Hydrogen Powered Long Haul Aircraft with Minimized Climate
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
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
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.
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
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
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. [
]. The mission statement of the project has been deﬁned as
•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
, 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
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 .
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
), cruise Mach numbers (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
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. 
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
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 [
], Sefain indicated the high potential of
the unconventional twin-boom conﬁguration using LH2. 
Within the CRYOPLANE project, a system analysis of LH
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
to fuel aircraft. Furthermore, analogously to the results stated in [
], the unconventional
conﬁgurations studied did not show advantages over the conventional ones. All examined LH
conﬁgurations showed a
higher operative empty weight (OEM) and a higher energy consumption when compared to kerosene fueled aircraft. [
Focusing on the environmental impact of LH
aircraft, Svensson assessed the global warming potential (GWP) of a
medium range (4000
) 180 PAX LH
aircraft designed using the design environment PIANO. In comparison to a
kerosene fueled reference aircraft the designed LH
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
and an increase in GWP for the kerosene fueled aircraft. 
Recently, Verstraete undertook design studies on long range (7500 nm) civil LH
aircraft. In 2009 he evaluated 380
and 550 PAX conventional conﬁgurations and a 550 PAX twin-fuselage conﬁguration. In line with the results from
], 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
], the results show that the wing area and required energy of a LH
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
introduces speciﬁc features. Especially the fact that a quantity of LH
equivalent energy content as kerosene has about four times the storage volume, strongly impacts the aircraft design
process. In addition, LH
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
from entering the gaseous phase. First studies on engine level modeling indicated equal
thrust speciﬁc power consumption (TSPC) for kerosene and LH
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
. 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
tanks to compare it with the conventional fuel tank. The
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
ground time at sea level and ISA deviation of
. 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
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.
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
]. 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 [
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 .
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
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
in laminar ﬂow. This can be achieved by e.g. decreasing the surface roughness, designing an airfoil
with a favourable pressure gradient or ﬂying at a low Reynolds number (high altitude, low airspeed, short wing chord
(characteristic length)). 
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. 
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
]. This makes an application of NLF on a wing designed for Mach 0
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
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
(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 [
]. Furthermore, a long haul
aircraft (design range of 6400
) was chosen as they contributed to 36% of global air transport CO
moved 10% of all air transport passengers in 2016 [
]. The aircraft design was conducted in the software Pacelab
APD . An overview of the technical speciﬁcations of the R2040+ is provided in Table 1.
Table 1 Key speciﬁcations of the R2040+reference aircraft
Wing span [m]67
Wing loading [kg/m2]713
Aspect ratio [−]12
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
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
greenhouse gas emissions. Additionally, the maturity in terms of the applied technologies and conﬁguration was taken
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
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
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
long haul aircraft in comparison to the R2040
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
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
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
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
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.
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
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 [
] 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 [
]. This study extended the aspect ratio to 19
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
(Nr. 4 in
Fig. 3). The LH
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
Design Nr. 4 suﬀers great penalties because of the low volume to surface ratio of the LH
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
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
split factor is applied. The LH
split factor is a measure of the fraction of the LH
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
split factor of zero all LH
is stored in front and aft the cabin. For a LH
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 [
]. The high penalty at split factors around 0.1 can be
explained easily: At this LH
split factor there is only a small amount of LH
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
Fig. 4 LH2tank integration eﬀects
Fig. 5 shows the design mission proﬁles for the diﬀerent aircraft designs. The intermediate LH
(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
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
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
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
, 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
seat pitch. This leads to the fact
that every economy class passenger has an increased leg room of 5
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.
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
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
, 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 [
] and obeys to CS25 certiﬁcation rules
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
) emissions due to the hydrogen combustion
(100% reduction). As expected, the water (H
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 [
]. Yet, studies show
that water vapor could have a similar severe climate impact as the eliminated CO
in the 90% likelihood range [
4]. Nitrogen oxide (NO
) emission is also strongly decreased. An additional beneﬁt of the hydrogen combustion is the
avoidance of aerosols [
]. 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
]. However, currently the level of scientiﬁc understanding of the climate impact of contrails and cirrus cloudiness is
low to very low , 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
]. 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
]. 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
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
aircraft has the potential to strongly reduce radiative forcing of contrails as
well as to completely eliminate in-ﬂight CO
emissions and thus meet ﬂight path 2050 goals regarding CO
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).
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.
Paul, A., Engelmann, M., Koops, L., Steinweg, D., Troeltsch, F., van Wensveen, J., and Hornung, M., “Emission reduction
potential across the long-haul air traﬃc network,” Deutscher Luft- und Raumfahrt Kongress 2019, edited by Deutsche
Gesellschaft für Luft- und Raumfahrt, Deutsche Gesellschaft für Luft- und Raumfahrt, 2019.
Penke, C., “Prospektive Ökobilanzierung einer erneuerbaren Wasserstoﬀversorgung für die Luftfahrt,” Ökobilanzwerkstatt
Brewer, G. D., Mor ris, R. E., Lange, R. H., and Moore, J. W., “Study of the Application of Hydrogen Fuel to Long-Range Subsonic
Transport Aircraft,” Volume II Final Report No. NASA CR-132559, Lockheed California Company and Lockheed-Georgia,
 Sefain, M. J., “Hydrogen Aircraft Concepts & Ground Support,” Ph.D. thesis, Cranﬁeld University, United Kingdom, 2000.
Westenberger, A., “Liquid Hydrogen Fuelled Aircraft - System Analysis (CRYOPLANE),” Final Technical Report (Publishable
Version), September 2003.
Svensson, F., “Potential of Reducing the Environmental Impact of Civil Subsonic Aviation by Using Liquid Hydrogen,” Ph.D.
thesis, Cranﬁeld University, United Kingdom, 2005.
Verstraete, D., “The Potential of Liquid Hydrogen for long range aircraft propulsion,” Ph.D. thesis, Cranﬁeld University, United
Verstraete, D., “Long range transport aircraft using hydrogen fuel,” International Journal of Hydrogen Energy, Vol. 38, No. 34,
2013, pp. 14824–14831. https://doi.org/10.1016/j.ijhydene.2013.09.021.
 Brewer, G. D., Hydrogen Aircraft Technology, 1st ed., CRC Press, Bosa Roca, 1991. URL https://ebookcentral.proquest.com/
Colazza, A. J., “Hydrogen Storage for Aircraft Applications Overview,” NASA Corporate Report No. CR-2002-211867,
National Aeronautics and Space Administration, 2002.
 Schultheiß, D., “Permeation Barrier for Lightweight Liquid Hydrogen Tanks,” Ph.D. thesis, University of Augsburg, 2007.
Sonnemann, D., “Strukturmechanische und thermodynamische Auslegung von Wasserstoﬀtanks für Flugzeuge,” Diploma
thesis, Universität der Bundeswehr München, 2006.
Oehlke, M., “Masseabschätzung eines Wasserstoﬀ-Außentanks für ein Turboprop-Verkehrsﬂugzeug,” Technical Report,
University of Applied Sciences Hamburg, 2009.
Bowers, A. H., Murillo, O. J., Jensen, R., Eslinger, B., and Gelzer, C., “On Wings of the Minimum Induced Drag: Spanload
Implications for Aircraft and Birds,” Technical Report No. 20160003578, National Aeronautics and Space Administration, 2016.
Phillips, W. F., Hunsaker, D. F., and Taylor, J. D., “Minimizing Induced Drag with Weight Distribution, Lift Distribution,
Wingspan, and Wing-Structure Weight,” AIAA Aviation 2019 Forum, American Institute of Aeronautics and Astronautics,
Reston, Virginia, 2019, p. 451. https://doi.org/10.2514/6.2019-3349.
Baur, R. v., “Tragwerk Transporter statistische Massenabschätzung,” Tech. Rep. MA 501 12-01, Luftfahrttechnisches Handbuch
- Massenanalyse, 2006.
Klein, A., and Viswanathan, S. P., “Approximate solution of minimum induced drag of wings with given structural weight,”
Journal of Aircraft, Vol. 12, No. 2, 1975, pp. 124–126. https://doi.org/10.2514/3.44425.
 Anderson, J. D., Fundamentals of Aerodynamics, 4th ed., McGraw-Hill, New York, 2007.
Bardewyck, T., “Treibstoﬀeinsparung zukünftiger Verkehrsﬂugzeuge durch Laminarität am transsonischen Flügel – das
Europäische Forschungsprojekt BLADE,” Plenary speech at Deutscher Luft- und Raumfahrtkongress, Friedrichshafen,
Germany, September 5th, 2018.
Michael Gubisch, “Why Airbus foresees laminar wings on next-gen aircraft,” Flight Global, 2018. URL https://www.ﬂightglobal.
International Air Transport Association, “IATA Forecasts Passenger Demand to Double Over 20 Years,” , October 2016.
Airbus Group, “Global Market Forecast 2018-2037,” , 2018. https://www.airbus.com/aircraft/market/global-market-forecast.
OAG Aviation Worldwide Limited, “OAG (Oﬃcial Airline Guide): Database of scheduled ﬂights,” , 2016. https://www.oag.
 Eurocontrol, “Base of aircraft data (BADA),” , 2015. https://www.eurocontrol.int/services/bada.
 TXT, “Pacelap APD,” , 2019. URL https://www.txtgroup.com/markets/solutions/pacelab-apd/.
Bradley, M. K., and Droney, C. K., “Subsonic Ultra Green Aircraft Research: Phase II – Volume II – Hybrid Electric Design
Exploration,” NASA Corporate Report No. CR–2015-218704, National Aeronautics and Space Administration, 2015.
Airbus Deutschland GmbH, “CRYOPLANE: Liquid Hydrogen Fuelled Aircraft – System Analysis: FINAL TECH-
NICAL REPORT,” , 2003. URL https://www.google.com/url?sa=t&rct=j&q=&esrc=s&source=web&cd=2&ved=
 Bauhaus Luftfahrt e.V., “PAXelerate Boarding Simulation,” , 2019. URL www.paxelerate.com.
Dahlmann, K., “Eine Methode zur eﬃzienten Bewertung von Maßnahmen zur Klimaoptimierung des Luftverkehrs,” Ph.D.
thesis, Ludwig-Maximilians-Universität München, Germany, 2011.
Lee, D. S., Fahey, D. W., Forster, P. M., Newton, P. J., Wit, R. C., Lim, L. L., Owen, B., and Sausen, R., “Aviation and global
climate change in the 21st century,” Atmospheric Environment, Vol. 43, No. 22, 2009, pp. 3520 – 3537.
Gauss, M., and Isaksen, I. S. A., “Impact of H
O emissions from cryoplanes and kerosene aircraft on the atmosphere,” Journal
of Geophysical Research, Vol. 108, No. D10, 2003, pp. 21,239. https://doi.org/10.1029/2002JD002623.
 Dallara, E., “Aircraft Design for Reduced Climate Impact,” Ph.D. thesis, Stanford University, California, United States, 2011.