2006-01-3092
Test Results for a Fuel Cell-Powered Demonstration Aircraft
Thomas H. Bradley
Georgia Institute of Technology
Blake A. Moffitt, Reid W. Thomas, Dimitri Mavris
Aerospace Systems Design Laboratory
David E. Parekh
Georgia Tech Research Institute
Copyright © 2006 SAE International
ABSTRACT
A fuel cell powered airplane has been designed and
constructed at the Georgia Insitute of Technology to
develop an understanding of the design and
implementation challenges of fuel cell-powered
unmanned aerial vehicles (UAVs). A custom 448W net
output proton exchange membrane fuel cell powerplant
has been constructed and tested. A demonstrator
aircraft was designed and built to accommodate this
powerplant and the fuel cell powered aircraft has
performed seven test flights to date. Test data show that
the aircraft performance validates the models used for
design and optimization and that the fuel cell aircraft is
capable of longer endurance, higher performance test
flights.
INTRODUCTION
Fuel cells are an attractive technology for implementation
as powerplants for aircraft because of their potential for
rechargability and high energy density. Where advanced
rechargeable batteries can reach energy densities of 150
Wh/kg at the module level, [1] regenerative fuel cells can
achieve >800 Wh/kg at the system level [2], and nonregenerative fuel cells can achieve >1000Wh/kg [3].
This study is concerned with the application of nonregenerative polymer electrolyte membrane (PEM) fuel
cells as the primary powerplant in a medium-scale
unmanned aerial vehicle (UAV). Implementation of nonregenerative fuel cells in UAVs is viewed as a first step
towards practical fuel cell-powered flight.
A few researchers have used small-scale model aircraft
to demonstrate fuel cell-powered UAVs [4-6]. Soban and
Upton [7] have performed preliminary design studies but
without hardware development or system-level
validation. AeroVironment has demonstrated a large
scale fuel cell UAV [8], but the design and development
challenges of fuel cell UAVs are still not well-understood
by the general UAV community.
In order to begin developing tools and technologies for
the design and implementation of fuel cells as
powerplants in aeronautical vehicles, a PEM fuel cell
UAV demonstration program was started in the Summer
of 2004 as a collaboration between the Georgia Institute
of Technology Aerospace Systems Design Laboratory
(ASDL) and the Georgia Tech Research Institute. A
portion of the design work is supported by the NASA
University Research Engineering Technology Institute
(URETI) grant to the Georgia Institute of Technology.
The primary research objectives of this program are: the
development of validated tools for fuel cell aircraft
design, analysis of the design tradeoffs present in fuel
cell aircraft, and the demonstration of a series of fuel cell
UAVs.
This paper presents a summary of the specifications of
the aircraft and its fuel cell powerplant. Laboratory and
field testing results are analyzed and used to validate the
aircraft propulsion model that was used for system
design. A summary of the flight tests completed to date
is also presented. A photograph of the fuel cell-powered
demonstrator aircraft is shown in Figure 1.
Figure 1. Fuel cell-powered demonstrator aircraft
FUEL CELL POWER PLANT
The powerplant for the fuel cell UAV is a 500W 32-cell
PEM stack (BCS Fuel Cells, Bryan, TX) with a custom
balance of plant. The fuel cell balance of plant was
designed for mobility, reliability and minimum weight.
The balance of plant (BOP) includes the air management
subsystem, hydrogen delivery subsystem, thermal
management subsystem, and system controller. The
fuel cell is self-humidified and requires an air manifold
pressure of 5 psig and a hydrogen manifold pressure of
3 psig. The fuel cell balance of plant provides a cathode
stoichiometry of >2.0 and 90% peak hydrogen utilization.
A schematic of the fuel cell powerplant system is shown
in Figure 2.
Figure 3. Fuel cell system hydrogen utilization
THERMAL MANAGEMENT SUBSYSTEM
Figure 2. Schematic of fuel cell powerplant system
AIR MANAGEMENT SUBSYSTEM
The air management subsystem maintains constant
pressure at the cathode of the fuel cell stack and
provides variable flow rate control. Variable flow rate
control is particularly important in a self-humidified fuel
cell system because of the risk of under-humidification at
low current densities. The 5 psig cathode pressure is
regulated with a calibrated ball valve (Microchek
14B14B-5 psi, Lodi, CA). Flow rate is closed-loop
controlled by pulse-width modulation of two diaphragm
compressors (T-Squared Manuf. T202, Lincoln Park,
NJ). These compressors are powered from the fuel cell
bus voltage. By using two compressors, and turning one
of the compressors off when low flow is required, higher
high flow rates and lower low flow rates are achievable
than is possible with a single compressor.
HYDROGEN DELIVERY SUBSYSTEM
Hydrogen is stored on board the aircraft using
compressed hydrogen stored within a composite tank
(Luxfer Gas Cylinders P07A, Riverside, CA). The
hydrogen storage system provides roughly 66 standard
liters (SL) of usable hydrogen. Dual regulators regulate
the hydrogen pressure to a manifold pressure of 3psi.
Hydrogen is purged from the cathode manifold for 0.2
seconds at roughly 8 second intervals. Hydrogen purge
is required to remove water and other contaminants from
the anode manifolds.
A liquid cooling circuit provides temperature regulation
for the fuel cell. A water pump (Laing Thermotech DDC,
Chula Vista, CA) circulates water through a custom
radiator of aluminum and carbon foam (Poco Graphite,
Decatur, TX) construction. Carbon foam is an advanced,
lightweight (0.56g/cc), thermally conductive material that
allowed for the construction of a light weight water-to-air
radiator. A number of design iterations were performed
to reduce the weight and power consumption of the
radiator.
In the final design iteration, the fuel cell radiator and fan
weigh 578g and rejects 242W over a temperature
difference of 20C with 6W of electrical power
consumption. Fuel cell product water is sprayed onto the
fins of the radiator to allow for evaporative cooling.
SYSTEM CONTROLLER
The fuel cell system is controlled by an ATMEGA32, 8-bit
AVR microcontroller (Atmega, San Jose, CA). The
functions of the fuel cell controller are:
1. Thermostatic Fan Control – This function turns on
the radiator fans when the fuel cell temperature is
greater than 45 C.
2. Closed-Loop Compressor Speed Control – An
inductive current sensor measures the current output
of the fuel cell. This function reads the current
sensor and calculates the required PWM conditions
for the compressors. The compressor speed is
closed-loop controlled with a PI controller.
3. Mode Control – There are three modes of fuel cell
controller operation: Idle, Purge, and Fly. In Idle
mode the fuel cell system is not producing power.
Purge mode allows the fuel cell controller to purge
the cathode manifold of the fuel cell at startup or
shutdown.
Fly mode is the normal mode of
operation that includes active, closed-loop control of
the fuel cell system.
4. Open Loop Compressor Purging Cycle – During
startup and shutdown, the compressors operate at
maximum flow rate to purge the cathode manifold of
the fuel cell.
52W/kg and a specific energy of 7.1Wh/kg, both
measured at conditions maximum net electrical power.
AIRCRAFT DESCRIPTION
The medium scale (Sw=188dm2, m=16.4kg) of the
aircraft platform is chosen for the demonstration aircraft
because of:
5. Purge Valve Control – The fuel cell controller
provides a signal to the anode purge valve to purge
the fuel cell anode manifold.
•
6. Data Acquisition – The fuel cell controller performs
data acquisition, data storage and transmission to a
ground station. The fuel cell controller functions as
both the fuel cell and aircraft data acquisition system.
The utility of a larger aircraft platform for
potential UAV missions including surveillance
and remote sensing.
•
Improved applicability of design methods and
validation data. The closer in scale that the
demonstration aircraft is to future utilitarian
aircraft, the more confidence can be granted to
the design and validation data.
•
The suitability in scale and cost for an academic
research program.
FUEL CELL SYSTEM PERFORMANCE
To validate the design and demonstrate the function of
the entire fuel cell system, the system was tested in the
laboratory before implementation in the fuel cell aircraft.
Fuel Cell Component Power Consumption, Watts
A total net output power of 448W is available from the
500W rated fuel cell system. The balance of the power
between the system net output power and the 523W
gross output from the fuel cell is consumed by accessory
loads. Figure 4 shows the breakdown of the power
consumed by the accessory loads as a function of fuel
cell system net output power. The 3W of power
consumed intermittently by the cooling system fan is not
included in this graph. The compressors make up the
largest part of the power consumed by the fuel cell
system. The water pump, controller and data acquisition
(DAQ) hardware consume a constant 13.3W.
Conceptual design of the aircraft was performed using
an integrated model of the powerplant, propulsion and
aircraft systems. The characteristics of this design
model were explored using a full factorial design of
experiments for the 12 design variables: wing area,
aspect ratio, tail type, electric motor, fuel cell, propeller,
fuselage shape, hydrogen storage type, number of
hydrogen storage tanks, and wing configuration. More
detail regarding the conceptual design task is available in
Moffitt et al. [9].
600
Compressor 1
Compressor 2
Control and DAQ
Coolant Pump
Electrical Output Power
500
400
300
200
Figure 5. Fuel cell demonstrator aircraft configuration
100
0
0
50
100 150 200 250 300 350 400 450
Net System Output Power, Watts
Figure 4. Breakdown of fuel cell powerplant power
consumption
Overall, the fuel cell system including balance of plant,
controls and hydrogen storage has a specific power of
The result of the conceptual and lower level design tasks
is the aircraft configuration shown in Figure 5. The
aircraft features a high aspect ratio, non-swept wing with
taper and dihedral in the outer spanwise sections. The
aircraft has a streamlined fuselage with an internal
volume of approximately 36.7L with a maximum cross
2
sectional area of 438cm and an overall length of 1.12m.
The aircraft uses a pusher configuration with a brushless
3-phase motor and a fixed pitch propeller. The aircraft
has twin tailbooms supporting an inverted V tail.
Table 1. Specifications for the fuel cell demonstrator
aicraft
Design Specification
Value
Wing Area
Aspect Ratio
Span
Taper Ratio
Taper Change Point
Outer Panel Washout
188 dm2
23
6.58 m
0.67
0.625
4 deg
Tail Area
Length (nose to tail)
Mass
Propeller Diameter
Propeller Pitch
Static Thrust / Weight
Cruise Airspeed
Dihedral at Tapered Section
45.5 dm2
2.38 m
16.4 kg
55.9 cm
50.8 cm
0.165
14.5 m/s
10 deg
The lower-level design of the aircraft to refine the aircraft
aerodynamics and propulsion system included additional
optimization metrics and constraints. When considering
stall performance, power limited turning performance,
minimum turning climb rate, minimum Reynolds number,
minimum bank angle and maximum allowed angle of
attack the wing geometry changed slightly from the
conceptual design results. The final aircraft design
resulted in a slightly higher wing area and aspect ratio
than the conceptual aircraft design. Specifications of the
aircraft design are given in Table 1.
Figure 6. Mass breakdown for the fuel cell demonstrator aircraft
AIRCRAFT CONSTRUCTION
Balancing the requirements of light weight and function
proved especially difficult for the demonstration aircraft.
Analysis performed at the design stage showed that the
aircraft performance is highly sensitive to the takeoff
gross weight of the aircraft. The large mass of the fuel
cell powerplant, necessitated structural analysis and
design not usually associated with small scale UAVs.
The construction methods and materials were also
constrained by budgetary requirements and an
aggressive build schedule.
The most critical component of the aircraft structural
design was the fuselage. The fuselage is designed to
support the 8.6 kg fuel cell system, wing loads and tail
moments under a minimum 3-g loading. The fuselage
consists of a structural internal frame with non-structural
fairing. The internal frame is designed to take all major
loads during both ground and flight operations. It is
made up of a roll-wrapped carbon fiber tubular spar and
two roll-wrapped carbon fiber composite tail booms
lugged and bonded to a tubular aluminum frame. The
fairing provides an aerodynamic shape to the aircraft,
and is constructed of a thin layer of fiberglass with
honeycomb stiffening in some panels to maintain the
desired shape. The main landing gear have a symmetric
airfoil cross-section and are machined from 6061-T6
aluminum billet. The nose gear is constructed of
composite tubing with an aluminum fork.
The wings are constructed of a composite spar and solid
foam cores sheeted with balsa wood. The empennage
surfaces are built from foam and balsa wood without an
internal spar.
The mass breakdown of the final aircraft is shown in
Figure 6. The weight of the aircraft is dominated by the
weight of the powerplant. The fuel cell and balance of
plant make up roughly 57% of the total weight of the
aircraft. The weight of the hydrogen fuel is included in
the hydrogen tank weight. For the test flights described
in this paper, the aircraft is fueled with 5g of hydrogen.
pilot riding in this vehicle. This arrangement is shown in
Figure 7.
AIRCRAFT FLIGHT TESTS
While the laboratory testing provide a significant amount
of data on the fuel cell power system, the ultimate
validation of this work comes from integration with the
aircraft and flight tests of the total system. The primary
goals of this research are to push the boundaries of
range and endurance for aircraft by leveraging the high
efficiencies of fuel cells. Working towards these goals,
the team has developed an incremental flight test plan,
focused on achieving low risk aircraft performance
milestones, which matches the aircraft’s capabilities and
the testing facilities available. The goal of this test plan
is to collect data regarding the aircraft performance and
demonstrate milestones of fuel cell flight, while
minimizing the risk to the aircraft.
Matching the aircraft’s performance capabilities with a
suitable flying location proved to be a difficult task. While
many sites exist around the country and throughout
Georgia for model aircraft flying, the size of this aircraft,
as well as the nature of the research provides only a
partial overlap with the available field options.
Furthermore, most hobby flying sites impose additional
performance constraints on the vehicle through their
geographic constraints and site obstacles such as trees
or power lines. The culmination of these factors does
not eliminate hobby sites from contention, but would
force the flight test program to take several initial leaps
before settling into the incremental schedule. As a
potential alternative, the research team investigated the
possibility of flying at smaller or private airstrips for full
size aircraft. Pursuing this option forced the aircraft up
against recent changes imposed by the Federal Aviation
Administration (FAA) on the use of unmanned aerial
vehicles in the vicinity of airports, which forbids this
option, given the timeline of the project. Determined to
proceed with the incremental approach, the team looked
outside these more traditional aeronautical options for a
facility that would allow for the planned initial test flights.
Ultimately, the Atlanta Dragway in Commerce, GA was
selected as the location for the initial demonstration
flights. The dragway provided nearly a thousand meters
of unobstructed paved facilities, without any regulatory
restrictions. This facility proved to be an excellent
location for the initial flight tests.
While the use of the Atlanta Dragway alleviated several
concerns and challenges, it also posed additional
constraints upon the flight test crew. The long straight
track, so ideal for the vehicle, made the control of the
aircraft much more difficult for the pilot and flight crew.
Ultimately it was decided that a chase vehicle would be
used to follow the flight tests down the track, with the
Figure 7. Fuel cell-powered demonstrator aircraft
undergoing flight test with chase vehicle
In order to perform longer-range and higher-performance
test flights, the aircraft was moved to the flying site of the
Georgia Jets radio controlled aircraft club. The Georgia
Jets facility, located in Gay, GA, features a 305m paved
runway, and a large cleared area.
TESTING SUMMARY
To date, the team has completed six test flights (Tests
#1-6) of the aircraft on hydrogen power at the Atlanta
Dragway and a single hydrogen powered circuit test flight
(Test #7) at the Georgia Jets facility.
Table 2. Test flight summary for fuel cell-powered flights
Flight Test
Duration
Max
Altitude
Test goals
Test #1
28 seconds
2.5m
Aircraft trim,
instrumentation
check and pilot
familiarization
Test #2
52 seconds
3.7m
Pilot practice and
aircraft trim
Initial test of flaps
Test #3
36 seconds
3.5m
Aircraft trim,
further flap testing
Test #4
12 seconds
2.5m
Final aircraft trim
flight
Test #5
50 seconds
3.5m
Achieve steadylevel flight
Test #6
25 seconds
3m
Achieve steadylevel flight
Test #7
120 seconds
25m
Circuit flight, rate
of climb test
The purpose of these flights has been trimming of the
aircraft controls, pilot familiarization with the aircraft,
initial flight testing of the fuel cell system and testing of
the flight data acquisition systems. The flight tests also
provide the first performance data sets for the fully
integrated aircraft. This data allows the team to validate
the design models for the vehicle and fuel cell system,
provides idle air flow. Figure 9 also shows that for this
test the fuel cell does not reach its operating temperature
of 50-60C over the course of this test.
8000
55
Fuel Cell Temperature
80
Voltage (V), Current (A),
Airspeed (ft/sec)
Takeoff
50
Compressor 2 Speed
50
6000
5000
45
4000
3000
40
2000
1000
35
0
20
40
60
Time, sec
80
0
100
Figure 9. Compressor speeds and fuel cell operating
temperature from straight-line flight (Flight Test #2)
Fuel Cell Output Current
Fuel Cell Output Voltage
Aircraft Airspeed
60
Fuel Cell Temperature, C
Figure 8 shows some results from Flight Test #2. The
fuel cell voltage and gross current (measured before
BOP loads) are plotted on the same axis as the aircraft
airspeed. As the aircraft begins takeoff the loading on
the fuel cell powerplant increases. The voltage of the
fuel cell decreases as the current draw decreases until
the maximum power operating condition is reached. At a
time of 26 seconds, the pilot rotates the aircraft and the
plane becomes airborne. The airspeed plateaus at 32
ft/sec for a few seconds as the pilot gains altitude, but
when an altitude of 3.5m is reached, the pilot accelerates
at full power until a time of 60 seconds. At that point the
pilot scales back the throttle, decreases airspeed, lowers
flaps and the aircraft touches the ground at a time of 78
seconds.
70
7000
Compressor 1 Speed
FLIGHT TEST RESULTS
Compressor Speed, rpm
and to verify the predicted flight performance. This
critical feedback will provide the team with significantly
greater confidence in the accuracy of future design
predictions, and will also aid in flight planning for future
test flights. A short summary of the tests completed to
date, their duration, maximum altitude, and the goal of
the flight is listed in Table 2.
Landing
40
The results of this test flight show that the aircraft
achieved full fuel cell output power, functional flap
deployment, improved aircraft control surface trim, and
functional data acquisition. The aircraft took 200m of
taxi distance and achieved straight-line flight and landing
without incident.
VALIDATION OF DESIGN MODEL
30
One of the main goals of both the building and testing
the demonstration aircraft is to obtain the data needed to
validate and calibrate the models used during the aircraft
design. This section describes the propulsion system
model and its partial validation using the preliminary test
results from the demonstration aircraft.
20
10
0
PROPULSION SYSTEM DESIGN MODEL
0
20
40
60
80
100
Time, sec
Figure 8. Sample results from straight-line flight (Flight
Test #2)
Figure 9 shows data from the same test flight. In this
figure, the two compressor speeds are shown on the
same plot as the fuel cell temperature. As shown in
Figure 9, during low fuel cell power demands, only one
compressor is used to provide the cathode air supply.
When the fuel cell power demand is increased as the
pilot takes off, both compressors work together. When
the power demand decreases, as the pilot begins to
land, Compressor 2 again shuts off and Compressor 1
Modeling of the fuel cell system is based on a static
polarization curve. During the conceptual design, linear
approximations to the polarization curves were used.
Although more refined models of the fuel cell can be
developed, the computational efficiency of the linear
approximation makes it particularly computationally
efficient when used in multidisciplinary optimization
techniques. Once the stack was purchased and the
balance of plant was designed and implemented,
refinements to the polarization curve were made using
data obtained from laboratory testing.
Propeller performance maps for the scale of this aircraft
only exist for a few propellers [10] and typical methods
based on classical momentum theory are not accurate
enough for designing a power limited aircraft. Thus a
custom propeller analysis code was developed based on
a mixture of momentum and blade element theory using
Goldstein’s vortex model [11,12].
Three different
propeller geometries were accurately measured for use
with the propeller code. In addition, scaling of the
propeller profiles and pitch distributions was used to help
predict the performance of other fixed pitch propellers
similar to the measured propellers.
The electric motor was modeled using a lumped
parameter equivalent circuit model of the motor, as
shown in Figure 10. This model uses a no-load current
(I0), motor internal resistance (Rm), controller resistance
(Rcont), and motor voltage constant (Kv = ωm/Vm). Values
of the lumped parameters were obtained from motor
manufacturers during conceptual design and were then
measured from motors actually purchased for the final
detailed design of the propulsive system.
25
250
20
200
15
150
10
100
5
The first data used for validation were obtained by bench
testing the full propulsion system. Current, voltage, and
motor speed were measured for all combinations of two
motors and two propellers. Figure 12 shows both the
predicted and measured data for the motor and propeller
selected for the final aircraft. Note that the propulsion
model offers a good estimate of the fuel cell current and
voltage up to about 23 amps. At currents above 23
amps, the voltage of the fuel cell begins to rapidly drop
causing the power output to plateau and then decrease.
As a result, the motor speed would also begin to
decrease slightly after 23 amps. For flight purposes, the
throttle was limited so that the maximum allowable
current was 23 amps, thus predicting performance at
higher currents was unnecessary. The model used to
calculate the values in Figure 12 is the linear fuel cell
polarization model used for conceptual design. More
details regarding the conceptual design are present in
Moffitt et al. [9].
Power (T*V), W
Thrust, N
300
MODEL VALIDATION
The most important data for validation in Figure 12 are
the motor speed data. Motor speed is dependent on all
components in the propulsion system. Note that the
predicted motor speed agrees with the measured data.
In order for this to happen, the code must correctly
predict the torque vs. angular speed relationship for the
propeller, fuel cell and motor in combination.
Figure 10. Electric motor model
30
By combining the measured fuel cell stack polarization
curve with the propeller and electric motor models, the
predicted propulsive performance of the aircraft could be
calculated for any number of flight conditions. For
example, Figure 11 shows both the maximum predicted
thrust and the maximum available propulsive power
generated by the optimized propulsion system.
50
Thrust
Power
0
0
0
5
10
15
20
25
30
Airspeed, m/s
Figure 11. Predicted maximum available propulsive thrust
and power
Figure 12. Propulsion system bench test results
The data in Figure 12 suggest that the propulsion model
is adequately predicting performance under static
50
400
Motorspeed Measured
Motorspeed Predicted
Thrust Measured
Thrust Predicted
350
300
45
40
35
250
30
200
25
20
150
Thrust, N
Data from the battery flight test is shown in Figure 13.
The flight lasted nearly 100 seconds and consisted of a
simple circuit around the field. The airspeed varied from
0 to 25 m/s. Using the measured flight airspeed in
conjunction with a linear battery model, the battery
current and propeller thrust was predicted and compared
to measured data. The predicted current is very close in
magnitude to measured data and closely predicts the
trends measured throughout the flight. The predicted
thrust also closely matches the measured data. This
comparison provides validation for the dependence of
the propulsion system model on airspeed.
the measured and predicted thrust for the other portions
of the flight.
Motorspeed, rad/s
conditions. However, to be useful, the model must also
be able to predict performance as airspeed is varied. To
verify that the model could predict performance over a
range of airspeeds, a test flight of the aircraft was
performed. A well-characterized nickel-cadmium battery
pack was used to power the flight test. Use of the
battery pack for initial flight testing allows for the
characterization of the aircraft independent of the fuel
cell powerplant and also allows the aircraft to be flown
without risk to the fuel cell system.
15
100
10
50
5
0
0
20
40
60
80
100
0
120
Time, sec
Measured Current
Predicted Current
Measured Airspeed
Measured Thrust
Predicted Thrust
45
40
35
70
60
50
30
40
25
20
Thrust, N
Current, A; Airspeed, m/s
Figure 14. Predicted vs. measured data for the circuit fuel
cell test flight (Flight #7)
80
50
30
15
20
10
10
5
0
0
0
25
50
75
100
125
150
Time, sec
Figure 13. Propulsion system flight test data for a battery
powered test flight
The final validation of the propulsion model was
performed by using the data from the fuel cell powered
flights. A non-linear polarization curve derived from
bench-top testing of the fuel cell system was used to
model the fuel cell.
Validation was performed by
comparing the predicted and measured motor speed and
thrust. Figure 14 shows this comparison for the fuel cellpowered circuit flight, Flight #7. Note that the propulsion
model accurately predicts the magnitude of the motor
speed and correctly predicts the increase in motor speed
near the time of 75 seconds. The thrust data for this
flight shows a dip in magnitude during climb out (10s30s) that may be an artifact of the strain-gauge
measurement setup. There is good correlation between
Together, these comparisons show that the behavior of
the fuel cell aircraft propulsion system can be predicted
to the accuracy required for design. With refinement and
further testing the accuracy of the model will be
improved.
CONCLUSIONS
Theoretical design models for fuel cell aircraft are
present in the field, but the modeling, optimization, and
implementation challenges of fuel cell aircraft were not
well understood. The design, construction and testing of
a fuel cell-powered demonstration aircraft is an important
step towards realization of fuel cells as a powerplant for
aviation.
The design of the fuel cell system has resulted in a
feasible fuel cell powerplant with >450W net output and
90% hydrogen utilization. The fuel cell demonstrator
aircraft is optimized around the use of a fuel cell
powerplant. The result is a highly effective platform for
extended testing and evaluation of the fuel cell-powered
aircraft concept.
Testing of the fuel cell demonstrator aircraft has followed
a conservative, incremental approach. To date, seven
flight tests have been performed to gather data regarding
the performance of both the fuel cell system and the
aircraft. These flight tests have included low-altitude
straight-line and higher-altitude circuit flights.
Data gathered from flight testing validates the conceptual
design models and the more refined system models.
Future tests will focus on evaluating the rate of climb,
endurance, maneuverability, and power requirements of
the aircraft with the goal of further refinement of design
techniques and modeling tools.
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