Polaris – Design of Liquid Hydrogen Turbo-Electric …

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POLARIS – DESIGN OF A LIQUID HYDROGEN TURBO-ELECTRIC
TRANSPORT AIRCRAFT

T. Dietl, J. Karger, K. Kaupe, A. Pfemeter, P. Weber, A. Zakrzewski, A. Strohmayer
Institute of Aircraft Design, Pfaffenwaldring 31, 70569 Stuttgart, Germany

Abstract
Polaris is a liquid hydrogen turbo-electric transport aircraft which is designed to fulfill the NASA N+3 goals for
entry into service in 2045. The main objective is to reduce the energy consumption by at least 60% compared
to an A320 of 2005. Secondary objectives are the reduction of the NOX emissions by 80% and minimizing the
noise emissions.
The concept combines the synergies of a liquid hydrogen fuel system with a HTS power transmission and an
intercooled gas turbine. By this a reduction of 62% in energy consumption and 80% lower NOx emissions are
achieved at a comparable design mission.

Aircraft; Liquid Hydrogen; Turbo-electric; Hybrid-electric, Superconducting; Open-rotor; Transport aircraft

Keywords

electric propulsion chain. The fuel tanks are stored in the
fuselage outside of the passenger cabin (FIG 3) using a
multifunctional fuselage concept. The cooling potential of
the LH2 is utilized to control the temperature in the
intercooler of the gas turbine and the HTS components. To
minimize noise emissions and drag the gas turbines and
generators are stored inside of the aft fuselage with minimal
power cable lengths. Air sucked in for the laminar flow
control at the leading edge of the wings is used for the air
blow out of the Coanda flap.

1. LIST OF ABBREVIATIONS

Alternating Current

AC
BSCCO Bismuth Strontium Calcium Copper Oxide
CFRP Carbon Fiber Reinforced Plastics
cL,max Maximum lift coefficient
CROR Contra-Rotating Open Rotor
Gaseous Hydrogen
GH2
High Pressure Compressor
HPC
Horizontal Tailplane
HTP
High Temperature Superconducting
HTS
Intercooled Recuperative Aeroengine
IRA
Laminar Flow Control
LFC
LH2
Liquid Hydrogen
OPR Overall Pressure Ratio
Standard Passenger Payload
SPP
TET
Turbine Entry Temperature
TSFC Thrust Specific Fuel Consumption
VTP
YBCO Yttrium Barium Copper Oxide

Vertical Tailplane

2. MOTIVATION

The NASA N+3 goals demand for aircraft with at least 60%
less energy consumption, 80% NOX and a significant noise
reduction. Furthermore, CO2 emissions can be avoided by
using hydrogen as energy storage.
To maximize the potential of the technologies a special
focus is laid on a synergetic combination and integration of
the used components.

3. DESIGN OVERVIEW

FIG 1. Overview of technologies with (1) cryogenic LH2 fuel
system, (2) Intercooled Recuperative Aeroengine (IRA) gas
turbines, (3) High Temperature Superconducting (HTS)
generators and electric motors and (4) wing with Laminar
Flow Control (LFC) and Coanda flaps

4. KEY TECHNOLOGIES

4.1. Propulsion Chain

Polaris focuses on the design of a future single-aisle
transport aircraft comparable to the best in class version of
an A320 in 2005. An analysis reveals that the largest
number of flights is conducted below a range of 1500 NM.
Therefore, the design point is set to a range of 1500 NM
with a payload of 13608 kg (150 passengers) [1].
The aircraft consists of a conventional tube and wing
configuration with forward swept wing, a U-tail and a turbo-

The propulsion chain consists of two IRA gas turbines each
coupled with a HTS generator powering its electric motors
to drive the contra-rotating open rotors (CROR).
The coupling between generator and electric motor acts as
electric transmission, which allows both the gas turbine and
the CRORs to run at their respective optimum speeds.
Electrical cross-wiring between the generators and the

©2018doi:10.25967/480344Deutscher Luft- und Raumfahrtkongress 2018DocumentID: 4803441 electric motors, as seen in FIG 2, enables all electric motors
to continue to operate in case of a generator or gas turbine
failure. To maintain the same speed ratio of electric motors
and gas turbine a variable-pitch propeller decreases the
power loading at the same speed to match the reduced
power provided by the remaining gas turbine.

conventional bypass architectures might not be sufficiently
provided. More synergies are found regarding the reduction
of bleed air temperature, therefore optimizing the cooling of
hot components and simultaneously enabling a reduction of
bleed air mass flow which raises core efficiency [4].

4.1.2. Superconducting Technology

and

coolant

both as

propellant

institutions have already

Cycle studies during the REVAP program proved the
necessity of a separation of propulsor and power
generation if IRA engine architecture shall be optimized [7]
– which is therefore realized in the Polaris concept.
Furthermore, a turbo-electric architecture enables an
independent positioning of propulsion chain components.
Incorporating conventional
technologies in the turbo-
electric propulsion chain architecture is not practical for the
Polaris concept, as power densities of electric motors and
generators are too low; but superconducting technology
becomes a key enabler for Polaris. Moreover, using liquid
hydrogen
for
superconducting wires, cooling can be realized without an
additional power demand as liquid hydrogen must be
evaporated before being burnt.
HTS technology exhibits high current densities at very low
resistance. Fully superconducting machine designs, using
HTS winding both on rotor and stator, show power densities
up to 40 kW/kg at rotational speeds of about 10000 rpm [7].
Several
realized partially
superconducting systems, thereunder General Electric’s
Homopolar Inductor Alternator with a power density of
[8]. Partially superconducting machines use
8 kW/kg
superconducting windings on the rotor where DC currents
induce a DC magnetic field, interacting with copper stator
windings which are excited with alternating current. Current
superconducting materials like BSCCO and YBCO show
AC losses which make their use as stator windings
impractical until today [9]. A lot of effort on research for low
AC loss HTS material is done by several research centers
and companies. According to the American Institute of
Physics, MgB2 with a critical temperature of 39 K and best
performance under 30 K, shows high potential to reduce AC
losses when arranged as fine, twisted filaments [9]. Liquid
hydrogen is on a temperature level well below the critical
temperature of MgB2 thus improving its current carrying
capacity [10].
Based on NASA’s technology roadmap, power densities of
HTS machines – including generators and motors – are
predicted to be as high as 33 kW/kg [11]. Further
calculations for the Polaris concept will use a more
conservative value of 20 kW/kg.

4.1.3. Contra-Rotating Open Rotor

Increasing efficiencies of the propulsor inhibits potential for
further improvement of the propulsion chain. Studies by
NASA, General Electrics and
the Federal Aviation
Administration have stated propulsive efficiencies of
96% [12] for open rotor concepts. In addition, an advantage
of open rotors is their compact integration.
Major concerns on the open rotor concept pertain to the
assumption of increased noise levels. However, studies as
from NASA [13] and field tests conducted by Safran in 2018
have proved a reduction of noise emissions compared to
enclosed engines. Moreover, CFD analysis [13] have
shown, that noise emissions of pusher configurations are
more uniform compared to tractor configurations. In order

FIG 2. Components of the propulsion chain

4.1.1. Gas Turbine

Present gas turbine cycles reach their limits when it comes
to an improvement of energy efficiency or thrust specific
fuel consumption (TSFC) along with a reduction of NOX
emissions. Designing a gas turbine at high load levels for
best core efficiencies causes high cycle temperatures.
Parametric optimization of a two-spool turboshaft in
GasTurb 13 shows, that high cycle temperatures require
high overall pressure ratios (OPR) to attain best core
efficiencies. An optimization of TSFC therefore pushes the
formation of NOX, as formation mechanisms show an
exponential dependency on cycle temperatures [2].
Regarding the 2045 time frame of Polaris, the IRA concept
shows the most promising cycle technology [3]. Intercooling
reduces the specific power demand of the high pressure
compressor (HPC), as the mass flow is cooled down
between compressor stages. The work needed by the HPC
to enhance OPR is decreased as the temperature at its
entry is falling [4]. Recuperation benefits from increasing
spread in temperature between exhaust mass flow and
compressor mass flow, thus enabling higher temperature
levels in the combustion chamber without manipulating the
fuel flow [4]. IRA cycles show the ability of higher core
efficiencies for an OPR of up to 40 [5].
For the performance calculation of Polaris the IRA cycle
performed by the TU Dresden as part of the REVAP
(REVolutionäre ArbeitsProzesse) program is chosen. The
cycle has a thermal efficiency of 50.8% at an OPR of 40
and a TET of 1590 K. Reaching equal thermal efficiencies
for a conventional Joule cycle, requires an OPR of 99 and
2000 K TET [6]. New combustion technology and the
reduction of OPR and TET are main drivers for low NOX
combustion [2].
Employing IRA into the Polaris concept yields some
additional advantages regarding intercooler technology.
Using LH2 as coolant exhibits high efficiencies of the
intercooler, allowing
to be minimized.
its surfaces
Intercooling during critical operating conditions, such as
take-off and climb, remains possible with a LH2 cooling
architecture, where otherwise the cooling air mass flow for

©2018Deutscher Luft- und Raumfahrtkongress 20182 to further reduce the noise emissions during flight, the
propellers are shielded by the U-Tail in both downward and
sideward direction. Flight Mach number has to be reduced
from Ma = 0.78 to Ma = 0.72 due to efficiency losses and
noise issues at higher Mach numbers which go along with
the rotor blade tip speed.

4.2. Multi-functional Fuselage Concept

A further reduction of the energy consumption can be
achieved by minimizing the structural weight of the aircraft
with a multi-functional aircraft structure. With carbon fiber
reinforced plastics (CFRP) an advanced technology is
applied to reduce the structural weight. The special feature
of the “Gondola Concept” as shown in FIG 3 is the further
partition of pressurized and unpressurized area. Polaris
uses this concept to minimize the structural weight and to
integrate the fuel tanks outside of the pressure cabin. The
present design is developed regarding weight, passenger
safety,
manufacturing
advancements.
According to the final report from the German Aerospace
Center (DLR) in 2003 a fuselage weight reduction of 28.7%
is achieved compared to an A320. This has been validated
with a demonstrator using
fulfill current
certification criteria [14].

worthiness

crash

tests

that

and

4.2.1.

Integral Shell Design

A significant advancement to current CFRP fuselages is the
use of a load-bearing skin. The structure is designed as
described in [14] to fulfill crash and fire resistance
requirements. The DLR analyzed and
tested a
demonstrator regarding:

Impact and residual strength behavior
3D-thermal-analysis
3D-tension and stress analysis
Stability against buckling
Fire safety
Crash safety
High velocity impact

•
•
•
•
•
•
•
• Manufacturing effort and costs

The integral shell design is developed to withstand crash
and impact. Using phenolic resin, a fire resistance is
achieved to block toxic vapors, smoke and a burn-through
for minimum 15 minutes at 1100 °C.

4.2.2. Gondola Concept

The upper part of the “Gondola Concept” includes the
pressure cabin to carry passengers and the crew. The cabin
layout of the A320 was chosen for Polaris. Unlike the A320
fuselage this pressure cabin does not include the cargo
area of the aircraft. With CFRP and a sandwich structure a
pressure cabin is designed without the need of a fully
circular shape. The advantage is a primary structure that
carries only the most necessary contents which require a
pressurized environment. Regarding a malfunction of the
fuel tanks, any security incident with respect to fire, smoke
or toxic vapors must take place outside of the passenger
area (FIG 3). To guarantee this requirement a placement of
the fuel tanks outside of the primary fuselage structure is
realized. LH2 fuel tanks, landing gear, wing, empennage
and propulsion system are attached to the primary structure
(FIG 3).

FIG 3. Gondola Concept with LH2 fuel tanks below and
outside the passenger cabin.

the secondary structure
structure”

The secondary fuselage structure is designed as a non-load
bearing and unpressurized area that houses the fuel tanks
is
and cargo. Furthermore,
regarding
“sacrificial
developed as a
crash/impact. While this area is not pressurized there are
less demands on the structural strength and a cargo door
is not needed to be as sealed as one in a pressure cabin. A
malfunction of the fuel tanks does not affect the passenger
area, as fire, smoke and toxic vapors cannot get into the
pressure cabin.

4.3. Aerodynamics

Polaris’ advanced aerodynamic
layout allows an
improvement of the glide ratio during cruise by 18% to
L/D = 20.2 . This improvement is mainly a result of a
minimization of turbulators on wings and fuselage while
focusing on long laminar airflows, which are achieved by
passive and active means. For this purpose the forward
swept wing, morphing wing technology and a special
surface finish [15], as well as a boundary layer control
system and Coanda flaps are installed.
The calculation of lift-dependent drag is performed in
XFLR5 and OpenVSP, whereas non-lift-dependent drag is
estimated using handbook methods. Addressing the non-
lift-dependent drag a special surface finish, creating a riblet
structure [15] can be used. This creates a similar turbulator
effect as in golf balls, reducing the drag of all components
in turbulent airflow by up to 8%.

4.3.1. Forward Swept Wing

Usage of a forward swept wing configuration allows to
reduce the wing’s aerodynamic sweep angle compared to
a conventional backward swept wing while maintaining the
same geometric sweep angle [16] .
The optimal sweep angle for Polaris is derived from
Krause [17] and Hepperle [18] as an exact estimation and
optimization of the sweep angle are out of scope of the
preliminary design of Polaris and have to be analyzed
separately. In combination with the selected airfoil a natural
laminar airflow on the lower surface of the wing is achieved.
As for the structural instabilities they are compensated by

©2018Deutscher Luft- und Raumfahrtkongress 20183 tailoring, which utilizes

an adequate aeroelastic
the
anisotropic twist-bending coupling of the carbon fiber layup
[19].
The airfoils of Polaris are of the same airfoil family as the
A320’s.

4.3.2. High Lift System

The requirement for a long laminar flow on the wing airfoil
in order to reduce drag during cruise prohibits the use of a
slat track, as this creates a gap in the airfoil and initiates the
laminar-turbulent transition. Nonetheless slats are required
for high lift coefficients, which are necessary for sufficiently
slow approach speeds. In order to prepare for this
challenge, multiple universities, especially the DLR in
Brunswick [20] research on the advantages of Coanda
flaps.
The Coanda flap uses a small air jet parallel to the airfoil in
its aft to control the boundary layer and sustain an attached
airflow over the surface and allow higher flap deflections
without flow separation. As Coanda flaps have the same
geometry as normal flaps, they can be both used for regular
flight without blowing air and as lift increasing devices
during the approach. The necessary air is partially
produced by the fans from the boundary layer control
suction system and additionally supplemented by two
redundant compressors that are located on the inboard
section of the wing and then distributed to the flaps.

FIG 4. Schematic figure of LFC (leading edge) and Coanda
flap (trailing edge) in approach configuration

to

results

coefficient

cL,max,LDG = 2.88

As the necessary lift coefficient for an approach is
cL,max = 2.8, a deflection of the flap by 20° is sufficient to
allow operation of the aircraft. Nonetheless, as the
technology of the Coanda flap is currently in development,
the maximum deflection angle is set to 40° with a maximum
theoretical lift coefficient cL,max = 3.2, a maximum angle of
attack of 8.5° with activated system and a technology
uncertainty factor of 0.9 used to reduce the estimated lift
coefficient. Using this uncertainty factor, the maximum local
lift
and
cL,max,T/O = 2.24 . Take-off and landing are possible with
inoperable Coanda system and higher angles of attack,
although the distances increase significantly. Due to the
absence of slats in the free stream, a decreased noise
emission can be assumed [21]. In addition, landing
approaches with Coanda flaps are performed at low angles
of attack of around 3° [20] increasing the visibility of the
pilots on the runway and increasing the overall safety.
The Coanda flap is designed as such, that during cruise,
when higher lift coefficients are limited by the onset of buffet
the air sucked in by the boundary layer control system on
the upper surface of the airfoil can be used to reduce the
induced drag of the airfoil. By ingesting an airflow into the
wake of the airfoil a reduction in the viscous dissipation is
achieved, reducing the induced airfoil drag by 1% [22].
According to Radespiel [20] in a functional Coanda flap
system the required mass flow equals to about 6 kg/(s*m)
adding up to a total mass flow of 206 kg/s in the Polaris

preliminary design. The necessary
approximately 5 kW [23].

fan power

is

4.3.3. Boundary Layer Control

To reduce disturbances and delay laminar turbulent
transition the boundary layer control system sucks in air
from the upper surface of the wing. Combined with the
forward swept wing and a low cambered airfoil laminar
airflow up to 40% of the airfoil are achieved that lead to a
drag reduction of 16% [24]. At this point the boundary layer
is no longer controlled and transition from laminar to
turbulent occurs. The energy necessary to drive the pumps
and propellers is diverted form the gas turbines in the rear
of the aircraft. Similar to the contra-rotating open rotors their
revolution speeds are tied to the gas turbine and use the
alternating current directly to avoid the disadvantage of
inverter losses.

4.3.4. Empennage

The U-tail is located directly below the gas turbines and has
a continuous transition from horizontal tailplane (HTP) to
vertical tailplane (VTP) to reduce interference drag in flight.
As shown in FIG 5 the wing wake does not have an
influence on the empennage or the propellers. The trailing
wakes at the transition from the HTP to the VTP shows that
the flow is straight and not disturbed.
In addition to the compensation of the moments generated
by the wings, the empennage is also designed to
encapsulate the propellers and shield their noise from the
downward direction. Due to the long profile in the propeller
area, a large part of the generated noise is deflected
upwards, thus reducing the noise to the ground. The
empennage covers the whole bottom side of the propeller
to get noise shielding, which is shown in FIG 6. Another
advantage of this configuration is that there is no possibility
of a shaded VTP through the HTP.

FIG 5. Influence of wing on the empennage
The morphing control surfaces on the wing and empennage
minimize drag and noise due to vortices in the deflected
state.

©2018Deutscher Luft- und Raumfahrtkongress 20184 FIG 6. Propeller noise emission shielding empennage with
morphing control surfaces
(2),
aerodynamic center and center of gravity.

(1), Coanda

flap

4.4. Technology Readiness Levels

Technology

TRL Source

4

4

5

8

6

4

7

4

7

5

6

[25]

[26]

[27]

[2]

[12]

[27]

[28]

[20]

[29]

[15]

[14]

HTS-Technology

Advanced Combustion
m
e
t
s
y
S

n
o
s
u
p
o
r
P

Low Noise CROR

LH2 Fuel Tank

Intercooler

l

i

Recuperator

Active Laminar Flow Control

i

c
m
a
n
y
d
o
r
e
A

Coanda Flap

Morphing Wing

Riblet Surface

Gondola-Fuselage Concept

TAB 1. TRLs for used technologies

5.

INTEGRATION OF SYSTEMS

5.1. Fuselage and Wing

In contrast to conventional designs this fuselage is
designed with the consideration of storing the fuel tanks yet
locating them outside the pressure cabin. For the maximum
fuel volume of 32 m3 most of the space below the
passenger cabin is used for fuel storage. For the basic
version of the aircraft the fuel system consists of two
parallel storage tanks at the front of the aircraft and a pair
of feeder tanks between the wing box and the gas turbines.
For a long-range version of the aircraft two additional tanks
can be installed. These optional fuel tanks can be stowed
in the cargo space and are connected to the storage tanks
in the front. The remaining cargo space remains usable to
load cargo containers and is detached from these additional
fuel tanks. Both versions provide enough cargo space for
the luggage of 150 passengers.

FIG 7. Polaris’ fuselage showing the fuel tanks (blue), cargo
space (brown) and structure

5.2. Landing Gear

The concentration of the propulsion components in the aft
shifts the center of gravity to the rear. This requires a main
landing gear that is located further back than in the
reference aircraft.
With the forward swept wing the main landing gear can
nonetheless be integrated in the wing root while considering
a proper position of aerodynamic center and center of
gravity during all flight conditions.

5.3. Empennage and Propulsion System

Contrary to the currently most commonly used gas-
turbines, that already have protective nacelles surrounding
the rotating parts, additional security measures must be
taken to ensure safety in case of a critical blade failure.
Therefore the propellers are mounted in the aft of the
aircraft, far away from the passenger cabin. Additionally, as
both propellers are close to one another, the rear section of
the aircraft is designed as such, that the engines are not in
the line of sight of one another and thus a ruptured blade
cannot collide with the propeller on the other side of the
aircraft (FIG 8).
The gas turbines are protected in case of a ruptured
propeller blade by the surrounding fuselage. Furthermore a
center frame between the two gas turbines is designed
considering the burst cone of each gas turbine (FIG 9).

FIG 8. Burstcone CROR

If the blade hits the morphing tail surface, only the limited
affected area becomes
inoperative. The unaffected
independent moving parts of the elevator can still be
controlled. The empennage is mounted to the main
structure under the turbines. The reinforced structure in this
area has sufficient stability to transmit the moments and
forces that occur.

©2018Deutscher Luft- und Raumfahrtkongress 20185 evaporated fuel is assumed to increase linear over time.
For maintenance or tank inspection a defueling procedure
is required. To inspect the tanks from inside the LH2 has to
be removed, purged and filled with breathable air. After
defueling and purging there is still gaseous hydrogen (GH2)
left in the tanks which has to be flushed out. The warm up
procedure can be started and after reaching 77.6 K the fuel
storage can be filled with dry nitrogen gas. This procedure
removes nearly all left hydrogen. After flushing them with
air to remove the nitrogen, the tank can be entered [30].
Refueling procedure after maintenance is similar but in
reverse order. The nitrogen flushes remaining air and CO2
out. Purging the tank from nitrogen is done by GH2. During
that the chill-down process of the tank starts by the use of
cold GH2. Fueling a warm tank with LH2 must be conducted
slowly at first to avoid over-pressurizing. The flow rate can
be increased with decreasing tank temperature. This whole
process shall be done overnight, to prevent absence form
service [30].

5.4.2. Fuel delivery lines

The fuel system consists of fuel delivery lines and pumps.
Due to hydrogen embrittlement a suitable material for the
fuel lines has to be selected. Steel and titanium are often
affected by
this problem. Ni-Cr based alloys show
similarities to austenitic steels, which shows a resistance to
this form of material degradation. They are already used in
space flight for cryogenic cooling of engine nozzles [32]
[33].
As material for the delivery lines Inconel 600 is used (see
FIG 10).

The validation of Polaris’ masses, aerodynamic and
mission performance is done with the data from CeRAS –
Central Reference Aircraft data System [34]. CSR-01 is a
short-range reference aircraft based on the A320 available
in 2005. With data
from e.g. propulsion system,
aerodynamics, masses and performance the necessary
information to validate the Polaris’ results are provided in a
single source published by the RWTH Aachen.
The iterations are carried out until masses, aerodynamic
and mission performance are modelled to a deviation of
< ∆1%. 6.1. Mass Estimation The masses of the different components are calculated using empirical formulas from Torenbeek [35] and the GD Method [36]. In a first step the reference aircraft is recalculated with the two methods. With the deviation of the calculated results a correction factor is determined to match both, the calculated and the provided CSR-01 data. In the second step Polaris’ masses are calculated with the same formulas and adjusted with the determined factor. A special attention is paid to the propulsion system, as HTS components and the LH2 fuel system cannot be calculated with empirical methods. 6.1.1. Electric Components Electric motors and generators are characterized by a power density of 20 kW/kg. With the supplied electric power FIG 9. Assembly of empennage, propulsion system and fuselage with gas turbine burst cone (1), ring frame (2) and center frame (3) 5.4. Fuel System In comparison to kerosene tanks LH2 tanks have to maintain the cryogen state of the fuel. This means the inner temperature must be kept at 21.7 K at a pressure of 1.4 bar [30]. The fuel system consists of up to six tanks. The main storage tanks can carry 600 kg of LH2 each. The feeder tanks and the additional storage tanks (long-range version) can carry 250 kg each. The fuel system is segmented into a left and a right side, each feeding the gas turbine on the same side. The segmentation into several tanks and two independent systems has on the one hand the safety reason that each gas turbine has its own fuel system and on the other hand weight and balance causes. The total fuel capacity is 1700 kg for the standard version and 2200 kg for the long-range version. The wall structure of an LH2 tank is closely linked to the selection of the insulation material. Due to safety concerns in case of e.g. a power outage, active cooling and vacuum insulation systems are disregarded for the LH2 fuel tanks of insulation consisting of Instead a passive Polaris. polyurethane foam is chosen. FIG 10. Fuel tank and fuel delivery line with their respective insulation The amount of diffused hydrogen by time mostly depends on the tank’s surface area and the insulation layer thickness. This amount can be estimated by equations based on [31]. Because of the surface area cylindrical shaped tanks are preferable over integral tanks (box- shaped). It is possible to leave the hydrogen in the fuel tanks on the ground or to defuel it during prolonged ground times. However it is recommended to leave the hydrogen in the fuel tanks due to defueling effort. During a mission of 3 hours 50.4 kg of LH2 evaporates. The amount of 5.4.1. Insulation of LH2 fuel tanks 6. AIRCRAFT PERFORMANCE ©2018Deutscher Luft- und Raumfahrtkongress 20186 of 22 MW, electric motor and generator weight 1100 kg per unit. lines are The superconducting power designed considering cross-feeding of the electric motors with a total length of 12 m at a specific cable mass of 10 kg/m. In total these lines have a mass of 120 kg. transmission 6.1.2. LH2 Fuel System The mass of the fuel tanks results out of the materials’ density and the tanks volume, insolation layer and anti- slosh walls. To regard degassing of LH2 the layer thickness varies between tanks. A detailed breakdown of the fuel tank calculation is shown in [31]. The fairing and vapor barrier are so light that they do not need to be considered. The small tanks can carry 250 kg and the large tanks can carry 600 kg each. the different fuel 6.2. Flight Performance To determine the saved energy of a new aircraft design an accurate calculation of the fuel consumption for both the reference aircraft and the new design is required. The CeRAS data base offer calculations for three different missions, each containing detailed information about the fuel for mission segments and reserve fuel. The three mission ranges (2750 NM and 500 NM with SPP, 2500 NM with max. payload) are recalculated to validate the calculation model with which the fuel consumption for a mission range of 1500 NM is determined. With the same calculation model the fuel consumption of Polaris is computed for its 1500 NM design mission and a 500 NM study mission. Parameter Unit Polaris CSR-01 Max. rate of climb ft/min 2285 Component Unit Small tank Large tank Cruise Mach number 0.72 Structure Insulation layer Anti-slosh walls Single tank All tanks kg kg kg kg kg 126 50 7 183 281 112 14 407 1547 Cruise Altitude L/D (Cruise) cL(max, LDG) Wing area Wing span 3698 0.78 35000 35000 20.17 17.43 2.88 2.80 100.0 122.4 34.0 34.1 - ft - - m2 m TAB 2. Tank component weight TAB 4. Flight performance data from Polaris and CSR-01 6.1.3. Mass Breakdown of Polaris and CSR-01 Table TAB 3. summarizes the masses of Polaris and CSR- 01 with breakdown of structural and propulsion system components. The values are given in kilogram. Description Polaris CSR-01 Max. Take-off Mass (MTOM) 53993 Operating Mass Empty (OME) 51967 Manufact. Mass Empty (MME) 33542 38153 77000 62100 7751 7520 Power Unit Equipped engines Electric motors Generators Propeller Air induction system Power cables LH2 fuel tanks 9185 1950 2200 2200 78 477 120 1547 573 39 LH2 fuel delivery system Engines control 231 n/a TAB 3. Mass breakdown of Polaris and