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Hydrogen Powered Long Haul Aircraft with Minimized Climate Impact

Authors:
Hydrogen Powered Long Haul Aircraft with Minimized Climate
Impact
Florian Troeltsch, Marc Engelmann, Fabian Peter, Jochen Kaiser§and Mirko Hornung
Bauhaus Luftfahrt e.V., Taufkirchen, 82024, Germany
Anna E. Scholz
Technical University of Munich, Garching, 85748, Germany
The present paper presents the design of a long range aircraft concept featuring liquid
hydrogen (LH2) as main energy source. This concept has been created during a Bauhaus
Luftfahrt e.V. internal design project and is named Hyliner (2.0). At first, the definition of
the top level aircraft requirements are described, including the motivation for the reduction
of the cruise Mach number to
0
.
7
and the increase of cabin space available per passenger.
Afterwards, the derivation of the Hyliner (2.0) from a conventional long range aircraft with
the same technology level representative for an entry into service date of 2040 is established.
Design decisions of the Hyliner (2.0) are explained and its performance and qualitative impact
on emission is discussed. While the energy consumption of the Hyliner (2.0) is
9%
higher
compared to the conventional reference aircraft with the same technologies integrated, the
combustion of hydrogen rather than kerosene offers a possible reduction of the climate impact
on long range operations.
I. Nomenclature
AR = aspect ratio
CL= lift coefficient
CFRP = carbon fiber reinforced plastics
Hyliner (2.0) = aircraft concept of Bauhaus Luftfahrt e.V. group design project 004
Df= friction drag
Dp= pressure drag
FL = flight level
FLOPS = flight optimisation system (software)
GWP = global warming potential
ISA = international standard atmosphere
LH2= liquid hydrogen
LTH = aviation technical handbook (German: Luftfahrttechnisches Handbuch)
MTOM = maximum take-off mass
MWI = multi-walled insulation
NLF = natural laminar flow
OEM = operative empty mass
PAX = passengers
PIANO = project interactive analysis and optimisation (software)
OEM = operating empty mass
TLAR = top level aircraft requirement
TSPC = thrust specific power consumption
PhD Student, Department Visionary Aircraft Concepts
PhD Student, Department Visionary Aircraft Concepts
Dipl.-Ing., Department Visionary Aircraft Concepts and Lead Research Focus Area Systems and Aircraft Technologies
§Dr.-Ing., Lead Department Visionary Aircraft Concepts
Prof. Dr., Executive Director Research and Technology
Research Assistant and PhD Student, Institute of Aircraft Design
1
II. Introduction
The goals set for the reduction of aviation greenhouse gas emissions require actions for all ranges of aircraft operation.
While several options are in active research for short and medium range operations such as electric and hybrid-electric
propulsion, challenges for the long range require different solutions. This is especially important due to the high amount
of energy needed for long range operations in comparison to that on short ranges and the lack of practicable alternatives
for long range travel in general. At Bauhaus Luftfahrt e.V. a group design project was undertaken with the specific
goal of analyzing possibilities to reduce the climate impact of long range aviation. During the course of the project, an
aircraft concept with a reduced cruise Mach number employing LH
2
as the primary energy storage for combustion
in gas turbines was designed according to the TLARs introduced in the following section. The design aspects of this
aircraft concept are the focus of this paper.
III. Top Level Aircraft Requirements
The top level aircraft requirements (TLARs) for the aircraft concept presented in this study are derived from the
goals of the group design project at Bauhaus Luftfahrt e.V. [
1
]. The mission statement of the project has been defined as
follows:
Design a long-haul traffic concept,
fulfil emission reduction goals,
incorporate measures to enhance both operational (on the air transport system level) and technical efficiency (on
the aircraft level),
keep in mind passenger comfort and requirements.
According to these goals, an interdisciplinary group of Bauhaus Luftfahrt e.V. worked on the three different focus areas
business model innovation, energy supply scenario and aircraft & cabin design in order to achieve a holistic approach
for the long-haul emission reduction. Concerning the aircraft and cabin design, important aspects included facilitating
the integration of new energy sources, increasing the passenger comfort as well as integrating new technologies.
During the development of the novel business model, which features elements regarding aircraft sharing, continuous
connecting banks to reduce on-ground time and novel on-board services, several inefficiencies in the current long-haul
network have been identified. These include a high amount of non-utilized time per aircraft due to scheduling and a
majority of current passenger travels being indirect connections. Project studies showed that aircraft sharing enables a
more efficient network structure and has a fleet reduction potential of roughly 25%. The result of this is an increased
load factor and a higher amount of passengers served per flight. A following assessment of today’s network structure
revealed a high potential for an aircraft designed for roughly 400 PAX and a design range of 6400
nm
, having the highest
impact on both fleet reduction as well as non-utilized time. Together with an enlarged cargo deck containing space for
novel business models during cruise, this results in a required total payload of 46,000 kg.
During the assessment of the business model, it was determined that a reduction in cruise speed for emission saving
purposes is feasible while at the same time maintaining the desired flight schedule with the reduced fleet size assessed
before. Therefore, the project’s design cruise Mach has been set to 0
.
7. Due to the slower speed and the longer travel
times (about 20% longer when compared to Mach 0
.
82 cruise), the cargo deck has been designed to fit modular sleeping
compartments and other interactive modules, depending on the flight type and time of day. Furthermore, due to the
increased time passengers spend on board, the cabin design had to be adapted to provide more comfort to passengers
during the extended journey.
Lastly, the desire for reduced emissions lead to the decision for LH
2
as energy storage, promising the highest benefits
compared to other technologies such as sustainable bio fuel. This is backed by an assessment done for the energy supply
scenario of the group design project [2].
IV. Other Hydrogen Aircraft Concepts
The first study on the application of hydrogen for long range civil aircraft known to the authors was undertaken in
1975 by Brewer et al. In their study on passenger aircraft they explored different configurations for varying design ranges
(3000
5500
nm
), cruise Mach numbers (0
.
8-0
.
9) and number of passengers (400-800). One of the main findings
was that none of the considered unconventional configurations (including a flying wing, twin-fuselage, canard and
a twin-boom configuration) showed superiority to a conventional one. A comparison with a Jet-A powered aircraft
furthermore revealed that the hydrogen aircraft was lighter, quieter, required less energy, minimized emissions, and had
2
a smaller span and wing area, but a larger fuselage. A feasibility analysis in terms of operations, maintenance and safety
identified no show stoppers, but indicated the crucial role of the hydrogen distribution network. [3]
In the year 2000, Sefain analysed hydrogen aircraft safety, operations and certification in detail, also taking into
account hydrogen production, storage and delivery. Additionally, he designed a twin-boom medium range (4000
nm
)
180 PAX aircraft. The main findings included that hydrogen is at least as safe as kerosene and that there are no show
stoppers regarding the twin-boom configuration. In contrast to the findings of [
3
], Sefain indicated the high potential of
the unconventional twin-boom configuration using LH2. [4]
Within the CRYOPLANE project, a system analysis of LH
2
aircraft was undertaken, investigating conventional and
unconventional configurations (including a blended wing body and a twin-boom) of various sizes and categories. The
main conclusions comprise that hydrogen aircraft are expected to be as safe as kerosene fueled aircraft and that it is
technically feasible to use LH
2
to fuel aircraft. Furthermore, analogously to the results stated in [
3
], the unconventional
configurations studied did not show advantages over the conventional ones. All examined LH
2
configurations showed a
higher operative empty weight (OEM) and a higher energy consumption when compared to kerosene fueled aircraft. [
5
]
Focusing on the environmental impact of LH
2
aircraft, Svensson assessed the global warming potential (GWP) of a
medium range (4000
nm
) 180 PAX LH
2
aircraft designed using the design environment PIANO. In comparison to a
kerosene fueled reference aircraft the designed LH
2
aircraft had a higher OEM and utilized 10% more energy. As for
the GWP, his main findings indicate that a reduction of cruise altitude leads to a reduction in GWP for the LH
2
aircraft
and an increase in GWP for the kerosene fueled aircraft. [6]
Recently, Verstraete undertook design studies on long range (7500 nm) civil LH
2
aircraft. In 2009 he evaluated 380
and 550 PAX conventional configurations and a 550 PAX twin-fuselage configuration. In line with the results from
[
3
], the results of the 380 PAX configuration showed a 25% MTOM reduction (at a similar OEM) and consequently a
20% wing area reduction when compared to a kerosene fueled aircraft. For the 550 PAX category however, Verstraete
identified the higher market potential for the twin-fuselage configuration. Four years later he refined the 380 PAX
conventional configuration study using the aircraft design environment FLOPS. Again aligned with the findings of
[
3
], the results show that the wing area and required energy of a LH
2
aircraft are smaller and the fuselage larger in
comparison to a kerosene fueled aircraft. The 30% MTOM reduction indicated in this study is even higher than the
MTOM reduction in the study of 2009. [7, 8]
V. Particularities of LH2Aircraft Design
The design of aircraft powered by LH
2
introduces specific features. Especially the fact that a quantity of LH
2
with
equivalent energy content as kerosene has about four times the storage volume, strongly impacts the aircraft design
process. In addition, LH
2
cannot be stored in integral tanks like kerosene, but must be stored in special tanks. These
tanks must be more pressure resistant than integral kerosene tanks and at the same time have excellent thermal insulation
in order to prevent the LH
2
from entering the gaseous phase. First studies on engine level modeling indicated equal
thrust specific power consumption (TSPC) for kerosene and LH
2
powered gas turbines. Thus in the study of this paper,
models for kerosene burning geared turbo fans (GTFs) are used and fuel flow values are adopted to provide equal TSPC
for LH
2
. The proposed conceptual tank structure is based on information concerning materials, material characteristics
and components found in literature [9–13].
These calculations serve as a first mass figure of the LH
2
tanks to compare it with the conventional fuel tank. The
LH
2
tank comprises an inner layer of carbon fiber reinforced plastics (CFRP) followed by an aluminium liner, which acts
as a hydrogen barrier in order to minimise hydrogen diffusion off the tank. These layers are succeeded by an insulation
of Freon blown polyurethane foam. The thickness of the thermal insulation layer is designed for 0
.
5% boil-off rate in
3
h
ground time at sea level and ISA deviation of
+
25
K
. Pressure-wise the tank is designed to withstand the pressure
difference at FL 410. For the tank size a cylindrical shape with two hemispheres or with two torispherical ends is
assumed [9, 11–13].
VI. Annexed Technologies
Along the choice of LH
2
as primary energy source and the resulting required changes to the aircraft design, further
technologies have been investigated within this study and the Hyliner (2.0) concept. They are explained in the following.
3
A. High Aspect Ratio Wing and Bell-Shaped Lift Distribution
The efficiency driven design calls for the application of drastic measures. To attain the emission reduction goals, the
Hyliner (2.0) features a high aspect ratio wing. For subsonic aircraft with a high altitude requirement, high aspect ratios
are often an incorporated feature, as can be seen on the Lockheed Martin U-2 or the Northrop Grumman RQ-4 Global
Hawk. Furthermore, the wing geometry of the Hyliner (2.0) does not need to accommodate volume for fuel storage,
thereby enabling planforms that otherwise would yield to little tank volume.
Another technology that has reemerged in aviation research is the concept of non elliptical, but bell-shaped lift
distributions [
14
,
15
]. Due to the high aspect ratio wing, the compromise between wing drag, weight and wingspan is
shifted from an elliptical to a bell-shaped lift distribution. This concept was applied on the Hyliner (2.0) and the results
are shown in Fig. 1. These results where derived using the LTH handbook method until an aspect ratio of twelve [
16
].
For greater aspect ratios the gradient of the mass increase was enlarged according to in-house numerical simulation
results. There is a smooth transition between these two models. Weight-wise, a wing with an optimized lift distribution
is as heavy as an elliptical loaded wing with 13.8% less wing span or 25.7% less aspect ratio [17].
Fig. 1 High aspect ratio wing mass estimation
On the left hand side of Fig. 1 the wing mass is set in relation to the mass of a wing with aspect ratio of eight and an
elliptical lift distribution. This ratio is presented in dependence of the aspect ratio (AR) for an elliptic lift distribution
(blue line) and a bell-shaped/optimal lift distribution (orange line). Before and aft the smoothed transition (at AR =
12 for the blue line, and AR = 16
.
2for the orange line) the mass is linearly dependant on the AR. On the right hand
side of the figure, the ratio of the two curves is shown. Up to an AR of ten the wing with the optimum lift distribution
has around 87% of the weight of a wing with elliptical lift distribution. In between AR twelve and eighteen both
models transition to the steeper mass increase calibrated to in-house numerical simulation results. After both models
finished the smoothed transition phase, the weight ratio of the bell-shaped lift distribution offers a weight saving of 31%
compared to the wing with the elliptical lift distribution. Thus, highest benefits for a an optimum lift distribution can be
achieved at aspect ratios greater than eighteen.
B. Natural Laminar Flow (NLF)
Compressible, viscous flow over bodies produces parasite drag which can be divided into skin friction drag (D
f
) and
pressure drag due to separation (Dp). The extent of this drag depends on the flow condition: laminar or turbulent.
To reduce drag it is advantageous to shift the transition point from laminar to turbulent flow downstream to benefit
from the lower D
f
in laminar flow. This can be achieved by e.g. decreasing the surface roughness, designing an airfoil
4
with a favourable pressure gradient or flying at a low Reynolds number (high altitude, low airspeed, short wing chord
(characteristic length)). [18]
In practice however, there are some challenges associated with laminarity. Among others these are contamination
(e.g. due to insects, de-icing agents, dust), atmospheric disturbances (e.g. due to rain), acoustic disturbances and
vibrations and material deformation (e.g. due to rivets, wing in-flight bending, reparations). This may lead to higher
maintenance efforts and more complex ground handling processes. In addition, new techniques to estimate degradation
and life time reliability are required to make this technology usable. [19]
However, the reduced cruise speed makes the application of laminar wing technologies very attractive, which is also
favoured by the high aspect ratio. The Hyliner(2.0) has a low sweep wing as it is very beneficial for NLF and the low
cruise Mach number enables a low sweep wing design. The preliminary results of the EC-funded BLADE demonstrator
showed potential for NLF even for Mach 0
.
78 [
20
]. This makes an application of NLF on a wing designed for Mach 0
.
7
and for the time horizon of 2040
+
very feasible. The reduced chord length, due to the high aspect ratio and non-existing
tank volume requirement positively supports large laminar wing surface areas.
VII. Reference Aircraft and Mission
As a baseline for this design study on LH
2
aircraft, a kerosene powered aircraft in the class of an A330-300 with an
advanced technology level of 2040
+
is chosen. This aircraft is called R2040
+
in the following. It is chosen to cruise at
a Mach number of 0
.
82 with a payload of 46
t
(as explained in Section III). This number of passengers was chosen as
future passenger growth rates are expected to be between 3
.
7% and 4
.
4% per year [
21
,
22
]. Furthermore, a long haul
aircraft (design range of 6400
nm
) was chosen as they contributed to 36% of global air transport CO
2
emissions and
moved 10% of all air transport passengers in 2016 [
23
,
24
]. The aircraft design was conducted in the software Pacelab
APD [25]. An overview of the technical specifications of the R2040+ is provided in Table 1.
Table 1 Key specifications of the R2040+reference aircraft
Parameter Value
Wing span [m]67
Wing loading [kg/m2]713
Aspect ratio []12
MTOM [t]264
OEM [t]138
Payload mass [t]46
Design range [nm]6400
Cruise Mach number []0.82
Fuel mass design mission [t]72.5
Tank volume [m3]128
VIII. Concept Collection and Down Selection
During the course of the group design project, various different aircraft concepts had been developed by the
participants. The existing concepts included a blended wing body, a twin fuselage, a default tube-and-wing option as
well as different LH
2
tank arrangements such as as placement in front and aft the cabin, a cylindrical layout above the
cabin and a multi-fuel aircraft.
In order to derive the most suitable aircraft concept for the given requirements, a down selection process has been
performed evaluating the concepts on more than 20 different criteria. The criteria were split into five different groups:
aerodynamic design, overall system integration, weights, aircraft & ground operations as well as emissions. Each
of the groups contained multiple different sub-criteria covering key aspects of the configurations. For example, the
aerodynamics aspects covered the different drag components, the system integration covered aspects such as the system
complexity and efficiency and the weights category consisted of criteria covering the different structural parts of the
aircraft. Considering the operations group, topics covered were airport compatibility, safety, passenger comfort and
maintainability whereas the emissions group consists of both internal and external noise, noise shielding as well as
5
greenhouse gas emissions. Additionally, the maturity in terms of the applied technologies and configuration was taken
into account.
At the end of the down selection procedure, a conventional tube and wing configuration being presented in the
following was identified as the most promising concept, as it was not only the most mature concept, but also received
a comparable score to the other configurations in each of the categories. The choice of a conventional configuration
layout is in-line with the findings of [3, 5, 7] as discussed in Section IV.
IX. Derivation of Target Design
The introduction of LH
2
and several new technologies to the aircraft design process is a challenge. In this section
the process of the implementation of new flight techniques and new technologies is described as well as the iterative
process to derive a well balanced design fulfilling high vehicle efficiencies, passenger comfort demands and providing
the possibility to enable airlines to incorporate new business models.
The final derived aircraft is named Hyliner (2.0) and is shown in Fig. 2. Especially noticeable are the blue LH
2
tanks in front and aft of the cabin. The tank in front of the cabin has a cat walk on one side to enable crew members to
walk from the cockpit to the cabin.
Fig. 2 Illustration of the Hyliner (2.0) aircraft
Table 2 shows characteristic data of the achieved LH
2
long haul aircraft in comparison to the R2040
+
aircraft
described in Section VII. Notice that two aircraft cruising at different Mach numbers are compared. This is done on
purpose to show the difference of this advanced LH
2
aircraft compared to an aircraft which would be expected to fly in
the same time regime if there is only evolutionary change to the aircraft and air transport system. In this comparison
particularly the change in fuel mass and corresponding tank volume is noticeable. The LH
2
aircraft has a MTOM which
is reduced by
25%. This allows for the installation of less thrust with the respective benefits. Another interesting
result is the relatively high wing loading for the LH
2
aircraft which cruises at Mach 0
.
7. This wing loading is still lower
than for the conventional aircraft cruising at Mach 0
.
82, but higher than expected for a conventional aircraft cruising at
Mach 0
.
7. This high wing loading is the result of the wing loading optimization. Field performance constraints can
easily be met and at the same time similar cruise altitudes as for conventional aircraft are reached. Combined with the
high aspect ratio wing, which favours high lift coefficients, optimal cruise efficiencies can be reached.
6
Table 2 Key specifications of the 2040+LH2(Hyliner (2.0)) aircraft compared to the reference aircraft R2040+
Parameter R2040+Hyliner (2.0)
Wing span [m]67 81
Wing loading [kg/m2]713 588
Aspect ratio []12 19.5
MTOM [t]264 196
OEM [t]138 128
Payload mass [t]46 46
Design range [nm]6400 6400
Cruise Mach number []0.82 0.7
Fuel mass design mission [t]72.5 18.6
Tank volume [m3]128 371
A. Design Choices and Cascade Effects
In order to derive a balanced aircraft design for the LH
2
powered aircraft meeting the TLARs several steps were
performed. The starting point is the R2040+ aircraft (Nr. 1 in Fig. 3). In a first step the cruise Mach number was
reduced to Mach 0
.
7(Nr. 2 in Fig. 3). Afterwards the wing geometry was changed to enable a bell-shaped/optimal lift
distribution for minimum induced drag for a given wing weight [
17
] as well as NLF and the aspect ratio was increased
(Nr. 3 in Fig. 3). To allow NLF modelling a transition Reynolds number of thirteen million was assumed. A prominent
example for an aircraft with high aspect ratio is the Boeing sugar study [
26
]. This study extended the aspect ratio to 19
.
5.
Hence, for the study in this paper this aspect ratio is adapted. Advantages and challenges for the chosen technologies are
discussed in Section VI. The aircraft including these new technologies then was adapted for the usage of LH
2
(Nr. 4 in
Fig. 3). The LH
2
tanks are integrated in front and aft the passenger cabin. The impact on the MTOM and OEM as well
as on used energy (heating value of used energy source) during the design mission of these design choices can be seen
in Fig. 3.
Fig. 3 Impact of design choices on energy and weights
7
Design Nr. 4 suffers great penalties because of the low volume to surface ratio of the LH
2
tanks as the fuselage
diameter limits the tank diameter. Consequently, a double deck configuration was investigated. With a double deck
configuration the volume to surface ratio for the LH
2
tanks can be increased significantly which results in a drastically
reduced OEM (see Nr. 5 in Fig. 3). This also lowers the energy consumption of the aircraft. The energy consumption
of this optimized aircraft is almost the same as the energy consumption of a conventional aircraft incorporating the
same advanced annexed technologies. To provide additional passenger comfort and enable new business models for the
airlines, the fuselage diameter is increased until the height of the cargo hold allows the usage of advanced containers
in which people can stand and walk to offer additional services to the passengers (see Section IX.B). This results in
approximately 9% higher design mission energy consumption (see Nr. 6 in Fig. 3). Additionally this cabin configuration
provides more seating and activity space to the passenger than cabins nowadays (see Section IX.B). The circular cross
section and the diameter of the chosen fuselage is a result of the iterative cabin, fuselage and LH2tank design.
Additionally, trade studies regarding different tank positions for this aircraft were conducted, which are shown in
Fig. 4. Here, the effect of over-cabin-tanks compared to tanks positioned in front and aft of the cabin is analyzed. The
y-axis shows relative block fuel consumption compared to the point with the least amount of block fuel in this trade
study. On the x-axis the LH
2
split factor is applied. The LH
2
split factor is a measure of the fraction of the LH
2
stored
in a tank above the cabin - which has the same length as the cabin - and the two tanks in front and aft the cabin.This
means that for a LH
2
split factor of zero all LH
2
is stored in front and aft the cabin. For a LH
2
split factor of one all fuel
is stored in an over-cabin-tank. It can be seen clearly that storage of the fuel in two tanks in front and aft the cabin is by
far the most efficient solution for this aircraft concept. This corresponds with the tank configuration chosen for the long
range aircraft concept analysed in the CRYOPLANE project [
27
]. The high penalty at split factors around 0.1 can be
explained easily: At this LH
2
split factor there is only a small amount of LH
2
stored in the over-cabin-tank. This results
in a very low volume to surface ratio which increases the needed amount of insulation for the tank tremendously. This
does not only result in a very heavy tank, but also in a very large fuselage cross section to contain passengers and the
tank including insulation.
The trade study was conducted for cylindrical tanks with torispherical end caps. Componentwise torispherical ends
result in a slightly heavier tank, but the overall length of the tank is shorter than a tank with spherical end caps containing
the same volume. The trade study trends are similar for tanks with spherical end caps. For this aircraft torispherical tank
end caps result in a lower design mission energy consumption for all points of the trade study compared to spherical
ends.
Fig. 4 LH2tank integration effects
Fig. 5 shows the design mission profiles for the different aircraft designs. The intermediate LH
2
(Nr. 4 and Nr. 5 in
Fig. 3) aircraft are not shown, since they show very similar behaviour as Nr. 6. The good performance of the high aspect
ratio wing flattens the drag polar for high C
L
values to an extend that a wing loading optimization for minimum fuel
burn for the optimum conventional aircraft (Nr. 3 in Fig. 3) results in a flat cruise at the highest allowed flight level of
41000 ft. Since the aerodynamic model in this study does not include a lift coefficient dependent wave drag component,
the glide ratio for high lift coefficients can be overestimated and thus the model might induce an overestimation of the
8
performance of an high aspect ratio wing. Compared to the MTOM block fuel fraction of 22% of Nr. 3, the Hyliner
(2.0) aircraft only has a 10% block fuel fraction. Due to this small weight change of the Hyliner (2.0) compared to the
other designs, its design mission does not require a flight level change and allows to fly in a smaller C
L
range during
cruise. This offers the possibility to implement efficiency increasing technologies that are only adoptable for a very
small range of operating conditions.
Fig. 5 Design mission profiles of designed aircraft concepts
B. Cabin Design
The cabin design of the aircraft is strongly driven by the existence of the two spherical hydrogen tanks in front of and
aft of the cabin, leading to a circular cross section with constant dimension throughout the cabin. Due to the fuselage
diameter of 8
.
46
m
, a two deck configuration with an enlarged cargo deck has been chosen as the most suitable design
(see Fig. 6). The available width enables a 2-4-4-2 seat abreast configuration with a 35 in seat pitch as an appropriate
layout for the main deck, providing a single-excuse design and an comfortable economy class experience thanks to the
three aisles. The upper deck is arranged with a 2-1-2 abreast premium layout and a 48
in
seat pitch. This leads to the fact
that every economy class passenger has an increased leg room of 5
in
compared to a typical Airbus A380 configuration.
The already vast personal space in the upper premium deck has been kept in line with current cabin designs.
9
Fig. 6 Schematic representation of the aircraft cabin layout and the cargo deck
The increased travel times caused by the lower cruise Mach number demand for an increased passenger comfort
during the flight. This is achieved by providing an economy class with an increased seat pitch as mentioned above
and multiple new dwell areas such as meeting, working or bar areas. Those additional spaces, accessible to everyone,
occupy an area of around 25
m2
and justify the increased cabin space per passenger of the aircraft. Additionally, the
cargo deck has been enlarged to fit interchangeable containers and is accessible for the passengers during cruise. The
height of the cargo deck is 2
.
5
m
, providing easy access for people standing upright. The containers and their content
are part of the novel business model and can contain sleeping compartments or food vending areas depending on the
time of day the flight is scheduled and the current provider of the containers. The business model for the cabin and
cargo deck concept will be highlighted in a separate paper from Bauhaus Luftfahrt e.V. at a later time.
The following figure depicts the cross section of the aircraft and highlights the three deck configuration with the
upper premium deck, the main economy deck as well as the enhanced cargo deck.
Fig. 7 Schematic representation of the aircraft’s cross section
The cabin layout was designed using the PAXelerate cabin design tool [
28
] and obeys to CS25 certification rules
10
regarding the amount and positioning of doors and seats. The current regulation does however not account for a three
aisle design as chosen for this aircraft.
X. Assessment of Emissions
Having designed the hydrogen aircraft, it can now be assessed in terms of its environmental impact. Since the
application of atmospheric models is beyond the scope of this paper, the in-flight emissions produced by the aircraft are
interpreted in a qualitative way. These values are given in Fig. 8, showing the emission changes of the Hyliner (2.0) in
comparison to the R2040+ aircraft. The comparison to the R2040+ aircraft, cruising with Mach 0
.
82, was conducted to
show the full potential of a radical change of the network and aircraft compared to which emissions are expected if only
evolutionary changes in the network and the aircraft take place.
Fig. 8 Change in in-flight emissions of the LH2Hyliner (2.0) vs. the conventional R2040+
The dominant change is the total elimination of carbon dioxide (CO
2
) emissions due to the hydrogen combustion
(100% reduction). As expected, the water (H
2
O) emissions have significantly increased. To date, the impact of water
vapor on the climate is not fully understood and the uncertainties in its determination are high [
29
]. Yet, studies show
that water vapor could have a similar severe climate impact as the eliminated CO
2
in the 90% likelihood range [
30
, Fig.
4]. Nitrogen oxide (NO
x
) emission is also strongly decreased. An additional benefit of the hydrogen combustion is the
avoidance of aerosols [
31
]. This has a strong impact on non-carbon dioxide effects, like e.g. possible reductions of
contrail radiative forcing due to less particles onto which water vapor can condense, build liquid droplets and then freeze
[
32
]. However, currently the level of scientific understanding of the climate impact of contrails and cirrus cloudiness is
low to very low [30], such that drawing a more detailed conclusion is difficult.
Due to uncertainties in climate modelling, a quantitative result can not be given in this paper and will be subject
to future work also within atmospheric physics research. However, a strong reduction of the climate impact of the
presented hydrogen aircraft concept in comparison to the conventional R2040+ might be expected.
XI. Conclusions and Outlook
This paper presented the aircraft and cabin design part of the Hyliner (2.0) group design project at Bauhaus Luftfahrt
e.V. [
1
]. Starting from the given TLARs, an overview of existing hydrogen aircraft studies is given. A key feature of
the design study is the reduced cruise Mach number of 0
.
7. This is the result of a detailed investigation of the flight
behaviour of aircraft on long-haul flights. It has been found that the average utilization per aircraft remains almost
unchanged when the flight Mach number is reduced to 0
.
7[
1
]. However, at the same time reducing the flight Mach
number from 0
.
82 to 0
.
7saves around 10% block fuel (see Fig. 3). After specification of particularities of aircraft design
for hydrogen powered aircraft, beneficial annexed technologies for the chosen flight regime as well as for hydrogen
aircraft are presented. Afterwards, the R2040+ platform is presented as well as the concept collection and down
selection process which results in the chosen Hyliner (2.0) configuration. Besides, the target design design choices and
11
the resulting cascade effects are addressed. As the cabin design is very important for the business model and has a
strong influence on the aircraft design process, several aspects of the cabin are described. As a major result the emission
assessment shows that the Hyliner (2.0) LH
2
aircraft has the potential to strongly reduce radiative forcing of contrails as
well as to completely eliminate in-flight CO
2
emissions and thus meet flight path 2050 goals regarding CO
2
emissions.
Further implementation of annexed technologies promise large improvement potential. As there is no fuel stored
in the wing structure, there should be more empty build volume than in a comparable conventional aircraft even
though the high aspect ratio wing’s structure will occupy more volume inside the wing. This could lead to a possibly
simpler implementation of aeroelastic tailored and controlled wings than in conventional aircraft. Another example of a
synergistic annexed technology is the application of the fuselage wake filling concept (see Fig. 9).
Fig. 9 LH2aircraft employing fuselage wake filling (future concept)
Recent studies by Bauhaus Luftfahrt e.V. indicate that this technology is especially well suited for fuselages with
lower slenderness ratio, as well as aircraft where the drag share of the fuselage is very high compared to the total drag,
which is the case for the presented Hyliner (2.0).
Acknowledgments
The authors would like to thank all colleagues from Bauhaus Luftfahrt e.V. who participated in the group design
project as well as Johannes Michelmann from Technical University of Munich for their support and valuable contributions
throughout the course of the project.
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14
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... It should be noted that the residual effective excess power of the FC power train P co,res,FC (cf. Equation (12)) is conservatively neglected when evaluating η co,e f f ,FC acc. to Equation (13). ...
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