performance and sizing of transport aircraft employing electrically-powered distributed propulsion
TRANSCRIPT
PERFORMANCE AND SIZING OF TRANSPORT AIRCRAFT EMPLOYING ELECTRICALLY-POWERED DISTRIBUTED PROPULSION
Hans-Jörg Steiner, Patrick C. Vratny, Corin Gologan, Kerstin Wieczorek, Askin T. Isikveren and Mirko Hornung, Bauhaus Luftfahrt e.V., Munich, Germany
Abstract
Realizing a significant reduction of enroute emissions with respect to greenhouse gases is one major challenge in aircraft design today. Conventional kerosene propulsion systems are going to reach their efficiency limits in near future and it will be very ambitious to fulfil the requirements for future aircraft transportation using conventional engines. Consequently, new approaches for propulsion system design and integration are required to further improve aircraft efficiency through synergy effects. In this paper, a universally electric, short-haul, medium-capacity aircraft utilizing electric motors and battery for motive power is used as datum, focusing on the impact of a distributed propulsion system on the aircraft design and flight performance. Initial studies were performed identifying that the critical design cases for electric motor sizing are the one-engine-inoperative (OEI) flight segments, i.e. the climb gradients required at take-off and landing as well as field length requirements. By increasing the number of installed engines (i.e. motor-fan-combinations) the OEI performance requirements may be satisfied with a reduced amount of installed motor and battery system power. An integrated aircraft performance analysis is conducted to estimate the possible net benefit in terms of increased aircraft range when increasing the number of installed engines. Aerodynamic efficiency degradation is considered as well as weight impacts due to electric motor scaling and necessary system architecture modifications. The analysis shows that a 6% increase in aircraft design range can be achieved when going from 2 to 4 installed propulsive devices.
1. NOMENCLATURE
1.1 Abbreviations
AEO All Engines Operative
BCU Battery Control Unit
EMS Electric Motor System
FL Flight Level
GPU Ground Power Unit
HTS High Temperature Superconducting
ISA International Standard Atmosphere
MTOW Maximum Take-Off Weight
OEI One Engine Inoperative
OEW Operating Empty Weight
PAX Passenger
SSC Second Segment Climb
SSPC Solid State Power Controller
T/O Take-Off
TOC Top of Climb
TOFL Take-Off Field Length
UESA Universally Electric Systems Architecture
1.2 Symbols
DFan Fan Diameter
FPR Fan Pressure Ratio
GR Gear Ratio
Climb Rate
L/D Aerodynamic Efficiency
m Mass (no subscript indicates aircraft mass)
n Number of Propulsive Devices
P Power
Q Torque
Ref Subscript indicating reference value
RPM Revolutions per Minute
T Thrust
V Flight Speed
V1 Take-Off Decision Speed
V2 Take-Off Climb Safety Speed
Motor Electric Motor System Efficiency
Prop Propulsor Efficiency
Flight Path Angle
Density
P Specific Power
2. INTRODUCTION
The European Union (EU) unveiled an array of ambitious emission reduction goals for implementation by the year 2050 going far beyond near-term objectives such as those espoused by the Advisory Council for Aeronautics Research in Europe (ACARE) in 2001. Although near-term objectives declared by the ACARE Vision 2020 [1] with 80% and 50% reduction in nitrous oxide (NOx) and carbon dioxide (CO2) emissions, respectively, have been adopted by the European research community at large for over a decade now, the EU “Flightpath 2050” agenda [2] stipulates a reduction of 90% in NOx-emissions, and of 75% in CO2 emissions. All quoted values are relative to the emissions of typical aircraft in-service during year 2000.
Consequently, new approaches for propulsion system design and integration are required to further improve the aircraft efficiency through synergy effects. One promising idea is to distribute the power of one core engine to two or even more fans, either by mechanical, pneumatic, or electric transmission [3]. This offers the possibility of increasing the propulsive efficiency due to a larger total fan area while maintaining a moderate increase in fan diameter and weight [4]. Another main aspect being investigated in terms of this concept, called “Distributed Propulsion”, is the possible benefit due to ingesting the airframe boundary layer into the propulsors and filling the wake of the aircraft. This, in theory, reduces the amount of power loss in the flow field due to dissipation and hence improves the efficiency of the vehicle [5],[6].
This central idea of distributing the propulsion power may be even driven further by using only electric energy as power source together with electric power distribution. Thus, the idea of reducing the energy demand by distributing the thrust is realized in a simpler way compared to that of mechanical or turbo-electric transmission systems and has the beneficial by-product of zero in-flight emissions. In this paper, a universally electric, short-haul, medium-capacity aircraft is used as datum, focusing on the impact of a distributed propulsion system on the aircraft design and flight performance.
Initial studies are performed in order to identify the critical design cases for propulsor and electric motor sizing and estimate the potential benefit of increasing the number of engines. It will be shown that with more propulsors available the critical one-engine-inoperative (OEI) requirements can be satisfied with a reduced amount of installed power for the electric motor system. In order to get a better estimation of the possible benefit, trends for an optimal number of propulsors with respect to range and aircraft mass will be shown based on an integrated aircraft performance simulation model. Besides this, the impact of aerodynamic degradation due to a higher number of propulsors will be discussed.
3. BASELINE AIRCRAFT DESCRIPTION
The study presents the benefits of distributed propulsion based on a medium-capacity short-range aircraft with two aft-fuselage mounted electric propulsive devices (Figure 1). This baseline aircraft is a novel concept for zero-emission regional transport developed at Bauhaus Luftfahrt [7],[8]. The concept, dubbed Ce-Liner, was conceived for an entry-into-service date of 2035, with technologies searching to comply with ever stricter
requirements in terms of noise and emissions. The aircraft is therefore equipped with a Universally Electric Systems Architecture (UESA), i.e. electric energy is used as sole form of energy for all aircraft systems as well as for the propulsion group. The battery-powered propulsion system ensures thus zero emissions for the gate-to-gate mission.
The baseline aircraft is designed for a range of 900 nm at a cruise Mach number of Ma 0.75 at flight level FL 330 and features a cabin layout for 189 PAX. The main dimensions can be summarized via the overall length of 43.0 m, the overall height of 12.9 m and the span of 36.0 m including non-planar components. The design is thus compatible to ICAO Annex 14 Code C requirements which limit the aircraft dimensions to a 36 m box. Fulfilling this requirement is obviously important, as the aircraft is designed for the mid-range market segment.
Figure 1: Baseline aircraft BHL Ce-Liner.
The baseline aircraft has a Maximum Take-Off Weight (MTOW) of 109.3 t, which corresponds at the same time to its maximum landing weight: as the aircraft is fully battery powered, no mass reduction takes place during the mission. With a reference wing area of 172 m
2, the aircraft
features a wing loading of 636 kg/m2. In the aircraft weight
budget, the Operating Empty Weight (OEW) accounts for 59280 kg (54.2% of MTOW), and the battery weight for 30170 kg (27.6% of MTOW).
As indicated above, the propulsion system is conceived as a fully electric battery powered architecture. The propulsors are realized as ducted fans that are mounted on pylons to the aft fuselage and provide a thrust of 124 kN each at take-off (T/O rating @ Ma 0.2) and 28.5 kN each in cruise. In the OEI case, the remaining propulsor can provide a take-off thrust of 147 kN (OEI rating @ Ma 0.2). The fans have a diameter of 2.70 m, and a very low specific thrust of 76.7 m/s. The design does, therefore, not only provide zero-emission transport in all phases including ground maneuvering, but contributes also to a reduction in perceived external noise.
The required shaft power during cruise is 15 MW. However, the electric motors driving the fans must be able to provide a maximum power of 22 MW each. This installed power is derived from the thrust requirement in the OEI case during Second Segment Climb (SSC). The UESA of the aircraft provides a nominal power of 28.8 MW to the motors. This can be increased to 34.6 MW for a brief period to satisfy the power requirement during take-off.
The performance of the propulsion system can be indicated by a power-to-weight ratio of 0.402 kW/kg in terms of installed motor power, and 0.263 kW/kg in terms of installed battery power. The baseline aircraft requires a Take-Off Field Length (TOFL) of 2120 m at sea level, ISA (2300 m at sea level, ISA+15); the climb phase to FL330 is realized in approx. 21 min.
In the presented study, this baseline aircraft will be equipped with a variable number of propulsors to demonstrate the impact of distributed propulsion.
4. BASIC CONSIDERATIONS AND MOTIVATION
This section discusses some basic considerations when applying distributed propulsion on the universally electric baseline aircraft. The technical aspects and the impact of distributing the propulsion power to a higher number of propulsive devices are treated in a preliminary manner using analytical derivations in order to show the key drivers and motivation for distributed propulsion.
For simplification, the terms “propulsive device”, or simply “engine”, are used in the context of this study to denote an integrated electric propulsive device as presented in Figure 2. The term “propulsor” indicates the ducted fan device. The assembly of electric motor, gearbox, controller and cooling is called the Electric Motor System (EMS).
Figure 2: Schematic presentation of the integrated electric propulsive device.
4.1 Illustration of the Potential Benefit
During the design of the propulsion system of the baseline aircraft configuration having two ducted fans driven by one EMS each, one major issue becoming apparent was the necessary oversizing of the EMS due to the mismatch of fan and motor power characteristics. The independency of the maximum available motor power with respect to the flight condition in contrast to the strongly varying power demand of the fans leads to a design where the EMS is operating in relatively low part-load setting during cruise. This situation becomes more pronounced due to the necessary sizing of the propulsion system for the OEI case. Besides the decreased efficiency due to part-load, which is less critical then it would be for gas turbines, this indicates a potential for improvement, for example, in terms of weight.
The situation is further illustrated with Figure 3, which shows two distinct sizing cases of the propulsion system
for the baseline aircraft (n = 2) and a system featuring four
propulsive devices. Generally, due to the operational characteristics of the ducted fan the propulsors are sized for the Top-Of-Climb (TOC) case, yielding a specified climb rate at a given cruise condition. For the reference aircraft a total thrust of 66 kN at FL330 and Ma 0.75 is
required, translating into a total shaft power demand of 18 MW. If second order effects like fan efficiency, drag and weight are neglected, this is independent of the number of engines.
Figure 3: Illustration showing the difference in sizing of the
EMS for two and four propulsive devices.
However, the sizing of the EMS is typically driven by an OEI case, e.g. the necessary climb gradient at SSC or approach-climb. For the baseline aircraft an OEI thrust of
127 kN at 5000 ft and V2 is required, translating into a total
shaft power demand of 22 MW. Here, the major difference due to the number of engines becomes apparent: While in the two engine case each EMS must be sized for the required 22 MW, leading to a total installed motor power of
44 MW, the distributed propulsion system with n = 4 only
requires a total installed motor power of 29.3 MW (= 4/3 times 22 MW). The assumption of the required OEI thrust being independent of the number of engines does not take
into account the different climb gradient requirements for n
> 2. This fact is considered in the next section. It should be recalled that this potential is given due to the separation of thrust producing device (propulsor) and the device supplying the necessary shaft power and only applies for the EMS. There is no benefit on side of the fans because they are sized for the TOC (AEO) case.
Additionally, the UESA of the aircraft must be designed to deliver the necessary amount of power for all flight conditions, i.e. AEO as well as OEI flight cases. Here, one of the main differences with respect to aircraft design of conventional and electric transport aircraft becomes apparent: while in conventional aircraft conceptual design the propulsion system may generally be treated as a single design variable (either thrust to weight ratio or power to weight ratio), in electric aircraft design a separate treatment of propulsor sizing, EMS sizing, and electric system architecture sizing is necessary.
4.2 Preliminary Sizing Considerations
There are several cases that might be critical for sizing the propulsion system, e.g.
Climb gradient for SSC (OEI)
Climb gradient for missed approach (OEI)
Climb rate at TOC (AEO)
Time to climb to Cruise Altitude
TOFL requirement (AEO, OEI)
For the investigated case of an electric propulsion system, the sizing case for the propulsor (thrust requirement) may be differing from the sizing case for the EMS and the UESA (power requirement). In the following, the dependency of these critical cases on the number of engines is derived based on first order analytical equations. For simplification, only SSC gradient (OEI) and TOC gradient (AEO) are considered in this section. This also reflects the findings for the design of the baseline aircraft, where these were found to be the critical sizing cases. However, further cases like e.g. missed approach at OEI may be treated in the same way as described in the following.
The available thrust of the propulsors relative to the thrust at a reference flight condition (thrust lapse) may be modeled as a function of flight speed and altitude using the following equation:
(1)
n
Ref
n
RefRef
V
V
V
T
T
where nV and n are regression exponents. The required
total thrust for a specified climb gradient or climb rate
is given relative to the aircraft weight as a function of aerodynamic efficiency:
(2)
V
h
DL
DLmg
T sin
1sin
1
Combining equation (1) and (2) and keeping in mind that the aircraft weight is constant, the installed thrust requirements with respect to TOC (AEO) and SSC (OEI) are derived as
(3)
n
Ref
TOC
n
Ref
TOC
TOCTOC
AEOTOC
Ref
V
V
V
V
h
DLmg
T
sin
1
,
(4)
n
Ref
Climb
n
Ref
ClimbClimb
Climb
Climb,OEI
Ref
V
V
V
DLn
n
mg
T
sin1
1
The result of equation (3) and (4) when applied to the baseline aircraft is shown in Figure 4. For the SSC, two cases for the required climb gradient are considered: the first one linearly extrapolates the existing certification
requirements [9] for n > 4 using the following equation
(5) n 003.0018.0sin
and the second one limits the required climb gradient at
the value for four engines (sin = 0.030).
It can be seen that for the baseline aircraft with two engines the critical thrust requirement is given by the TOC case. This is due to the large thrust lapse of the fans having a very low specific thrust. Increasing the number of engines decreases the required thrust due to OEI climb, but as this is not the critical case, this has no influence on the propulsor sizing.
Figure 4: Installed thrust requirement with respect to TOC (AEO) and SSC (OEI) as a function of number of engines.
In order to investigate a possible benefit in terms of EMS weight, the required shaft power for the critical flight cases is estimated. The power can be calculated with the required thrust as given in Equation (2) and the propulsor
efficiency Prop of the fan. Propulsive device efficiency is
further explained in Section 5.1. The required shaft power with respect to TOC (AEO) and SSC (OEI) can be derived as
(6)
TOCProp
TOC
TOCAEOTOC
Motor gV
mg
T
m
P
,
(7)
ClimbProp
Climb
ClimbClimb,OEI
Motor gV
mg
T
n
n
m
P
1
The required installed EMS power for the reference aircraft according to equations (6) and (7) is plotted in Figure 5 as a function of the number of engines. The propulsive device efficiency at TOC and at SSC is 78% and 50%, respectively. Additionally, the required installed EMS power for an AEO TOFL of 2300 m is shown in the figure. This value has been taken from the detailed studies shown later on in Section 6. The magnitude is independent of the number of engines and becomes the critical sizing case for more than four engines.
Figure 5: Installed EMS power requirement with respect to TOC (AEO) and SSC (OEI) as a function of number of
engines.
2 3 4 5 6 7 8 9 100
0.05
0.1
0.15
0.2
0.25
0.3
0.35
0.4
0.45
0.5
Number of engines
Pro
puls
or
thru
st
to a
ircra
ft w
eig
ht
ratio
2nd segment climb (OEI), extrapolated
2nd segment climb (OEI), limited at 3.0%
Top Of Climb 300ft/min (AEO)
2 3 4 5 6 7 8 9 100
0.05
0.1
0.15
0.2
0.25
0.3
0.35
0.4
0.45
0.5
Number of engines
E-M
oto
r pow
er
to a
ircra
ft w
eig
ht
ratio [
kW
/kg]
2nd segment climb (OEI), extrapolated
2nd segment climb (OEI), limited at 3.0%
Top Of Climb 300ft/min (AEO)
Take-Off Field Length 2300m (AEO)
It can be seen that the power requirement for TOC with all engines operative is below the power requirement for the necessary climb gradient at OEI and the AEO TOFL. Hence, with the OEI case being the critical sizing case for the EMS, an increase of the number of engines has the potential of considerably reducing the installed motor power, and thus, the weight of the EMS. The weight saving potential can be roughly estimated by assuming a linear relationship between power and weight. The components of the electric system impacted by the installed motor power include the motor itself, the motor controller, the gearbox, the motor Solid State Power Controller (SSPC), and the according cryocoolers to achieve High Temperature Superconducting (HTS) capabilities, i.e. approximately 10% of the aircraft OEW. Thus, a 33% reduction of installed motor power can save up to 3.3% of aircraft empty weight.
The required installed UESA power is a result of detailed mission and take-off analysis yielding the climb time to cruise altitude and the TOFL. For the baseline aircraft, a possible reduction in UESA power may be expected due to the UESA of the baseline aircraft being sized for the necessary power during the AEO phase of the take-off
with one engine failing at V1. With the OEI phase of take-
off becoming less critical for more engines, also the required power for the AEO phase may be reduced.
The previously shown potential for improving the performance of the baseline aircraft with a distributed propulsion system gives reason for a more sophisticated analysis. The goal of the following section is to estimate the possible net benefit when negative effects of increasing the number of engines are considered. This includes degraded aerodynamics as well as weight increase due to the electrical power distribution and the scaling of the electric components. However, this study did not take into account a possible benefit achievable by a stronger integration of the propulsion system within the airframe.
5. DISTRIBUTED PROPULSION SYSTEM MODELING
This section describes the modeling approach used in the aircraft performance study shown in Section 6. The modeling covers the ducted fan performance, the electric motor and gearbox model, as well as the necessary modifications to the UESA.
The basic design rules for increasing the number of engines are given by the assumption of constant total fan area, i.e. the fan pressure ratio is assumed to be constant for any number of engines. The fan diameter is then derived based on the baseline fan diameter as a function of number of engines:
(8)
nn
n
D
D Ref
RefFan
Fan 2
,
Assuming a constant tip speed of the fans, the rotational speed scales according to the following equation:
(9)
2
n
n
n
RPM
RPM
RefRef
Further, assuming that the total installed EMS power is
distributed equally to all motors, the torque Q acting on the
gearbox scales according to
(10) 33
38
nn
n
Q
Q Ref
Ref
Equation (9) and (10) directly show that a reduction of gearbox weight can be expected when increasing the number of engines.
5.1 Ducted Fan Model
A zero-dimensional performance model of a ducted fan with the ability to predict design and off-design performance was used in order to estimate the maximum
available thrust and the propulsor efficiency Prop. Here,
propulsor efficiency is defined as the ratio of usable
propulsion power (net thrust times flight speed, TV) to fan
shaft power, hence covering transmission efficiency and propulsive efficiency. Transmission efficiency is composed of fan polytropic efficiency, as well as intake, ducting and nozzle losses [10]. The model is based on basic gas-dynamic relationships and standard compressor theory [11].
The fan performance model was used to calculate the maximum available thrust and necessary shaft power at a given flight condition (shown in Figure 6 for the baseline aircraft fan having a design thrust of 33 kN at 33000 ft and Ma 0.75). However, it is not feasible to supply the fan with the depicted maximum power of up to 40 MW. Rather than that, the motor power is rated to a certain limit (22 MW for the baseline aircraft). This infers a drop of fan rotational speed and thrust, accordingly. Generally, it can be said that the maximum thrust of the ducted fan at high altitudes is driven by fan geometry and fan pressure ratio, while the maximum thrust at low altitudes is limited by the supplied shaft power through the electric system.
The mass of the ducted fan for the baseline aircraft was estimated based on flow path geometry sizing and empirical relations. For this preliminary study on the impact of distributed propulsion, a constant thrust to weight ratio of the ducted fan is assumed. This is supported by the fact that the total nacelle surface area is constant with respect to the number of engines, when assuming a linear relationship between nacelle length and nacelle diameter given mainly by the fan pressure ratio [12].
Figure 6: Maximum thrust and according shaft power (contoured) of the baseline aircraft ducted fan (design
point at FL330, Ma 0.75).
0
5000
10000
15000 0
0.2
0.4
0.6
0.80
50
100
150
200
250
Mach NumberAltitude [m]
Thru
st
[kN
]
5 MW
10 MW
15 MW
20 MW
25 MW
30 MW
35 MW
40 MW
Shaft Power
Design Point
5.2 Electric Motor Model
A major characteristic of a HTS motor compared to a conventional electric motor is that the specific power increases with increasing power [13]. A conventional motor has nearly a constant specific power [13]. This section deals with an approach to model the mass and the specific power of a HTS motor depending on its design power and maximum rotational speed. Equation (11), also derived from [13], determines the specific power of a HTS
motor (given in kW/kg) depending on the motor power P
(given in kW) within a power range of 200 kW to 800 kW at 3000 rpm. Finally, in order to determine the mass of the electric motor Equation (12) is used:
(11) 4208.05731.3)( PPP
(12)
P
Motor
Pm
Extrapolating these values to higher motor powers than in the above defined range leads to very high specific powers, which are not in line with values found in current literature. By changing the extrapolation method from the power function used in Equation (11) to a logarithmic function, good agreement with specific powers at higher electric motor power is achieved.
The trend curves defined with this method are also in good accordance with curves found in [14] (at least in a low rpm range). For the dependency of the rotational speed on the mass impact also values from literature are used to derive a dependency function. As baseline, 5 MW HTS motors are taken at different rotational speed, namely 230 rpm
with a specific power ρP,Motor of 0.20 kW/kg [15], a
16000 rpm motor with a specific power ρP,Motor of
7.78 kW/kg [16] and a 35000 rpm motor with 9.17 kW/kg [17]. It is assumed that the values of [15] and [16] do not include the cooling equipment. Therefore the original specific powers are adapted with Equation (13) assuming the same efficiencies for the electric motor as mentioned
in [1] with a cryocooler specific power ρP,Cool of 0.33 kW/kg
[18].
(13)
1
,,
,
11
CoolP
Motor
MotorP
gplusCoolinP
This equation only includes the efficiency losses of the
motor ηMotor, the efficiency loss of the cryocooler was not
considered in this investigation. Such an approach is described in [10]. Analysis has shown that also a logarithmic curve fitting of these values predicts the values found in literature well. It is assumed that this behavior of the rotational speed on the specific power can be projected to other power ranges.
The result of the combination of both methods (rotational speed and power dependency) is shown in Figure 7. Here, the specific power of a HTS Motor is shown for different rotational speeds and at different design powers. The trends represented here are also in a good accordance with the results presented in [14], where a similar approach is shown. The values in this chart are already corrected with a technology factor to meet the motor performance as defined in the baseline aircraft [1]. The technology factor includes a further development and optimization of the cooling system as well as an improved motor design including HTS materials for rotor and stators.
Figure 7: Estimated development of the specific power of a HTS motor depending on its rotational speed and design
power.
5.3 Gearbox Model
The rotational speed of a ducted fan depends on the fan pressure ratio and fan diameter, which has to be taken into account, when distributing the thrust onto several propulsors. The mass of a HTS motor also shows a dependency with respect to design power and rotational speed as shown in the previous section. Because the rotational speed of the fan is fixed, the rotational speed of the electric motor gives an additional degree of freedom to optimize the mass of the propulsion system. Due to the electric motor being more weight efficient at higher rotational speeds, a gearbox may be required, which reduces the rotational speed of the motor to the required rotational speed of the fan. For the gearbox, a planetary reduction gear system is assumed, which is also installed on the baseline aircraft. For the estimation of the gearbox mass an equation derived by NASA [19] is adopted. It is also assumed that a single planetary stage can withstand a maximum gear ratio (GR) of 4.4. A higher reduction gear system would require a second stage. The GR is calculated with the following equation (14):
(14)
Fan
Motor
RPM
RPMGR
Finally with a reference gear ratio GRRef, mass mRef
(corrected with a technology factor k), torque QRef and the
required output torque Q, the mass of one planetary gear
stage can be estimated with Equation (15) taken from [19] incl. the reference values:
(15)
ff
fGearQ
Q
GR
GRkmm
Re
15.0
Re
Re
If the GR exceeds the mentioned rate of 4.4 it is assumed that stage 1 is designed for the maximum GR and stage 2 is designed for the remaining GR to meet the required output torque and rotational speed at the propulsors. The stage allocation is illustrated in Figure 8.
0 0.5 1 1.5 2 2.5 3 3.5
x 104
0
2
4
6
8
10
12
14
16
18
20
Motor speed [rpm]
Sp
ecif
ic P
ow
er
[kW
/kg
]
5 MW
10 MW
15 MW
20 MW
Basline
5 MW today
Values from literature
Figure 9 shows the resulting specific power ρP,Total for the
combination of electric motor and gearbox, as calculated with Equation (16) for different power levels and rotational speeds of the fans.
(16)
GearMotor
TotalPmm
P
,
Figure 9: Specific power of electric motor and gearbox combination as a function of rotational speed of fan and
motor and maximum motor power.
It can be recognized that with increasing rotational speed of the fan, the breakpoint of a one stage gear to a two stage gear moves to higher rotational speeds of the electric motor. Furthermore the mass of the gearbox system also decreases slightly with higher rotational speeds of the fan, which can be recognized by the smaller drop in specific power due to the gearbox installation. This decreased gearbox mass is a result of the reduced output torque due to the higher rotational speed of the fan. Furthermore it can be shown that with higher fan speeds the electric motor can also operate at lower speeds with nearly the same weight impact.
5.4 Adaption of Universally Electric System Architecture
The UESA of the baseline aircraft has to be adapted to handle the increased number of propulsive devices. The major change of the UESA is given by the shift of an aft mounted engine configuration towards an under-wing mounted configuration with high-power busses installed in each wing. On each bus a maximum of eight battery packs can be installed, which are powering the electric motors. Both systems are connected together via a cross-link SSPC, which allows a power transfer from one propulsion system side to the other in an OEI case. Furthermore, the location of the battery containers is also kept as in the reference aircraft. The new system architecture is shown in Figure 10. Because only the propulsion system has changed, the subsystems stay the same as defined in the reference aircraft.
Figure 10: Adapted UESA for wing mounted fans based on [7].
The next step is to adapt the cable lengths, and in turn the cable masses and the position of the motor controllers and protection switches to the new design. For that reason, the cables and power control units are completely removed from the aft section. The cables and busses are now extending to both sides of the wing, depending on the position of the most outboard located propulsor. Also the SSPCs are now mounted in the wing, as well as the motor controllers. The SSPC and cable, which connect both propulsion busses, are designed for the OEI case, which allows a power transfer from one system to the other. This maximum power of the cross-link SSPC can be calculated with Equation (17):
(17) )1(2
n
nPP Motor
LinkCross
0 0.5 1 1.5 2 2.5 3 3.5
x 104
6
8
10
12
14
16
Fan speed 2000 rpm
Motor speed [rpm]
Mo
tor
plu
s G
earb
ox
Sp
ecif
ic P
ow
er
[kW
/kg
]
5 MW
10 MW
15 MW
20 MW
0 0.5 1 1.5 2 2.5 3 3.5
x 104
5
10
15
20
Fan speed 3000 rpm
Motor speed [rpm]
Mo
tor
plu
s G
earb
ox
Sp
ecif
ic P
ow
er
[kW
/kg
]
5 MW
10 MW
15 MW
20 MW
0 0.5 1 1.5 2 2.5 3 3.5
x 104
5
10
15
20
Fan speed 4000 rpm
Motor speed [rpm]
Mo
tor
plu
s G
earb
ox
Sp
ecif
ic P
ow
er
[kW
/kg
]
5 MW
10 MW
15 MW
20 MW
0 0.5 1 1.5 2 2.5 3 3.5
x 104
5
10
15
20
Fan speed 5000 rpm
Motor speed [rpm]
Mo
tor
plu
s G
earb
ox
Sp
ecif
ic P
ow
er
[kW
/kg
]
5 MW
10 MW
15 MW
20 MW
Motor Bus Wing Right
BCU BCU
ACDC
ACDC
Hot. Bat
BCU BCU
GPU.
SS
PC
SS
PC
SS
PC
SS
PC
SS
PC
SS
PC
SS
PC
SS
PC
Motor Bus Wing Left
BCU BCU
ACDC
ACDC
BCU BCU
SS
PC
SS
PC
SS
PC
SS
PC
SS
PC
SS
PC
Subsystems
SSPC
SSPC
SSPC
2
n
2
n
E-Motor
Sta
ge
1
Sta
ge
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Gearbox
Figure 8: Definition of reduction gearbox stages between propulsor and electric motor.
2 stages 1 stage
6. INTEGRATED AIRCRAFT STUDY
This section shows the results on aircraft level mainly focusing on the changes in necessary installed EMS and UESA power depending on the number of propulsive devices. In order to make the configurations comparable to the baseline aircraft, all performance requirements, i.e. TOFL and LFL requirements as well as SSC gradients, missed approach climb gradients and the maximum climb time to cruise altitude were retained. For the simulation of the aircraft performance, the aircraft conceptual design tool PaceLab APD [20] was extended in order to model the unique features of the investigated aircraft:
Mission performance adoption for universally electric aircraft with constant mass and zero fuel burn [21]
Adoption of wing aerodynamics methods for C-Wing morphology
Replacement of wing weight method in order to model wing inertial relief sensitivity according to Luftfahrttechnisches Handbuch (LTH)
Due to the high accuracy required for the low-speed performance, the empiric method for TOFL calculation was replaced by numerical low-speed performance methods [22].
For the identification of the reduction potential due to distributed propulsion on the installed power of the entire propulsion system on aircraft level, configurations with 2, 4, 6, 8 and 10 engines mounted under the wing are investigated. These configurations are derived from the baseline aircraft and are schematically shown in Figure 11. The under-wing mounted engine integration was chosen in order to allow for a simple comparison of configurations over a large range of number of engines. As an additional side effect, this configuration allows a reduction of wing mass by decreasing the structural stresses due to the wing bending moment. As negative effect for this configuration a mass increase acc. to Raymer [23] is considered by extending the landing gear length to retain a feasible ground clearance. Also an aerodynamic degradation is considered by assuming 20% of the baseline nacelle drag being due to interference effects, and retaining that absolute value for each of the installed engines.
Figure 11: Investigated configurations.
The results of the integrated performance analysis are listed in Table 1 below. For each of the investigated concepts the performance requirements being critical for
aircraft sizing are highlighted. For n = 2 the installed motor
power is sized by the SSC gradient (OEI), and installed
UESA power is driven by the TOFL (OEI). The short time to cruise altitude of 20.6 min compared to the requirement of 25 min indicates that the UESA is over-powered in terms of climb performance. When going from 2 to 4 engines, the installed motor power as well as the UESA power can be reduced. The better OEI performance during take-off allows also for a reduction in UESA power resulting in the climb time requirement becoming critical. Increasing the number of engines further, does not allow for a further reduction of motor power since the UESA power at take-off has still to be processed by the motors. Here, the AEO TOFL becomes the critical sizing condition.
The shift of critical sizing constraints for the investigated variation in number of engines is additionally visualized in Figure 12 for the TOFL. It can be seen that up to 4 engines the OEI case is the critical one. For 6 engines and more the AEO field length (including 15% margin) becomes the critical sizing case.
Figure 12: Take-off field length (AEO and OEI case) of the investigated concepts.
Table 1: Results of the integrated performance analysis.
Figure 13 shows the results of the power loading of the different configurations of the integrated aircraft study. Additionally, the power loading characteristics as being analytically estimated in Section 4.2 is depicted. It can be seen that the analytical results agree well with the integrated aircraft study for 4 engines. The further predicted power reduction could not be realized due to a minimum installed motor power required for the take-off with all engines operative.
Baseline
Number of Engines 2 2 4 6 8 10
Installed Power
Power Loading (E-Motor) kW/kg 0.402 0.403 0.271 0.263 0.264 0.264
Motor System Power (AEO) MW 43.9 44.0 29.6 28.8 28.8 28.9
UESA Power (nominal) MW 28.8 28.9 24.7 25.0 25.3 25.6
UESA Power (Take-Off) MW 34.6 34.7 29.6 28.8 28.8 28.9
Aircraft Performance
Mission (ISA+10)
Range nm 900 864 955 945 929 909
Time to ICA min 20.6 20.6 25.0 25.0 25.0 25.0
Take-Off (ISA+15, SL)
Field Length (AEO) m 1950 1950 2230 2300 2300 2300
Field Length (OEI) m 2300 2300 2300 2210 2150 2120
Field Length (FAR) m 2300 2300 2300 2300 2300 2300
Climb (ISA+20, 5000ft)
2nd Segment Climb
Gradient (AEO) % 9.39 9.38 6.90 6.43 6.45 6.47
2nd Segment Climb
Gradient (OEI) % 2.43 2.43 3.78 4.88 5.64 5.88
Missed Approach (ISA)
Climb Gradient (AEO) % 14.4 14.4 7.26 6.80 6.82 6.59
Climb Gradient (OEI) % 2.34 2.34 3.64 4.89 5.71 5.97
Distributed Propulsion (wing-mounted)
Figure 13: EMS power loading as a function of number of installed engines (shown are the results of the integrated aircraft performance study and the analytically predicted
values acc. to equation (7)).
Figure 14 gives an overview of the mass impact of the distributed propulsion system on the aircraft OEW and the contributing shares, i.e. electrical system architecture, propulsion system (including electric motor, gearbox, and controller), structure, and operational items and equipment. Due to the decreased motor power loading and the decreased UESA power, the weight of these components can be significantly reduced. The UESA is found to have a minimum weight with four engines. Exceeding this number leads to a higher number of electrical components, like protection switches, which would have a higher negative mass impact than the decreased power loading. The optimum of the integrated electric propulsor mass is identified at six propulsors, which is also the optimum in terms of the total aircraft OEW. Exceeding this number would also lead to a higher electrical motor mass as shown in Section 5.2, with the specific power of a HTS motor decreasing with decreasing design power.
Figure 14: OEW mass breakdown for the investigated distributed propulsion concepts.
The possible reduction in aircraft OEW directly translates into an increase in design range when keeping the aircraft MTOW constant (cf. Figure 15). The weight delta in OEW is used for enlarging the battery system and thus increasing the energy capacity of the aircraft. However, this potential increase in design range is counteracted by a reduction of aerodynamic efficiency assumed for the distributed propulsion system. For this preliminary study, the aerodynamic degradation is modeled by a constant interference drag of each installed engine independent of the engine size.
The best result in terms of aircraft design range was found for the configuration with 4 engines: Here, a range increase of 6% over the baseline aircraft was predicted. This benefit diminishes when increasing the number of engines to more than four due to the additional drag and weight introduced into the aircraft system.
Figure 15: Aircraft design range as a function of number of installed engines.
7. CONCLUSION AND FUTURE WORK
In this paper the potential of improving the performance of a baseline universally electric transport aircraft by employing distributed propulsion was investigated. This research was motivated by preliminary analysis, showing that the critical sizing cases for the EMS of the baseline aircraft are one-engine-inoperative cases. Therefore, there is significant potential for weight reduction due to the large fraction of electric component weight for the given electric aircraft application. A reduction of installed electric motor power also mitigates one major issue of electric propulsion – namely the motors operating in deep part load during cruise. This is a result of the large variation in power demand of the propulsors between high altitudes for cruise and low altitudes for operations like, for example, take-off or second segment climb, which are even required during a one-engine-inoperative condition.
The potential benefit of increasing the number of propulsive devices has been further investigated using a more detailed mission performance simulation. Here, also negative impacts with respect to aerodynamics and weight of the electrical system architecture have been taken into account to estimate the possible net benefit. However, this study did not consider a stronger coupling of airframe and propulsion system, which might further increase the benefit of distributed propulsion systems. The pure down-scaling of the propulsion devices on the one hand increases the electric motor weight due to a decrease in specific power. On the other hand the higher rotational fan speeds due to the smaller fan diameter and the consequently reduced torque decrease the gearbox weight. The increased weight of the necessary power management and distribution system was estimated based on geometric and power load considerations. The performance studies showed a possible increase of aircraft design range up to 6% compared to the baseline aircraft when installing a total number of 4 engines.
Ongoing and future work comprises on the one hand side several modifications to the baseline aircraft, e.g. investigations on a self-trimming C-wing using morphing technology [24]. On the other hand side, the concept of distributed propulsion will be further examined, especially
Baseline
with focus on a stronger coupling between airframe and propulsors. First results from the authors are published at the ICAS 2012, investigating a novel integrated propulsion concept called Propulsive Fuselage [6].
8. ACKNOWLEDGEMENTS
The authors like to acknowledge the work of all Bauhaus Luftfahrt team members participating in the conceptual design of the baseline aircraft.
9. REFERENCES
[1] Advisory Council for Aeronautical Research in Europe (ACARE), “European Aeronautics: A Vision for 2020,” January 2001.
[2] European Commission (EC), “Flightpath 2050: Europe’s Vision for Aviation,” Report of the High Level Group on Aviation Research. Publications Office of the European Union, Luxembourg, 2011.
[3] J.L. Felder, H.D. Kim, and G.V. Brown, “Turboelectric Distributed Propulsion Engine Cycle Analysis for Hybrid-Wing-Body Aircraft,” 47th AIAA Aerospace Sciences Meeting Including The New Horizons Forum and Aerospace Exposition, Orlando, 2009.
[4] C.A. Hall, and D. Crichton, “Engine and Installation Configurations for a Silent Aircraft,” 17th International Symposium on Airbreathing Engines,
Munich, 2005.
[5] S. Sato, P.C. Mody, D.K. Hall, E.R. Blanco, and J.I. Hileman, “Assessment of Propulsion System Configuration and Fuel Composition on Hybrid Wing Body Fuel Efficiency,” 49th AIAA Aerospace Sciences Meeting, Orlando, 2011.
[6] H.-J. Steiner, A. Seitz, K. Wieczorek, K. Plötner, A.T. Isikveren, and M. Hornung, “Multi-Disciplinary Design and Feasibility Study of Distributed Propulsion Systems,” International Congress of the Aeronautical Sciences ICAS, Brisbane, 2012.
[7] Visionary Aircraft Concepts Group. “Concept 002: Initial Technical Assessment of an Electrically-Powered, Medium-Capacity, Short-Haul Transport Aircraft,” Internal Report IB-12021, Bauhaus
Luftfahrt e.V., 2012.
[8] A.T. Isikveren, A. Seitz, P.C. Vratny, C. Pornet, K.O. Plötner, and M. Hornung, “Conceptual Studies of Universally-Electric Systems Architectures Suitable for Transport Aircraft,” Deutscher Luft- und Raumfahrtkongress DLRK, Berlin, Germany, 2012.
[9] EASA, “Decision No. 2003/2/RM on certification specifications, including airworthiness codes and acceptable means of compliance, for large aeroplanes («CS-25»),” Report No. 2003/2/RM, Brussels, Belgium, 2003.
[10] A. Seitz, O. Schmitz, A.T. Isikveren, and M. Hornung, “Electrically Powered Propulsion: Comparison and Contrast to Gas Turbines,”
Deutscher Luft- und Raumfahrtkongress DLRK,
Berlin, Germany, 2012.
[11] H.-J. Steiner, and O. Schmitz, “Ducted Fan Model,” Internal Report IB-11013, Bauhaus Luftfahrt e.V., 2011.
[12] A. Seitz, “Advanced Methods for Propulsion System Integration in Aircraft Conceptual Design,” Dissertation, Institut für Luft- und Raumfahrt, Technische Universität München, 2012.
[13] P. Masson, “Wind Turbine Generators: Beyond the 10MW Frontier,” Symposium on Superconducting Devices for Wind Energy Systems, Barcelona, Spain, 2011.
[14] S. Stückl, “The All Electric Aircraft – Emissionsfreies Fliegen im Jahr 2035?,” Deutscher Luft- und Raumfahrtkongress DLRK, Bremen,
Germany, 2011.
[15] S. Greg, G. Bruce, and S.K. Swarn, “The Performance of a 5 MW High Temperature Superconductor Ship Propulsion Motor,” IEEE Transactions on Applied Superconductivity, Vol. 15, No. 2, pp. 2206-2209, 2005.
[16] P. Masson and J. Pienkos, “High Power Density Superconducting Electric Motors,” NASA/DoD UAPT Program Annual Review, NASA Glenn Research Center, 2007.
[17] K. Sivasubramaniam, T. Zhang, M. Lokhandwalla, E.T. Laskaris, J.W. Bray, B. Gerstler, M.R. Shah, and J.P. Alexander, “Development of a High Speed HTS Generator for Airborne Applications,” IEEE Transactions on Applied Superconductivity, Vol. 19, No. 3, pp. 1-6, 2009.
[18] G.V. Brown, “Weights and Efficiencies of Electric Components of a Turboelectric Aircraft Propulsion System,” 49th AIAA Aerospace Sciences Meeting, Orlando, 2011.
[19] C. Reynolds "Advanced Propfan Engine Technology (APET) Single- and Counterrotation Gearbox / Pitch Change Mechanism, Final Report", Pratt & Whitney United Technologies Corporation, NASA-CR-168114, Vol. 1 & 2, 1985.
[20] Pace GmbH, PaceLab Suite 2.0, 2010.
[21] Vratny, P.C., A Battery Powered Transport Aircraft – A Concept Study of a Real All Electric Aircraft, AV Akademikerverlag, Saarbrücken, 2012.
[22] Gologan, C., A Method for the Comparison of Transport Aircraft with Blown Flaps, Doctoral Thesis, Technische Universität München, 2010.
[23] Raymer, D.P., Aircraft Design, American Institute of Aeronautics and Astronautics, Washington, DC,1999.
[24] M. Trapani, M. Pleißner, A. T. Isikveren, and K. Wieczorek, “Preliminary Investigation of a Self-Trimming Non-Planar Wing Using Adaptive Utilities,” Deutscher Luft- und Raumfahrtkongress DLRK, Berlin, Germany, 2012.