W/Pax, as given by [43], is used for estimation.

${\dot{Q}}_{solar}$ describes the amount of heat generated by sunlight hitting the cabin windows. This heat flow can be estimated according to [24,34] using the following equation:

$${\dot{Q}}_{solar}=q_{sun}·A_{window}·n_{window}$$

(15)

where $q_{sun}$ is the specific heat load, assumed to be 1367 W/m², which is the solar energy emitted by the sun and measured on Earth Lampl.2021. The size of an aircraft window $A_{window}$ is assumed in this study to be 0.08 m², regardless of the aircraft type. We came up with this size by measuring an A320-200 window. And the number of aircraft windows $n_{window}$ is determined from aircraft images. In addition to solar radiation, there is also conductive heat transfer between the cabin and the aircraft environment due to temperature differences. This conductive heat flow ${\dot{Q}}_{conduction}$ can be determined according to [23] using the following equation:

$$\begin{array}{cc}\hfill {\dot{Q}}_{conduction}& ={\displaystyle \frac{1}{R}}·A_{skin}·(T_{skin}-T_{Cabin})\hfill \\ \hfill T_{skin}& =T_{amb}·(1+0.18·M{a}^2)\hfill \\ \hfill R& ={\displaystyle \frac{1}{{\alpha }_0}}+R_{skin}+{\displaystyle \frac{1}{{\alpha }_{cabin}}}\hfill \end{array}$$

(16)

where $T_{skin}$ is the surface temperature of the aircraft, which depends on the ambient temperature and the Mach number. The fuselage heat transfer resistance is described by the variable R and depends on the aircraft skin heat transfer coefficient ${\alpha }_0$, the aircraft skin heat transfer resistance, and the cabin wall heat transfer coefficient ${\alpha }_{Cabin}$. The heat transfer coefficient $\alpha$ describes the heat transport in the boundary layer as a function of the temperature and velocity fields [44]. ${\alpha }_0$ represents the convective heat transport on the outer surface of the aircraft and ${\alpha }_{Cabin}$ on the inner surface of the aircraft skin. $R_{skin}$ is the average heat transfer resistance of the fuselage and depends on the materials used. For the aircraft models considered, a heat transfer resistance of $R_{skin}=3.5$ m²K/W is assumed [45].

As previously described, the cabin air mass flow comes from the mixing unit, where the air from the climate packs is mixed with the recirculated air. In modern commercial aircraft, the recirculation ratio $ϵ$ is between 40 and 60%. The fresh air mass flow rate conditioned by the climate packs can be calculated using the following equation:

$${\dot{m}}_{fresh}=(1-ϵ)·{\dot{m}}_{cabin}$$

(17)

Since the certification directive requires a mass flow of ${\dot{m}}_{nom,Pax}=0.25$ kg/s per passenger minimum fresh air, as stated in Equation (8), the air mass flow of the climate packs ${\dot{m}}_{fresh}$ may need to be adjusted. This adjustment is determined as follows:

$${\dot{m}}_{Packs}=max({\dot{m}}_{fresh},{\dot{m}}_{nom})$$

(18)

5.2. Estimating the Electrical Power Required for the Recirculation Fan

Fresh air is introduced into the cabin and extracted from under the seats [46]. Some of this extracted air is recirculated and cleaned by various filtering systems. The recirculated air is driven by fans. The fan power is calculated using the following formula:

$$P_{Fan}={\displaystyle \frac{w_{t,12}·\dot{m}}{\eta }}$$

(19)

where $\dot{m}$ is the air mass flow to be delivered to the mixer. In this paper, it is assumed that the recirculation rate is 50%. $\eta$ describes the fan efficiency, which can be estimated to be 0.7 based on manufacturer’s data [47]. $w_{t,12}$ is the required work. Assuming isentropic compression, it can be calculated using the following equation, according to [48]:

$$w_{t,12}={\displaystyle \frac{n}{n-1}}·{\displaystyle \frac{p_1·V_1}{m}}·\left({{\displaystyle \frac{p_2}{p_1}}}^{{\displaystyle \frac{\kappa -1}{\kappa }}}-1\right)$$

(20)

According to this formula, the work $w_{t,12}$ must be performed for the volume change from state 1 to state 2. Assuming isentropic compression, $n=\kappa$ and, thus, the adiabatic index ($\kappa =1.4$). $p_1$ and $V_1$ are the pressure and volume before compression and $p_2$ is the pressure after compression. Assuming that the recirculating air can be treated as an ideal gas, Equation (20) is simplified to the ideal gas equation $pV=mRT$, as follows:

$$w_{t,12}={\displaystyle \frac{\kappa }{\kappa -1}}·R·T_1\left({{\displaystyle \frac{p_2}{p_1}}}^{{\displaystyle \frac{\kappa -1}{\kappa }}}-1\right)$$

(21)

where $T_1$ is the cabin temperature. The pressure increase $p_2/p_1$ is estimated in this work with a factor of 1.1.

This results in a power model that can calculate the required power based on the given aircraft configuration and flight phase.

5.3. ATA 22, 23, 27, 31, 34: Avionic Systems

The electrical power consumption of different avionics systems can be calculated together. The electrical power consumption is presumably not dependent on the size of the aircraft but on the generation of the aircraft. Newer commercial aircraft use different avionics systems than older generations of aircraft. In [49], the electrical power requirements of an A380-800 are listed. Assuming that an electrified A320-200 uses the same concept as the A380-800, the power requirement would be as follows:

$$\begin{array}{cc}\hfill P_{Avionik}& =P_{ATA-22}+P_{ATA-23}+P_{ATA-27}+P_{ATA-31}+P_{ATA-34}\hfill \\ \hfill & =255+1097+16+1406+917\hfill \\ \hfill & =3691\mathrm{W}\hfill \end{array}$$

(22)

Therefore, the electrical power requirement for avionics systems is estimated to be approximately 4 kW, regardless of aircraft size.

5.4. ATA 25: Equipment

The electrical power consumption of galleys is primarily determined by the number of passengers $n_{pax}$. A fixed value for power consumption per passenger $P_{W/pax}$ can be assumed. [50] assume a power requirement of 320 W per passenger, while [24] assumes an electrical power of 250 W. For a conservative electrical power estimation, the larger of the two values is used in this paper. The power consumption is calculated in this paper by using the following formula:

$$P_{galley}=n_{pax}·P_{W/pax}·\text{Usage-Factor}$$

(23)

The usage factor depends on the flight phase and is defined by [33] for the first time. In her work, she calculated the galley power consumption by multiplying the nominal power by so-called usage factors. The usage factor can range from 0 to 1, based on statistical data. A value of 0 indicates that the system is not in use, while a value of 1 indicates that the system is operating at nominal power. The following factors were used for each flight phase:

• Ground + Taxi: 0.14
• Take-Off: 0.18
• Climb + Cruise: 0.5
• Landing: 0.18

5.5. ATA 28: Fuel System

To estimate the required electrical power, the formula provided by [9,32] is used. According to this formula, the power for each pump can be estimated as follows:

$$P_{Pump}={\displaystyle \frac{\Delta p·\dot{m}}{{\rho }_K·\eta }}$$

(24)

where $\Delta p$ is the pressure increase to be applied by the pump, calculated as the difference between the required nominal pressure $p_{Nom}$ and the tank pressure $p_{Tank}$ in the tank, $\dot{m}$ is the fuel mass flow to be pumped, ${\rho }_K$ is the fuel density and $\eta$ is the pump efficiency. The required pressure differences and mass flows can be found in [51] and are listed in

Table 5. Parameters for calculating the electrical power of the fuel pumps from [26].

Variable	Transfer Pump Value	Booster Pump Value	Unit
$\Delta p$	69	69–103	kPa
$\dot{m}$	3	2.5–5	kg/s
${\rho }_k$	0.8	0.8	kg/m3
$\eta$	0.6	0.6	-

. The density of kerosene is assumed to be a constant $0.8$ kg/dm³ and the pump efficiency is assumed to be 60%. Both assumptions are derived from the literature, as given in [9,26]. As shown in Table 5, the minimum and maximum values for mass flow and pressure rise are given for the booster pump. We assumed that the booster pumps must be able to operate at maximum capacity at all times, hence calculations are consistently based on the maximum specified values.

The total power requirement $P_{total}$ for the fuel pumps is calculated using the following equation

$$P_{total}=P_{Boosterpump,max}·n_{Boosterpumps}+P_{Transferpump}·n_{Transferpumps}$$

(25)

where $n_{Boosterpump}$ is the number of booster pumps in the fuel system and $n_{Transferpumps}$ is the number of transfer pumps. Since no information on the number of different pumps was found in the literature, the number was estimated in this paper based on the following assumptions:

• Booster pumps: Each engine is supplied with fuel throughout the flight by an active pump located either in the center tank or in one of the wing tanks.
• Transfer pumps: The number is based on the number of fuel tanks, excluding bleed tanks. It is assumed that fuel is pumped from the outer tanks to the engine feed tanks throughout the flight.

• The assumed number of pumps can be found in

  Table 6. Assumed number of pumps for the analyzed aircraft models from [51].

  	Number of Engines	Number of Booster Pumps	Number of Tanks	Number of Transfer Pumps
  A320-200	2	2	5	4
  A330-200	2	2	8	5
  A380-800	4	4	12	9
  B737-800	2	2	4	3
  B757-300	2	2	4	3
  B777-300	2	2	4	3

  .

5.6. ATA 29: Hydraulic System

As described in Section 4.5, this paper assumes that the hydraulic system is electrified and the hydraulic pumps are electrically driven. The calculated electrical pump power is dependent on the hydraulic power required for the various actuation systems such as flight control and landing gear. Several parameters are required. The required drive power is calculated using the following equation:

$$P=M_e·\omega ={\displaystyle \frac{1}{{\eta }_{hm}·{\eta }_{vol}}}·\Delta p·Q_e$$

(26)

where ${\eta }_{hm}$ is the hydromechanical efficiency, caused by friction losses, ${\eta }_{vol}$ is the volumetric efficiency, which accounts for leakage and compression losses, $\Delta p$ is the pressure difference to be provided by the pump, and $Q_e$ is the effective flow [43]. A detailed description can be found in [52,53]. As the required data are not freely available, the following approach is taken to estimate the electrical power requirement, where the total required power is obtained by summing the partial powers of the hydraulic subsystems:

$$P_{total}=\sum P_{el,Hydraulicsubsystem}$$

(27)

The electrical power required for a subsystem can be estimated using a simple equation that takes into account the number of active pumps n per system, the nominal flow rate $V_{flow}$ of each pump, the pressure difference $\Delta p$ between the required hydraulic pressure in the system and the reservoir pressure, and the pump efficiency $\eta$. Thus, for each hydraulic system, we have the following:

$$P_{el,Hydraulicsubsystem}=\sum _{i=1}^n{\displaystyle \frac{V_{i,flow}·\Delta p_{hydr}}{{\eta }_{Pump}}},$$

(28)

The pump efficiency is assumed to be 85.5% according to the manufacturer’s specifications [47]. The number of pumps, nominal flows, and pressure differentials can be obtained from AIR5005A [54] for each aircraft type considered and are listed in

Table A1. Data required to calculate the electrical power consumption for the hydraulic system [54].

	Unit	A320-200	A330-200	A380-800	B737-800	B757-300	B777-300
$V_{green}$	L/min	140.00	175.00	162.00	140.00	142.00	142.00
$V_{yellow}$	L/min	140.00	175.00	162.00	140.00	142.00	142.00
$V_{blue}$	L/min	23.00	175.00	0.00	0.00	26.50	22.70
$p_{hydr,green}$	bar	204.00	203.45	330.00	204.00	196.55	196.55
$p_{hydr,yellow}$	bar	204.00	203.45	330.00	204.00	196.55	196.55
$p_{hydr,blue}$	bar	196.00	203.45	0.00	0.00	19.66	196.55
$p_{reservoir}$	bar	3.52	3.52	3.52	3.10	3.44	3.45
$et{a}_{green}$	%	85.50	85.50	85.50	85.50	85.50	85.50
$et{a}_{yellow}$	%	85.50	85.50	85.50	85.50	85.50	85.50
$et{a}_{blue}$	%	85.50	85.50	85.50	85.50	85.50	85.50
$n_{pump,green}$	-	1	2	4	1	1	1
$n_{pump,yellow}$	-	1	1	4	1	1	1
$n_{pump,blue}$	-	1	1	0	1 (standby)	2	2

.

Some of the aircraft models studied have more than one pump per hydraulic system that is used cooperatively. They are typically operated at partial load under normal conditions. This means that the pumps are not operating at their maximum capacity and are only delivering a fraction of their full power. The required hydraulic power is shared between the pumps in the system. Estimations of the necessary hydraulic flow in systems with multiple pumps require knowledge of the hydraulic redundancy concept. This is especially influential for mixed supply system architectures, such as Airbus’s 2H2E supply architecture. For example, in the case of the A380-800, which has four pumps per hydraulic system, all four pumps operate at partial load under normal circumstances. If one of the four pumps fails, the three remaining pumps must provide more power. The required electrical power depends on the minimum power required by the system, which depends on the minimum number of pumps in operation. Therefore, it is necessary to estimate the minimum number of required pumps.

Since the architecture of the hydraulic system is known, we perform a reliability analysis to estimate the minimum number of pumps. The reliability analysis uses a fault tree analysis to identify and evaluate potential failures or faults in complex systems. It establishes the relationship between the failure or fault of the system and the identified causes [8,55,56].

The analysis begins by identifying a specific undesirable event or state that needs to be avoided. For the estimation of necessary pumps, the “Total Loss of Flight Control” is used for its clarity and public availability. To calculate the probability of a specific undesirable event, probability theory and existing failure rates of components are used. The results for different hypothetical redundancy concepts are compared to established safety targets. We estimate the used redundancy by identifying the version with the highest number of required pumps that still meets the safety target.

Performing this for the Airbus A380-800 results in Figure 6. (Note that combinations of pressure loss in one system and loss of sufficient flow in the other has been omitted for simplicity. The complete fault trees would have to consider those extensively.) Central to this is the assumption that the simultaneous loss of overall pressure on both hydraulic systems and the loss of sufficient hydraulic flow by pump loss should be in the same order of magnitude. (This is due to the fact that the system is generally only as reliable as its weakest element.) We calculate the average probability per flight hour for the considered failures. The probability of simultaneous loss of pressure in both systems is calculated by the following:

$${\lambda }_{doublepressureloss}={\displaystyle \frac{{\left(t·{\lambda }_{circ}\right)}^2}{t}}={\displaystyle \frac{t^2·{10}^{-8}}{t}}{\mathrm{FH}}^{-1}$$

(29)

based on our assumption that the loss of hydraulic occurs every 10,000 h, which corresponds to a failure rate of ${\lambda }_{circ}\approx 1·{10}^{-4}{\displaystyle \frac{1}{FH}}$. For the exposure time t a flight time of ten hours is assumed. This represents the right branch of the fault tree’s third row in Figure 6.

The left branch is arranged so that engines (noted from E1 to E4) are grouped into sub-branches with their respective pumps in the yellow system (noted as Y1 to Y4) and green system (noted as G1 to G4). The failure rate for the engine is assumed to be ${\lambda }_{eng}\approx 2.5·{10}^{-5}{\displaystyle \frac{1}{FH}}$. The failure rate for a single pump is assumed to be ${\lambda }_{pump}\approx 1·{10}^{-4}{\displaystyle \frac{1}{FH}}$. Iteratively increasing the number of pumps that can fail in this calculation results in different estimated failure rates for the simultaneous loss of multiple hydraulic pumps.

The results are given in Equations (32)–(34) for the allowable loss of three, two, and one pump per system, respectively.

The probability of loss of four hydraulic pumps can be calculated as follows:

$$\begin{array}{cc}\hfill {\lambda }_{Lossof4Pumps}& ={\displaystyle \frac{1}{t}}·{(2.5·t·{10}^{-5}+t^2·{10}^{-8})}^4\hfill \\ \hfill & ={\displaystyle \frac{1}{t}}·{(2.5·10·{10}^{-5}+{10}^2·{10}^{-8})}^4\hfill \\ \hfill & =3.97·{10}^{-16}{\mathrm{FH}}^{-1}\hfill \end{array}$$

(30)

Combining Equation (30) at the OR-Gate in Figure 6 with the probability of complete pressure loss, Equation (29), results in the following:

$$\begin{array}{cc}\hfill {\lambda }_{Doublehydfail}& \approx {\lambda }_{Lossof4Pumps}+{\lambda }_{doublepressureloss}\hfill \\ \hfill & \approx 3.97·{10}^{-16}+{10}^{-7}\approx 1·{10}^{-7}{\mathrm{FH}}^{-1}\hfill \end{array}$$

(31)

• Simultaneous loss of three hydraulic pumps:

  $${\lambda }_{Lossof3Pumps}={\displaystyle \frac{{(2.5·t·{10}^{-5}+t^2·{10}^{-8})}^3}{t}}·4\approx 6.3·{10}^{-12}{\mathrm{FH}}^{-1}$$

  (32)
• Simultaneous loss of two hydraulic pumps:

  $${\lambda }_{Lossof2Pumps}={\displaystyle \frac{{(2.5·t·{10}^{-5}+t^2·{10}^{-8})}^2}{t}}·6\approx 3.7·{10}^{-8}{\mathrm{FH}}^{-1}$$

  (33)
• Loss of one hydraulic pump:

  $${\lambda }_{Lossof1Pump}={\displaystyle \frac{2.5·t·{10}^{-5}+t^2·{10}^{-8}}{t}}·4\approx 1·{10}^{-4}{\mathrm{FH}}^{-1}$$

  (34)

The loss of two pumps (Equation (33)) matches the desired order of magnitude, as estimated in Equation (29), and is selected as the considered redundancy for the aircraft.

Similar calculations can be performed for triplex hydraulic systems. For the lack of electrical backup supply, such hydraulic systems have been analyzed under the assumption that one hydraulic circuit has to be sufficient for normal operations even at the loss of a single engine.

5.7. ATA 30: Ice and Rain Protection System

As described in Section 4.6, most of the electrical power is required to protect the leading edge of the wing and the windshield. Consequently, the focus for estimating the total electrical power consumption is on these two components.

Calculation Method for Leading Edge De-Icing

Necessary power for ice protection is commonly described per area [35]. In the MEA, the leading edge of the wing is thermally de-iced using an electrical resistance system similar to that used on the B787-800. The total power is estimated in this paper using the calculation methods of [23,24,34,36].The necessary electrical power can be estimated using the given equation:

$$P_{WIPS,el}={\displaystyle \frac{A·{\dot{q}}_{WIPS,el}}{{\eta }_{WIPS}}}$$

(35)

The electrical thermal de-icing effectiveness, denoted as ${\eta }_{WIPS}$ and assumed to be 70% according to [36], is an important factor to be taken into account. The wing’s surface area to be protected, denoted as A, must be estimated. Usable estimates for fractions of protected leading edge areas are described by [23,33].

The heat flow ${\dot{q}}_{WIPS,el}$ required for de-icing the protected wing area can be determined using the following formula:

$${\dot{q}}_{WIPS,el}={\dot{q}}_{cyc}·(1-{\kappa }_{ps})·{\kappa }_{cyc}+{\dot{q}}_{cont}·{\kappa }_{ps}$$

(36)

The protected area is divided into cyclically heated areas and continuously heated areas, with ${\kappa }_{ps}$ describing the relative amount of continuously heated area. For the model, it is assumed that, as a rule, 20% of the protected area according to [36], is continuously heated, so ${\kappa }_{ps}$ = 0.2.

The first term of Equation (36) describes the amount of heat needed for cyclical heating of the remaining regions (1 −${\kappa }_{ps}$), with ${\dot{q}}_{cyc}$ computed using Equation (37). During each cycle that lasts for $t_{cyc}$, the region is heated for a specific duration of time $t_{heat}$. The activity ratio ${\kappa }_{cyc}$ can be calculated with the formula, where ${\kappa }_{cyc}=t_{heat}/t_{cyc}$. It is assumed that the activity ratio is ${\kappa }_{cyc}=5$ %, according to [36]. The heat flow ${\dot{q}}_{cyc}$ necessary for de-icing the cyclically heated zones is determined by applying the subsequent equation:

$${\dot{q}}_{cyc}={\displaystyle \frac{{\rho }_{ice}·l_{ice}}{t_{heat}}}·c_{ice}·(T_0-T_{ambient})+L_f$$

(37)

Here, ${\rho }_{ice}$ is the density of ice, which has a value of 920 kg/m². $l_{ice}$ constitutes the presumed thickness of the ice that requires melting during each cycle. A thickness of $l_{ice}$ = 0.5 mm for ice has been assumed within this paper to calculate the total heat flow required to de-ice the protected wing area. $t_{heat}$ denotes the duration of a cycle and $c_{ice}$ = 2060 J/kgK is the specific heat of ice. $L_f$ is the latent heat of ice, which is 332,500 J/kg. This value represents the amount of heat energy required to change a given amount of a substance from the solid to the liquid state at a constant temperature.

The second term of Equation (36) describes the required power per surface area for de-icing. In general, the electrical power requirement ${\dot{q}}_{cont}$ depends on the state and the number of impinging super-cooled water droplets. The mass flow rate per wing area of water droplets impinging on the leading edge of the wing is calculated by the following formula:

$${\dot{m}}_{local}=v_{TAS}·{\rho }_{LWC}·E_m$$

(38)

Mass flow rate depends on true airspeed $v_{TAS}$ and mass of super-cooled water droplets per cubic meter of air, expressed as the liquid water content (LWC) ${\rho }_{LWC}$. LWC depends on air temperature and the average diameter of the water droplets. As suggested by [36], assuming that the water droplets have a mean diameter of 20 µm, the LWC can be obtained from a diagram in the CS-25 certification regulations [41] as follows:

$$\begin{array}{cccc}\hfill {\rho }_{LWC}& =0.615\mathrm{g}/{\mathrm{m}}^3,\hfill & \hfill \mathrm{at}& T=0°\mathrm{C}\hfill \\ \hfill & =0.42\mathrm{g}/{\mathrm{m}}^3,\hfill & \hfill \mathrm{at}& T=-10°\mathrm{C}\hfill \\ \hfill & =0.2\mathrm{g}/{\mathrm{m}}^3,\hfill & \hfill \mathrm{at}& T=-20°\mathrm{C}\hfill \\ \hfill & =0.14\mathrm{g}/{\mathrm{m}}^3,\hfill & \hfill \mathrm{at}& T=-30°\mathrm{C}\hfill \end{array}$$

(39)

Interpolation is used to determine the LWC for other temperatures. The mass flow rate depends on the water trapping efficiency, denoted as $E_m$. $E_m$ is a function of flight speed, water droplet size, viscosity, and air density, and can be calculated using a detailed formula as described in AIR 1168/4 [36]. A simplified formula is often given in the literature, as follows:

$$E_m=0.00324·{\left({\displaystyle \frac{v}{t}}\right)}^{0.613}$$

(40)

This formula depends on the flight speed v and the maximum wing thickness t. According to [36] this formula is only valid under certain conditions:

• The relative thickness of the wing is between 6 and 16%.
• The angle of attack is $\alpha =$ 4°.
• The mean diameter of the water droplets is $d_{med}=$ 20 µm.
• The flight altitude is h = 10,000 ft = 3048 m.

• At other altitudes, the water collection efficiency deviates by less than 10% from the actual result [24]. For a first estimate, this deviation is tolerated.

As previously described, aircraft surfaces can only freeze under certain conditions, one of which is that the wing surface temperature must be below 0 °C. Therefore, to prevent icing, a desired surface temperature $T_{skin}$ is set in advance for estimation. According to AIR1168 [35], a temperature between 2 and 10 °C seems reasonable for electrothermal de-icing. In this paper, the midpoint of this temperature range is chosen, giving a surface temperature of $T_{skin}$ = 6 °C. To ensure that the protected leading edge of the wing reaches this specific surface temperature, a heat flux ${\dot{q}}_{cont}$ is required, which is estimated using the following equation:

$${\dot{q}}_{cont}={\dot{q}}_{sens}+{\dot{q}}_{conv}+{\dot{q}}_{evap}+{\dot{q}}_{kin}$$

(41)

This heat flux has four components resulting from different heat transfer processes, as shown in Figure 7.

The sensible heat flux ${\dot{q}}_{sens}$ refers to heat transfer due to temperature differences between a fluid and a surface, involving conduction, convection, or radiation. The impact of super-cooled water droplets on the aircraft surface causes the surface temperature to decrease. This heat flux consists of two terms and is calculated as follows:

$${\dot{q}}_{sens}=\underset{1.\mathrm{Summand}}{\underbrace{{\dot{m}}_{local}·(T_{skin}-T_{aw})·[(1-n)·c_{liq}+n·c_{ice}]}}+\underset{2.\mathrm{Summand}}{\underbrace{n·L_f}}$$

(42)

The first term calculates the heat flux required to bring the super-cooled water droplets to the desired temperature. Here, ${\dot{m}}_{local}$ is the mass flow rate of super-cooled water droplets per unit area on the wing surface as calculated in Equation (38). $T_{skin}$ represents the target temperature for the wing’s leading edge surface (6 °C), and $T_{aw}$ denotes the adiabatic wall temperature. This temperature is an ideal assumption where there is no heat transfer between the surface and the environment by conduction, convection, or radiation. The adiabatic wall temperature depends on the recovery factor $R_c$, the adiabatic index $\kappa =1.4$, the Mach speed $Ma$, the dynamic viscosity of air $\eta$ at an assumed ambient temperature of −30 °C, the specific heat capacity $c_p$ of air, and the thermal conductivity $k_0$ of air, calculated as follows [23,24,36]:

$$\begin{array}{ccccc}\hfill \mathrm{Adiabatic}\mathrm{wall}\mathrm{temperature}:& \hfill & \hfill T_{aw}& =T_{ambient}·(1+R_c·{\displaystyle \frac{\kappa -1}{2}}·M{a}^2)\hfill & \end{array}$$

(43)

$$\begin{array}{ccccc}\hfill \mathrm{Recovery}\mathrm{factor}:& \hfill & \hfill R_c& =1-0.99·(1-P{r}^{0.5})\hfill & \end{array}$$

(44)

$$\begin{array}{cccc}\hfill \mathrm{Prandtl}\mathrm{number}:& \hfill & \hfill Pr& ={\displaystyle \frac{\eta ·c_p}{k_0}}\hfill \end{array}$$

(45)

$$\begin{array}{cccc}\hfill \mathrm{Dynamic}\mathrm{viscosity}:& \hfill & \hfill \eta & =1.5636·{10}^{-5}\mathrm{kg}/\mathrm{ms}\hfill \end{array}$$

(46)

$$\begin{array}{ccccc}\hfill \mathrm{Specific}\mathrm{heat}\mathrm{capacity}:& \hfill & \hfill c_p& =1003.5\mathrm{J}/\mathrm{kgK}\hfill & \end{array}$$

(47)

$$\begin{array}{cccc}\hfill \mathrm{Thermal}\mathrm{conductivity}:& \hfill & \hfill k_0& =0.0277\mathrm{W}/\mathrm{mK}\hfill \end{array}$$

(48)

The required heat flux is also affected by the composition of the super-cooled water droplets. The freezing fraction n in the droplets is determined by the following equation:

$$n={\displaystyle \frac{c_{liq}·(T_0-T_{ambient})}{L_f}}$$

(49)

Here $T_0$ = 273.15 K corresponds to the freezing temperature of the water, and $T_{ambient}$ varies with altitude. The specific heat capacity of water $c_{liq}$ and ice $c_{ice}$ is the amount of heat energy required to raise the temperature of one kilogram of water or ice by one degree Kelvin. For water, $c_{liq}$ equals 4190 J/kgK and for ice $c_{ice}$ equals 2060 J/kgK. The second term calculates the heat required to melt the ice component in the droplets. The convective heat flux ${\dot{q}}_{conv}$ refers to heat transfer by convection, a mechanism by which heat is transferred by the movement of a fluid. Super-cooled water droplets heat up by convection, resulting in a reduction in the leading edge temperature. A heat flux is required to counteract the surface cooling and can be calculated using the following equation [23,24]:

$${\dot{q}}_{conv}=h_0·(T_{skin}-T_{aw})$$

(50)

$h_0$ is the local heat transfer coefficient, which indicates how effectively heat energy is transferred from the surface to a fluid, such as air. The coefficient depends on several factors, including the velocity of the fluid and the temperature difference between the surface and the fluid. The heat transfer coefficient is determined by the following formula:

$$h_0=Nu·{\displaystyle \frac{k_0}{0.5·c_{slat}}}$$

(51)

Here, $k_0$ is the thermal conductivity of air at 255.3 K and has a value of 0.0227 $\mathrm{W}/\mathrm{mK}$ [36]. $c_{slat}$ is the average depth of the slats on the wing, more details can be found in [23]. The Nusselt number $Nu$ is a dimensionless parameter describing the convective heat transfer between a surface and a flowing fluid. It is determined by the equation

$$Nu=0.0296·R{e}^{4/5}·P{r}^{1/3}$$

(52)

where $Pr$ is the Prandtl number, see Equation (45), and $Re$ is the Reynolds number, another dimensionless parameter that indicates the ratio of inertial to viscous forces in the fluid. Since the flow type fundamentally influences heat transfer, this number is an important parameter. Using the air density $\rho$, the true airspeed v, the average slat depth $c_{slat}$, and the dynamic viscosity $\eta$ (see Equation (46)), the Reynolds number is calculated as follows:

$$Re={\displaystyle \frac{\rho ·v·\left(0.5·c_{slat}\right)}{\eta }}$$

(53)

The evaporative heat transfer ${\dot{q}}_{evap}$ is a phenomenon that happens during the process of evaporation, where liquid water transforms into a gaseous state. This phase change requires heat energy. The heat loss resulting from evaporation at the wing’s leading edge can be quantified using the subsequent equation:

$$\begin{array}{c}\hfill {\dot{q}}_{evap}=L_e·{\dot{m}}_{evap}\end{array}$$

(54)

Assuming that the water droplets fully evaporate, a specific latent heat of $L_e$ = 2257 kJ/kg is required for water evaporation [24]. The mass flow rate ${\dot{m}}_{evap}$ that needs to be heated depends on several factors and can be calculated using the following equation [23]:

$${\dot{m}}_{evap}=min\left(0.7·{\displaystyle \frac{h_0}{c_p·p_{ambient}}}·(p_v\left(T_s\right)-RH·p_v\left(T_{ambient}\right),{\dot{m}}_{local}\right)$$

(55)

$h_0$ describes the local heat transfer coefficient as calculated using Equation (51). $c_p$ denotes the specific heat capacity of air, and $p_{ambient}$ is the ambient pressure. $RH$ is assumed to be 1 (100% humidity). The saturation vapor pressure $p_v$ is calculated based on temperature T as described in [23]:

$$p_v\left(T\right)=2337·exp\left\{6789·\left({\displaystyle \frac{1}{293.15}}-{\displaystyle \frac{1}{T}}\right)-5.031·ln\left({\displaystyle \frac{T}{293.15}}\right)\right\}$$

(56)

Kinetic heat transfer ${\dot{q}}_{kin}$ refers to heat transfer resulting from motion. It arises due to the interaction of fluid molecules with the surface, leading to the transfer of heat energy. The amount of heat flow is influenced by several factors, including the airspeed of the aircraft referred to as $v_{TAS}$. Therefore, as a consequence of acceleration, water droplets hitting the leading edge of the aircraft encounter kinetic heating. The following equation can be used to calculate the heat flow:

$${\dot{q}}_{kin}=-{\dot{m}}_{local}·{\displaystyle \frac{v_{TAS}^2}{2}}$$

(57)

Calculation method for ice and fog protection of cockpit windows

• Since most commercial aircraft use the electrical resistance system, it is further examined in this work. The electrical power needed for anti-icing the cockpit window can be approximated based on AIR 1168 [35] by using the following:

$$P_{Frontshield}=P_{AIR1168}·A_{window}$$

(58)

$A_{window}$ denotes the area to be anti-iced measured in square inches, and $P_{AIR1168}$ signifies the necessary electrical power per anti-iced area according to the SAE regulation. To estimate the required electrical power, an average power consumption of 3.75 W/in² is assumed, and the area to be anti-iced is estimated from publicly available aircraft geometry parameters, as no public data on window sizes are available. The areas can be found in

Table 7. Estimated size of the cockpit windows (all side windows plus windshields).

Aircraft Types	Assumed Cockpit Windows Size in m2
A320-200	1.48
A330-200	1.64
A380-800	1.54
B737-800	0.94
B757-300	1.16
B777-300	1.22

. This results in the electrical power required for ice and fog protection of the cockpit window during the entire flight.

5.8. ATA 33: Lighting System

As outlined in Section 4.7, the lighting system is divided into internal and external lighting. For general cabin lighting as part of internal lighting, LED strips are used to illuminate both the cabin’s ceiling and walls. The quantity of LED strips varies depending on the seating arrangement. For a single-aisle configuration, like that of the A320-200, four LED strips are installed. In a two-aisle seating configuration, six LED strips are installed accordingly. The required power for this can be determined using the following equation:

$$P_{Cabin,light}=n_{strips}·P_{led,tube}·{\displaystyle \frac{l_{cabin}}{l_{led,tube}}}$$

(59)

According to the manufacturer’s specifications, a frequently used LED tube measures $l_{led,tube}=1.83$ $\mathrm{m}$ in length and requires an electrical power of $P_{led,tube}=31.5$ $\mathrm{W}$ [57]. The power estimation for reading lamps can be computed using the following:

$$P_{reading,light}=n_{Pax}·P_{reading,light,pax}$$

(60)

Every passenger has an individual reading lamp. It requires an electrical power $P_{reading,light,pax}$ of $3.2$ W according to manufacturer’s information [58]. Indicator lights are necessary for the “fasten seat belt” and “no smoking” signals for each row of seats. Assuming three passengers sit in each row of seats. For the indicator lights, an electrical power of $P_{indicatorsign,pax}=1.4 W$ is assumed [59]. This leads to the subsequent electric power demand, as follows:

$$P_{indicatorsigns,light}={\displaystyle \frac{n_{pax,max}}{3}}·P_{indicatorsign,pax}$$

(61)

External lighting can be divided into two types: lights used throughout the entire flight, and lights used solely during takeoff and landing. The first category includes various positions and warning lights, which are described as follows:

• 3 strobe lights, located at each wingtip and the aircraft tail. The power requirement can be estimated using the following equation:

$$P_{Strobe,lights}=n·P_{Strobe}=3·4\mathrm{W}=0.012\mathrm{kW}$$

(62)

A strobe light requires a maximum of 4 watts, resulting in a constant consumption of 0.012 kW.

• 2 anti-collision beacon lights, which are mounted on both the top and bottom of the aircraft and blink at a specific frequency throughout the entire flight. The power can be estimated using the following equation:

$$P_{Beacon,lights}=\sum f_{flash,pulse}·E_{Beacon}$$

(63)

On the top of the aircraft, there is a red warning light that requires 35 J per flash. According to CS 25.1404c, a flash rate between 40 and 100 times per minute is required. For the electrical power estimation in this paper, an average of $f_{flash,pulse}$ = 70 per minute is assumed. The white warning light mounted on the bottom of the aircraft requires 35 joules per flash and specifies a flash rate of 60 flashes per minute [60]. This results in the following consumption:

$$\begin{array}{cc}\hfill P_{\mathrm{beacon},\mathrm{lights}}& =70{\min }^{-1}·35\mathrm{J}+60{\min }^{-1}·35\mathrm{J}=4550\mathrm{J}/\min =0.0758\mathrm{kW}\hfill \end{array}$$

(64)

For the wing and engine scan lights, a total of two lamps are required. These lights are used for visual identification of ice on the wing and engine. Each light requires an electrical power of 600 W [16]. Their power requirement can be estimated using the following equation:

$$P_{wing,engine,lights}=n·P_{wing,engine,light}=2·600\mathrm{W}=1.2\mathrm{kW}$$

(65)

The lighting used exclusively during the takeoff and landing phase includes landing lights, take-off lights, taxi lights, and lights that illuminate the airline logo. For the headlights and logo lights, there are fixed values in the literature [16,61], as follows:

• 2 landing lights

  $$P_{landing,lights}=n·P_{landing,light}=2·600\mathrm{W}=1.2\mathrm{kW}$$

  (66)
• 1 take-off light

  $$P_{take-off,lights}=n·P_{take-off,light}=1·600\mathrm{W}=0.6\mathrm{kW}$$

  (67)
• 1 taxi light

  $$P_{taxi,lights}=n·P_{taxi,light}=1·400\mathrm{W}=0.4\mathrm{kW}$$

  (68)
• 2 logo lights

  $$P_{logo,lights}=n·P_{logo,light}=2·60\mathrm{W}=0.12\mathrm{kW}$$

  (69)

5.9. ATA 38: Water and Waste System

As outlined in Section 4.8, the water and wastewater system is subdivided into three subsystems. The potable water system uses electrical power to warm the water for the lavatories. A boiler, one for each water faucet, is used to accomplish this. Each lavatory boiler needs an electrical power of approximately $P_{boiler,lavatory}=230$ W (Lavatory Heater Airbus P/N 24E507009G03). The number of boilers required depends on the number of lavatories installed, which is determined by the number of passengers and airline specifications. The estimated lavatory count in

Table 8. Assumed number of lavatories.

Aircraft Type	Assumed Number of Lavatories
A320-200	3
A330-200	6
A380-800	17
B737-800	3
B757-300	4
B777-300	10

is based on Lufthansa and United’s cabin arrangements. The necessary power can be calculated as follows:

$$\begin{array}{cc}\hfill P_{Water,Waste}& =n_{boiler,lavatory}·P_{boiler,lavatory}\hfill \end{array}$$

(70)

The wastewater system consists of drainage lines and heated drain openings. A determination of the necessary electrical power was not possible due to a lack of data.

The toilet system requires electrical power for the vacuum generator. It creates a vacuum in the toilet tank up to a cruising altitude of 16,000 ft. Above flight altitudes of 16,000 ft, the vacuum is supplied using differential pressure. Due to the short flight time to FL160 and the possible use during parking, it is assumed in this paper that the energy consumption can be neglected.

5.10. ATA 44: Cabin Systems

As outlined in Section 4.9, this ATA chapter encompasses different systems such as CIDS and IFE. As there are only calculation methods available for IFE in the literature, this paper concentrates solely on the electrical power needs of this system. Theoretical, frequently referenced current demands for IFE vary from 80 to 110 W for each passenger, as demonstrated in

Table 9. Literature-based power requirements.

Source	Aircraft	El. Demand in kW	Power/Pax in W/Pax	Year of Publication
[62]	A380	50–60	90–110	2016
[26]	300 Pax	30	100	2002
[50]	B787	20	80	2011

.

Updated estimations from 2019 indicate that the electricity needed for the in-flight entertainment system and charging electronic devices is notably reduced. The authors of [63] employed data from passenger surveys and airline statistics to examine the media preferences of passengers during their flights. The findings indicated that 20% of respondents reported traveling for business purposes. For this group, it was assumed in the study that they would use their laptops during the flight and recharge them on board. The rest of the passengers spent 44% of the flight time using IFE and 46% using personal electronic devices (PEDs). In flight, 10% of passengers switched between IFE and PEDs. The study also presumed that passengers utilizing PEDs would charge them during the flight.

To determine the electrical power demand of the IFE, it is necessary to obtain the power consumption data during the charging process. The study shows that laptops consume 70 W, while PEDs consume 18 W. As for IFE, only the screen integrated into the back of the seats requires power as IFE data are transmitted wirelessly. It is assumed that screens consume 15 W. For a prudent approximation, we assume that the displays draw 15 W throughout the flight, irrespective of whether the passenger utilizes IFE. Moreover, a 15% safety margin is accounted for to ensure a system availability of more than 99% [64].

As a result, the power per passenger ratio can be calculated using the following equation:

$$\begin{array}{cc}\hfill P_{IFE,Pax}& =(0.2·70\mathrm{W}/\mathrm{P}\mathrm{a}\mathrm{x}+0.8·0.46·18\mathrm{W}/\mathrm{P}\mathrm{a}\mathrm{x}+15\mathrm{W}/\mathrm{P}\mathrm{a}\mathrm{x})·1.15\hfill \\ \hfill & \approx 41\mathrm{W}/\mathrm{P}\mathrm{a}\mathrm{x}\hfill \end{array}$$

(71)

Ref. [24] also assumes a power consumption value of this magnitude for IFE (50 W/Pax).

6. Results

The electrical power consumption estimation for individual aircraft systems is presented here, contingent on the aircraft model and flight phase. The calculated results are later compared with values obtained from the literature.

6.1. ATA 21: Environmental Control System

The power demand for the Aircraft Conditioning System (ACS) is shown in Figure 8 over the different flight phases. The power consumption ranges from around 25 kW during Taxi Out for the A320-200 to 650 kW during cruise for the A380-800. As shown, the computed power varies according to the flight stage, with the highest electric power usage occurring during the cruise phase for all studied aircraft models. The electric power used, as specified by Equation (4), is reliant on mass flow and pressure ratio. Figure 9 plots these two parameters throughout the flight mission. The pressure varies significantly according to the flight phase, whilst the mass flow that the ACS must provide is more or less constant and varies slightly upon both the flight phase and the type of aircraft.

Figure 10 illustrates the power consumption of the recirculation fan during different flight phases. The power usage varies from approximately 18 kW during “Climb 3” for the A320-200 to nearly 65 kW during “Descent 3” for the A380-800. The minimum electrical power is needed during the “Climb 3” stage. As seen in the figure, the electrical power requirement for the fan increases as the aircraft descends.

The total required electrical power for the ECS is provided in Figure 11. To enhance comparability, the total required power is normalized per passenger. This ratio is shown in Figure 12 across different flight phases. The power/pax ratio ranges from approximately 0.25 kW per passenger during “Taxi Out” for the A320-200 to 1.28 kW per passenger during cruise for the B757-300.

A literature review was conducted to compare these results with available research data. The results of the literature review are presented in

Table 10. Values given for ECS power consumption in the literature.

Source	Aircraft	Number of Pax	Power Specification [kW]	Power/Pax [kw/Pax]	Electrified?
[24]				0.135	Yes
[66]		100	150	1.5 *	Yes
[67]		350	400	1.14 *	Yes
[65]		300	597	1.99 *	Yes
[65]		600	632	1.05 *	Yes
[68]	∼A320		129–204	0.76–1.2 *	Yes
[23]				1.14	Yes

* Derived values.

. All values marked with an asterisk in the table are derived values that can be calculated from values found in the literature. None of the researched papers provided information on the specific flight phases for the given values. As shown in the table, power/pax ratios vary from 135 W/Pax according to [24] to 1.99 kW/Pax according to [65]. The majority (57%) of literature values fall within the range of 1 to 1.5 kW/Pax. Compared to the power/pax ratios calculated in this paper, a considerable match exists with the flight phases of “Climb 3” and “Cruise”. The alignment between calculations and published data suggests the consistency of the applied methodology.

6.2. ATA 25: Equipment

The power usage for the galley is shown in Figure 13. The data illustrates the electrical power requirement during each flight phase. The needed power ranges from approximately 8 kW during “Taxi Out” to nearly 86 kW while cruising on an A380-800. This calculation is only an approximation. Actual power usage relies on a range of hard-to-predict factors, including the hot and cold food and beverages provided, the level of in-flight service, as well as operational factors such as the duration and stage of the flight.

The authors of Ref. [50] specified an electrical requirement of 120 kW for an aircraft with 375 passengers. Compared to the electrical demands calculated in this paper, the literature value was comparable to that of a B777-300 (383 passengers). For the B777-300, an electrical power requirement of 62 kW was calculated. The difference amounted to 58 kW.

Due to the scarcity of available data and the previously described numerous factors influencing the actual power, it is not possible for this study to offer any assertion regarding the precision of the computed electrical power requirements for the equipment systems of ATA 25.

6.3. ATA 28: Fuel System

The power requirements for the fuel system are presented in Figure 14. The power output ranges from 3.45 kW to 8.2 kW. The power demand is assumed to remain constant across all flight phases.

Upon comparing these calculated values to those found in the literature, there is a significant deviation. The usage of uniform pump types across aircraft models does not yield very accurate results. The modeling shows significant weaknesses that fail to account for relevant parameters. Also, the calculations are based on assumptions about the number of active pumps in the system, which may prove mistaken. For example, [65] specifies an electrical power consumption of 12.8 kW for an A320 aircraft. However, the calculations in this paper deviate from this literature value by almost 70%. These differences highlight the necessity of a thorough review of modeling and assumptions to more accurately assess the actual electrical power demands for fuel supply in aircraft within the detailed design phase.

6.4. ATA 29: Hydraulic System

The graph in Figure 15 illustrates the total electrical power necessary to operate the electrified hydraulic pumps in various airplane models. As demonstrated, the power demand ranges from 109 kW for the B757-300 to 412 kW for the A380-800. It is assumed that the condition remains constant throughout all flight phases.

The computed electrical power requirement to provide the hydraulic system in Airbus aircraft concurs with values found in the literature. For the Airbus A320, a value of 127 kW was obtained from the calculation, while the literature reported a value of 125 kW, as presented in

Table 11. Required pump capacity from the literature.

Source	Aircraft	Demand	Year of Publication
[62]	A320-200	125 kW	2016
[62]	A340	240 kW	2016
[62]	A380	368 kW	2016

. On the other hand, for the Airbus A380, the calculated value was 412 kW, whereas the literature cited a value of 368 kW. The results illustrate a tendency for the calculated values to lie within the same power range as the literature values, with an agreement of $\pm 10\%$.

In contrast, the calculated electrical power requirements for Boeing aircraft, particularly the B757-300 and B777-300, are remarkably low. The power magnitude of the B737-800, which is nearly the same size as the A320-200, is comparable to the A320 (B737-800: 110 kW, A320: 126 kW). However, in contrast, the B777-300, similar in length to the A340, displays a significantly lower value (B777-300: 124 kW) when compared to a value found in the literature for the A340 (240 kW).

The deviations of hydraulic power from the anticipated dependency on aircraft size in Boeing planes, combined with the absence of comparable values in the literature, raise questions about whether the assumptions for parameter determination or the modeling approach are inaccurate, suggesting potential deficiencies in the modeling. Accurately computing the electrical power for hydraulic pumps, and clarifying these inconsistencies may necessitate the consideration of all hydraulic loads applied, including actuator loads throughout diverse flight stages.

Additionally, differing redundancy concepts of aircraft manufacturers may also cause a significant distortion of the power estimations.

6.5. ATA 30: Ice and Rain Protection System

Assuming the presence of “Icing Conditions”, the electrical power needed to de-ice the wing leading edges can be obtained from Figure 16. The required electrical power in kW is shown for different flight phases. It is noteworthy that the de-icing system remains inactive during the “Climb 3”, “Cruise”, and “Descent 1” phases. As defined in Section 3.5, it was assumed that the aircraft operates at cruising altitudes where “icing conditions” are not present. All other flight phases demonstrate a constant power requirement. The power range was calculated to span from 25 kW for the A320-200 to 149 kW for the B777-300 [9]. Figure 17 shows the relation between passenger number and maximum required power for de-icing in kW.

For an A330 or B787 aircraft, [9] stated in his dissertation that an electrical power requirement of 100 kW is needed for the electro-thermal de-icing of the wing leading edge, without specifying the size of the wing leading edge to be de-iced. In this paper, an electrical power requirement of nearly 123 kW was calculated for the A330-200. This results in a relative deviation of about 19%.

For the de-icing of a B787-800 using the electrical–thermal de-icing system, [50] reported that the electrical power requirement ranges from 30 to 85 kW per wing. [69] provided a range of 45–75 kW for the same type of aircraft. The protected wing area of approximately 13.5 m² for a B787-800 is comparable to that of an A380-800 (13.3 m²). Although the A380 is significantly larger in terms of wing area than the B787, the area to be de-iced is roughly the same. This is due to the fact that only the slat adjacent to the outer engine is de-iced on the A380, whereas a total of four slats per wing are de-iced on the B787. Further details can be found at [23]. The power requirement estimated from our model for the A380-800 is about 52 kW for both wings, which can be seen in Figure 16 and falls into the literature range.

For de-icing the wing leading edge of an A320-200 aircraft, [68] provides an estimate of the electrical power required. The ground, takeoff, and landing phases require 17 kW, while the climb, cruise, and descent phases require 25 kW. According to our models, the “Taxi Out—limb 2” and “Descent 2—Taxi-In” phases require an estimated power of around 25 kW. The estimated power requirements differ by about 0–32%. A noteworthy difference exists between the assumed system activity in our model and [68]. In [68] anti-icing system is assumed to be active during cruise flight, while our model assumes that commercial aircraft fly at altitudes where icing conditions are not likely present and can, therefore, be dismissed. The implied activation of the anti-icing system during cruise flight, as depicted in the literature source, raises questions. The calculated power range is mostly consistent with the literature’s researched values, but differences persist, especially in the assumed system activity over the flight mission. The power requirements calculated in this study can serve as an initial anticipated power range.

The required electrical power for ice and fog protection of the cockpit windows is shown in Figure 18. Depending on the aircraft type, the calculated and rounded power requirements range from 5.5 kW for the B737-800 to nearly 10 kW for the A330-200.

The number of references discussing the amount of electrical power necessary for ice and fog protection of cockpit windows in the existing literature is limited. Only two references could be identified. According to [65], the electrical power requirement for the DC10 is 7.2 kW. The area of the protected windows in the cockpit is estimated to be 1.64 m² based on aircraft geometry data, which is comparable to that of the A330-200 as depicted in Table 7. For an A330-200 our model estimates a power requirement of 9.5 kW. [9] provided similar results. In Schlabe’s dissertation, an electrical power requirement of 6 kW was assumed for de-icing an unspecified-sized aircraft windscreen. The apparent deviation of the calculated values from the literature is likely due to the estimation of cockpit window size. As no public data on window sizes are available, the area used in this paper was determined based on freely available aircraft images. The window size drives the required electrical power. Therefore, the calculated values only serve as an initial assumption.

6.6. ATA 33: Lighting System

The total electrical power necessary for the lighting system during all flight phases is demonstrated in Figure 19. The power requirements vary from under 4 kW for the B737-800 to 15 kW for the A380-800. As depicted, the power requirements fluctuate according to the flight stage, with the highest electrical power deemed necessary during the “Approach Initial” to “Touchdown” phases for every aircraft evaluated. This variation transpires due to the existence of externally switchable lighting. As outlined in Section 5.8, there exist lights specifically allocated for takeoff and landing procedures. The categories of these lights include landing, takeoff, taxi, and logo lights, which serve their respective purposes.

Only a few references regarding the nominal power requirements were found in the literature. Ref. [65] looked into the electric demand for two aircraft. For an aircraft capable of transporting 600 passengers, a power requirement of 22.7 kW was reported, resulting in 37 W per passenger. The power requirements in this paper were normalized for comparison with the literature value. The results are displayed in Figure 20. The power required per passenger ranges from just under 22 W/Pax for a B757-300 in the “Climb 1—Descent 3” phases, to nearly 33 W/Pax for an A320-200 in the “Approach Initial - Touchdown” phases. The calculated values differ between 10% and 42% compared to the literature value.

Furthermore, [65] provided the electrical power demand for a B777-300 during cruising, which amounts to 10 kW. This literature value aligns well with the calculated power requirement. The calculation, shown in Figure 19, indicates a requirement of 9 kW for the B777-300 during cruise. The difference from the literature is 1 kW or 10%.

The second literature source, ref. [68], specifies an electrical power requirement of 10 kW for all flight phases. The commercial aircraft size is approximately equivalent to that of an A320-200. A maximum requirement of 5.5 kW was calculated for the A320-200 in this paper, with a relative deviation of 4.5 kW.

Despite the deviation from literature values, it can be inferred that both values are within the same order of magnitude. Thus, the determined values can be considered a suitable estimate.

6.7. ATA 38: Water and Waste System

The electrical power consumption during various flight phases is illustrated in Figure 21, ranging from 0 to 4 kW. The profiles for the A320-200 and the B737-800 are identical due to the correlation between the number of toilets and boilers. Both planes necessitate an electrical power of 0.7 kW. As there is no literature available on power consumption, a comparison cannot be made.

6.8. ATA 44: Cabin System

The electrical power requirement for the In-Flight Entertainment (IFE) is calculated using Equation (71) and can be seen in Figure 22. In this figure, the required electrical power in kW is plotted across the examined aircraft models and the number of passengers. The power range is between 7 kW for the A320-200 and 23 kW for the A380-800. Since the power requirement is solely dependent on the number of passengers, there is continuous consumption in all phases. The calculated values should be considered as approximations that require further investigation and verification to provide a more accurate assessment.

The authors of Ref. [50] stated that a B787-800, configured to accommodate 375 passengers, has an electrical power requirement of 20 kW. Since the power requirement depends on the number of passengers, the researched value can be compared with the power consumption calculated in this paper for a B777-300 (383 passengers) using Equation (71). For the B777-300, an electrical power requirement of 16.5 kW has been calculated. The calculated power class correlates with the value researched in the literature. In summary, the power requirements calculated in this paper can be used as an initial estimate for the expected power class.

6.9. Electrical Power Demand for Each Aircraft Type

After calculating the electrical power demand for individual airplane systems and presenting the results in detail, the next step is to determine the total electrical power demand for every airplane type during different flight phases. The demand is calculated as a function of the flight phase via the summation of the individual loads.

The total power required for the A320-200 is depicted in Figure 23. The graph presents the power demand across different flight phases. It varies between 230 kW while taxiing and 391 kW during cruise. The primary consumers of power during the flight phases are the ECS and hydraulic system.

The total electrical power requirements for the other examined aircraft can be found in Figure 24, Figure 25, Figure 26, Figure 27 and Figure 28. It is noticeable that the power requirement over the flight mission follows the same pattern for all aircraft. What varies, however, is the power requirement depending on the aircraft type. Taking cruise flight as an example, the analyzed aircraft models require the following electrical powers:

• A320-200: 391 kW;
• A330-200: 724 kW;
• A380-800: 1250 kW;
• B737-800: 402 kW;
• B757-300: 514 kW;
• B777-300: 720 kW.

According to our calculations, the A380-800 requires the most electrical power, followed by the A330-200 and the B777-300. While the B777-300 is larger than the A330-200 in terms of typical seating capacity, cabin length, and maximum takeoff weight, the calculated power requirements are similar at 724 kW and 720 kW, respectively. With regard to the largest consumers, it is notable that the ECS of the A330-200 requires nearly 380 kW compared to 510 kW for the B777-300. The power demand of the ECS depends on environmental conditions, the number of passengers, aircraft geometry, recirculation rate, and cabin temperature. Since only the number of passengers and aircraft geometry varies in the two considered aircraft models, the higher requirement for the B777-300 can be explained by the higher seating capacity and the larger aircraft geometry (number of windows and fuselage surface) compared to the A330-200. Another notable feature is the disparity in electrical power demand for the hydraulic system between the A330-200 and B777-300, with the former requiring almost 273 kW and the latter only 124 kW. This contradicts expectations, according to which bigger aircraft would typically entail more hydraulic power. Due to the difference in hydraulic power for the B777, the low power requirement for the hydraulic system compared to the A330 is also noticeable at this point. To qualify the results, it may be necessary to account for all hydraulic loads, including actuator loads during various flight phases. These more precise calculations of electrical power could help to elucidate differences. Moreover, the electrical power requirements for B737-800 and A320-200 are comparable at 400 and 390 kW, respectively. These aircraft models have similar dimensions, maximum takeoff weight, and seating capacity, as presented in

Table 12. Comparison of the aircraft models.

	B737-800	A320-200
Number of seats	189	168
Cabin length [m]	29.4	30
MTOM [kg]	79,016	75,166

. The most significant consumers of electricity in both aircraft operate within a similar power range. The A320-200 requires 217 kW for the ECS, the B737-800 requires 244 kW. The hydraulic system’s power demand is also comparable, with the A320 requiring 126 kW and the B737-800 110 kW.

Overall, the power requirements for the modeled aircraft systems are in the same range as the literature values. This suggests that the applied calculation methods and assumptions for these systems are initially representative and can be used for the sizing of electrical supply systems in aircraft.

7. Conclusions

The calculation methods presented for estimating the electrical power requirement have demonstrated that they are well suited for certain systems, while they are only applicable to a limited extent for others. For instance, the outcomes of the ECS are in accordance with the literature values, as evidenced in Section 6.1. Similar observations can be made with the cabin system, for which the calculated power requirement correlates with the literature values (see Section 6.8). The values determined in this study may be regarded as an initial estimate of the expected power requirements. Furthermore, the calculated values for WIPS demonstrate a high degree of correlation with the values obtained from the research process, as evidenced by the findings presented in Section 6.5.

In contrast, the power requirement of the equipment (Section 6.2) exhibited a deviation of around 50% in comparison to the literature values. This can be attributed to the multitude of variables that are inherently challenging to anticipate, such as catering and in-flight service, which are subject to variation based on the specific airline and flight route in question. Furthermore, the calculated values deviate by approximately 70% from the literature data with regard to the estimation of the fuel system. These discrepancies highlight deficiencies in the model. One potential source of error may be an erroneous assumption regarding the number of active pumps, as illustrated in Section 6.3. Additionally, a discrepancy was observed between the calculated and literature-derived values for the electrical power requirement for the de-icing of the cockpit windows. Such discrepancies could be attributed to an erroneous estimation of the window dimensions. In the absence of publicly available data on the window size, an estimation was derived from freely available images, which introduces an element of uncertainty into the power calculation.

In estimating the hydraulic system, there is a discrepancy between calculated values and literature values for some of the analyzed aircraft types. While the calculated values for the Airbus aircraft are within a range of ±10% of the literature values, the calculated performance requirements for the two Boeing aircraft, the B757-300, and the B777-300, are significantly lower. This indicates potential deficiencies in the modeling approach or in the assumptions employed to ascertain the parameters. In the absence of available literature values, a comparison could not be made with regard to the estimation of the water waste system, as detailed in Section 6.7.

In conclusion, the presented estimation methods provide a robust foundation for determining electrical power requirements in an initial analysis, with ample room for further improvements.

8. Summary and Outlook

This research paper aims to estimate the electrical power demands of distinct aircraft systems during different flight phases and for different aircraft types, assuming that conventional systems are replaced by electrically powered systems. The background to this research is that aircraft manufacturers treat the precise power requirements of individual systems as proprietary information, which is not publicly available. In the literature, there is a lack of data on the electrical power requirements of individual aircraft systems. If available, it is often normalized, making it impossible to derive the actual electrical power requirements.

The primary objective was to identify the total electrical power consumption of different aircraft types. This information will help determine the overall power requirements of aircraft systems. To achieve this goal, existing calculation methods from scientific literature were used. Since the results of these methods are regularly presented in normalized form, we focused on identifying necessary variables for estimating the order of magnitude of electrical power requirements. Several aircraft models, including A320, A330, A380, B757, B737, and B777, were considered, under the assumption of electrification of aircraft systems previously operated hydraulically or pneumatically. A thorough description of different aircraft systems, their structures, and functions was given. Various calculation methods have been presented for estimating the electrical power demands. In the third part, the results for the electrical power requirements of individual aircraft systems were presented and analyzed.

In order to compare the calculated results, a thorough literature search was carried out to identify existing data. This search unveiled a scarcity of data available. Compared to the identified literature, the calculated results seem to be at least in the correct order of magnitude. This indicates that the calculations and assumptions can be provisionally deemed useful.

While this research is an important step in estimating the electrical power requirements of individual systems, it is equally important to consider the design of fuel cell systems. Understanding the duration of each power demand is essential to properly size the fuel cells. In addition, the cooling requirements for the fuel cells must be considered; for example, a 400 kW electrical demand would generate approximately 140 kW of thermal losses at an assumed 65% efficiency, requiring additional electrical power for cooling that is not currently included in the overall power estimates. In addition, factors such as hydrogen storage, air supply, and increased wiring weight are critical to a comprehensive assessment of power requirements but are not discussed in detail. Future research should focus on these aspects to provide a more complete understanding of the challenges and opportunities associated with integrating fuel cell systems into modern aircraft systems.

The results of this work contribute to a better understanding of the electrical power requirements of aircraft systems while highlighting the need for further investigation into the wider implications of fuel cell technology in aviation.