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Hydrogen Powered Long Haul Aircraft with Minimized Climate Impact

  • Deutsche Aircraft
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.
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
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
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 [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
), 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. [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
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. [4]
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. [6]
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
for LH
. 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
distributions [
]. 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 [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
) 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
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)). [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 [
]. 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
emissions and
moved 10% of all air transport passengers in 2016 [
]. 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
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
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
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
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.
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
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
, 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 [
, Fig.
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 [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. [
]. 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.
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... Moreover, studies by Huete et al. (Huete et al., 2022(Huete et al., , 2021, Verstraete (Verstraete, 2015), Brewer (Brewer, 2017), and the Airbus Cryoplane study (Faass, 2001), model the performance of a long-range large aircraft completely powered by LH 2 but using a double-decker tube-wing airframe. Studies by Troeltsch et al. (Troeltsch et al., 2020) and Proesmans et al. (Proesmans and Vos, 2022) simulate the performance of a long-range large LH 2 aircraft which cruises at a lower ...
... Studies on combustionbased LH 2 aircraft (Hoelzen et al., 2022;Lammen et al., 2022;Onorato et al., 2022;Prewitz et al., 2020;Silberhorn et al., 2019;Svensson, 2005;Yang et al., 2022) model the energy performance of a regional/small-to mid-size aircraft for a short-to medium-range application. Other literature evaluating combustion based LH 2 aircraft includes studies that have: (i) evaluated smaller payload and range combinations below what is typical for long-range LTA aircraft (Cipolla et al., 2022;Gomez and Smith, 2019); (ii) significantly change the cruise Mach number and/or altitude, leading to different aircraft design characteristics compared to the typical baseline long-range LTA aircraft (Proesmans and Vos, 2022;Troeltsch et al., 2020); and (iii) investigated bi-fuel aircraft where LH 2 provides a proportion of the required fuel (Grewe et al., 2017). ...
... However, in contrast to these two studies, the Airbus Cryoplane study (Bauen et al., 2020;Faass, 2001) showed a 9 % increase in SEC due to larger wetted area to accommodate LH 2 tank, in a double-decker longrange large aircraft (14,000 km design range seating approximately 300 PAX) (Airbus, 2003). Troeltsch et al. (Troeltsch et al., 2020) designed an LH 2 aircraft (range 11,853 km with 400 passengers) with reduced cruise Mach speed of 0.7 (from 0.82 of the reference Jet-A aircraft), which shows a 9 % increase in SEC compared to the reference aircraft. Proesmans et al. (Proesmans and Vos, 2022) conducted a pareto-optimal design examination of LH 2 powered aircraft for long-range aircraft, where each fuel and aircraft type was designed separately to minimise climate impact or operating costs. ...
Decarbonising long-range aviation is challenging. This study evaluates the performance of six low-carbon fuels and their realistic impacts on aircraft design for a large long-range passenger aircraft using Breguet’s range equation. Liquid hydrogen (LH2) and 100 % synthetic paraffin kerosene (SPK) are the only two alternative fuels found to be viable. Using present-day technology, we find that the design-point specific energy consumption (SEC, MJ/tonne-km) of tube-wing aircraft powered by LH2 and 100 % SPK are 11 % higher and 0.2 % lower relative to Jet-A, respectively. At off-design points, SEC of 100 % SPK and LH2 are always similar to and greater than Jet-A, respectively. LH2 aircraft SEC decreases with increasing range and is less sensitive beyond 10,000 km. In a first, we develop an equation that enables LH2 aircraft weight-sizing. Our results should inform studies on LH2 and 100 % SPK aircraft operating costs and lifecycle emissions.
... Design range (nmi) Brewer long range [7] H2FLY and Deutsche Aircraft Do 328 demonstrator ZeroAvia HyFlyer II H2Fly HY4 (Flown) Airbus ZEROe turbofan [30] FlyZero midsize [31] Verstraete very large long range [46] Clean Sky 2 short range [20] Clean Sky 2 long range [20] Bauhaus Luftfahrt Hyliner (2.0) [82] Cryoplane (CMR1-200) [83] Seitz et al. [84] CHEETA concept [85] Combustion Hybrid Fuel cell GH 2 LH 2 Figure 16: Hydrogen combustion is currently the preferred propulsion system for aircraft larger than small turboprops. ...
... The most commonly proposed configurations involve integrating the tanks into the fuselage of a conventional tube-and-wing configuration, a variety of which are shown in Figure 27. Many reserve the whole cross section in a portion of the fuselage for one or more tanks [7,24,30,31,82] . This approach offers the highest tank volume to surface area ratio because the tanks use the entire fuselage width. ...
... These factors carry a weight penalty, which results in degraded performance compared to the whole cross-section tank configuration. Troeltsch et al. [82] perform a trade study that compares the fuel consumption of a configuration that stores fuel in large tanks fore and aft of the cabin in the whole cross-section to one that stores fuel above the cabin, and also a combination of the two. The configuration with tanks at the front and rear of the cabin that fill the fuselage's cross-section has over 10% lower block fuel burn. ...
... Nowadays, after an hiatus, there is a renewed interest in these type of aircraft due to environmental concerns as noted by more recent flight tests conducted using hydrogen-based fuel cells [180,181]. Direct pollutant emissions can be substantially reduced, in its use phase, for an aircraft using hydrogen-fueled combustion, as combustion of hydrogen mainly releases water vapor (100% for CO 2 , SO and soot emissions; and more than 70% reduction in NO ) [182] or even to zero in the case of using fuel-cells [183]. Although water vapor is a powerful GHG with a global warming potential two or three times higher than carbon dioxide [184], aircraft emissions have limited interference in the overall natural water cycle. ...
... Studies carried out in the Cryoplane project [197] showed that large hydrogen tanks can lead to an increase of energy consumption around 10% when compared with a conventional aircraft due to higher aerodynamic drag and structural weight. Similar values were achieved by Troeltsch et al. [182], where they draw a scenario for 2040 and compared their HPA design also with a conventional aircraft. Verstraete [211] also estimated similar values for small short range aircraft (up to 18%). ...
As climate change is exacerbated and existing resources are depleted, the need for sustainable industries becomes ever so important. Aviation is not an exception. Despite the overall carbon dioxide emissions related to the aviation sector accounts for 2%–4% currently, forecasts for air travel indicate an annual growth of 3%–5% and other industries present more potential to reduce carbon emissions once they recur to an increasing use of renewable energies. This option is more difficult in aeronautics since an efficient and lighter energy storage system is required and the current state of the art in battery technology is far from the specific energy densities of fossil fuels and its production is not friendly to the environment. Thus, a herculean effort to integrate several promising mitigation strategies in an efficient way is required. In this paper, a review of the most upfront solutions towards greener aviation is presented and categorized as follows: concepts of operations, energy storage, propulsion systems, aerodynamics, structures, materials, and manufacturing processes. In the end, potential synergies between the different technologies to achieve green aviation are proposed.
... However, their usage leads to a net and not true zero CO 2 emission reduction [5,12]. Second, new aircraft propulsion and fuel systems using green hydrogen (H 2 ) are promising options to achieve true zero CO 2 emissions in aviation, but also to reduce other climate impacts from NO x emissions or from contrail and cirrus cloud creation [1,[13][14][15][16][17][18][19][20][21][22][23][24]. Such systems might be hydrogen-electric using fuel cells and electric motors or direct hydrogen combustion jet engines. ...
... Considering also non-CO2 emissions, H2 combustion could reduce climate impact in flight by 50 to 75%, and fuel-cell propulsion by 75 to 90% [17]. Pure hydrogen may be stored as a compressed gas in a pressurized tank or as a liquid in a cryogenic tank [18]. It may be also safely stored as a metal hydride or extracted from a hydro-carbon like jet fuel, a process known as reformation. ...
Conference Paper
This paper is framed in the context of the GENESIS Project (Gauging the ENvironmEntal Sustainability of electrIc and hybrid aircraft Systems), which complies with the European Union topic JTI-CS2-2020-CFP11-THT-13 (Sustainability of Hybrid-Electric Aircraft System Architectures) as part of the Clean Sky 2 programme for Horizon 2020. The research work is focused on gauging the environmental sustainability of electric aircraft in a life-cycle-based, foresight perspective to support the development of a technology roadmap for transitioning towards sustainable and competitive electric aircraft systems. The analyzed aircraft segment is regional aircraft, to identify, design and assess prospectively the best energy storage and transmission topology. Different alternatives including batteries, fuel cells, hybrid and conventional powertrain technologies are evaluated and compared over different time horizons. In particular, the paper is focused on the description of the workflow implemented to define the Top-Level Aircraft Requirements for a non-conventional regional class hybrid-electric aircraft with 50 passengers, and on the identification of key specifications in terms of on-board energy storage, shaft power level and weight.
... In the context of hydrogen storage on aircraft, several problems need to be addressed. First, low density of pure hydrogen at ambient conditions involves the need for it to be compressed or liquefied (LH2) in order to safely store a significant amount [8]. The extraction of electricity from hydrogen can be efficiently achieved by means of continuously supplied fuel cells, power conversion units which are sized based on the power requirement. ...
Conference Paper
The growing sensitivity to the problem of sustainability requires a rethinking of how aviation is typically conceived by modern society. The aim of the research today must be to analyze the feasibility of disruptive solutions, which drastically reduce consumption and make it possible to meet the growing demand in the commercial aviation sector. The current level of technological maturity does not allow direct implementation on large commercial aircraft, which are responsible for most of the emissions from aviation. In this context, the Clean Sky 2 ELICA project aims to trace a technological roadmap towards green aviation, using the Small Air Transport as a test bed. Two different 19-seat commuter aircraft are presented in this work. The first one, with entry into service in 2025, presents a hybrid-electric architecture with batteries. The second configuration, with entry into service in 2035, is entirely propelled with hydrogen fuel cells, allowing the direct emissions of carbon and nitrogen oxides to be totally eliminated. Both configurations benefit from distributed electric propulsion.
... Preliminary studies on overall LH2 aircraft design involved airplanes of different categories. In [20][21] [22] the liquid hydrogen propulsion is integrated in long range aircraft, while in [23] a LH2 medium-range airliner is investigated; [24] proposed a research on LH2 regional aircraft. A summary of these aircraft designs is proposed in [25], in which the sizing process is applied to LH2 transport aircraft of regional, medium-range, and longrange respectively, providing a general comparison of the advantages/disadvantages of each case-study. ...
Liquid hydrogen (LH2) may enable the decarbonisation of long-haul aviation. However, its low volumetric energy density and subsequent tank space and weight requirements could penalise an aircraft's specific energy consumption (SEC, MJ/tonne-km). We evaluate the impacts of developments in four technology areas – aerodynamics, structures, cryo-tank gravimetric index (η), and overall efficiency (η_o) – on the design-point performance of a large subsonic tube-wing LH2 aircraft. We characterise the critical value of η, which must be exceeded to enable a given design range. For a design range of 14,000 km, η must exceed 0.52 today but only 0.35 with expected 2030 airframe and engine efficiency improvements. Using the most optimistic technology development estimates we observe that SEC could reduce by ∼25% via improvements in η_o and aerodynamics and by 33% via improvements in all four areas. Developments in technologies to improve and reduce drag are critical to enabling zero-carbon long-haul air travel.
Conference Paper
Full-text available
Because the wing-structure weight required to support the critical wing section bending moments is a function of wingspan, net weight, weight distribution, and lift distribution, there exists an optimum wingspan and wing-structure weight for any fixed net weight, weight distribution, and lift distribution, which minimizes the induced drag in steady level flight. Analytic solutions for the optimum wingspan and wing-structure weight are presented for rectangular wings with four different sets of design constraints. These design constraints are fixed lift distribution and net weight combined with 1) fixed maximum stress and wing loading, 2) fixed maximum deflection and wing loading, 3) fixed maximum stress and stall speed, and 4) fixed maximum deflection and stall speed. For each of these analytic solutions, the optimum wing-structure weight is found to depend only on the net weight, independent of the arbitrary fixed lift distribution. Analytic solutions for optimum weight and lift distributions are also presented for the same four sets of design constraints. Depending on the design constraints, the optimum lift distribution can differ significantly from the elliptic lift distribution. Solutions for two example wing designs are presented, which demonstrate how the induced drag varies with lift distribution, wingspan, and wing-structure weight in the design space near the optimum solution. Although the analytic solutions presented here are restricted to rectangular wings, these solutions provide excellent test cases for verifying numerical algorithms used for more general multidisciplinary analysis and optimization.
For the future usage of hydrogen as an automotive fuel, its on-board storage is crucial. One approach is the storage of liquid hydrogen (LH2, 20 K) in double-walled, vacuum insulated tanks. The introduction of carbon fiber reinforced plastics (CFRP) as structural material enables a high potential of reducing the weight in comparison to the state-of-the-art stainless steel tanks. The generally high permeability of hydrogen through plastics, however, can lead to long-term degradation of the insulating vacuum. The derived objective of this dissertation was to find and apply an adequate permeation barrier (liner) on CFRP. The investigated liners were either foils adhered on CFRP specimens or coatings deposited on CFRP specimens. The coatings were produced by means of thermal spraying, metal plating or physical vapor deposition (PVD). The materials of the liners included Al, Au, Cu, Ni and Sn as well as stainless steel and diamond-like carbon. The produced liners were tested for their permeation behavior, thermal shock resistance and adherence to the CFRP substrate. Additionally, SEM micrographs were used to characterize and qualify the liners. The foils, although being a good permeation barrier, adhered weakly to the substrate. Furthermore, leak-free joining of foil segments is a challenge still to be solved. The metal plating liners exhibited the best properties. For instance, no permeation could be detected through a 50 µm thick Cu coating within the accuracy of the measuring apparatus. This corresponds to a reduction of the permeation gas flow by more than factor 7400 compared to uncoated CFRP. In addition, the metal platings revealed a high adherence and thermal shock resistance. The coatings produced by means of thermal spraying and PVD did not show a sufficient permeation barrier effect. After having investigated the specimens, a 170 liter CFRP tank was fully coated with 50 µm Cu by means of metal plating.
Hydrogen is since long seen as an outstanding candidate for an environmentally accept- able, future aviation fuel. Given that most comprehensive studies on its use in aviation were performed over two decades ago, the current article evaluates its potential as a fuel for long range transport aircraft at current and future technology levels. The investigations show that hydrogen has the potential to reduce the energy utilisation of long range transport aircraft by approximately 11%. The use of hydrogen namely allows a much smaller wing area and span since the wing size is not restricted by its fuel storage capacity. At a given price per unit energy content, the smaller wings lead to a reduction of around 30% in take-off gross weight and 3% in direct operating costs for a given fuel price per energy content. The hydrogen-fuelled aircraft are furthermore slightly more sensitive to a possible reduction in operating empty weight in the future and 20% less sensitive to further improvements in engine thrust specific fuel consumption.
Conference Paper
xemissions are less severe and persistent contrail formation is less likely. By considering these altitude eects and additionally applying climate impact reduction technologies, impacts can be reduced by 45-70% with simultaneous direct operating cost savings of 0.6%. Uncertainty is assessed, demonstrating that relative climate impact savings can be expected despite large scientic uncertainties. Additionally, operating cost savings are predicted over a wide range of fuel prices. Strategies for improving climate performance of existing aircraft are also explored, revealing potential climate impact savings of 20-40%, traded for a 2% increase in total operating costs and reduced maximum range.
Perturbations in H2O caused by aircraft in the year 2015 are calculated with a chemical transport model (CTM) and used as input for radiative forcing calculations. The focus is on a hypothetical fleet of cryoplanes, i.e., liquid-hydrogen-powered aircraft. For comparison, the effects of subsonic and supersonic kerosene aircraft are assessed. The CTM applies the accurate second-order moment scheme for advective transport and is run in T42 resolution (2.8° × 2.8°) with 40 layers between the surface and 10 hPa. Aircraft emissions are taken from NASA inventories for the year 2015. In the cryoplane experiments the projected subsonic kerosene fleet is replaced completely by cryoplanes, which emit 2.55 times as much H2O. Longwave and shortwave components of radiative forcing due to the modeled H2O increases are calculated for different seasons. In northern midlatitudes near the tropopause, the fleet of cryoplanes is calculated to increase zonal-mean H2O by more than 250 ppbv on an annual average. The resulting radiative forcing at the tropopause strongly depends on season, ranging from 0.0027 W/m2 in October to 0.0135 W/m2 in April on a global average. Subsonic kerosene aircraft are found to have a rather small impact on H2O levels and lead to an annually averaged global-mean radiative forcing of 0.0026 W/m2. Supersonic kerosene-powered aircraft have a more pronounced impact on H2O concentrations than subsonic cryoplanes and cause a radiative forcing of nearly 0.05 W/m2. Several sensitivity studies are performed for cryoplanes, dealing with cruising altitude, tropopause height, tropospheric lifetime, and stratospheric sinks of aircraft-emitted water vapor.
The feasibility, practicability, and potential advantages and disadvantages of using liquid hydrogen as fuel in long range, subsonic transport aircraft of advanced design were studied. Both passenger and cargo-type aircraft were investigated. To provide a valid basis for comparison, conventional hydrocarbon (Jet A) fueled aircraft were designed to perform identical missions using the same advanced technology and meeting the same operational constraints. The liquid hydrogen and Jet A fueled aircraft were compared on the basis of weight, size, energy utilization, cost, noise, emissions, safety, and operational characteristics. A program of technology development was formulated.