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Comparative Assessment of Parallel-Hybrid-Electric
Propulsion Systems for Four Different Aircraft
D. Felix Finger
*
Department of Aerospace Engineering, FH Aachen University of Applied Sciences, Aachen, Germany
and
School of Engineering, RMIT University, Melbourne, Australia
Carsten Braun
†
Department of Aerospace Engineering, FH Aachen University of Applied Sciences, Aachen, Germany
Cees Bil
‡
School of Engineering, RMIT University, Melbourne, Australia
As battery technologies advance, electric propulsion concepts are on the edge of disrupting
aviation markets. However, until electric energy storage systems are ready to allow fully
electric aircraft, the combination of combustion engine and electric motor as a hybrid-electric
propulsion system seems to be a promising intermediate solution. Consequently, the design
space for future aircraft is expanded considerably, as serial-hybrid-, parallel-hybrid-, fully-
electric, and conventional propulsion systems must all be considered. While the best
propulsion system depends on a multitude of requirements and considerations, trends can be
observed for certain types of aircraft and certain types of missions. This paper provides insight
into some factors that drive a new design towards either conventional or hybrid propulsion
systems. General aviation aircraft, VTOL air taxis, transport aircraft, and UAVs are chosen
as case studies. Typical missions for each class are considered, and the aircraft are analyzed
regarding their take-off mass and primary energy consumption. For these case studies, a high-
level approach is chosen, using an initial sizing methodology. Results indicate that hybrid-
electric propulsion systems should be considered if the propulsion system is sized by short-
duration power constraints (e.g. take-off, climb). However, if the propulsion system is sized by
a continuous power requirement (e.g. cruise), hybrid-electric systems offer hardly any benefit.
Nomenclature
C = cruise
CD,0 = zero lift drag coefficient
E = energy
E* = mass specific energy
H = hybridization
L = lift
L/D = lift-to-drag ratio
m = mass
P = power
P* = mass specific power
P/W = power-to-weight ratio
T/W = thrust-to-weight ratio
W/S = wing loading
ε = stopping criteria
*
PhD Candidate, FH Aachen UAS / RMIT University, f.finger@fh-aachen.de, AIAA Student Member
†
Professor, FH Aachen UAS, c.braun@fh-aachen.de, AIAA Member
‡
Professor, RMIT University, cees.bil@rmit.edu.au, AIAA Associate Fellow
Acronyms
BSFC = brake-specific fuel consumption
EM = electric motor
ICE = internal combustion engine
MALE = medium altitude long endurance
mf = mass factor
MSL = mean sea-level
MTOM = maximum take-off mass
PH = parallel-hybrid
TAS = true airspeed
TLAR = top-level aircraft requirement
TPE = turboprop engine
UAV = unmanned aerial vehicle
VTOL = vertical take-off and landing
I. Introduction
LECTRIC motors offer very high efficiency and low mass per unit power. However, their application in aircraft
is held back by heavy and bulky energy storage systems (batteries) [1], [2]. For combustion engines, the efficiency
is low, and the mass per unit power is high. Nevertheless, their compact and light energy storage systems (carbon-
based fuel) still makes them highly desirable for aircraft applications.
In automotive applications, hybrid-electric drivetrains are used to boost the propulsion system’s efficiency. Each
part of the propulsion systems is designed in such a way that the overall performance is maximized, and each part
operates at its best performance point [3]. Thus, combustion engines provide continuous power, and electric systems
allow short periods of boosted power during acceleration. It is likely that this mode of operation is also beneficial for
hybrid-electric aircraft.
The authors pose the following hypothesis: Hybrid-electric propulsion systems are best suited for aircraft with
fluctuating power requirements. It is best to provide a constant baseline load for a combustion engine. Short durations
of high power requirements can then be absorbed by an electric propulsion system. This optimization will either result
in an aircraft with a reduced MTOM, or an aircraft with reduced energy consumption, depending on the optimization
objective. Aircraft that are flown at a constant power setting and are not throttled back significantly during the flight
are unlikely to benefit from hybrid-electric propulsion systems.
These assumptions apply only to traditional configurations that do not gain additional benefits that electric
propulsion systems can offer. Therefore, aircraft that use distributed propulsion schemes to reduce drag and/or increase
lift with active high-lift systems (comp. e.g. Ref. [4]) are exempt.
To test the hypothesis, aircraft of different classes and for different missions must be compared under the
consideration of hybrid-electric propulsion systems. Unfortunately, it is difficult to compare different publications on
hybrid-electric aircraft. Because the technology is still emerging and best practices for design have yet to be developed
[5], technology assumptions of different studies (e.g. Refs. [6], [7] or [8]) can be very different and hard to compare.
The same holds true for the general design philosophy. Therefore, the aim of this paper is to compare different classes
of aircraft with the same initial sizing methodology for conventional and hybrid-electric aircraft, using a common set
of technology assumptions.
By comparing general aviation aircraft (Part-23), transport aircraft (Part-25), UAVs, and VTOL air taxis, a wide
design space is considered, and meaningful conclusions can be drawn. The expectations, how hybrid-electric
propulsion systems for these kinds of aircraft will perform, are described below.
General Aviation Aircraft
Typically, general aviation aircraft are characterized by speed, rather than efficiency. A general aviation aircraft
is paid for by “the hour” and, therefore, often flown at the maximum cruise speed. However, because certification
specifications require a slow landing speed for these aircraft, they are often designed with a relatively low wing loading
W/S. Thus, their best lift-to-drag ratio L/D is achieved at low speeds, while the aircraft cruises at high speed. This
requires a large amount of power, which in turn also helps the aircraft achieve good climb rates and short take-off
distances. In performance sizing terms, such an aircraft is cruise limited, as the performance during cruising flight
dominates all other performance requirements. Because the most time is spent in cruise, this phase is also dominating
energy consumption, as it is flown at a constant high power setting.
Because the constant high load in cruise flight is dominant, little improvements are expected from hybrid-electric
propulsion technology for general aviation aircraft.
Regional Transport Aircraft
While regional turboprop transport aircraft fly much faster than general aviation aircraft, their focus is more on
cost-effectiveness than on speed. At the same time, the flown distances are higher. Consequently, even more than for
general aviation aircraft, cruising flight dominates the energy consumption of transport aircraft. The typical load is
constant at a high thrust setting.
Because the constant load in cruise flight is prevailing, little improvements are expected from hybrid-electric
propulsion technology for transport aircraft.
E
VTOL Air Taxis
Vertical Take-Off and Landing (VTOL) aircraft are designed with special focus on the take-off and landing phase.
High power is needed for VTOL aircraft, as only powered lift is available in the hover flight phases. This class of
aircraft is (vertical) take-off limited, as their design is dominated by the hover phase. While high cruise speeds are
possible, due to the large installed power, economic considerations forbid the use of full power during cruise. The
power requirement during take-off can make this phase dominant in terms of energy consumption, if flight distances
are short, as for urban air mobility. If longer ranges are considered, most of the energy is consumed in cruise flight.
Finding a good balance between those extremes is quite challenging.
Because high peak loads must be absorbed during take-off and the average load during the flight is comparatively
low, hybrid-electric propulsion technology is expected to improve the performance of VTOL Air Taxis.
UAVs
Unmanned aircraft are often designed for maximum endurance, as they are mostly used for surveillance purposes.
This requires the ability to loiter efficiently at a low power setting. At the same time, the ability to provide quick in-
and egress to and from the target area, as well as the ability to operate from small airfields, requires high power for
short periods. This means that energy consumption is dominated by the loiter flight phase, at a constant low power
setting, but the installed power is dictated by secondary flight phases. Just like air taxis, this aircraft class is challenging
to balance, as different flight phases have very different power requirements.
Because high peak loads must be absorbed during take-off, climb, and during the dash to the target, and the average
load while loitering is comparatively low, hybrid-electric propulsion technology is expected to improve performance
of long-endurance UAVs.
Only parallel-hybrid-electric propulsion systems (compare Fig. 1) will be investigated in this paper. This allows
covering three general propulsion architectures because a parallel-hybrid layout with 0% hybridization is effectively
a conventional propulsion system, and a parallel-hybrid layout with 100% hybridization is effectively a fully electric
propulsion system.
Serial-hybrid-electric propulsion layouts are not considered in this study, because, as shown in Refs. [9] and [10],
the serial-hybrid will always perform worse than the parallel-hybrid for a similar set of parameters. This is caused by
the additional mass that the generator system will add to the aircraft, the corresponding reduction on propulsive
efficiency, and the fact that the electric motor must be sized to the maximum P/W. The advantage of serial-hybrid
systems is their geometric flexibility. The electric motor(s) can be installed independently of the location of the
combustion engine(s). This gives way to distributed propulsion layouts (see e.g. Refs. [11], [12] or [13]), which can
take advantage of favorable aero-propulsion interaction. Thereby, the weight increase of the propulsion system is
traded against improved aerodynamic efficiency. The assessment of those effects is beyond the scope of this paper,
however.
Further information on hybrid-electric propulsion architectures is provided in Ref. [14].
Fig. 1 Parallel-hybrid-electric propulsion architecture. [14]
This paper is structured the following way: After this introduction, the methodology for the sizing and technology
factors are discussed in Sec. II. In Sec. III, an overview of the notional aircraft design concepts is given. Then, in Sec
IV, the technology levels for the study are described. The sizing results are presented in Sec. V, followed by the
discussion in Sec. VI. Finally, Sec. VII gives a comprehensive conclusion.
Gearbox
Combustion
Engine Fuel
Electric Motor
Gearbox Electric
Motor Generator Combustion
Engine
Battery Fuel
Parallel Hybrid
Variant 1
Serial Hybrid
Parallel Hybrid
Variant 2
Combustion
Engine Fuel
Electric
Motor Battery
Gearbox
Battery
II. Methodology for Initial Sizing
A. Initial Sizing Method for Hybrid-Electric Aircraft
In aircraft conceptual design, one of the first tasks is the initial sizing process. Once the top-level aircraft
requirements (TLARs) have been defined (how far, how fast, what payload) and a first concept idea is defined, the
sizing process can be begun.
The authors’ methodology for sizing hybrid-electric aircraft is documented in [9]. Conventional take-off and
landing general aviation aircraft were studied, as shown in Refs. [10], [15], [16] and [17]. The sizing of VTOL aircraft
is discussed in Ref. [18]. The method covers the sizing of aircraft with conventional propulsors, hybrid-electric
systems, or fully electric propulsion systems. A thorough validation study of the methodology was conducted and
published in Ref. [19].
The primary goal of the method is the identification of the optimal design point (P/W and W/S) of such aircraft
and, in addition to this, the corresponding degree of hybridization. Analog to the classical methods, the methodology
is separated into two major parts: Point performance, also referred to as the matching diagram, and mission
performance, also known as the weight estimation. For both parts, certain input parameters are necessary, representing
the TLARs, which are defined for the individual aircraft. These requirements specify the TLARs, including the flight
mission, the aerodynamics, and the propulsion system (number of engines, conventional, serial- or parallel-hybrid,
etc.) and its corresponding efficiencies.
In the first step, the point performance tool determines the matching diagram. The required P/W, with respect to
the W/S, is determined for constraints like the desired rate of climb, the desired take-off distance, or the desired cruise
airspeed. The mission performance analysis is based on a classical iterative process presented in most aircraft design
books (see e.g. Refs. [20] [21]). To cover the mix of consumable (fuel) and non-consumable (batteries) energy sources
on board, the masses are not treated as fractions, as it is done in other sizing algorithms (e.g. Ref. [22]), but as absolute
values. Additionally, the classical endurance and range equations of Breguet are not used. Instead, the mission is
divided into short segments and simulated, using a universally valid, energy-based approach. The mission is defined
explicitly by the requirements. Within this approach, fuel weight and battery weight are determined by calculating the
required energy for each time step of the flight phase. This energy is split into its consumable and non-consumable
parts through the degree of hybridization of energy – HE – of the corresponding flight phase and converted into the
necessary thrust or power.
Finally, based on a first estimate for the MTOM, all masses that make up the gross weight are calculated. This can
be done by using Class-I or Class-II mass estimation methods (see Refs. [20] [21], [23], [24]. Using the results of the
point performance analysis, the propulsion system is sized, as its engine, motor, and integration weights are estimated.
In the final step, the empty weight is determined. Because the propulsion system mass is calculated separately, this
weight fraction covers the usual operating empty weight without fuel, battery, and engine weights. Based on the new
MTOM, the next iteration step can be started. The iteration stops when a certain mass convergence, defined by the
stopping criteria ε, is reached. The process is shown in Fig. 2.
B. Measures of Merit
The converged aircraft can be analyzed with respect to several measures of merit. The most obvious is MTOM, as
aircraft designers traditionally use weight as a surrogate for cost [20]. It was found that for a given set of TLARs, the
lightest aircraft performed best over a range of operating conditions [25]. While the design tool supports more detailed
cost estimation methods for hybrid-electric general aviation aircraft (DAPCA IV from Ref. [22], modified as described
in Ref. [26]), this analysis is not yet fully implemented and tested for transport aircraft, UAVs, or VTOL aircraft.
Therefore, this functionality will not be used in this paper.
Another suitable measure for assessing aircraft is their energy consumption. The required energy for the design
mission is readily available, as it is determined by the mission analysis and then used to determine the aircraft’s energy
mass fraction. However, if only the energy consumption during the flight is assessed, the environmental impact of
each flight is not fully captured. The reduction in efficiency that is caused by sourcing energy and delivering it to the
aircraft must be accounted for. This is done by primary energy factors. Primary energy is a measure for the total energy
that was expended to extract the energy from natural resources and to provide the extracted energy to the consumer.
Gasoline fuel, for example, has to be refined from raw oil, which needs to be extracted from oilfields. The energy to
produce the fuel is summed up in the primary energy factor. The factors for Germany are 1.10 for fossil carbon-based
fuel and 2.80 for electricity (data from Ref. [27], dated 2016). The factor for electricity is that high, because of the
composition of electricity. Coal-burning and nuclear power plants have a significant share, which have a high primary
energy factor as the thermal efficiency of the power stations needs to be accounted for. The factor will decrease as the
use of renewable energy sources increases.
Fig. 2 Sizing process.
C. Optimization Approach
If required, optimization with respect to the measures of merit discussed above can be carried out according to the
process in Fig. 3. Because the methodology is implemented in Matlab, the Matlab Optimization Toolbox is used for
this task. Global optimization schemes like the genetic algorithm or the particle swarm method can be employed to
find an optimal design for a given set of TLARs and constraints. However, for certain optimization problems, it can
be faster to use a gradient-based optimization scheme under consideration of nonlinear constraints (fmincon). To avoid
local optimums, the optimization process is started multiple times at different initial points for the same set of inputs
(multistart).
Fig. 3 Optimization process.
Specifications
Baseline Design
Mission Performance /
Weight Estimation
Evaluate Objective Evaluate Constraints
(incl. Point Performance)
Optimal
Design?
Final Design
Change
Design
To obtain the results of Sec. V, optimization is used to select the best possible combination of W/S, P/W, wing aspect
ratio AR, and the split point between the engine and electric motor, HP. By using this set of optimization parameters,
the most important design variables for the sizing of subsonic aircraft are covered [28]. The results are all computed
using Matlab’s genetic optimization algorithm with a convergence criterion of 10-7 of the solver, and ε = 0.1 kg for the
mass iteration.
The influence of W/S and AR on the induced drag of the aircraft is captured, using methods from Ref. [20].
However, the zero-lift drag coefficient CD0 is held constant, regardless of the sizing result. This approach favors
aircraft with higher wing loadings since drag coefficients are referenced to the wing area. However, as the concept
aircraft, as drawn and as analyzed, are reasonably close to the sizing results, this systematical error is not significant
in the overall scope of accuracy. Especially for initial sizing, this methodology is industry practice and well
established.
D. Wing Mass Estimation
To estimate the operating empty mass of all aircraft in this study, Class-II mass estimation equations from Nicolai
(Ref. [21]) are employed. These equations are used in unmodified form, except for the wing mass equation. A
description and rationale of the modification is given below.
Mass estimation equations use exponential functions to account for the additional mass due to increasing AR. A
wing mass estimate is multiplied with the AR and an exponent, smaller than one. Here, this is called the AR mass
factor mfAR. Nicolai’s equation (Eq. 1) suggests an exponent of 0.6, which is directly related to the additional wing
root bending moment of the added span. However, while bending moment constraints will drive the wing mass for
AR values of up to 10 or 12. Beyond this, aeroelastic effects will dominate and are the cause of an increased wing
mass. This effect is not captured by the basic Eq. 1. To remedy this, the authors introduce a crude modification to
Nicolai’s approach in Eq. 2.
(1)
(2)
This way, mfAR is virtually unchanged at AR below 10, and then exponentially increases, as very high AR are
considered. Fig. 4 shows both Eq. 1, and its modification, Eq. 2, for ARs up to 30.
Fig. 4 Mass penalty for high aspect ratio wings.
This measure was introduced to avoid boundary optima. If the unmodified equations are used, the optimizer will,
for some cases, select unrealistically high ARs (over 60), as the aerodynamic performance outweighs the additional
mass. However, such high ARs are well beyond the prediction capability of Eq. 1, and, in almost all cases, not a
meaningful design solution. With this modification, much better agreement with real-world aircraft is obtained.
0
2
4
6
8
10
12
0 5 10 15 20 25 30
AR mass factor [-]
Wing aspect ratio [-]
Nicolai
High AR Penalty
III. Notional Design Concepts
For this sizing study, four concept aircraft were designed. The concepts will be presented in the following
subsections, where the respective, top-level design requirements, and design missions will be outlined. All
performance requirements are stated for standard day conditions.
E. General Aviation Aircraft
The Cirrus SR-22 is chosen as a baseline for mission specification and top-level design requirements used for the
general aviation sizing study. This aircraft is described in numerous references (e.g. Refs. [22] [29] [30]) and well
known in the aviation community. It is a single propeller, general aviation aircraft, sized for a typical 1150 km cruise
mission, with 45 minutes of reserve energy. A 380 kg payload must be carried at this range. A graphical representation
of the mission is shown in Fig. 5. This concept will be analyzed to see if a hybrid-electric replacement of the SR-22
makes sense. Top-level requirements and mission description are shown in Table 1.
Table 1 Top level requirements general aviation aircraft.
Requirements
Mission
General Aviation Aircraft
Take-off Ground Roll [m]
340
Taxi & Take-off
at MSL
Rate of Climb at MSL [m/s]
5
Climb
to 3000 m
Stall Speed [m/s]
32
Cruise
for 1150 km
Cruise Speed (TAS) [m/s]
90
Loiter
for 45 min
Payload [kg]
380
Descend, Landing, Taxi
MSL
Aerodynamics
CD0 [counts]
254
F. Regional Transport Aircraft
A notional regional turboprop airliner, similar to the Fokker 50 and the ATR 72, is selected for the regional
transport sizing study. Consequently, the chosen concept is a twin-turboprop transport aircraft designed to carry a
7000 kg payload over a distance of 1750 km at 450 km/h. For this concept, the top-level requirements and mission
description are shown in Table 2. A graphical representation of the mission is shown in Fig. 6.
The general mission set-up is very similar to the general aviation mission, even though range is lightly increased.
However, for this transport aircraft, turboprop engines are considered as the baseline technology. Turboprop engines
are relatively lightweight and reach mass specific power levels that are more comparable to electric motors than
combustion engines (about three to four times higher than a four-stroke engine). Consequently, as the difference in
P/W between a turboprop and an electric motor is not as large, it is expected that performance improvements for
turboprop-electric hybrids will be harder to realize than for piston engine-electric hybrids.
Table 2 Top level requirements regional transport aircraft.
Requirements
Mission
Regional Transport
Take-off Ground Roll [m]
1400
Taxi & Take-off
at MSL
Rate of Climb at MSL [m/s]
7
Climb
to 5000 m
Stall Speed [m/s]
50
Cruise for
1750 km
Cruise Speed (TAS) [m/s]
125
Loiter
for 45 min
Payload [kg]
7000
Descend, Landing, Taxi
MSL
Aerodynamics
CD0 [counts]
275
45 min
1150 km at 90 m/s
Climb to
3000m
Descend
to MSL
Fig. 5 General aviation design mission.
G. VTOL Air Taxi
A four-seat tilt-wing design with a configuration similar to A³’s Vahana
1
is used for the VTOL Air Taxi study.
The authors presented a similar concept in Ref. [18]. Eight propellers provide thrust in hover and forward flight. Four
propellers are mounted on the forward wing, and four on the rear wing. The aircraft is designed as an L+C VTOL
design but can also be considered an L+L/C design, if the motors are switched off during cruise flight [11].
The aircraft is designed to fly Uber’s urban air taxi sizing mission, which is described in Ref. [31]. This mission
is illustrated in Fig. 7. Cruise range is 100 km at 240 km/h. Time in hover is limited to 90 s for each vertical take-off
and each landing. The mission requires a full transition cycle for the main mission and another full transition cycle for
the reserve mission. Top-level requirements and mission description are shown in Table 3.
Table 3 Top level requirements VTOL air taxi.
Requirements
Mission
4 Seat VTOL Air Taxi
T/W in Hover at MSL
1.4
Taxi & Vertical Take-off
at MSL
Rate of Climb at MSL [m/s]
2.5
Climb
to 300 m
Stall Speed [m/s]
33.5
Cruise for
100 km
Cruise Speed (TAS) [m/s]
67
Cruise to alternate
15 km
Payload [kg]
400
Descend, Vertical Landing, Taxi
MSL
Aerodynamics
CD0 [counts]
250
H. MALE UAV
A notional twin-boom reconnaissance medium altitude long endurance (MALE) unmanned aircraft is selected for
the UAV sizing study. The concept features a single pusher propeller and twin tail booms. This is a typical
configuration for unmanned aircraft [32]. The UAV is designed for a typical long term surveillance mission, which is
shown in Fig. 8: 12 hours of endurance at best endurance speed is required, as well as a quick 50 km dash at twice
that speed into- and out of the target area. A 200 kg payload of surveillance and communication equipment must be
carried. Because the electro-optical equipment disturbs the outer mold line of the aircraft, a high CD0 of 350 drag
counts is assumed (compare Ref. [33]). Top-level requirements and mission description are shown in Table 4.
1
A³, "Vahana," [Online]. http://www.airbus-sv.com/projects/1
45 min
1750 km at 125 m/s
Climb to
5000m
Descend
to MSL
100 km at 67 m/s
Vertical Take-Off
Transition
Climb to
300m
Descend
to MSL
Transition
Vertical Landing
Descend
to MSL
Transition
Rejected Landing
Transition
15 km at 67 m/s
Fig. 6 Regional transport design mission.
Fig. 7 VTOL Air Taxi design mission.
Table 4 Top level requirements MALE UAV.
Requirements
Mission
MALE UAV
Take-off Ground Roll [m]
500
Taxi & Take-off
at MSL
Rate of Climb at MSL [m/s]
5
Climb
to 3000 m
Stall Speed [m/s]
30
Ingress to Target Area
50 km
Cruise Speed (TAS) [m/s]
80
Loiter over Target
for 12 h
Loiter Speed (TAS) [m/s]
40
Egress from Target Area
50 km
Payload [kg]
200
Descend, Landing, Taxi
MSL
Aerodynamics
CD0 [counts]
350
IV. Technology Levels
Because of constant and rapid improvements in the field of electrical system technologies, a reference for the
technological assumptions must be set. All sizing studies are conducted for four different technology levels, where
different characteristics of the electric propulsion system’s components are considered. Level 1 and level 2 are
representatives of near term technology, while levels 3 and 4 represent more optimistic assumptions. Battery, electric
motor, and combustion engine technology assumptions are explained in more detail below, while Table 5 summarizes
the technology levels for all aircraft concepts.
Battery Technology
The battery specific energy E* is one of the key drivers of hybrid-electric aircraft performance. In this study, E*
values of 250 Wh/kg for the near term technology levels 1 and 2 are assumed. E* = 500 Wh/kg is used for the advanced
technology levels 3 and 4. These specific energy values are pack level values. Doubling the specific energy is a very
optimistic assumption, but allows to assess if electric systems will have a significant effect in the future, in case that
near term technology does not have any impact on each of the different aircraft classes.
Technology level 1 and 3 considers batteries with a 4C discharge rating. This corresponds to a minimum discharge
time of 15 minutes, and is representative for today’s high energy cells. Technology levels 2 and 4 consider batteries
with a 20C discharge rating. This allows a very rapid full discharge in 3 minutes. This allows using these batteries for
boost applications, like additional take-off power.
The operational strategy of the hybrid system is biased toward using the combustion engines. In each flight phase,
the required power is compared to the available power. If the required power can be delivered by the ICE, then the
electric system is not used. Only if the required power is higher than the ICE’s maximum power (e.g. during take-off
and climb – and once the ICE has failed), the EM is activated and delivers the necessary additional power.
All concepts’ electrical systems are sized with respect to the following considerations:
The full design mission must be flown with 80% of the capacity. The bottom 20% capacity is considered
unavailable. This prevents the battery from deep discharge and reduces the voltage drop-off at the end of the
mission. This is a measure intended to improve both safety and battery life.
All missions start with the battery 100% charged, and the battery is not recharged over the course of the flight.
Electric Motor Technology
For electric motors, a specific power P* of 5 kW/kg is considered at technology levels 1 and 2. This value is
achieved by Siemens with their motor SP260 [34]. Just as for the battery technology, that value is doubled to P* = 10
720 min
at 40 m/s
50 km at 80 m/s
Climb to
3000m
Descend
to MSL
50 km at 80 m/s
Fig. 8 MALE UAV design mission.
kW/kg for technology levels 3 and 4. To describe the losses of the entire electrical system, a constant equivalent motor
efficiency of 80% is used, and all other electrical component’s efficiencies are set to 100%. This approach is
considered conservative. This equivalent motor efficiency remains constant for all technology levels.
Combustion Engine / Turboprop Technology
Because this paper explores the influence of electric systems on aircraft performance, the state of technology of
combustion engines and turboprop engines is not varied to avoid too many parameter variations. Therefore, specific
power and efficiency (brake specific fuel consumption BSFC) are held constant.
Table 5 Technology levels.
Concept
Technology Assumption
Technology Level
1
2
3
4
All concepts
Battery Specific Energy [Wh/kg]
250
250
500
500
Battery Discharge Rate [-]
4
20
4
20
Motor Specific Power [kW/kg]
5
5
10
10
General Aviation
Combustion Engine Technology
4-stroke ICE
ICE Specific Power [kW/kg]
1.18
ICE best BSFC [g/kW/h]
315
Regional Turboprop
Combustion Engine Technology
Turboprop
TPE Specific Power [kW/kg]
3.30
TPE best BSFC [g/kW/h]
420
VTOL Air Taxi
and
MALE UAV
Combustion Engine Technology
4-stroke ICE
ICE Specific Power [kW/kg]
1.00
ICE best BSFC [g/kW/h]
315
V. Results
In this section, the sizing results for the four aircraft concepts are presented. All aircraft are designed under
consideration of parallel-hybrid-electric propulsion systems. This allows the hybridization ratio to vary between 0
(conventional, fuel-based propulsion) and 1 (fully electric propulsion).
The sizing results are presented in tables – e.g. in Table 6. This table – and every subsequent data table – is
structured in the following way: The columns contain the sizing results for different technology levels (1-4). As stated
in the previous section, higher technology levels correspond to more advanced technology. Each technology level is
listed twice because of the two different optimization objectives. The first optimization objective is to achieve a
minimal MTOM, and the second objective is to minimize primary energy consumption.
A. General Aviation Aircraft Study – Cirrus SR-22 Replacement
As outlined in the introduction, little benefit was expected from using hybrid-electric propulsion for general
aviation aircraft. This hypothesis is confirmed by the results shown in Table 6. Only two conventional design points
without any hybridization are found by the optimizer, regardless of the battery and motor performance. The sizing
result is the same for all technology levels, even though different results are reached for the two optimization
objectives. Consequently, hybrid-electric propulsion systems do not offer any benefit for an SR-22 type general
aviation aircraft. Even at the highest technology level, hybrid-electric propulsion does not offer a reduction in take-
off mass or primary energy consumption, when compared to a conventional propulsion system.
The MTOM optimization converged to a smaller AR and lower P/W, while the primary energy optimization
converged to an aspect ratio more appropriate for a motor glider (AR = 14.5) and a higher P/W. The impact of AR on
wing mass is captured by the modified version of Nicolai’s method for empty mass estimation. The higher P/W of the
aircraft optimized for minimum primary energy consumption is caused by the ICE’s BSFC model: the best efficiency
is assumed at 80% power, and the engine is slightly oversized to operate at its best efficiency. Thus, the aircraft that
is optimized for minimal primary energy consumption is about 6% heavier but consumes about 10% less fuel than the
aircraft, which is optimized for minimum take-off mass. W/S remains constant and is constrained by the stall speed
requirement.
Table 6 Sizing results of the general aviation aircraft.
Tech Level
1
2
3
4
1
2
3
4
Objective
MTOM
MTOM
MTOM
MTOM
Primary E
Primary E
Primary E
Primary E
MTOM [kg]
1486
1486
1486
1486
1578
1578
1578
1578
Primary E [J]
1.111E+10
1.111E+10
1.111E+10
1.111E+10
1.005E+10
1.005E+10
1.005E+10
1.005E+10
mfuel [kg]
234.6
234.6
234.6
234.6
212.3
212.3
212.3
212.3
mbat [kg]
0.0
0.0
0.0
0.0
0.0
0.0
0.0
0.0
W/S [N/m²]
1177
1177
1177
1177
1177
1177
1177
1177
P/W [W/kg]
146.5
146.5
146.5
146.5
168.2
168.2
168.2
168.2
AR [-]
8.8
8.8
8.8
8.8
14.5
14.5
14.5
14.5
HPPH [-]
0.0%
0.0%
0.0%
0.0%
0.0%
0.0%
0.0%
0.0%
HEaverage [-]
0.0%
0.0%
0.0%
0.0%
0.0%
0.0%
0.0%
0.0%
delta to Level 1 results
delta MTOM
0.0%
0.0%
0.0%
0.0%
0.0%
0.0%
delta Primary E
0.0%
0.0%
0.0%
0.0%
0.0%
0.0%
B. Transport Aircraft Study – Regional Turboprop
The benefit of hybrid-electric propulsion systems is also very limited for such a regional transport configuration.
The hypothesis is confirmed, again. No case was found, where hybrid-electric propulsion offers a benefit, even if
optimistic technology levels are assumed.
When optimization results for minimum MTOM and minimum primary energy consumption are compared, the
tendencies are very similar to those of the general aviation study. A lower aspect ratio wing reduces the MTOM but
increases fuel consumption slightly. The MTOM optimized aircraft are 3% lighter, but consume 8% more primary
energy. The actual energy saving will be lower than 14% because the minimal primary energy consumption is achieved
by moving to a heavier aircraft with a very high AR wing. Here, the wing’s AR is driven to 20.1, a value more typically
found with gliders or UAVs. This value will be very challenging to achieve without external bracing. Therefore, the
wing’s mass will probably be even higher than predicted and therefore, the optimum AR will be reduced, which will
result in higher primary energy consumption. Evidently, this aircraft class is not particularly suited for hybrid-electric
propulsion. Again, W/S remains constant and is constrained by the stall speed requirement.
Table 7 Sizing results regional transport aircraft.
Tech Level
1
2
3
4
1
2
3
4
Objective
MTOM
MTOM
MTOM
MTOM
Primary E
Primary E
Primary E
Primary E
MTOM [kg]
20924
20924
20924
20924
21633
21633
21633
21633
Primary E [J]
1.848E+11
1.848E+11
1.848E+11
1.848E+11
1.716E+11
1.716E+11
1.716E+11
1.716E+11
mfuel [kg]
3902.6
3902.6
3902.6
3902.6
3625.1
3625.1
3625.1
3625.1
mbat [kg]
0.0
0.0
0.0
0.0
0.0
0.0
0.0
0.0
W/S [N/m²]
3827
3827
3827
3827
3827
3827
3827
3827
P/W [W/kg]
177.7
177.7
177.7
177.7
170.9
170.9
170.9
170.9
AR [-]
13.7
13.7
13.7
13.7
20.1
20.1
20.1
20.1
HPPH [-]
0.0%
0.0%
0.0%
0.0%
0.0%
0.0%
0.0%
0.0%
HEaverage [-]
0.0%
0.0%
0.0%
0.0%
0.0%
0.0%
0.0%
0.0%
delta to Level 1 results
delta MTOM
0.0%
0.0%
0.0%
0.0%
0.0%
0.0%
delta Primary E
0.0%
0.0%
0.0%
0.0%
0.0%
0.0%
C. VTOL Aircraft Study – VTOL Air Taxi
For the notional VTOL air taxi, hybrid-electric propulsion systems can offer benefits. Still, data (Table 8) shows
that the conventional combustion engine is the most viable propulsion system if near term technology (level 1) is
considered. However, it is questionable if such a distributed propulsion configuration should be chosen if eight
combustion engines are used as powerplants. If a conventional propulsion system is the best propulsion option, a
rotorcraft configuration appears to be more promising. If a technology level of 2 or higher is considered, significant
savings in MTOM and primary energy consumption are possible. Technology levels 2-4 enable mass savings between
1/3 and almost 1/2 of the MTOM. The highest savings are naturally achieved at the highest technology level. Energy
consumption is reduced between 39 % and almost 65 %. The VTOL aircraft segment shows a high potential for
successfully applying hybrid-electric propulsion technology.
The high power requirement during the short hover periods is sizing the electric part of the propulsion system,
while cruise power is supplied by a combustion engine. The hybridization ratio is high, with over 70% of the installed
power supplied by the electric system, if MTOM is minimized, and even 80% hybridization of power, if primary
energy consumption is minimized. This requires a battery that is capable of high discharge rates. For this reason, the
4C discharge rating of technology level 3 is a significant handicap, even though the assumed specific energy of
500 Wh/kg is quite high. Thus, the level 3 designs consume more energy than the level 2 designs, even though the
entire aircraft converges at a slightly lower MTOM for both optimization objectives at level 3.
As for the other case studies, higher aspect ratio wings are selected if primary energy consumption is minimized.
For the air taxi designs, AR is also increased as hybridization is introduced. As the technology levels progress, AR
increases as well. Clearly, hybrid-electric propulsion is very sensitive to the aerodynamic performance, even more so
than aircraft with conventional propulsion systems. Minimum energy consumption is always achieved with a hybrid-
electric propulsion system. This indicates that the combustion engine is sized for efficient loiter, and the high power
flight segments (take-off, climb, dash) use the support of the electric system.
Table 8 Sizing results VTOL air taxi.
Tech Level
1
2
3
4
1
2
3
4
Objective
MTOM
MTOM
MTOM
MTOM
Primary E
Primary E
Primary E
Primary E
MTOM [kg]
2704
1807
1737
1439
2813
1889
1780
1494
Primary E [J]
4.403E+09
2.236E+09
2.701E+09
1.571E+09
4.261E+09
1.949E+09
2.548E+09
1.511E+09
mfuel [kg]
93.0
35.7
32.6
23.2
90.0
27.9
27.0
21.4
mbat [kg]
0.0
269.8
286.8
116.9
0.0
311.6
314.7
123.9
W/S [N/m²]
1200
1235
1236
1236
1236
1236
1236
1236
P/W [W/kg]
367.3
367.3
367.3
367.3
367.3
367.3
367.3
367.3
AR [-]
5.0
7.2
8.0
11.0
7.3
14.8
13.3
17.2
HPPH [-]
0.0%
73.1%
74.9%
79.5%
0.0%
80.6%
80.2%
81.0%
HEaverage [-]
0.0%
35.9%
37.3%
41.4%
0.0%
43.8%
43.0%
44.7%
delta to Level 1 results
delta MTOM
-33.2%
-35.8%
-46.8%
-32.8%
-36.7%
-46.9%
delta Primary E
-49.2%
-38.7%
-64.3%
-54.3%
-40.2%
-64.5%
D. UAV Study – MALE UAV
For the MALE UAV concepts, hybrid-electric propulsion systems can offer benefits. If current- or near term
technology (levels 1 and 2) are considered, the conventional baseline performs best, if minimal mass is selected as the
optimization objective. All other cases favor hybrid-electric propulsion.
Minimum mass is reached, as expected for MALE UAVs, with high aspect ratio wings. As batteries with higher
specific energy become available (levels 3 and 4), the MTOM can be reduced by 14%, and primary energy usage goes
down by over 36%. Compared to the baseline, aspect ratio must be very slightly increased from 17.7 to 17.8, which
also results in a very slight reduction of P/W.
Minimum energy consumption is always achieved with a hybrid-electric propulsion system. The hybridization
ratio is high, with over 71% of the installed power, and more than 11% of the total energy is supplied by the electric
system. This indicates that the combustion engine is sized for efficient loiter, and the high power flight segments (take-
off, climb, dash) use the support of the electric system. To achieve the lowest primary energy consumption, the wing’s
aspect ratio is driven to very high values (23.9 and 26.3). Still, these are appropriate aspect ratio values for long-
endurance UAVs. As for the general aviation and transport aircraft examples, W/S remains constant for all
optimizations and is constrained by the stall speed requirement.
Table 9 Sizing results MALE UAV.
Tech Level
1
2
3
4
1
2
3
4
Objective
MTOM
MTOM
MTOM
MTOM
Primary E
Primary E
Primary E
Primary E
MTOM [kg]
1201
1201
1032
1032
1458
1458
1100
1100
Primary E [J]
9.357E+09
9.357E+09
5.939E+09
5.939E+09
7.299E+09
7.299E+09
4.529+09
4.529+09
mfuel [kg]
197.6
197.6
116.4
116.4
141.2
141.2
101.8
101.8
mbat [kg]
0.0
0.0
105.7
105.7
304.3
304.3
116.1
116.1
W/S [N/m²]
1102
1102
1102
1102
1102
1102
1102
1102
P/W [W/kg]
139.3
139.3
138.3
138.3
135.4
135.4
134.5
134.5
AR [-]
17.7
17.7
17.8
17.8
23.9
23.9
26.3
26.3
HPPH [-]
0.0%
0.0%
67.6%
67.6%
71.1%
71.1%
72.5%
72.5%
HEaverage [-]
0.0%
0.0%
9.6%
9.6%
11.1%
11.1%
11.7%
11.7%
delta to Level 1 results
delta MTOM
0.0%
-14.1%
-14.1%
0.0%
-24.6%
-24.6%
delta Primary E
0.0%
-36.5%
-36.5%
0.0%
-38.0%
-38.0%
VI. Discussion
Parallel-hybrid powertrains appear particularly suited to improve the performance of aircraft that are driven by
(vertical-) take-off and climb performance requirements, as long as the cruise power requirements do not drive the
design. Sizing the electric part of the propulsion system to peak power requirements allows for small engines, and the
battery weight is kept at a minimum because the battery is used only for a small fraction of the total flight time.
Because the main advantages of hybrid-electric aircraft arise from the design of the ICE to cruise conditions, mission
parameters, which increase the difference between the maximum required power and the power needed in cruise, very
much favor the hybrid-electric propulsion layout.
In contrast, general aviation and turboprop transport aircraft are dominated by cruise power requirements. Even if
optimistic technology assumptions are used, it is unlikely that the implementation of a hybrid-electric propulsion
system will yield benefits in terms of MTOM or primary energy consumption.
While the additional complexity of the powertrain is an issue, this case study shows that hybrid electric powertrains
must be seriously considered for UAVs and VTOL aircraft. If short bursts of high power are required, hybrid-electric
propulsion systems can improve both MTOM and primary energy consumption.
Last but not least, the design of hybrid-electric aircraft is both highly mission-critical and very sensitive to
technology variations. Results can change considerably, for different sets of input parameters, even in the same aircraft
class. For instance, the suitability of hybrid-electric propulsion systems for UAVs depends very much on the two
mission parameters: the length of the high-speed in-and egress flight phase, and if such a dash segment is required in
the first place. Once this stage gets too long, the battery will grow beyond reasonable mass fraction, and the system
becomes very inefficient. At this point, a fully conventional system is more appropriate.
VII. Conclusion
In this study, four different aircraft concepts are compared with respect to possible benefits from hybrid-electric
propulsion systems. These four concepts cover entirely different classes of aircraft. A general aviation aircraft and a
regional transport aircraft are considered, with both concepts driven primarily by long-duration (cruise) performance
requirements. Also, a VTOL air taxi and a MALE UAV are analyzed. Both of these concepts are driven primarily by
short-duration (take-off, dash) requirements. Two optimization objectives are used to assess the four concepts:
minimum MTOM and minimum primary energy consumption. To include future technology projections in the results,
four different levels of technology are applied.
The authors’ hypothesis is confirmed: Hybrid-electric propulsion systems are best suited for aircraft with
fluctuating power requirements. Short durations of high power requirements can be absorbed by an electric propulsion
system. Aircraft that are flown at a constant power setting and are not throttled back significantly during the flight are
unlikely to benefit from hybrid-electric propulsion systems.
For both optimization objectives, and even the most optimistic technology levels, the general aviation aircraft, and
the regional transport aircraft converged to conventionally powered fuel burning propulsion configurations. Neither a
reduced MTOM nor reduced energy consumption could be realized. However, the VTOL air taxi concept shows a
significant performance improvement, if hybrid-electric propulsion systems are used. High levels of hybridization
(>70% of the total installed power) will cause both a reduction in MTOM (>30%) and in primary energy consumption
(up to 65%). Significant improvements for both objectives were also found for the UAV concept. Here, hybrid-electric
propulsion has the largest impact if the concept is optimized for minimum energy consumption.
While these results were only calculated for four aircraft and their respective TLARs, it is expected that the general
trends can be extrapolated to similar designs. However, the design of hybrid-electric aircraft is highly mission-critical,
and even small parameter changes can have a large impact on the sizing outcome. Also, the additional complexity of
hybrid-electric powertrains is an issue that must be considered. Nevertheless, this study shows that hybrid-electric
propulsion system can greatly improve the performance of aircraft, but that this kind of propulsion system is not the
best solution for all types of aircraft and all missions.
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