Experts in teams
TPK4852 - Hydrogen in transportation for a safe and
sustainable future
Feasibility Study of Water Retention on
An Airbus A320 Hydrogen Retrofit for
Short Haul Flight
Authors:
Mads Dørum
Matthias Ransau
Mikael Bourmsitrov
Mikael Tolo
Thomas Aleksander Jonsson
January - April, 2025
Nomenclature
∆H
Enthalpy of evaporation
ṁ
Mass flow
Q̇
Heat flow
η
Efficiency factor
η
Fuel cell efficiency
A
Area
C
Specific heat capacity
Cu
Copper
E
Energy
fc
Fuel cell
GHG Greenhouse gas
H2
Hydrogen gas
k
Thermal conductivity of a material
LH2
Liquid hydrogen
LHV
Lower heating value
m
Mass
P
Power
P EM Proton exchange membrane
Q
Heat
T
Temperature
V
Volume
Contents
List of Figures
iii
List of Tables
iv
Abstract
1
1 Introduction
1
i
2 Theory
1
2.1
Hydrogen Fuel Cells . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
1
2.2
Hydrogen Tank . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
2
2.3
Water vapour . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
3
2.4
Thermal management . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
3
2.5
Weight of LH2
. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
4
2.6
Electric Propulsion . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
4
3 Methodology
4
3.1
Fuel Cells and Amount of LH2 . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
5
3.2
LH2 tank . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
5
3.3
Climact impact of Water vapour and technical feasibility of water storage . . . . .
5
3.4
Thermal management of fuel cell . . . . . . . . . . . . . . . . . . . . . . . . . . . .
6
3.5
Propulsion . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
6
3.6
Cutting of cargo space . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
6
3.7
Cost analysis . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
7
3.7.1
Cost Structure . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
7
3.7.2
Key Assumptions . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
7
4 Results
7
4.1
Amount of Hydrogen . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
7
4.2
LH2 Tank . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
8
4.3
Fuel cells . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
8
4.3.1
Fuel cell array placement . . . . . . . . . . . . . . . . . . . . . . . . . . . .
9
4.4
Propulsion . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
9
4.5
Water Vapour . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
10
4.5.1
Water Volume Estimation . . . . . . . . . . . . . . . . . . . . . . . . . . . .
10
4.5.2
Water stored in wing fuel tanks . . . . . . . . . . . . . . . . . . . . . . . . .
10
4.6
Thermal management . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
10
4.7
Cost analysis . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
11
4.7.1
Retrofit Cost Estimation . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
11
4.7.2
Comparative Context and Economic Consideration . . . . . . . . . . . . . .
12
4.7.3
Airport Infrastructure Costs . . . . . . . . . . . . . . . . . . . . . . . . . . .
12
4.7.4
Maintenance . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
12
4.7.5
Liquid Hydrogen Cost . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
12
4.7.6
Cost comparison and impact . . . . . . . . . . . . . . . . . . . . . . . . . .
12
ii
4.8
4.7.7
Cost per passenger . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
12
4.7.8
Sensitivity analysis . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
13
4.7.9
Conclusion . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
13
Final weight of retrofit . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
13
5 Discussion
14
5.1
Calculations . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
14
5.2
Fuel Tank . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
14
5.3
Fuel Cells . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
15
5.4
Water storage . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
15
5.5
Thermal management . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
15
6 Conclusion
16
Bibliography
17
Appendix
20
A Matlab code for thermal management
20
B Simulink diagram for thermal management
20
C Amount of LH2 and fuel cells required for the trip calculations
21
D Take off and landing energy used calculation datasheet
22
E SuperStack-Flex-Datasheet
23
F Motor-HyperCore-Datasheet
24
G Final volume and amount of LH2 needed calculations
25
H Tank volume and weight calculation
26
I
27
Tail cone space calculation
J Motor and cargo space calculations
27
List of Figures
1
Schematic of a foam-insulated hydrogen storage tank with LH2 [1]. . . . . . . . . .
2
2
Stages of the plane ride . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
5
iii
3
A visual representation of how the LH2 tank would fit in an Airbus A320 [2]. . . .
8
4
Details of lower fuselage of an Airbus A320 [3]
. . . . . . . . . . . . . . . . . . . .
9
5
Result from simulating thermal control . . . . . . . . . . . . . . . . . . . . . . . . .
11
List of Tables
1
Operating cost breakdown per one-hour flight . . . . . . . . . . . . . . . . . . . . .
7
2
Amount og LH2 required for each stage of flight . . . . . . . . . . . . . . . . . . . .
8
3
Calculated fuel cell power demand and number needed per relevant flight stage . .
8
4
Weight and size of the fuel cell array . . . . . . . . . . . . . . . . . . . . . . . . . .
8
5
Available space in lower fuselage of the aircraft . . . . . . . . . . . . . . . . . . . .
9
6
Number of HyperCore Engines . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
9
7
Total weight of motors . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
9
8
Cost breakdown for a one-hour flight . . . . . . . . . . . . . . . . . . . . . . . . . .
13
9
Weight comparison of Airbus A320s . . . . . . . . . . . . . . . . . . . . . . . . . .
14
iv
Abstract
As the aviation industry moves toward decarbonization, hydrogen propulsion emerges as a promising zero-emission alternative to fossil fuels. This report investigates the feasibility of retrofitting
an Airbus A320 for one-hour short-haul flights using liquid hydrogen (LH2 ) and hydrogen fuel
cells. In addition to addressing power, storage and energy requirements, the study examines water vapor emissions—a significant environmental concern due to contrail formation and radiative
forcing—and evaluates the technical viability of retaining water produced in-flight. Using simplified energy models and realistic aircraft specifications, the retrofit design incorporates cryogenic
LH2 tanks, PEM fuel cells, and thermal management systems utilizing waste water and LH2 for
cooling. It is concluded that while the retrofit is not feasible with current technology, where significant challenges are fuel cell weight, and economic viability. However cost analysis suggests
hydrogen aviation may become competitive by 2050, contingent on infrastructure development,
reduced hydrogen prices, and regulatory incentives.
1
Introduction
Climate change is one of the most pressing global challenges of our time. In response, international
agreements like the Paris Accord have set ambitious targets to reduce greenhouse gas (GHG)
emissions and limit global warming. As part of its commitment, the European Union has pledged
to reduce GHG emissions and to achieve carbon neutrality by 2050 [4].
The aviation sector plays a significant role in this context. Currently, it accounts for approximately
3.8–4% of total EU GHG emissions and around 2.5% of global emissions [5]. As aviation continues
to grow, decarbonizing this sector will be essential for achieving climate targets and transitioning
to a carbon-neutral society.
Norway emits two times more CO2 per capita than the European average. In addition, the Oslo
Trondheim and Oslo Bergen flights are the third and fourth busiest in Europe [5, 6]. These flights
take around an hour and are about 400 km long.
This report investigates the necessary modifications to retrofitted an Airbus A320 for approximatley
one-hour (∼ 400 km) flights using liquid hydrogen (LH2 ) as a fuel source.
Hydrogen is a promising alternative to conventional jet fuel, offering the potential for zero-emission
flight, only emitting water as a by-product. However, if hydrogen aviation becomes the norm, the
large addition of water vapour into the atmosphere would ironically start heavily contributing to
global warming.
In addition to the retrofit, this study considers the feasibility of retaining the produced water.
Lastly, a cost analysis is included to assess the financial viability of the retrofit.
2
Theory
2.1
Hydrogen Fuel Cells
A hydrogen Fuel Cell is an electrochemical device that converts hydrogen and oxygen into electricity. The specific type of fuel cell used in this project uses the PEM (Proton Exchange Membrane)
architecture. Its only by-product is water and heat. The burn value of hydrogen used in the report
is 120 kJ/kg [7], which assumes heat is being lost [8]. One can calculate the power that goes in
(Pin ) to the fuel cell from the flow of liquid hydrogen ṁH2 :
Pin = ṁH2 · LHVH2
1
(1)
PEM fuel cells have an electrical efficiency η of around 40-60% [9], giving rise to the formula for
power out (Pout ) of the fuel cell:
Pout = η · Pin
(2)
Given the efficiency of the fuel cell and the required power output, Equation (1) and (2) can be
used together to get the required mass flow of hydrogen:
ṁH2 =
Pout
η · LHVH2
(3)
Optimal operating temperatures for these kind of fuel cells are around 80◦ C [10]. The amount
of waste water produced can be derived from the chemical reaction, which is approximately 9kg
water produced per kg hydrogen consumed [11].
2.2
Hydrogen Tank
LH2 is stored at cryogenic temperatures near 20 K and takes up to 4 times the space of conventional
Jet-A fuel (kerosene) per unit of energy [1]. Therefore the design of the tank is crucial.
Thermal insulation is essential to minimize boil-off losses. Several insulation strategies exist,
like vacuum-jacketed designs and foam-based insulation. While a vacuum offers superior thermal
performance, it is often a complex, expensive, heavy structure. Foam insulation, by contrast,
provides an effective and more reliable and robust alternative suitable for systems where some
boil-off (in our case 2% per hour) is acceptable [1, 11, 12, 13].
The tank will be completely filled between each flight. Therefore, the foam based tank is an
acceptable solution, as the boil-off that will happen in the short hour of flight is not significant
enough to motivate a vacuum based solution [13]. The construction of such a foam based solution
can be seen in Figure (1).
Figure 1: Schematic of a foam-insulated hydrogen storage tank with LH2 [1].
Steady-state heat conduction Q through a flat wall follows the following formula [14]:
Q=
k · A · ∆T
,
d
(4)
where k is the thermal conductivity of the material, A is the surface area perpendicular to heat
flow, ∆T is the temperature difference across the material and d is the thickness of the material.
The tank foam insulation must have a very low heat conductivity. A popular material used is
W
polyurethane foam (PUF) [11] which has a thermal conductivity k = 0.25 mK
[15] and a density
kg
of approximately 32 m3 [16]. The wall material must be resistant to cryogenic conditions and to
2
hydrogen embrittlement. In aerospace applications, aluminium-lithium alloys are common due to
g
their favourable strength-to-weight ratios (weighing 2.47 cm
3 [17]) and thermal properties [11, 18].
These materials enable tanks to operate safely at moderate internal pressures while minimizing
wall thickness and mass.
2.3
Water vapour
While hydrogen-power aircrafts eliminate carbon emissions, water vapour emissions at high altitudes present a distinct environmental concern. The upper atmosphere is extremely dry, and even
small increases in moisture can lead to the formation of contrails and cirrus clouds. These cloud
formations trap outgoing infrared radiation, amplifying the greenhouse effect [19]. Unlike how at
lower altitudes where water vapour cycles quickly, moisture in the upper troposphere can persist
for extended periods, due to slow vertical transport and lower temperatures, and have a disproportionately large impact on radiative forcing [20]. Although only 5–10% of atmospheric water
vapour exists above the 500 hPa pressure level, approximately 5.5–6 km (where most flights take
place), it contributes to nearly 50% of the total greenhouse effect of water vapour [19].
The produced water also poses potential risks if released in or near airports. For instance, increased
humidity could contribute to fog formation, reducing visibility, or create slippery conditions on the
tarmac. In colder climates like Norway, this moisture may freeze, forming ice on runways. Water
storage onboard would help mitigate these risks and is therefore an important consideration.
However, storing water produced by hydrogen fuel cells presents several technical challenges. Water
has a high exit temperature (approximately 80°C) from the fuel cells and ambient temperatures
at cruise altitudes can reach as low as -50°C [21], posing a risk of freezing. Additionally, water
is chemically more reactive than jet fuel and can cause corrosion [22], particularly in aluminium
structures typically used for fuel tanks. These factors necessitate specialized materials, insulation,
and system adaptations for safe and effective storage.
2.4
Thermal management
The fuel cells produce a lot of heat inside the plane, which is an enclosed environment and therefore
needs cooling. A numerical method will be used later to find how much energy and cooling the
system needs.
In a thermal system, most of the excess heat goes into heat sinks which are often made of copper
[23]. The total heat capacity Ctot of the heat sink is approximated by using the coppers mass mCu ,
and heat capacity CCu :
Ctot = CCu · mCu
(5)
It is assumed the heat of the fuel cells Qf c is represented by the heat of the heat sinks. To get an
intuitive measurement of the fuel cell temperature Tf c , the heat is convert by using the formula
for heat capacity:
Tf c (t) =
1
Qf c (t)
CCu tot
(6)
It was assumed that all of the heat produced by the fuel cell comes from its loss fraction (1 − η) [8].
While operating at higher altitudes the fuel cell is also less efficient [24]. This means that instead
of using the efficiency factor η used in Equation (2), a somewhat higher efficiency factor ηw has to
be used to calculate the waste heat (Q̇w ) produced by the fuel cell:
Q̇w = (1 − ηw )Pin
3
(7)
A model for heat flow using mass flow with vaporization of a cooling liquid is discussed here; [25],
which will be used as a base in the simulation. In the case of using liquid hydrogen stored close
to its vaporization point as a coolant for the fuel cell, the heat flow during the liquid phase can be
ignored, and the model for the heat flow of the hydrogen Q̇H2 is:
Q̇H2 = −((Tf c − TH2 )CH2 + ∆HH2 )ṁH2 ,
(8)
where ∆HH2 is the enthalpy of evaporation of hydrogen. Simulating the same for water (without
the added effort of vaporization) would result in a similar equation:
Q̇H2 O = −((Tf c − TH2 O )CH2 O )ṁH2 O
2.5
(9)
Weight of LH2
Using the total energy an aircraft uses through a trip (Etrip ), it is possible to find the total amount
of chemical energy needed in the LH2 fuel (ELH2 ) by multiplying Etrip by the fraction of the LH2
efficiency η and the combustion engines efficiency ηcom :
ELH2 = Etrip
η
ηcom
(10)
Then dividing by the burn value of hydrogen, the total weight of hydrogen (kgLH2 ) will be:
kgLH2 =
2.6
ELH2
J
120 M
kg
(11)
Electric Propulsion
An electric propulsion system is needed to fly the plane. With an electric propulsion system it is
necessary to use propellers on the plane. The engines used on the Airbus today are quite powerful
and can push out over 100 kN per engine [26]. It is necessary to achieve the same force, but in an
electric system.
3
Methodology
The process of conducting an aircraft retrofit requires a systematic approach of inspecting relevant systems of the Airbus A320 and replacing them. Here, components identified as needing
replacement will have an explanation of how calculations were done. Further, how factors like
weight, needed power and LH2 was found, is explored. The physical restraints of the aircraft is an
important consideration, meaning re-design or moving some surrounding structures was needed to
accommodate new systems, as it was decided the fuselage was to remain unchanged.
Establishing clear limitations for this retrofit project was an important step in making this report
possible. It was therefore decided that only retrofitting of the plane will be done, as it is assumed
the airport already has the infrastructure needed to support this plane. Further, many assumptions
and simplifications have been made. The point of this report is to find possible barriers in making
such a short haul retrofit plausible.
4
3.1
Fuel Cells and Amount of LH2
To calculate the amount of fuel cells the plane would need, an approximation using the engines of
the original A320 was used. It is equipped with two IAE V2500-A5 engines. Thrust and cruising
performance data was used to calculate the total power output of the engines [26], and therefore
the power output needed by the fuel cells (See Appendix C and D), thus giving total amount of
fuel cells needed.
A flight is separated by three basic stages, as represented in Figure (2). By using the aircraft’s
speed during take-off, landing and cruise, with the power output of the engines in each phase, the
total energy of the flight was found.
Figure 2: Stages of the plane ride
When calculating the amount of LH2 needed for the flight, this energy is used in Equation (10)
and (11), where it is assumed η is 60% at sea level and 40% at high altitude (See Appendix E).
3.2
LH2 tank
To safely contain cryogenic LH2 while minimizing weight and preserving aircraft operational efficiency, a Li-Al metal compound was used for the hard walls between the PUF insulation.
With the amount of hydrogen needed for the trip (see section 4.1) and extra reserve fuel according
to regulations [27, 28, 29], the amount of LH2 accepted for the plane to fly was found (see appendix
kg
G). This is then used to find the volume of the LH2 using 0.0141 m
3 [30]. Then the amount of
material needed for the tank walls were found using Equation (4) assuming a 2% boil-off per hours
as being acceptable.
To ease the calculations, the shape of the tank was approximated to be spherical. Lastly it was
found whether the tank could fit in the aircraft using the maintenance manual of the Airbus A320
[31] and some rudimentary calculations.
3.3
Climact impact of Water vapour and technical feasibility of water
storage
To evaluate the environmental impact of high-altitude water emissions, it was estimated that the
total mass of water vapour within the 6,000–8,000 meter cruise altitude band, using typical water
vapour densities at these heights. This is then compared to the projected water emissions from a
global hydrogen aircraft fleet, assuming a residence time of water up to one year
To assess the technical feasibility of storing the water produced, the existing fuel storage space
within the wings was used. In conventional configurations on the Airbus A320, jet fuel is stored in
tanks located inside the wings [31]. With hydrogen propulsion, this space would remain unused,
presenting an opportunity to repurpose the wing tanks for storing liquid water generated during
flight.
5
To assess the feasibility of water storage in existing wing tanks, the volume of water produced during
a typical one-hour flight was found, using stoichiometric calculations based on expected hydrogen
consumption. This value is compared to the available internal tank volume and structural load
capacities. The technical implications of high water temperatures, freezing risk, and corrosivity
were analysed.
3.4
Thermal management of fuel cell
The thermal management was modelled and simulated in order to get an overview of the effectiveness of different cooling methods that are of particular relevancy. The heat management
calculations focused exclusively on the cruise stage, as this was deemed the most important phase
of flight and has much simpler calculations than the other stages.
It was assumed that all emitted heat from each fuel cell goes into their own copper heat sink of
about 1 kg per fuel cell, and that this is a good enough estimation of the fuel cell’s own temperature.
The liquid hydrogen will be warmed up to operating temperatures by sending it through the heat
exchanger with the coolant of the fuel cell. Based on the storage temperature of hydrogen discussed
in section 2.2, the temperature of fresh hydrogen fuel was set to a constant 20 K. Realizing this is
close to the vaporization temperature of hydrogen, formula 8 was used directly. A simplification
was made that omits the intermediate cooling of the liquid that is actually used to cool the fuel
cell; this was done because it would mostly add complications to the dynamics, and the group
wanted to focus on a constant cooling requirement.
The produced water which would be stored in the wings was used for cooling for the heat management as well. As the water likely will have cooled down to a temperature lower than when it was
produced by the fuel cell, it was assumed to drop to about 60◦ C = 333.15K. Assuming that this
water goes through a heat exchanger with the fuel cell, Equation (9) was used for the simulations.
The ideal operating temperature of a PEM cell is around 80◦ C as mentioned in section 2.1, so the
next step was finding the required cooling u(t) to keep this temperature. Combining Equation (8)
and (9) with Equation (7) and adding the cooling input results in the final heat dynamics:
Q̇f c (t) = Q̇w + Q̇LH2 + Q̇H2 O + u(t)
(12)
The system was modelled and simulated using Matlab and Simulink. u(t) was determined by the
means of a PID-controller, which for this study was more of a mathematical tool to help find the
cooling required to keep a steady temperature; Exactly how this controller works is considered
beyond the scope of this study. The implementation of the model and simulator can be found in
Appendix A and B.
3.5
Propulsion
To calculate the number of engines needed, the nominal power and the peak power presented in
appendix F was used for cruising and take-off respectively. Then comparing that to the force
exerted by the two IAE V2500-A5 engines on the original Airbus A320, the number of engines
needed was found.
3.6
Cutting of cargo space
To fit all the hydrogen components a lot of space was needed, especially for the fuel cells. In the
Airbus A320 the main kerosene fuel tank is located between the wings of the aircraft. This space
will remain unused in the retrofit, and thus will be removed. Further, some cutting of the cargo
space was deemed necessary, which was done with the help of the Airbus A320 maintinance manual
[31], and cargo space details [3].
6
3.7
Cost analysis
This section presents an economic feasibility assessment of retrofitting an Airbus A320 for hydrogen
propulsion. The focus is on comparing cost differences with conventional jet fuel-based operations.
The analysis explores whether a hydrogen-powered retrofit could become economically viable by
the year 2050.
3.7.1
Cost Structure
The methodology for determining the cost distribution is based on the approach presented by Steer
[32]. In this analysis, the operating cost of a one-hour flight is broken down into specific categories
relevant to airline operations, including fuel, maintenance, crew, airport and handling charges,
navigation, depreciation, marketing, and administrative costs. Each category has been assessed
for its expected cost impact in a hydrogen retrofit scenario, based on industry evaluations.
Table 1: Operating cost breakdown per one-hour flight
Cost Category
Share of Costs
Operations Impact
Cost Impact
Fuel
24%
Requirement to use new fuel type
Significant increase
Maintenance
5%
Requirement to maintain new aircraft
type
No material impact
Crew
14%
Requirement for pilots to train on new
aircraft type; none for cabin crew
No material impact
Airport & Handling
Charges
31%
Requirement for new airport and
groundhandling infrastructure (which
airports will increase charges to fund)
Increase
Navigation Charges
7%
None
None
Depreciation
8%
Requirement to purchase new type of
aircraft
Increase
Marketing
3%
None
None
Administrative Costs
8%
None
None
3.7.2
Key Assumptions
The baseline cost for a conventional one-hour flight between Oslo and Trondheim is estimated at
A
C10,000. This figure is based on reported average operating costs for Airbus A320 family aircraft,
which are commonly used on short-haul routes. According to Simple Flying [33], the typical hourly
operating cost of an A320 falls within this range under standard commercial conditions.
4
Results
Here the necessary calculated parameters and specific components used in the retrofit will be
presented. Further this section goes into detail of their weight and where in the plane it is placed.
4.1
Amount of Hydrogen
Using the method described in section 3.1, the Amount of LH2 needed for each stage of flight was
found, and is presented in Table (2) (calculations in appendix C and D).
7
Table 2: Amount og LH2 required for each stage of flight
Phase
Takeoff
Cruising
Landing
Total
4.2
Energy required
18865.0 MJ
32641.9 MJ
6151.6 MJ
57658.5 MJ
Amount of LH2
157.21 kg
272.02 kg
51.26 kg
480.49 kg
LH2 Tank
The tank volume found as 9.88 m3 , while its weight is calculated to be 1018.4 kg. This results
was calculated with the assumption of 2% boil-off being acceptable in a one hour window (See
appendix H). The tank will be placed in the tail of the Airbus A320 the, as this space is mostly
unused [11] and would conceptually look like figure 3.
Figure 3: A visual representation of how the LH2 tank would fit in an Airbus A320 [2].
As there is a volume of 18.5 m3 in the tail cone of the aircraft (see appendix I), the tank has space
to fit in the volume and will follow the structure of the plane. Due to there being more space
than needed, these extra margins allow to fit the APU and other electronical units of the original
aircraft.
4.3
Fuel cells
The final Fuel cell used in the study is the SuperStack fuel cell from ZeroAvia [34].
Table 3: Calculated fuel cell power demand and number needed per relevant flight stage
Phase
Takeoff
Cruising
Power Demand
16.25 MW
7.25 MW
Number of fuel cells needed
108.3
85.3
As seen in table (3), the aircraft would require 16.24 MW during take-off and 7.25 MW during
cruising. The Power demand is much higher for the take-off phase, but the power output from the
fuel cells are much lower with increased altitude, therefore the demand is higher for the take-off
phase. This means that the retrofit of the plane would need 108.3 ≈ 109 fuel cells, then:
Table 4: Weight and size of the fuel cell array
Weight
Volume
Per SuperStack fuel cell
170 kg
619 mm · 670 mm · 414 mm (Dual Stack)
8
Total
18530 kg
9.36 m3
4.3.1
Fuel cell array placement
The fuel cell array is placed in between the wings and cargo compartments in the lower fuselage
of the plane, as seen in Figure (4).
Figure 4: Details of lower fuselage of an Airbus A320 [3]
Table (5) is based on the calculations in appendix J. 12.64 m3 is more than enough space for the
fuel cells and there is a possibility of cutting into the existing cargo space if needed. Utilizing this
space ensures that the fuel cells can be installed without impacting passenger capacity and is an
ideal location in terms of weight distribution.
Table 5: Available space in lower fuselage of the aircraft
Total Available Cargo Space
Space Between Compartment 1 & 2
4.4
50.05 m3
12.64 m3
Propulsion
The retrofit uses the Hypercore Motordrive from ZeroAvia [35].
Table 6: Number of HyperCore Engines
Calculated with Nominal Power (750 kW)
Calculated with Peak power (900 kW)
21.65 ≈ 22
18.04 ≈ 19
As seen in table (6), the system requires approximately 18.04 engines to meet the power demand
of 16.24 MW. Since partial engines are not feasible and symmetry is essential, a minimum of 20
engines is placed on the retrofitted plane. Thus the weight of the propulsion system:
Table 7: Total weight of motors
18 Motors
19 Motors
20 Motors
1656 kg
1748 kg
1840 kg
9
4.5
Water Vapour
4.5.1
Water Volume Estimation
The volume V of the 6km - 8km atmospheric shell is calculated as the difference between two
spheres:
V =
4
π(80003 − 60003 ) ≈ 1.24 × 1012 m3
3
At an altitude of approximately 7,000 meters, the typical water vapour concentration ranges from
3
3
0.01 to 1.0 g/m [36]. For conservative estimation, a mid-range value of 0.5 g/m was used. Thus,
the total mass of water vapor currently present in this layer is:
mwater = 1.24 × 1012 × 0.0005 = 6.2 × 108 kg
Next, the potential global water emissions from a fully hydrogen-powered commercial aircraft fleet
were considered. Assuming approximately 50 million flight hours annually [37], and using the
estimated hydrogen consumption rate of 480.5 kg per flight hour (as calculated in Section 4.1),
the total annual water production was derived based on the stoichiometric relationship between
hydrogen and water is:
Water mass = 480.5 × 8.936 ≈ 4294 kg
This value is used to estimate the hourly water emissions from hydrogen-powered aircraft.
Using an emission factor of 4294 kg of water per kg of hydrogen consumed, the total annual water
emissions are:
Water emitted = (50 × 106 hours) × (4294 kg water/hour) = 214.7 × 109 kg
The global average residence time of atmospheric water vapour is approximately 9 days [38]. In
this study, the same value is conservatively assumed for water vapour at high altitudes, despite
evidence indicating that residence times at these elevations may be significantly longer [39].
Therefore the extra amount of water vapour in the atmosphere is given by:
9
9
= 5.29 × 109 kg
Extra water = 214.7 × 10 ×
365
4.5.2
Water stored in wing fuel tanks
By using a thin layer of Closed-Cell Spray Polyurethane foam [40], the fuel water inside the fuel
tanks will be less prone to freezing. Furthermore, by using a Ethylene Chlorotrifluoroethylene
lining on the inside of the wing fuel tanks, it can be ensured that tank integrity will be held [41].
Further, as seen from the previous section, approximately 4 tonns of water is produced throughout
the Oslo Trondheim flight.
4.6
Thermal management
Figure 5 presents the simulation result of the thermal management system. The plot on the left
illustrates the temperature evolution of the fuel cell over time, while the plot on the right shows the
excess heat generated. This excess heat represents the amount of cooling required, which in this
10
case is being abstracted as an input from a controller. The fluctuations come from the controller
adjusting its output and is not the focus of this result.
Figure 5: Result from simulating thermal control
As shown in Figure 5, when neither the produced water nor the liquid hydrogen is used for cooling,
approximately 5 MW of heat must be removed from the fuel cell. Utilizing the waste water that
has cooled down a bit reduces this to around 4.8 MW. When the excess heat is used to warm the
liquid hydrogen to operating temperatures, required cooling can be reduced to approximately 4.5
MW. Applying both techniques at once reduces this further down to about 4.3 MW. These results
demonstrate that preheating the liquid hydrogen not only addresses the challenge of bringing it to
operating temperature, but also significantly aids in cooling the fuel cell. While less effective on
its own, using the cooled water still provides a meaningful contribution to the system’s thermal
management.
4.7
Cost analysis
4.7.1
Retrofit Cost Estimation
Public data on full hydrogen conversions is limited, so the estimate is based on assumptions
supported by relevant benchmarks. A new CFM LEAP-1A or PW1100G engine, typically used
on the A320neo, is priced at approximately US$12 to 13 million [42]. A hydrogen retrofit would
require alternative engines, such as those developed by ZeroAvia [43], although no official pricing
has been released for these systems.
A complete cabin refurbishment for a narrow-body aircraft costs around A
C10 million, based on
figures from Air India’s recent fleet upgrade (US$400 million for 40 aircraft) [44].
In addition to replacing the engines, the hydrogen conversion would require major changes to the
aircraft’s infrastructure. This includes installing cryogenic hydrogen tanks, reinforcing the fuselage
and wings to handle new loads and accommodate liquid water storage from fuel cell systems, and
adapting systems for ventilation, insulation, and safety. Certification-related modifications would
also be necessary. Considering all these factors, the estimate of a total retrofit cost is approximately
€40 million for converting a mid-life A320 to liquid hydrogen propulsion. As supplier quotes are
not yet publicly available, this figure is based on assumptions. A sensitivity analysis later in the
paper evaluates how different retrofit costs could affect the overall business case.
Depreciation is calculated based on a mid-life Airbus A320 with an estimated value of A
C30 million,
plus the retrofit cost. Assuming the aircraft has 35,000 flight hours remaining, this results in a
depreciation cost of A
C2,000 per flight hour. This estimate is included in the overall cost analysis.
11
4.7.2
Comparative Context and Economic Consideration
Developing a completely new short- to medium-haul hydrogen aircraft is estimated to cost around
A
C15 billion, including design, testing, and certification [32]. Compared to these high initial development costs, retrofitting existing aircraft appears more economically favorable, especially for
those already in operation with substantial remaining operational life. Economic viability may
improve over time due to evolving carbon taxation policies and other governmental environmental
measures. Increasing taxes on conventional jet fuel could progressively improve the competitiveness
of hydrogen-powered aviation, potentially offsetting the higher upfront costs.
4.7.3
Airport Infrastructure Costs
The establishment of liquid hydrogen (LH2) infrastructure in airports is expected to significantly
impact airport handling costs. Large airports adopting hydrogen could require significant infrastructure investments, such as liquefaction and storage facilities, and potentially pipeline systems.
The costs of hydrogen refueling infrastructure could initially be significantly higher than those
of conventional systems, being approximately five times more expensive than conventional hydrant systems [45]. Thus, a 20% increase in airport and handling fees is considered a conservative
estimate.
4.7.4
Maintenance
Altough the original table showed no increase in maintenance cost, a conservative 10% increase is
applied due to the uncertainties associated with the use of new technology in the retrofit
4.7.5
Liquid Hydrogen Cost
According to a study by Steer [32], the unit cost of hydrogen fuel, including production, distribution, and liquefaction, is projected to decline from A
C3.90 per kilogram in 2035 to A
C3.45 per
kilogram by 2050. This 10% reduction is primarily driven by improvements in production efficiency and a shift towards greater use of pipeline-based gaseous hydrogen distribution, which is
more cost-effective than decentralized liquefaction. As a result, the future cost competitiveness of
hydrogen aviation will be closely linked to infrastructure development and technological advancements across the hydrogen value chain. The total fuel cost is calculated by multiplying the unit
price by 480, which corresponds to the hourly hydrogen consumption identified in Section 4.1.
4.7.6
Cost comparison and impact
A one-hour flight with a hydrogen-retrofit A320 is estimated to cost about 11 342 A
C in 2035—around
13% more than the 10 000 A
C needed for a standard jet. By 2050, cheaper hydrogen fuel could lower
this to roughly 11 126 A
C, reducing the extra cost to about 11%.
4.7.7
Cost per passenger
If the hydrogen tanks end up taking cabin space, some seats may need to be removed. To keep the
estimates conservative, the seating-capacity is reduced from 180 to 160. With fewer passengers to
share the operating costs, the unit cost rises from 55.56 A
C on a conventional A320 to 70.89 A
C in
2035 and 69.54 A
C in 2050—roughly 28 % and 25 % higher.
12
Table 8: Cost breakdown for a one-hour flight
Category
Conventional
H2 RF 2035
H2 RF 2050
Fuel (1-h flight)
Maintenance
Crew
Airport & handling
Navigation charges
Depreciation
Marketing
Admin
2,400.00
500.00
1,400.00
3,100.00
700.00
800.00
300.00
800.00
1,872.00
550.00
1,400.00
3,720.00
700.00
2,000.00
300.00
800.00
1,656.00
550.00
1,400.00
3,720.00
700.00
2,000.00
300.00
800.00
Total cost
10,000.00
11,342.00
11,126.00
180
55.56
160
70.89
160
69.54
Seats
Cost per passenger [€]
4.7.8
Sensitivity analysis
This analysis examines how three factors change the passenger cost: (1) the price of liquid hydrogen,
(2) the retrofit cost, and (3) how many seats remain in the cabin.
Base case. With hydrogen at 3.9 A
C/kg, a 40 millionA
C retrofit and 160 seats, the cost is 70.89 A
C
per passenger.
Best scenario. Cheap hydrogen (2 A
C/kg), a lower retrofit bill (25 millionA
C) and keeping all
180 seats cut the cost to 55.56 A
C—about 22 % below the base case.
Worst scenario. Expensive hydrogen (5 A
C/kg), a pricey retrofit (80 millionA
C) and only 150 seats
push the cost up to 86.75 A
C, roughly 22 % higher.
If just one factor changes:
• Hydrogen at 2 A
C/kg: cost falls 8 %. Hydrogen at 5 A
C/kg: cost rises 5 %.
• Retrofit between 25 and 80 millionA
C: cost moves from –4 % to +10 %.
• Seats from 180 to 150: cost swings –11 % to +7 %.
The analysis shows that passenger cost can vary by up to 22 %, with seat count having the largest
impact, followed by retrofit cost and hydrogen price.
4.7.9
Conclusion
Our cost-analysis shows that a hydrogen-retrofit A320 is still more expensive to operate than a
kerosene jet in the near term. Yet by 2050, helped by cheaper green hydrogen, better ground
infrastructure, and carbon taxing, the cost gap could narrow or even disappear. In fact, a recent
briefing argues that with full EU carbon taxes, hydrogen aircraft could already beat fossil jets at
cost as early as 2035 [46]. Reaching that point will require sustained progress on three fronts:
Lower hydrogen fuel prices, cheaper retrofits through scale and standardization, and operational
efficiencies will be essential.
4.8
Final weight of retrofit
In table (9), the final weight of the retrofitted aircraft is compared to two Airbus A320. Operating
Empty Weight (OEW) is the weight of a plane that has everything except passengers, cargo
13
and usable fuel on board. While Maximum Take-Off Weight (MTOW) represents the maximum
certified weight at which the aircraft is allowed to take off [31]. The retrofit weight includes the 4
tonns of produced water as well.
Table 9: Weight comparison of Airbus A320s
Airbus A320 (OEW + fuel, short haul)
Final LH2 retrofit (OEW + fuel + water)
Airbus A320 (MTOW)
∼ 53 000 kg
∼ 67 500 kg
∼ 78 000 kg
The final retrofit weighs more than and Airbus A320 Taking the same trip, but is well below
MTOW of the Airbus A320.
5
Discussion
The retrofit of an Airbus A320 to a hydrogen-powered configuration presented a wide spectrum
of technical, operational, and environmental challenges, each with its own set of complexities and
interdependencies. Therefore a lot of assumptions throughout the report have been made (but has
been pointed out when applicable). The nature of this report is thus imprecise.
It was not explored, but the necessary change in airport infrastructure to support liquid hydrogen,
which differs significantly in handling, will be a barrier that makes such a retrofit more unfeasible.
5.1
Calculations
As seen in Table (9), the retrofit weighs more than the airbus A320 normally would on this route.
This is not a problem for our calculations however, as they have been based on flight data of
aircraft weights heavier than of the retrofit. Thus the calculations are realistic still.
A significant part of the project time focused on performing calculations to estimate the energy
requirements and fuel consumption for a one-hour hydrogen-powered flight. The more accurate
way of doing the flight calculations, would instead consist of a detailed aerodynamic lift and drag
analyses. But due to the lack of exact post-retrofit aircraft specifications, a simplified but practical
approach was adopted.
Since the data on efficiency of existing Fuel cell was very sparse, an assumption was made to
use an efficiency between 40% and 60% [9]. This assumption is not an ideal approach since the
efficiency can vary a lot based on the Thermal management of the Fuel cell and the Fuel cell model
itself. Further, the flight trajectory was simplified for modelling purposes. During take-off, aircraft
typically exhibit variable climb rates; however, our calculations assume a constant rate of climb,
resulting in a linear ascent profile2.
5.2
Fuel Tank
The decision to allow a 2% per hour boil-off rate for the LH2 tank is rooted in operational practicality. This rate provides a sufficient buffer in scenarios where the aircraft experiences delays
before takeoff, such as queueing on the runway. In such cases, the contingency fuel stored aboard
ensures that the plane can remain grounded for several hours without compromising flight safety
or requiring additional fueling.
The size of the LH2 tank, however, is constrained by the aircraft’s internal space limitations, as the
retrofitted A320 does not offer abundant space for large cryogenic tanks. As a result, the design
prioritizes minimal tank volume—just enough to cover the energy demands of a single short-haul
flight plus reserves.
14
This reflects a deliberate trade-off: to make the retrofit concept as technically and spatially feasible
as possible, the tank is kept small, and the fuel load tightly optimized. Consequently, the aircraft
would require refueling after every flight. While this may seem restrictive, it aligns with typical turnaround procedures, during which boarding and deplaning occur—processes that provide
enough time for rapid LH2 refueling.
5.3
Fuel Cells
Operationally, the retrofit raises important questions about the impact on flight dynamics, range,
maintenance, and ground infrastructure. For example, the increased weight and volume of fuel cell
systems and hydrogen storage can affect cargo capacity.
The selection of the SuperStack PEM fuel cells from ZeroAvia was guided by their compact size and
relevance to aviation. There could exist better fuel cells that could be more suited for the retrofit,
but such fuel cells were not found. The market for fuel cells is relatively small, but hopefully, this
will change in the future.
Nevertheless, our calculations show a high number of fuel cells are needed: approximately 109. The
number of fuel cells required for the retrofit is a significant concern, primarily due to their combined
weight. While the total volume of the fuel cells is relatively manageable within the available space
of the aircraft, their placement is a critical factor. Therefore they have to be installed equally on
each side, if the cargo space is to be used. This is not unrealistic, as for trips like Oslo-Trondheim,
a lot of passengers don’t bring baggage [47, 48]. But even with this, the weight is still too much
for the floor of the lower fuselage to handle. It can only support 600 kg/m2 [3]. Then the lower
fuselage would need structural support, which would increase weight. This is a huge blow to the
feasibility of the project. Better fuel cells would have to be developed until something like this can
be done, or:
An option to decrease the number of fuel cells was to use a battery hydrogen hybrid to Power
the take-off phase of the flight. This option can not only decrease the weight but also make the
required amount of LH2 for the trip smaller. Unfortunately, the group was not able to make the
hydrogen/battery optimizer programme to work correctly, so a battery weight was not found.
5.4
Water storage
If all aircraft were converted to hydrogen planes, the amount of water vapour in the upper atmosphere could increase by approximately 9 times, significantly impacting radiative forcing.
Water storage onboard is therefore critical to avoid offsetting the environmental benefits of hydrogen aviation. Some of the captured water could also be reused onboard for non-potable applications,
such as toilet systems and perhaps introducing free, safe to drink water on board in the future,
reducing the need for additional water supplies.
As water is captured during flight, the aircraft becomes progressively heavier, impacting overall
performance and fuel efficiency. But this does not exceed the weight of the fuel that used to be
there, and as stated earlier, thus does not impact our calculations.
5.5
Thermal management
The method of thermal management that has been described will help with the problem that is
keeping the fuel cell array cool, while also being possible to heating the LH2 before usage. However,
how one would resolve the problem of the rest of the required cooling was not in the scope of this
report, and shall remain as a problem for future work.
15
6
Conclusion
This project investigated the feasibility of retrofitting an Airbus A320 for short-haul hydrogenpowered flight by replacing the conventional kerosene-based propulsion system with a hydrogen fuel
cell array, a cryogenic LH2 storage tank, electric motors, and an integrated thermal management
system.
The technical and economic analyses highlight both the promise and the significant challenges of
such a transition. While the zero-emission potential of hydrogen fuel cells is compelling, critical
barriers—including the weight of the fuel cell system and elevated operational costs—currently
limit feasibility.
Nevertheless, ongoing advancements in fuel cell technology, the development of supporting infrastructure, declining LH2 fuel prices, and the introduction of climate-focused regulations such
as carbon taxation, could collectively enable hydrogen-powered short-haul aviation to become a
practical and sustainable alternative in the coming decades.
16
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19
Appendix
A
Matlab code for thermal management
eta = 0.4; % efficiency fuel cell at high elevation
eta_w = 0.6; % efficiency fuel cell in general
r = 80; % temperature in celcius we want to keep the heat sink of
the fuel cell at
T_lh2 = 20; % temperature LH2 in kelvin
T_h20 = 40+273; % temperature water in kelvin
H_h2 = 446; %
C_h2 = 8.1; %
C_h2o = 4.18;
C_cu = 0.385;
kJ / kg heat to change phase of hydrogen
kJ /( kg * K ) specific heat capacity of hydrogen
% kJ /( kg * K ) specific heat capacity of water
% kJ /( kg * K ) specific heat capacity of copper
m_cu = 1*106; % kg total mass of copper heat sink
C_cu_tot = ( C_cu * m_cu ) ; % kJ / K copper heat sink heat capacity
P_in = 8300/ eta ; % kW power supplied to fuel cell to power the
plane
lhv_h2 = 120000; % kWs / kg lower burn energy of hydrogen
% kg /( W * s ) kg hydrogen consumed per second to produce one watt
flow_h2 = P_in / lhv_h2 ; % kg / s
flow_water = 9* flow_h2 ; % kg / s
Q_waste = (1 - eta_w ) * P_in ; % kW waste heat
Q_lh2 = ( T_lh2 * C_h2 - H_h2 ) * flow_h2 ; % kW cooling from lh2
Q_h2o = T_h20 * C_h2o * flow_water ; % kW heat from water produced
A = -( C_h2 * flow_h2 + C_h2o * flow_water ) / C_cu_tot ; % rate of internal
system dynamics
B = C_cu_tot ; % conversion factor from Kelvin to kW
C = 1/ C_cu_tot ; % conversion factor from kJ to Kelvin
w = ( Q_waste + Q_lh2 + Q_h2o ) ; % heat from fuel cell , water and
liquid hydrogen
init = ( r +273) / C ; % initial value close to reference
B
Simulink diagram for thermal management
This is the Simulink diagram for the thermal management of the aircraft
20
C
Amount of LH2 and fuel cells required for the trip calculations
A
1
2 Airbus A320 :
3
Enignes:2 X IAE V2500-A5
4
5 IAE_V2500
6 useful links
7 ZA2000
8 Engine used in the retrofit ^
9
10
11
12 Fuel Cell:
Zero Avia SuperStack Flex Fuel Cell Power
13 System
14
15 Power output at sea level
16 power at 14000 feet
17 Power output at 25000 feet
18 Weight
19 Dimensions
20
21 Fuel Cells required at thrust
22 Fuel Cells required at cruising
23
24 Total Weight
25 Total Dimensions
26
27 Takeoff:(sheet 2 for calculations)
28 Energi inn fra takoff til 25000 feet
29 KG LH2 takeoff
30
31 Cruising
32 Effekt of Fuel cell
33 MWh på cruising
34 Brennverdi av Hydrogen
35 MWh/effekt
36 Energi inn
37 KG LH2 i cruising 0.5timer
38
39 Landing:
40 Energi inn fra landing
41 KG LH2 landing
42 SUM of kg LH2
B
C
D
E
F
G
H
Thrust
thrust per engine
estimated thust
estimated speed
110.31 kN
220.62 kN
265 km/h
73.61 m/s
16240083.3 W
16.24 MW
16240.08333 kW
P(watt)
150 kW
115 kW
85 kW
170 kg
0.085849 m^3
108.267
85.33824
18530 kg
9.36 m^3
18865.02 MJ
157.2085 kg
150000 W
115000 W
85000 W
109 =E9/D15
86 =I9/D17
=B18*C21
=B19*C21
=Datasheet!F17*3.16*1000
=B28/B34
0.4
40%
3.626875 MWh
=I10*0.5
120 MJ/kg
9.067188 MWh
=B33/B32
32641.88 MJ
=B35*3.6*1000
272.0156 kg
=B36/B34
6151.62 MJ
51.2635 kg
480.4877
=Datasheet!N17*3.6*1000
=B40/B34
=B29+B37+B41
21
I
J
Cruising
thrust per engine
estimated thust
estimated speed
P(watt)
15.75 kN
31.5 kN
829 km/h
230.28 m/s
7253750.000 W
7.254 MW
K
D
Take off and landing energy used calculation datasheet
A
1 takeoff:
2
3
4
5
6
7
8
9
10
11
12
13
14
15
16
17
B
Minutes
1
2
3
4
5
6
7
8
9
10
11
12
13
14
15
C
Effekt
D
E
F
MWh
MWh/effekt
16 0.266667 0.44444445
15.429 0.25715 0.42858333
14.858 0.247633 0.41272222
14.287 0.238117 0.39686111
13.716
0.2286
0.381
13.145 0.219083 0.43816667
12.574 0.209567 0.41913333
12.003 0.20005
0.4001
11.432 0.190533 0.38106667
10.861 0.181017 0.36203333
10.29
0.1715
0.42875
9.719 0.161983 0.40495833
9.148 0.152467 0.38116667
8.577 0.14295
0.357375
8.006 0.133433 0.33358333
sum=
5.96994446
G
H
MW
0.6
0.6
0.6
0.6
0.6
0.5
0.5
0.5
0.5
0.5
0.4
0.4
0.4
0.4
0.4
0.016666667
22
I
Landing:
J
Minutes
1
2
3
4
5
6
7
8
9
10
11
12
13
14
15
K
Effekt
L
M
N
MW
MWh
MWh/effekt
0.4
3.3252 0.05542 0.13855
0.4
3.3252 0.05542 0.13855
0.4
3.3252 0.05542 0.13855
0.4
3.3252 0.05542 0.13855
0.4
3.3252 0.05542 0.13855
0.5
3.3252 0.05542 0.11084
0.5
3.3252 0.05542 0.11084
0.5
3.3252 0.05542 0.11084
0.5
3.3252 0.05542 0.11084
0.5
3.3252 0.05542 0.11084
0.6
3.3252 0.05542 0.092367
0.6
3.3252 0.05542 0.092367
0.6
3.3252 0.05542 0.092367
0.6
3.3252 0.05542 0.092367
0.6
3.3252 0.05542 0.092367
sum=
1.708783
O
E
SuperStack-Flex-Datasheet
SuperStack Flex Fuel Cell Power System
AVAILABLE NOW
The ZeroAvia SuperStack Flex is a customizable fuel cell power generation
platform with 150 kW net total continuous power and lightweight, optimized
balance-of-plant, developed as a best-fit approach for Experimental and
Unmanned Air Vehicles (UAV) with further potential applicability in marine
and ground vehicle use cases.
The systems have been designed for maximum flexibility allowing for
rearchitecting to meet a range of different power, performance, and
packaging requirements, while maintaining exceptional pressure, mass flow
rate, humidity, and temperature control.
The baseline SuperStack Flex configuration is comprised of 3 modules, that
can be installed in physically separate locations on an airframe.
Single 275 Cell to Dual 455 Cell Stack options allow for various Voltages
(200 – 600 Vdc Nominal), Powers (75 – 250 kW), and Fault Tolerant
configurations.
F E AT U R E S
PHYSICAL
• Peak Gross Power 175 kW @ Sea Level ISA
• Weight: 170 kg
Standard
• Dimension: 619 mm x 670 mm x 414 mm
• Continuous Net Power:
(Dual Stack)
• 150 kW @ Sea Level ISA Standard
• Connectors HV: Configurable
• 115 kW @ 14,000 feet ISA Standard
• Connectors LVDC: Configurable
• 85 kW @ 25,000 feet ISA Standard
• Mounting: + / - 9 G Horizontal,
• Specific Power 1 kW/kg
+ / - 3.75 G Lateral, + 9 / - 4.5 G Vertical
• Low Voltage Control Power +28 Vdc per
E N V I R O N M E N TA L
MIL-STD-1275
• Altitude: up to 25,000 feet
• CAN bus communication for firmware
upgrades, configuration programming and
(compressor limited)
status reporting
• Temperature: -30°C – 50°C
• Coolant: FC EG and Water-Glycol
• Humidity: 0 – 100%
E L E C T R I C A L I N T E R FA C E S — I / O
Signal Connector: D38999/26WF32SN
HX control: PWM output / 5 v input
E-Stop: Digital
P/N
HVDC
LVDC
HVAC
CURRENT
OUTPUT
POWER
SERVICE LIFE
COMMS
ZA 150-PGS
400 Vdc (Max Load)
18 – 35 Vdc
400 Amps
150 kW
>5000 Hours
CAN bus
Continuous
Continuous
470 Vdc (Nominal)
800 Vdc (Min load)
23
F
Motor-HyperCore-Datasheet
900 kW 4x3 Phase Permanent Magnet
Synchronous Machine 600–820 Vdc
HyperCore is a leading-edge technology, with a high speed, high performance
integrated motor drive. The fully Integrated Motor-Inverter system is designed
to be powerful, compact, and ultra-efficient. The motor is a permanent magnet
electric machine combined with 4 Silicon Carbide inverters. The full integrated
edrive enhances an optimized integration of aircraft, engine compartment.
Designed and produced for high altitude and unpressurized areas, optimized
thermal, mechanical, and electrical integration allows the HyperCore to
achieve high reliability and long life time by reducing cable weight, interfaces,
and coupling, connectors. The inverters are fully equipped for high-speed bus
communication and a high degree of redundancy and availability.
F E AT U R E S
• Comply with SC part 33 – FAA, and E19 EASA
• Continuous Power 750 kW @ 75°C Coolant continuous
PHYSICAL
• Max Power 900 kW for 5 min
• Weight: 93 kg
• DC voltage: 600–820 Vdc
• Dimension: 600 mm x 600 mm x 253 mm
• Continuous torque: 360 Nm
• Connectors HV: AMPH Voltarius
• Max Torque: 430 Nm for 5 min
• Connectors LVDC: MIL-DTL-38999
• Power Density full integrated motor with inverter: 10 kW/kg
• 4x3 phase coil winding pattern for added redundancy
• 4 resolvers to give independent speed and direction to each inverter
E N V I R O N M E N TA L
• MTBF per MIL-HDBK-217F: 25,000 hr.
• Operation temperature: -45°C to 100°C
• Compliant: DO160-G
• Coolant temperature: -45°C to 75°C
• 12 temperature sensors for reliable monitoring of condition
• Altitude: up to 35,000 feet (unpressurised area)
• Efficiency: 94–96% vs load and speed
• Temperature and Altitude: DO160-G CAT B2
• Low torque ripple below 2%
• Temperature variation: DO160-G CAT A
• Max speed: 24,000 RPM
• Vibration: DO160-G – S curve U
• Coolant: Turbine oil and Water-Glycol (performances upon request)
• Water proofness: DO160-G – CAT S
• Fault-tolerant and fail-safe following MIL-PRF-GCS600A
• Fire flammability: DO160-G – CAT B
• CAN bus communication for firmware upgrades, configuration
• Emission RF: DO160-G – CAT LMH
programming and status reporting
• Analogue and digitally controlled torque, speed and DC power
• Lightening: DO160-G – CAT A3K3L3
(Upon request for higher CAT~B4)
• Dual redundant CAN interfaces
• ESD: DO160-G – CAT A
E L E C T R I C A L I N T E R FA C E S — I / O
Resolver, Encoder and Encoder SIN/COS
8 DI/DO/Analog inputs
2 Throttle inputs
8 PT1000
P/N
HVDC
LVDC
HVAC
CURRENT
OUTPUT
POWER
FUNDAMENTAL
FREQUENCY
LV CONNECTOR
ZA 900-IMI
600–820 Vdc
18–35 Vdc
300 Arms
750 kW
Up to 2 kHz
Ethernet Bus
1000 Vdc Unloaded
Transient 50 Vdc
330 Arms for 1 min
continuous
RS485
RS232
2 CAN buses
24
G
Final volume and amount of LH2 needed calculations
25
H
Tank volume and weight calculation
26
I
Tail cone space calculation
J
Motor and cargo space calculations
A
1
2 Motor For electric Drivetrain:
3 HyperCore Motor
4
5 MW Required
6 weight #1 motor
7 # Motors Required 900kW =
8 # Motors Required 750kW =
9 Weight with 19 Motors
10 Weight with 22 Motors
11 Weight with 20 Motors
12
13
14 Cargo space
15 cargo 1 from nose
16 cargo 2 from nose
17
18 Space Between cargo 1 -2
19
B
750 kW
0.9 MW
16.24
92 Kg
18.04454
21.65344
1748 kg
2024 kg
C
D
E
F
G
19 approx
22 approx
1840 kg
7.25 m
18.426 m
=8.16-(1.82/2)
=22.69-(1.82/2)-3.354
6.326 m
=B16-B15-4.85
20 cargo volume 2
21 lenght cargo 2
22 face area cargo 2
23
13.1 m^3
6.554
1.998779
24 Extra volume
25 Existing cargo volume
26 total cargo volume
27 total cargo lenght
12.64428 m^3
37.4 m^3
50.04428 m^3
25.03742 m
=B20/B21
There is extra compartment
space between cargo 1 and 2.
This can be used as space for the
fuel cells. Since
=B22*B18
=B25+B24
=B26/B22
27