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Figure 1. Number of People Affected by Aircraft Noise in the U.S6
Figure 2. History of Demand, Efficiency, and Fuel Burn6
regulation. It may seem that the upcoming standard is not so constraining on a per-aircraft basis given that the
centerline takeoff noise level has been reducing at rate of 3 dB per decade.6 Nevertheless, Fig. 1 shows that further
reductions in the number of people affected by aircraft noise is expected to be small over the next 20 years, because
the anticipated long-term
growth in commercial airline
demand is projected to
outpace that of evolutionary
noise-reduction technology
improvements.7
Similarly, the rate of
evolutionary emissions-
reduction technology advances
has been outpaced by that of
past and projected growth in
demand for air transportation,
as shown in Fig. 2. In order to
maintain the current level of
emissions into the future, it
can be deduced from the
figure that fuel consumption
per revenue-passenger-
kilometer would have to
reduced by half.4 Another
source of concern is high-
altitude aircraft emissions.
Although regulations that
restrict harmful aircraft
emissions at local levels
(landing and takeoff altitudes
below 3000 ft), such as
hazardous air pollutants,
oxides of nitrogen (NOX),
carbon monoxide (CO),
aerosols, unburnt particulates
and hydrocarbons (HC), have
been in place since 1973 by
the U.S. Environmental
Protection Agency (EPA),8 a
movement to establish
regulations against aircraft
emissions near the
stratosphere did not exist until
recently. Better understanding
of the global effects of CO2,
water vapor, and NOX, have
prompted the ICAO to consider various penalization options that include an emission trading system, emission
related-levies, and voluntary measures for high-altitude emissions since 2001.5 There is evidence that CO2, water
vapor (through the contrails it induces), and NOX emissions, even in small quantities, can significantly affect global
ozone depletion and climate changes.9 Therefore, it seems that continued growth of aviation cannot be guaranteed
unless a very significant reduction or even a complete elimination of high-altitude aircraft emissions is achieved
through technological breakthroughs.
At the time of this writing, the sudden rise in fuel costs has forced the third10
and fourth11
largest airlines of the
United States to file for bankruptcy. It is still premature to conclude whether this sudden climb to record-high prices
indicates the world has now reached peak-oil,12
the critical point at which the scarcity of oil starts becoming a
legitimate concern. With or without oil peaking, a continual rise towards demand-driven, higher fuel prices allow
alternative fuels or energy technologies to penetrate the energy market more easily, especially when consuming
American Institute of Aeronautics and Astronautics
3
nations have the desire to decrease their dependence on foreign oil. The recent commercial success of hybrid electric
vehicles demonstrates that the automotive industry’s investment in alternative energy technologies is accelerating
the pace of development. For example, a seven-fold increase in fuel cell power-density has been achieved in the
later half of the 1990's.13
Since then an additional three-fold increase has been reported,14
while improvements in the
fuel consumption and efficiency of conventional gas turbine engines have remained relatively constant during the
same time period.9 It is generally expected that aviation will follow the lead of other industries in adopting
alternative energy technologies as the maturity and economic viability of those technologies improve.15
Latest research efforts towards electrically powered and propelled aircraft suggest that a revolution in aircraft
propulsion system technologies is viewed as the most promising approach of addressing the issues discussed above.
Key technologies, such as fuel cells, electric motors, batteries, power electronics, etc., are especially attractive since
they allow the inter-connected issues of noise, emission and energy to be addressed simultaneously. A revolutionary,
electric aircraft propulsion system would virtually eliminate the source of engine core noise, potentially produce
zero emissions (excluding water), and operate at higher efficiencies due to its independence from the Carnot
limitation.
Such appealing potentials of electric propulsion technologies are currently receiving an unprecedented amount of
interest from all sectors of the aerospace industry. By 2009, NASA aims to “develop and validate technologies that
would enable a 10-decibel reduction in aviation noise (from the level of 1997 subsonic aircraft)” and “flight
demonstrate an aircraft that produces no CO2 or NOx to reduce smog and lower atmospheric ozone” by 2010.2 The
latter project, as known as the “Zero Emissions Demonstration” specifies that the “revolutionary zero emissions
aircraft” will be “a hydrogen powered fuel-cell aircraft with cryogenic electronic motors embedded in the wings.”
These goals are consistent with earlier findings by internal NASA feasibility studies of fuel-cell based electric
aircraft concepts during the 1998-2004 time-frame.
Furthermore, the Boeing Company’s More Electric Aircraft (MEA) initiative16
and the U.S. Naval Air System
Command's recent collaboration with Georgia Tech17
are examples of top down approaches that aim to accelerate
the transition to an electric propulsion paradigm. Bottom-up approaches that aim to prototype electrically powered
air-vehicles as technology demonstrators are being undertaken by AeroVironment,18
Aviation Tomorrow,19
and the
Georgia Tech Research Institute (GTRI).20
III. Technical Challenges
The revolutionary paradigm shift towards electric flight inevitably introduces new technical challenges to the
field of aircraft design. These challenges with respect to electric aircraft concepts can largely be categorized as the
lack of analysis capability both at the propulsion design level and vehicle design level, lack of methodological
capability to adequately link the two levels, and introduction of new sources of uncertainty.
At the propulsion design level, there is currently a lack of trusted and validated analysis capability for estimating
the performance and size of electric aircraft propulsion systems. Designers of conventional aircraft engines have
access to legacy tools, such as NASA Engine Performance Program21
(NEPP) and Weight Analysis of Turbine
Engines22
(WATE), which contain several decades of engine development experience and knowledge. Moreover,
ample historical data on variety of engine types and cycles allow the designers to estimate a notional engine’s
performance characteristics, such as thrust lapse, power lapse, and fuel consumption behavior, as well as scaling
laws without having to track every detail of the energy conversion process inside the engine. Such a level of
expertise in analysis capability does not yet exist for electric aircraft propulsion systems, whose enabling
technologies and historical database are still evolving rapidly.
A similar challenge exists with traditional aircraft sizing and synthesis tools at the vehicle design level. Legacy
vehicle sizing codes, such as FLight OPtimization System23
(FLOPS) and AirCraft SYNThesis24
(ACSYNT) are
structured around the principle assumption that an aircraft’s weight decreases with flight. This assumption that the
rate of change in aircraft weight equals the fuel flow rate can become invalid for a number of possible revolutionary
aircraft concepts. For example, the weight of an aircraft such as AeroVironment’s Helios,25
designed for perpetual
flight, remains constant due to the regenerative propulsion ideology. Another revolutionary aerospace concept is a
zero-emissions or emissionless aircraft that stores all harmful by-products of the propulsion system’s energy
conversion process, thereby gaining weight as fuel is consumed. The combined lack of available propulsion system
design tools and aircraft sizing capability acts as a significant hindrance to the early phases of aircraft design when
the exploitation of sizing and synthesis tools and empirical knowledge is utilized extensively.
The lack of methodological capability to link the propulsion and vehicle level analyses can result from various
types of couplings, which do not exist for conventional architectures, between the two levels. The traditional method
of interfacing the propulsion design level and vehicle design level through an engine deck and an engine scaling law
American Institute of Aeronautics and Astronautics
4
is appropriate only when there are negligible couplings between the elements of the propulsion system and those of
the vehicle level analyses. Nevertheless, certain emerging electric aircraft propulsion concepts introduce more
ambiguous boundaries between the airframe, propulsion system, and energy storage. Examples include an ambient
energy-harvesting aircraft26
whose energy receiving components cannot be sized without considering the airframe’s
geometry and a structurally integrated, fuel cell wing concept.15
Furthermore, a coupling between the propulsion
mechanism and mission is possible, as is the case of solar powered, high altitude, long endurance concepts.27
The new sources of uncertainty introduced by electric aircraft are regulatory uncertainty and uncertainty of
revolutionary technologies. Regulatory uncertainty refers to the risks inherent in obtaining any necessary licenses to
construct or operate the project from the appropriate regulatory authority. Regulatory requirements include safety
regulations, environmental regulations, maintenance regulations, etc. Historically, regulatory uncertainty has created
significant consternation to airframe and engine designers because any future updates to those standards could
possibly force a premature retirement of the product. Although the transition towards electric aircraft technologies is
motivated in part by the uncertainty in future noise and emission regulations, this new breed of aircraft cannot be
immune to regulatory uncertainty forever. In light of revolutionary technology improvements, existing regulatory
requirements regarding safety, emission, noise, etc. are likely to be re-examined and modified accordingly. A likely
candidate may be a completely new one-engine-out performance requirement at take-off for electric aircraft
propulsion systems. In short, revolutionary aircraft may be just as likely to be forced into early retirement as
conventional aircraft unless possible scenarios for new standards that are more tailored towards electric aircraft (e.g.,
zero water emissions above 30,000 ft) are adequately accounted for in advance.
“For complex systems, the search for feasible and viable solutions often requires the application of multiple new
technologies.”28
Infusing technologies generally incur penalties in other disciplines as the “price” of the benefits.
The impact of a technology, the “benefit” and the “price” cannot be precisely predicted at the conceptual design
phase, particularly if the technology is ranked at a low technology readiness level (TRL) and if its impact propagates
through many disciplines, infusing the technology introduces a significant source of design uncertainty.
IV. Research Strategy
In order to overcome the technical challenges that lay ahead, it seems desirable for the aerospace community as a
whole to vigorously implement a coordinated strategy of addressing the challenges through increased levels of inter-
agency collaborations. Many difficult decisions will need to be made so that realistic technology milestones and
proper allocation of resources can be established to reduce the risk, uncertainty, and lengthy development-to-
adaptation times of revolutionary electric propulsion and aircraft technologies. In making those kinds of large-scale
decisions, a quantitative assessment environment for electric propulsion and aircraft architectures that provides
insight into the system-wide responses of alternative technology and policy evolution scenarios would be valuable,
thereby increasing the possibility of arriving at more informed decisions.
Figure 3 notionally represents the authors’ vision of such a design and decision-support environment, which
consists of three inter-connected layers: analysis, design, and decision-support. At the core of this environment are
advanced design methods capable of properly synthesizing and sizing revolutionary aircraft with revolutionary
propulsion architectures. Supporting these methods are physics-based analysis tools or models for each electric
technology component. The top-most layer is where various decision-support or assessment techniques – supported
by the advanced design methods – take place to guide the highest-level decision makers and thus facilitate the
overall decision-making process.
Currently, research efforts on novel synthesis and analysis methods for revolutionary electric propulsion and
aircraft architectures are on-going at Georgia Tech and its collaborative partners, Florida A & M University
(FAMU), the Ohio State University (OSU) and GTRI, as part of a broad-scale aeropropulsion and power
technologies project sponsored by NASA and the Department of Defense (DoD).29
The following sections present
summaries of progress made in the two key research thrusts towards the creation of the environment: electric
propulsion modeling and synthesis, and generalized, energy-based aircraft sizing method.
American Institute of Aeronautics and Astronautics
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V. Electric Propulsion Modeling and Synthesis
The research objectives behind the thrust area of electric propulsion modeling and synthesis are twofold. First,
the modeling aspect of this work aims to investigate and then extract the fundamental physical principles and sizing
relationships of each enabling technology as useful mathematical equations. The intention is to address the lack of
analysis capability identified in the previous section by accumulating a physics-based analysis and sizing capability
that will constitute the bottom layer of Fig. 3
Second, the synthesis aspect of the research thrust endeavors to create a top-level design framework for electric
aircraft propulsion systems or architectures. Such a framework can further serve to address the lack of analysis
capability by allowing the estimation of a notional electric aircraft propulsion architecture’s thrust behaviors, fuel
consumption trends, weight, and volume. The framework also serves as a learning ground for identifying any
methodological improvement opportunities in the field of aircraft propulsion system design, thus addressing the lack
of methodological capability discussed in the previous section by contributing to the core layer – advanced design
methods – shown in Fig. 3.
A. Technical Approach
Electric aircraft propulsion architecture can be characterized by the following key technology components:
power generation device or source, balance-of-plant (BOP), power management and distribution (PMAD) system,
transducer, and propulsor. In order to estimate the performance and scale of these individual technologies, research
work on the following representative models have been completed.
The recent popularity of fuel cells as the primary power source in vehicular applications prompted the
development of a generic, steady-state fuel cell model, which takes into account the three main sources of loss
(activation over-voltage, ohmic loss, and mass transport loss) at the single-cell level.30
Furthermore, a fuel cell stack
sizing model that computes the stack weight and volume of through a handful of geometric and material properties,
such as the number of cells, cell active area, thickness of each element, and either the density or loading (mass per
unit area) of each element, was developed.
All required models of ancillary components, such as a compressor, humidifier/intercooler, heat exchanger,
pumps, etc., that are separate from the primary power generation device are collectively known as the BOP. For the
compressor, an in-house analysis capability was chosen for its model. CMGEN is a parametric compressor map
generating program developed under the supervision of NASA Lewis Research Center.31
The code is capable of
Advanced Design Method
Electric
Propulsion
System Sizing
and Analysis
Energy Based
Aircraft Sizing
Volumetric
Sizing
Design and Decision-Support Environmentfor Electric Aircraft Concepts
Physics-Based Analysis
Hydrogen Tanks
Electric Battery
Fuel Cells
Electric Motor
HTS Motor
Cooling System
Compressor
Humidifier
Water Management
Power Management
Decision-support or assessment techniques
Energy Transformation
Decision-Support Methodology
Control and Conditioning
Energy Storage
Advanced Design Method
Electric
Propulsion
System Sizing
and Analysis
Energy Based
Aircraft Sizing
Volumetric
Sizing
Design and Decision-Support Environmentfor Electric Aircraft Concepts
Physics-Based Analysis
Hydrogen Tanks
Electric Battery
Fuel Cells
Electric Motor
HTS Motor
Cooling System
Compressor
Humidifier
Water Management
Power Management
Decision-support or assessment techniques
Energy Transformation
Decision-Support Methodology
Control and Conditioning
Energy Storage
Figure 3. Vision of Novel Design and Decision-Support Environment
American Institute of Aeronautics and Astronautics
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creating compressor maps for four different types of compressor configurations with design point pressure ratios
between 1.2 and 24. A separate model was developed for carrying out the balance-of-plant analyses such as water
and thermal management. The power electronics for the PMAD system are currently modeled in less detail than
other components through their representative efficiency values. Thus far, a simplified sizing approach that
estimates the weight of the BOP and PMAD system via characteristic specific power (throughput power per unit
mass) has been implemented.
A generic electric motor model which follows the method of Larminie32
to approximate a two dimensional motor
efficiency map is utilized as the transducer model. Instead of assigning a single value for a motor’s specific power
and making an equally speculative assumption about its power density (throughput power per unit volume), a
regression-based approach was adopted in sizing electric motors.
GTPROP, an in-house code originally developed and validated by Hamilton Standard33
was chosen as the
propulsor design / analysis tool. This FORTRAN-based model outputs all the typical performance metrics (e.g.,
torque, shaft rotational speed, propeller efficiency, and brake shaft horsepower as well as optimum blade geometry,
cost, and weight.
The top-level design framework for electric aircraft propulsion architectures was created using Phoenix
Integration's ModelCenter.34
The process integration software package allows analysis models that do not share a
native platform to be seamlessly synthesized under a cohesive simulation environment. Therefore, some of the more
recently developed in-house battery and hydrogen storage models or the superconducting motor and cryogenic
cooler models developed at FAMU can be easily incorporated to the synthesis framework as they become necessary.
B. Formulation: Fuel-cell-based Electric Aircraft Propulsion Architecture
As an initial investigative step towards identifying those areas in which new methodological contributions can be
made, a design approach that emulates the traditional engine design process of on and off design analyses was
formulated for a notional fuel-cell-based electric aircraft propulsion architecture shown in Fig. 4.
Main Motor
PMAD
C. MotorC.
Humidifier/
intercooler
Hyd
roge
n
Ta
nk
Air
H2O
HX
H2
Inverter/Controller
P. MotorPump
PEMFCStack
Cathode
Anode
P. Motor PumpAir
Mechanical Power Transmission
Electrical Power Transmission
Cold Coolant Flow
Hot Coolant Flow
Main Motor
PMAD
C. MotorC.
Humidifier/
intercooler
Hyd
roge
n
Ta
nk
Air
H2O
HX
H2
Inverter/Controller
P. MotorPumpPump
PEMFCStack
Cathode
Anode
PEMFCStack
Cathode
Anode
P. Motor PumpAir
Mechanical Power Transmission
Electrical Power Transmission
Cold Coolant Flow
Hot Coolant Flow
Mechanical Power Transmission
Electrical Power Transmission
Cold Coolant Flow
Hot Coolant Flow
Figure 4. Schematic of Notional Fuel-cell-based Electric Aircraft Propulsion Architecture
The goal of on-design analysis is to parametrically design and size a propulsion architecture to a single
propulsive power requirement defined by uninstalled reference thrust (FR), free-stream Mach number (M∞), and
altitude (h). For fuel-cell-based propulsion architectures, the required sequence must be a power matching process
that balances the steady-state output power from the fuel cell stack (Pe) with the sum total of the power draws of all
electrical loads while considering the various sources of loss into account. Thus, as shown in Fig. 5, the on-design
analysis process begins with the propeller and electric motor models that output shaft power (PSH) and motor
efficiency (ηmm), both of which are needed to find the power balance at the given reference point. An iterative
solution scheme that employs an optimizer is necessary for the remainder of the analysis process due to the inherent
coupling between the compressor, heat exchanger and the fuel cell stack. The optimizer iterates on current density
(i) and Pe until the steady-state stack output power equals the combined power draws of the propulsion motor (PSH),
compressor motor (Pc), and heat exchanger (Phx), plus sufficient margin to compensate for the power losses through
the fuel cells (ηfc), motors (ηmm, ηcm), compressor (ηc), heat exchanger (ηhx), and PMAD (ηs,ηdcdc) system.
Constraints g1, g2, and g3 ensure that the water management of PEM fuel cells is properly done, the current balance
of the architecture is satisfied in addition to the power balance, and the stack voltage (Vstack) stays within the
American Institute of Aeronautics and Astronautics
7
prescribed lower and upper limits, respectively. The parentheses around g2 and g3 indicate that these constraints are
optional. Lastly, the weight breakdown of the architecture is obtained by aggregating the results of the component-
level sizing models that estimate the weight of the propeller, propulsion motor, compressor motor, fuel cell stack,
PMAD system, and BOP.
Atmosphere
Model
Propeller
Main
Electric
Motor
PEM
Fuel Cell
Compressor
& Aux.
Motor
BOP
Optimizer Optimizer
Reference
Point
Power
Balance
EndEnd
Heat
Exchanger
h ∞MFR ,
sQ,
mmSHP η,
Output File
atmatmatm PT φ,,atmatm PT , a,∞ρ
i
infc PV ,
cP
hxP
eP
Minimize ε≤− )'
1(e
e
P
P
):(
)0)(:(
3
2
UBVLBg
IIIIg
stack
hxcmmme
≤≤
≥++−
'eP
infc PV ,
cη
fcV
0':1 ≥−ψψg
'eP
Figure 5. On-Design Analysis Sequence
Atmosphere
Model
Motor-Prop
Equilibrium
Off-Design
Point
“Engine
Deck”
On-Design
Results
h PPSM ,∞
atmTa,,∞ρ
mmSHP η, F
PPShM ,,∞
Power Draw
Current Draw
Specific Fuel Consumption
Component Efficiencies
Minimize
SHP
ε≤− )'
1(e
e
P
P
PEM
Fuel Cell
Compressor
Map & Aux.
Motor
BOP
Optimizer
Power
Balance
Heat
Exchanger
'eP
infc PV ,
cP
hxP
eP
):(
)0)(:(
3
2
UBVLBg
IIIIg
stack
hxcmmme
≤≤
≥++−
'eP
infc PV ,
cη
fcV
0':1 ≥−ψψg
'eP
g4 : hardware limitations
atmatmatm PT φ,,atmatm PT ,
AN ,
Atmosphere
Model
Motor-Prop
Equilibrium
Off-Design
Point
“Engine
Deck”
“Engine
Deck”
On-Design
Results
h PPSM ,∞
atmTa,,∞ρ
mmSHP η, F
PPShM ,,∞
Power Draw
Current Draw
Specific Fuel Consumption
Component Efficiencies
Minimize
SHP
ε≤− )'
1(e
e
P
P
PEM
Fuel Cell
Compressor
Map & Aux.
Motor
BOP
Optimizer Optimizer
Power
Balance
Heat
Exchanger
'eP
infc PV ,
cP
hxP
eP
):(
)0)(:(
3
2
UBVLBg
IIIIg
stack
hxcmmme
≤≤
≥++−
'eP
infc PV ,
cη
fcV
0':1 ≥−ψψg
'eP
g4 : hardware limitations
atmatmatm PT φ,,atmatm PT ,
AN ,
Figure 6. Off-Design Analysis Sequence
The performance characteristics of the sized architecture can be evaluated at a wide range of off-design points,
i.e., operating conditions beside the design point, once the on-design analysis is complete. The objective of off-
design analysis is to discover the operational envelope of the propulsion architecture. This is why in Fig. 6, thrust is
shown as a fall-out, and shaft power is instead shown as a given. The variable partial power (throttle) setting, PPS,
controls at what fraction of the maximum continuous output power the propulsion motor operates for a given off-
design point. Again, a constrained optimization approach is taken to iteratively solve for the power balance. Because
the size of the fuel cell stack is known from on-design analysis, the optimizer only needs to iterate on Pe until
convergence is achieved per the power balance, while satisfying all constraints (g1 - g4). Having hardware
limitations, g4, as a constraint prevents a designer from wrongfully concluding component-damaging off-design
points (e.g., motor's maximum speed is exceeded) as feasible operating conditions. The sequence of Fig. 6 can be
American Institute of Aeronautics and Astronautics
8
repeated for a number of off-design points and throttle settings to map out the feasible performance envelope of the
fuel-cell-based electric aircraft propulsion architecture.
C. Preliminary Investigation: Electric Aircraft Propulsion for General Aviation
A prior study by Choi, Soban, and Mavris reports on the application of the above on and off-design formulation
to a General Aviation (GA) class propulsive power requirement.35
The off-design analysis results of a fuel-cell-
based electric aircraft propulsion architecture, initially designed to a General Aviation (GA) class propulsive power
requirement of 138. 4 horsepower (FR of 300.286 pounds at M∞ of 0.1952 and h of 8000 feet), reveal performance
trends that highlight the advantages of electric propulsion over conventional, air-breathing propulsion. Figure 7
shows the lack of power-lapse, which is undesirable but nonetheless unavoidable for internally reciprocating engines.
Furthermore, Fig. 8 illustrates that electric propulsion allows component-level efficiencies to be much greater and
more robust against different flight regimes than a conventional propulsion method constrained by the Carnot limit.
0
100
200
300
400
500
600
700
0 0.02 0.04 0.06 0.08 0.1 0.12 0.14 0.16 0.18 0.2
Mach
Th
rus
t (l
bf)
h=0 ft
h=4000 ft
h=8000 ft
h=12000 ft
0.116
0.118
0.12
0.122
0.124
0.126
0.128
0.13
0 0.02 0.04 0.06 0.08 0.1 0.12 0.14 0.16 0.18 0.2
Mach
SF
C (
lb/h
p/h
r)
h=0 ft h=4000 ft h=8000 ft h=12000 ft
0
100
200
300
400
500
600
700
0 0.02 0.04 0.06 0.08 0.1 0.12 0.14 0.16 0.18 0.2
Mach
Th
rus
t (l
bf)
h=0 ft
h=4000 ft
h=8000 ft
h=12000 ft
0.116
0.118
0.12
0.122
0.124
0.126
0.128
0.13
0 0.02 0.04 0.06 0.08 0.1 0.12 0.14 0.16 0.18 0.2
Mach
SF
C (
lb/h
p/h
r)
h=0 ft h=4000 ft h=8000 ft h=12000 ft
Figure 7. Thrust and Fuel Cell Consumption Behaviors at Full Throttle
0
0.1
0.2
0.3
0.4
0.5
0.6
0.7
0.8
0.9
1
0 0.02 0.04 0.06 0.08 0.1 0.12 0.14 0.16
Mach
Eff
icie
ncy
Main Motor
Compressor Motor
Compressor
Propeller
0
0.1
0.2
0.3
0.4
0.5
0.6
0.7
0.8
0.9
1
0 0.02 0.04 0.06 0.08 0.1 0.12 0.14 0.16
Mach
Eff
icie
ncy
Main Motor
Compressor Motor
Compressor
Propeller0
0.1
0.2
0.3
0.4
0.5
0.6
0.7
0.8
0.9
1
0 0.02 0.04 0.06 0.08 0.1 0.12 0.14 0.16
Mach
Eff
icie
ncy
Main Motor
Compressor Motor
Compressor
Propeller
0
0.1
0.2
0.3
0.4
0.5
0.6
0.7
0.8
0.9
1
0 0.02 0.04 0.06 0.08 0.1 0.12 0.14 0.16
Mach
Eff
icie
ncy
Main Motor
Compressor Motor
Compressor
Propeller
Figure 8. Variations of Component-level Efficiencies at Full Throttle with Mach Number at 0 ft and 8000 ft
D. Work-in-progress
Continuing research in this thrust area is expected to further enhance the modeling and synthesis capability for
electric aircraft propulsion architectures. The ongoing modeling efforts, including work on the High Temperature
Superconductor (HTS) motor36
and more physics-based sizing relationships for other types of transducers and power
electronics,37
are already showing a great deal of promise. Currently, the synthesis aspect of the research thrust is
interested in examining the coupling between the propulsion-level and vehicle-level analyses as a potential area for
identifying new methodological improvement opportunities.
VI. Generalized, Energy-based Aircraft Sizing Method
Aircraft sizing is a critical aspect of system-level study because the aircraft sizing process is a prerequisite task
of most design and analysis activities, including internal layout, cost analysis, and system effectiveness analysis.
American Institute of Aeronautics and Astronautics
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However, traditional aircraft sizing methods are significantly specialized for aircraft powered by internal
combustion engines. Electric propulsion concepts produce propulsive thrust through their energy conversion process,
completely different from the way conventional internal combustion engines produce power, which prevents
designers from applying traditional aircraft sizing methods to the design of electric propulsion aircraft and
necessitates a novel aircraft sizing method.
A. Technical Approach
Efforts to develop such a method have already been initiated. Jonathan R. Smith et al. developed an electric
aircraft sizing method suitable for battery powered electric aircraft.38
Harmats and Weihs proposed a sizing method
for a high-altitude long-endurance remotely piloted vehicle powered by a hybrid propulsion system combining solar
and internal-combustion based on a cohesive mathematical formulation.39
However, such previous research for the
development of electric aircraft sizing method was limited to certain types of electric propulsion system and
missions. In contrast, the authors have continued to hold a more general approach, since different motivations
toward electric aircraft have yielded various concepts of the applications of electric or electromagnetic devices into
aero-propulsion systems. In addition, the authors envision a method is applicable for a wide range of applications
beyond electric propulsion aircraft.
The formulation of the method was initiated by identifying shortcomings of traditional aircraft sizing methods.
First, certain types of electric propulsion systems are not compatible with the traditional way of synthesizing
propulsion system design into aircraft sizing. A propulsion system is a device that produces propulsive thrust
through a series of energy conversions. This process is typically affected by the design parameters of the system and
operating conditions such as Mach number, altitude, and ambient temperature. Because the technology behind the
traditional air-breathing combustion engines has matured, thrust lapses and specific fuel consumption (SFC)
behaviors of traditional air-breathing combustion engines are well understood. Thus, aircraft design engineers need
not track every detail of the internal energy conversion process inside the engine to size aircraft. Instead, most
traditional aircraft sizing methods generally require thrust and SFC data, the resultant quantities of the involved
energy conversion process. This information can be obtained directly from either engine companies or well-
established historical data and can also be created by engine performance/weight analysis codes such as NEPP.
However, such conventional forms of data may not be available for some of the emerging electric propulsion
systems due to their substantial differences in their energy conversion processes. For instance, one would not try to
construct SFC data for an electric battery-powered or solar-powered aircraft. In addition, some of the electric
propulsion systems requires component-level sizing in conjunction with aircraft level sizing. For example, the size
of energy receiving or collecting components of beam-powered, solar-powered, or ambient energy-harvesting
aircraft is generally coupled with aircraft geometry. Furthermore, in contrast to conventional aero-propulsion
systems, emerging electric propulsion systems rapidly evolve, continuously incorporating new material, innovative
components, and cutting-edge technology. Thus, a comprehensive electric propulsion system analysis tool has not
yet been developed. All of the above discussion indicates that traditional interfaces between the propulsion system
sizing and the aircraft sizing, thrust lapses and specific fuel consumption (SFC) behaviors, as well as engine scaling
laws, will not fit in electric propulsion aircraft sizing. This technical issue can be solved only by integrating
propulsion system sizing into the aircraft sizing process. The proposed method attempts to resolve this technical
issue by modeling a propulsion system as a series of energy conversion devices and sizing them within the aircraft
sizing process.
Another deficiency of the traditional sizing method is its inflexibility in dealing with combined propulsion
systems or hybrid propulsion systems. Most existing aircraft are equipped with a single engine or multiple identical
engines. Therefore, the thrust available and fuel consumption for most conventional aircraft can be established by
one engine deck. However, if different types of propulsion systems and energy sources are equipped, and the power
contribution rate of each propulsion system varies with flight conditions, then the traditional sizing formulation
cannot handle this situation properly. Unlike internal combustion engines, electric propulsion systems will take
advantage of integrating a combination of different types of power devices into single propulsion system thanks to
the versatility of electric power. For example, the combination of high specific-power devices such as lithium-
polymer battery and ultra-capacitors and high specific-energy devices such as fuel cells may provide an optimum
solution for an electric aircraft whose power profile has high peaks for short periods. As a means to analyze such a
hybrid or heterogeneous power-generation system, the energy based sizing method introduces a concept of “power
path,” which consists of power devices along the same stream of energy conversion.
In order to fly the mission, an aircraft must store sufficient energy onboard. Conventional stored energy is
hydrocarbon fuel, whose weight is consumable during flight. However, energy sources for electric aircraft such as
an electric battery and nuclear fission cells maintain virtually constant weight during energy conversion processes,
American Institute of Aeronautics and Astronautics
10
which introduces another complication to aircraft sizing. Particularly when the propulsion system has both energy
sources: one whose weight is consumable and the other whose weight is non-consumable, the traditional aircraft
sizing method is not able to correctly estimate weight variation. In the energy based sizing method, such an
unconventional energy source is defined as non-consumable energy, and it is treated separately in the development
of the formulation.
The most widely used mission analysis technique for conventional aircraft is based on the assumption that the
rate of change in aircraft weight equals the fuel flow, which leads to a historical equation, the Breguet range
equation. However, aircraft such as the Helios, which is equipped with regenerative power systems, maintains the
same weight during the entire mission. Furthermore, more stringent emissions regulations of the future may force
aerospace engineers to innovate propulsion systems so that the system can separate specific by-product components
from engine emissions and store them onboard during flight. For example, zero-emission aircraft, which take in
external air to oxidize hydrogen fuel and stores water onboard, will gain weight as fuel burns. This behavior cannot
be analyzed by the traditional mission analysis based on Breguet range equation. Therefore, more generalized
weight decomposition and weight differential equations are implemented in the new sizing method.
B. Formulation Aircraft sizing is an analytical process that determines the best combination of two scales of a baseline
configuration, a geometric scale that is dictated by the wing area and a propulsive scale that is dictated by the
amount of thrust of the engine by establishing two balances: power balance and energy balance, which are achieved
by the constraint analysis and the mission analysis respectively. Power balance is referred to as matching the
available power, Pavailable, to the required power, Prequired. Similarly, energy matching is referred to as matching the
available energy, Eavailable to the required energy, Erequired. Then, the most fundamental equations of aircraft sizing are
given as:
requiredavailable PP = (1)
requiredavailable EE = (2)
Table 1. Comparison of Traditional Formulation and Proposed Formulation for Aircraft Mission Analysis
and Weight Estimation
−−−−====
−−−−
−−−−
)1(
)(
)(
)1()(
1s
s
s
s
TO
s
CE
W
W
kW
W β
)()()1()(
s
CE
sss
CE
W
WΞΥ−−−−==== β
(((( )))))()()(
)1(
)(
exp s
CE
ss
s
s
kW
WΞΥ−−−−====
−−−−
PFE WWWW ++=
FdWdW =CEkdWdW =
TO
F
PTO
W
W
WW
−−−−−−−−
====
Γ1
Old FormulationOld Formulation New FormulationNew Formulation
ENEPROPEE WWWWW δ+++′=RCEPE WWWWW +++=
TOF WW )1)(1( Π−−−−++++==== ε
NECE
PTO
WW
ΩΩΦ∆Γ −−−−−−−−−−−−−−−−′′′′−−−−====
1
0=kWhen
0≠kWhen
∑=
+=Ωm
s TO
s
CECECE
W
Wwhere
1
)(
)1( εTOCECE WW Ω=
∑=
− ΞΥ+=Ωm
s
s
NE
ss
NENEwhere1
)()()1()1( βεTONENE WW Ω=
The available power is determined by the maximum power produced by the propulsion system. The required
power is dictated by the point performance requirements that specify the ability of performing specific maneuvering
motions such as take-off, climb, sustained turn, instantaneous turn, acceleration, cruise, approach, and landing.
Combining Eq. (1) and Newton’s Second Law yields the following power constraint equation for each power path,
American Institute of Aeronautics and Astronautics
11
which describes the ratio of power to the take-off gross weight as a function of wing loading for given aerodynamic
properties and the rate of change in energy height. Reference 40 presents the detailed derivation process.
Vg
Vh
dt
d
VqS
RC
S
W
q
nK
S
W
q
nK
W
qS
W
p
o
DTOTO
TOi
i
TO
i
oSL
o
i
++
++
+
Π=
2
1 2
2
2
1
ββ
βα
βτ
η
(3)
An individual constraint equation for each performance requirement can be derived from Eq. (3). A set of
constraints equation as a function of the ratio of power to the take-off gross weight and wing loading create a
feasible solution area in which the proper combination can be found.
The available energy is the amount energy from energy sources including onboard sources (stored energy) or
outboard sources (transmitted or harvested energy). The required energy is dictated by the mission performance
requirements that specify the ability of performing a series of motions. The required energy is estimated by
summing the required energy for all mission segments. The way of calculating required energy for each mission
segment differs depending on type of energy source: consumable energy or non-consumable. In addition,
generalized weight decomposition and weight differential equation are implemented in the estimation of energy
weight and aircraft weight. Comparison between traditional formulation and new formulation are provided in .
Reference 40 lists the detail process of the derivation and notation of symbols.
The new formulation is general, so it can be applicable to a wide range of unconventionally powered aircraft as
well as electrically powered aircraft. If one applies the new formulation to the design of conventional aircraft, the
associated equations will reduce to those of the traditional formulation, which indicates that the proposed method is
an extension and generalization of the traditional method.
C. Preliminary Investigation: High Speed Electric High Altitude Long Endurance UAV
As a proof of concept, the formulation was applied to a Global Hawk-like HALE configuration powered by an
all-electric propulsion system.40
The aircraft performs surveillance missions for 24 hours at a station located about
3,000 nm away from the base. It carries 1,900 lbs. of payload, including synthetic aperture radar, a digital charge-
coupled device (CCD) camera, and a third-generation infrared sensor system. The mission also requires high-power
extraction (40 Hp at 65,000 feet) to drive the electric payloads. The electric HALE aircraft is powered by multiple
propellers driven by electric motors mounted on the bottom of the wing. The electric propulsion system consists of
proton exchange membrane (PEM) fuel cells, a power management and distribution system (PMAD), main motors,
and other accessories. The aircraft is sized twice: first with off-shelf technology and second with advanced
technology. The advanced technology portfolio includes high temperature super-conducting motor, 10%
improvement in fuel cell efficiency, 50% improvement in fuel cell specific power, and 38% PMAD weight
reduction.
The results of constraint analysis are depicted in Fig. 9. The design point for the aircraft with the off-the-shelf
technology is selected as 40 lbs/ft2 of wing loading and 57 Hp/lbs of power-to-weight ratio. Wing loading of the
aircraft with the advanced technology remains the same and the power-to-weight ratio decreases to 55 Hp/lbs,
thanks to the benefits of increased efficiency of fuel cells and the motors.
Advanced TechnologyOff-the-self technology
Take-off Roll
Cruise
Climb
Approach
Design Point
Take-off Roll
Cruise
Climb
Approach
Design Point 20.0
40.0
60.0
80.0
100.0
120.0
140.0
160.0
180.0
200.0
0 10 20 30 40 50 60 70
W ing Laoding
Po
SL/W
20.0
40.0
60.0
80.0
100.0
120.0
140.0
160.0
180.0
200.0
0 10 20 30 40 50 60 70
W ing Laoding
Po
SL/WW/S =40
PoSL/W=57W/S =40PoSL/W=55
Figure 9. Constraint Analyses Results
American Institute of Aeronautics and Astronautics
12
The weight comparison and detail weight breakdown of electric propulsion systems are shown in Fig. 10. The
sized aircraft with off-the-shelf technology is approximately three times as heavy as the Global Hawk. Since the
aircraft weight is far beyond the reliable range of the regression equation, these weight values are not reliable.
However, the results provide sufficient information to prove that the off-the-shelf technology has not yet matured
enough to the point powering high-speed HALE aircraft. However, infusing advanced technology can reduce the
aircraft weight significantly. This dramatic reduction in aircraft weight is well beyond the level of components
improvement. The reason for such substantial improvement is that the impact of component level technology on
aircraft sizing is recursive and cumulative. In the case of this study, if motor weight reduces, aircraft weight reduces,
and thus, the required thrust and wing area must reduce, which in turn reduce the motor size. This chain of impact
propagation will continue until a thrust balance and fuel balance are achieved. Therefore, evaluation of new
technology must be based on the propagated impact on aircraft-level design.
0
10000
20000
30000
40000
50000
60000
70000
80000
90000
Global Hawk Off-the-shelf Advanced Technology
We
igh
t (l
bs
)
Payload
Wf
Airfram e and Subsys tem s
H2 Tanker
Propuls ion
Off-the-shelf
Advanced Technology
Weight Breakdowns of Propulsion Systems
Motor, 350, 10%
PMAD, 1037,
28%
Humidif ier and
Intercooler,
326, 9%
Compressor,
245, 7% Fuel Cell, 1283,
34%
Prop, 431, 12%
Motor, 5245,
22%
PMAD, 6073,
26%
Humidif ier and
Intercooler,
1380, 6%
Compressor,
1036, 4%Fuel Cell, 8140,
36%
Prop, 1478, 6%
Off-the-shelf
Advanced Technology
Weight Breakdowns of Propulsion Systems
Motor, 350, 10%
PMAD, 1037,
28%
Humidif ier and
Intercooler,
326, 9%
Compressor,
245, 7% Fuel Cell, 1283,
34%
Prop, 431, 12%
Motor, 5245,
22%
PMAD, 6073,
26%
Humidif ier and
Intercooler,
1380, 6%
Compressor,
1036, 4%Fuel Cell, 8140,
36%
Prop, 1478, 6%
Figure 10. Comparison of Weight Breakdowns
D. Comprehensive Sizing Environment
The proposed method mostly concerns balancing power and energy. However, a successfully sized configuration
must achieve one more criterion: a balance between the required aircraft volume and the available aircraft volume,
as illustrated in Fig. 11. In general, volume balance is verified through more detailed studies of the internal
arrangement after the initial aircraft configuration is fully established through the aircraft sizing process. In the case
of traditional aircraft design, however, the volume balance is implicitly secured to a certain degree via the
application of historical regression rules to weight estimation without the direct assessment of volume balance,
simply because all existing aircraft whose weight data are used to construct the regressed equations contain all
subsystems, structures, and fuel inside the aircraft. In addition, it is not too far-fetched to regard the aircraft as being
designed as a small perturbation from the historical trend.
Nevertheless, in the design of a revolutionary aircraft that uses unconventional propulsion systems and consumes
unconventional energy sources, such an implicit volume balance will not work. For instance, the converged
configurations in the sample sizing study, in which volume balance was not considered, may not provide sufficient
volume for all required systems. If they do not, the external configuration may be modified so that the efficiency of
internal packaging increases or the aircraft may need to be scaled up beyond the minimum size at which both power
balance and energy balance are achieved. Therefore, early consideration of the impact of unconventional energy
and/or the propulsion system on the required volume will be essential so that rework due to volume issues can be
avoided in the following design phase. All three criteria will not always play equal roles in the aircraft sizing process.
The relative importance of the three depends on various parameters such as sizing requirements, the shape of the
configuration, and the characteristics of the propulsion systems and energy sources, including the following four
parameters: the specific energy and energy density of the energy source, and the specific power and power density
of the propulsion system. For instance, if the aircraft is powered by an energy source whose specific energy is
American Institute of Aeronautics and Astronautics
13
General Aviation
Small
UAV
Range
Velocity Fighter
Airline Jet
Helios
General Aviation
Small
UAV
Range
Velocity Fighter
Airline Jet
Helios
Figure 12. Notional Mission Space Exploration
exceptionably higher than that of conventional fuel, the weight of the energy source may be trivial, and thus, energy
balance will not be a major concern in the aircraft sizing process any longer.
Propulsion
Weight
Aerodynamics
Notional
Concept
Volume
Mission Analysis
Constraints Analysis
Mission PerformanceRequirements
Point PerformanceRequirements
Volume Analysis
A/C Scaling Factor and
Subsystem Weight
Weight Estimation
Propulsion
Weight
Aerodynamics
Notional
Concept
Volume
Mission Analysis
Constraints Analysis
Mission PerformanceRequirements
Point PerformanceRequirements
Volume Analysis
A/C Scaling Factor and
Subsystem Weight
Weight Estimation
Figure 11. Notional Process of a Comprehensive Aircraft Sizing Method
VII. Discussions of Future Work
The emerging design and decision-support environment will enable one to size and optimize airframe and
electric propulsion system simultaneously, which leads a wide range of application to the research regarding
revolutionary propulsion aircraft including electrically powered aircraft. First of all, the environment may aid the
decision maker in selecting the best propulsion system architecture for a given mission. As various propulsion
devices are available, the selection of the correct propulsion system architecture becomes more challenging than
ever unless it is proved that one specific type of architecture substantially outperforms for most missions. The
architecture selection will require a wide range of decision making, including energy sources, power generation,
conversion, and means of producing thrust. For example, even if fuel cells are selected as the primary power source,
a variety options are still available: type of fuel cells (PEM or SoFC, etc.), type of fuel (methane, hydrogen, or
petroleum, etc.), type of fuel storage (pressurized gaseous tank, cryogenic liquid tank, etc.), and type of electric
motors (general AC motor, brushless DC motor, and superconducting motor, etc.). Some of the down-selections can
be made by simple quantitative decision support
analyses. Nevertheless, capability of quantitative
assessment for different types of propulsion
system architecture is crucial in the selection of
the best combination, particularly when one tries
to find a correct “power mix” for hybrid
propulsion aircraft.
As well as such cardinal research in which
specific mission identifies the best propulsion
system, explorative research in which specific
propulsion system expands available mission
space will be also of importance (Fig. 12).
Human technology history has proven an axiom
that needs promote new technology and, in turn,
the technology reveals new needs. As
American Institute of Aeronautics and Astronautics
14
revolutionary propulsion systems become tangible, the technology may open up new feasible and viable mission
spaces that the aerospace community has abandoned previously, fettered by ineluctable logic due to obvious
limitation of conventional propulsion systems. Furthermore, some electric propulsion systems may be more
attractive for such an unconventional mission. Therefore, evaluating various concepts for electric propulsion systems
must come along with exploring mission space in the light of their current capability as well as projected capability
in near/long term future. The integrated environment may facilitate such research that identifies possible
applications of emerging electric propulsion systems, combined with operational analysis that is able to map the
mission capability and the cost of the indivisual electric propulsion aircraft concept to its commercial or military
value.
Technologies associated with most emerging electric propulsion components are premature. As identified in
Reference 41, the off-the-shelf fuel cell and electric power components are barely good enough to make a flyable
fuel cell airplane, mainly because of their low specific power and power density. Dramatic improvements in major
components to higher technology readiness levels must be achieved to yield a feasible and viable solution. A
potential application of the proposed method is to support resource allocation studies that select the best mix of
investments under a limited budget constraint.
Acknowledgments
The authors would like to thank NASA and the Department of Defense for financially supporting this research as
part the URETI for Aeropropulsion and Power Technology initiative. We would also like to recognize our
collaborative research associates at the Georgia Tech Research Institute and Florida A&M University, as well as our
colleagues at the Aerospace Systems Design Laboratory who contributed to portions of the research reported in this
publication.
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