52nd AIAA/ASME/ASCE/AHS/ASC Structures, Structural Dynamics and Materials Conference<BR> 19th
4 - 7 April 2011, Denver, Colorado
AIAA 2011-1986
Design of a Hybrid Electrical Propulsion System
Chad Campbell, Eric Dittman, Jennifer MacRae, Claudia Ehringer, Alex Stone, Kevin Lake, and Spencer Haskins1
and
R.R. Mankbadi2
Embry-Riddle Aeronautical University and the Florida Center for Advanced Aero-Propulsion (FCAAP), Daytona
Beach, FL, 32114
The focus of this work is on exploring the design of an aircraft hybrid propulsion system
by integrating a gas turbine engine with an electric motor assisted by solar power. We
discuss how the two systems can work in tandem while optimizing the efficiency. The Pilatus
PC-12, a single engine, mid-sized turboprop aircraft, was chosen as the platform for this
study. The hybrid system is optimized for maximum efficiency, and further significant fuel
savings were found to be possible by including thermoelectric generators.
Nomenclature
V
SA
SB
T2
T1
=
=
=
=
=
voltage difference
Seebeck coefficient of material A
Seebeck coefficient of material B
Temperature 2, hot side
Temperature 1, cool side
I. Introduction
The hybrid electric car is now a practical alternative to gas-powered cars. Therefore, it is time to explore the
feasibility of a hybrid-electrical aircraft. We propose herein the design of a turboprop electrical-gas turbine
propulsion system. We discuss a system for integrating an electrical motor with a gas engine, and the optimum
selection of the motor and engine. We explore several secondary sources of energy, such as solar energy and
thermo-electric generators. While our focus is on the economical feasibility of the design, such a system can also be
helpful in saving our environment. Combustion of fossil fuels releases harmful greenhouse gasses, smog, acid rain,
climate change (such as global warming), and health concerns in humans. It is claimed that 10% of all pollution in
the atmosphere above the United States is caused by aviation. Thus, the proposed system reduces fuel consumption
and minimizes the pollution effects.
The outline of the paper is as follows. In section 2 we discuss the electrical system. In section 3 we research the
gas-turbine engine, in section 4 we discuss solar-power assist, in section 5 we discuss secondary energy sources to
optimize the system’s overall efficiency. And in section 6 we discuss the final selection of the system.
II. The Electrical System
The following section will cover the approach used to select a general system for hybrid propulsion; options
available for electric motors, and the battery system.
A. Series and Parallel Hybrid Systems Integration
There are two main choices for hybrid system: A series or Parallel system. In the Series Hybrid Propulsion System
(Figure 2), the propeller is powered exclusively by an electric motor mounted to the propeller shaft. The electric
motor may be powered from the battery, a capacitor, a flywheel, or directly from a generator that is run from the
turbine. Efficiencies in this system are gained by the fact that the turbine need not run continuously, and when it is
run, it only runs at its most efficient speed and power setting. While the turbine is running it may power the
propeller motor with any excess power being applied to charge the battery or capacitor. The battery may be fully
1
2
Students at ERAU, College of Engineering.
Distinguished Professor at ERAU, College of Engineering.
1
Copyright © 2011 by the American Institute of Aeronautics and Astronautics, Inc. All rights reserved.
charged on the ground before flight, allowing the plane to be taxied on battery power alone. The option of wind
milling the propeller when descending, and using this power to charge the battery also exists. A simple working
example of an aviation series hybrid propulsion system already exists on the 737NG. It is called the Wheel Tug, and
it allows the plane to taxi with an electrical drive motor mounted in its nose landing gear. The motor is powered
directly from the planes auxiliary power unit. This saves fuel and improves aircraft efficiency by delaying the need
to start the main turbine engine until the aircraft is ready for takeoff. This system requires the motor to be rated at
the full take-off power of the aircraft. The combined weight of this added equipment makes the series hybrid system
impractical for primary propulsion in aviation.
In the Parallel Hybrid Propulsion System, the electric motor works in conjunction with the turbine to spin the
propeller. The electric motor is mounted to the existing propeller shaft and may either add power to the propeller
shaft from the battery, or charge the battery by acting as a generator and capturing power from the turbine. Figure 3
shows a diagram of a parallel hybrid system. A parallel hybrid propulsion system may be classified as either a full
hybrid or a power assist hybrid. A full hybrid system has the advantage of being able to run only from battery
power for short periods of time. In order to do this the system requires a clutch to disengage the turbine from the
propeller shaft, adding complexity and weight to the system. Another option is for the turbine to spin with the
propeller while operating off the electric motor; this way of operating adds a significant amount of parasitic
frictional drag to the propulsion system and hurts efficiency. In a power assist hybrid system, the propulsion system
is always powered by the turbine, with the electric motor only providing a boost in power during certain phases of
flight. With both systems, a motor-generator can be used to provide power during a certain phase of the flight, and
then use the excess power in another phase of the flight to recharge the batteries.
Figure 1. Series Hybrid System [1]
Figure 2. Parallel Hybrid System [1]
2
B. Electric Motor Options
In selecting the electrical motor suitable for propulsion
applications, we need to keep the weight to a minimum. A base
line electric motor size of 100kw was used for motor selection
and comparison. The most basic of these choices is between
using an alternating current (AC) or a direct current (DC)
motor. AC motors are usually lighter than DC motors, but they
require complex power conversion hardware to convert the
direct current of the battery into a useful power supply for the
AC motor. Three-Phase AC motors are usually lighter than a
single-phase AC motor, but require components to create the
three phase signal. DC motors tend to be larger than an
equivalently rated AC motor and generally require more
maintenance. The advantage to a Permanent Magnet (PM) DC
motor is that, unlike a separately excited DC motor, they do not
require an excitation power source to establish a magnetic field.
PM motors therefore are always ready to function as a motor or
generator. The disadvantages to PM motors are that they are
heavy and costly because they require windings composed of Figure 3. Symetron P-100
rare Earth magnets. We considered various motors by various
manufacturers (Hansen, Premag, and Razor). From this list of
currently available motors, the Symetron P-2 was chosen
because it is air-cooled, low weight and did not require
permanent magnets in its construction.
In order for the hybrid electric propulsion system to be
feasible, it will have to weigh as little as possible. One AC and
one DC motor were found that are in development and exceed
all current motors in terms of power produced vs. weight of
the motor. The AC motor was the Symetron P-100 (Figure
4), which is designed specifically to produce our baseline
power of 100kw. The P-100 is able to function as a generator
and is designed to fit around a drive shaft. The P-100 also
operates at peak efficiency around 1700 RPM which is the
Figure 4. DHARMA Motor Prototype [3, 4]
propeller speed of the PC-12.
This greatly simplifies
installation of the motor because it can be mounted in the nose of the aircraft and fit around the propeller shaft
without a heavy transmission. The DC motor considered for this project is the DHARMA (Double Halbach
Aferrous Radial Airgap Motor), and it is currently in development. This motor was specifically designed for
aviation use and is therefore constructed to be as light as possible while still maintaining a useable efficiency. It is
constructed from a filament wound high modulus carbon in order to save weight. It can also be constructed directly
around the propeller shaft and designed to match the dimensions of the nose of the aircraft.
The final motor design choices were based primarily on their power vs. weight. The DHARMA motor excelled
in these criteria and was clearly the best choice. It is however still in the prototype phase and has not yet exceeded
15 kW in power. The Symetron P-100 is nearly in production, but this motor is optimized for use in a ground
vehicle and therefore weighs considerably more than the DHARMA concept. The top three electric motor choices
are overviewed in Table 1.
Table 1. Facts on Top Three Motor Choices
Motor
KW/Lb Generator?
Type
Symetron P-2
0.43
Yes
AC
Symetron P-100
0.52
Yes
AC
DHARMA
0.81
Yes
DC
3
C. Battery Selection
The three types of batteries researched were Lithium-Ion (Li-ion), Nickel-Cadmium (NiCd), and Nickel-Metal
Hydride (NiMH). Lithium-ion batteries [2] work by using a lithium ion which moves between a cathode and an
anode. These batteries have a relatively high energy density of 91 Watts-Hour/lb (WH/Lb). They are commonly
used in aerospace applications for a variety of reason. They can be formed into a wide range of sizes and therefore
can be formed to fill any available space. Lithium-ion batteries also do not suffer from the same memory effect that
many other batteries suffer from, which leads other batteries to hold less charge each time they are charged.
Additionally, they have a very slow self-discharge rate, of approximately 5% per month.
These batteries typically cost about $1.50/Wh. Since approximately 15 minutes of power are needed for takeoff,
at 100 KW, 25 KW-hr are needed. Assuming
91 WH/lb, the total weight of the batteries can
be estimated as 275lb. Therefore, using
lithium-ion batteries without considering any
additional power sources during flight would
require the use of 275 lbs of batteries to
supply the power needed for takeoff and
climb alone. Lithium-ion batteries do have
some significant drawbacks. They are very
dependent on their age, and as a result the
older the battery is the more internal
resistance it will have. For this reason, it is
recommended that the batteries are bought as
close to use time as possible, to maximize the
useful life once purchased. Additionally,
lithium-ion batteries are not as durable as
many other types of batteries, and can be
extremely dangerous if they are mistreated. Figure 5. Energy Density of Batteries
They also have a small but present chance of
catching on fire.
Nickel-cadmium batteries are rechargeable batteries that use nickel oxide hydroxide and metallic cadmium as
electrodes. These batteries are very damage resistant and can supply a high surge current if needed due to their low
internal resistance. They have an energy density of 27.3 WH/lb, which is much lower than that for Li-ion batteries.
Additionally, Nickel-cadmium batteries cost about $7/Wh, which is almost five times as much as what the typical
lithium-ion battery costs. These batteries are able to be used individually or in battery packs, but are sealed due to
their extreme toxicity. Using nickel-cadmium on this aircraft would require the use of 917 lbs of batteries to supply
the power needed.
Nickel-metal hydride batteries [5] use a
hydrogen-absorbing alloy as the negative
electrode in place of the cadmium in nickelcadmium batteries. These batteries are cheaper
than NiCd batteries, but at $2.7/Wh, are still
almost twice as expensive as Li-ion batteries.
They also have a better specific energy density
than NiCd batteries at 36.2 WH/lb, but this is still
less than half of what is possible with Li-ion
batteries. Using these batteries would require the
use of 690 lbs of batteries to supply the power
needed. Comparison of energy storage per unit
weight is given in Figure 5. Cost per energy
storage for the various batteries is shown in Figure
6. Due to its low cost and high energy density,
Lithium-ion is the clear energy choice for this
application. Although care will need to be taken
to minimize the risk of using this battery, it was Figure 6. Cost per Watt-hour of Batteries
determined that it was still the best choice for this
4
system.
III. The Gas Turbine Engine
We discuss herein the possibility of
downsizing the PC-12’s current engine. A
typical turbo-prop engine is shown in Figure 7.
Since an electric motor will be added to the
propulsion system, less power output is
required from the turboprop engine (Figure 8).
A new engine should be lighter and cheaper
with a lower power output yielding lower
specific fuel consumption.
The selection was narrowed down to three
Engine choices, the current engine (PT6A67B), a smaller engine in the same family
(PT6A-36) [7], and a smaller engine from a Figure 7. Schematic of the Turboprop Engine [6]
different company (Walter M601F). The PT6A
is designed for medium altitudes and rugged
conditions. New aerodynamic and material
technologies have enabled the engine to gain
more power without significantly increasing in
size. Other innovations have reduced emissions,
increased maintenance intervals and further
enhanced ease of operation with the
introduction of digital electronic control to
small gas turbine engines. The engine takes
advantage of its rearward, reverse flow inlet
and forward facing turbine section that provides
fast maintenance turn-around through on-wing
hot section refurbishment in most aircraft
installations. The PT6A-67B is a two-shaft
engine with a multi-stage compressor driven by
Figure 8. Flight Conditions When Electric Motor Will be
a single-stage compressor turbine and an
used to Augment Thrust
independent shaft coupling the power turbine to
the propeller through an epicyclic concentric
reduction gearbox. The engine uses 4 axial
stages and 1 centrifugal stage in the compressor
and a 2 stage axial turbine; which is an
independent power turbine with shrouded
blades. With the use of the epicyclic speed
reduction gearbox the output speed is optimized
for the highest power and the lowest propeller
noise and enables compact installation, which is
ideal for our hybrid because of the extra
components being added to the aircraft’s
fuselage.
The Walter M601 is a dual-shaft reverseflow free turbine turboprop engine. The gas
generator section consists of 2 axial and 1
Figure 9. Power vs. Weight of Engine Options
centrifugal compressor stages, an annular
combustion chamber, and a single stage axial
compressor turbine. The power section consists of a single stage axial power turbine, exhaust system, and a, 2 stage
planetary reduction gearbox with torque meter. An accessory gearbox with integral oil tank is mounted on the rear of
the engine. The Walter M601 was originally designed for use in remote parts of Russia and Siberia where rugged
5
durability with minimal field maintenance requirements were top priorities. The engines have been widely used in
twin-engine LET 410 aircraft for commuter and cargo operations. Unlike some turbo-prop engines, "hot section
inspections" between overhauls are not required with the Walter M601. The engine is equipped with limiters which
prevent over temperature and over speeding during starting and reverse thrust operating conditions. The limiters
indicate when certain operating parameters are exceeded, such as temperature, torque, and engine rotating speeds of
the gas generator and propeller.
Specifications of these engines are given
in reference [8] The main comparison
between each engine was based on power
output and total weight of engine and
gearbox. Figure 9 shows the power output vs.
weight, while Figure 10 shows the output
RPM vs. weight. Notice that as the propellor
RPM is reduced, weight increases. It stands
to reason that if the lightest engine were
geared down to the lowest RPM (1700), its
weight
would
increase
significantly.
Therefore, it was decided to keep the original
engine for use with the hybrid system since
there is no apparent weight or cost benefit
with a new engine.
Figure 10. Propeller RPM vs. Weight
IV. Solar Power
Typically, solar panels use semiconducting materials to
capture photons of light. When the light photon has the
proper energy, the panel creates an exaction, which is an
electron paired to a positively charged electron vacancy called
a hole. This pair is then separated, and these separate entities
are collected at the electrodes. There are three main types of
solar panels in use today: Monocrystalline, polycrystalline,
and amorphous. Monocrystalline solar panels are made from
a single large crystal and are the most efficient, yet the most
expensive. Polycrystalline panels [8] are very similar except
instead of being made from a single crystal, they are made Figure 11. Prospective Solar Panel Locations
from several small crystals. These are the most common
types of panels; however, they are slightly less efficient than
the Monocrystalline panels [6]. Amorphous or “thin film”
solar panels are made by spreading silicon on large plates of
stainless steel and are cheaper to produce and much lighter
than the other makes of solar panels. After investigating the
properties and characteristics of each type of solar panel, it
was determined that amorphous solar panels are the most
feasible to work on an aircraft.
This decision was made
based on light weight, shape, and durability. Average
specifications for these types of solar panels are given in
Table 2. The practical efficiency of solar panels ranges from
Figure 12. Prospective Solar Paint Locations
0.05 to 0.2.
Solar paint is a new technology that implements nanoparticles called
quantum dots into polymers thus producing a material able to detect and Table 2. Estimated
generate energy from light in the infrared spectrum. These quantum dots can Specifications for Solar Power
be made from several materials including lead sulfide and cadmium solenoid. Options Solar Panel Solar Paint
There are several benefits of solar paint over solar panels. The ability to Efficiency
10%
14%
2
Weight
3.5 Kg/m 4.54kg/gal
$100/.155m2 $100/gal
Cost
6
capture light in the infrared spectrum allows solar paint to produce energy on cloudy days, in the shade, and even at
night. This means that the extra weight and costs of implementing this technology won’t go to waste during night
flights as it would with solar panels. Also, since it is applied like normal paint, it allows for easy coverage of the
fuselage and other obscurely shaped areas where solar panels cannot be applied. An overview of the specifications
of solar paint is tabulated in Table 2. Today’s practical efficiencies stand around 10-20% while the near future
capabilities are projected to be around 35-45%.
Figure 13 shows the power output that can be
expected based on varying efficiencies.
Weight in the aircraft industry is extremely
important. Therefore, a weight analysis of each
technology was completed and graphically
displayed in figure 14. The solar paint is able to
cover the entire plane (figure 12) whereas the
panels cover only the wings and horizontal tail
(figure 11). But, the solar paint does not add much
more weight than that of the solar panels. Also, for
this small addition of weight compared to the solar
panels, solar paint generates significantly more
power than the solar panels. However, since cost is
the number one priority of most companies, the
Figure 13. Power vs. Efficiency of Solar Options
analysis would not be complete without a cost
examination, displayed in Figure 15.
SELECTION: Amorphous or “thin celled” solar
panels were the only perceivable type of solar
panel able to be used on an aircraft. Other panels
were too thick to be mounted on a wing and would
have disrupted the flow and therefore the lift
generation over the wing. Amorphous panels are
capable of being de-iced without any problems
which is a necessity for higher altitude flying;
however, dirt/soot build-up can negatively impact
their power generation capabilities. Amorphous
panels can also withstand the changes in pressure
and temperature as the plane climbs to cruise and,
since they are not made of crystal that can easily
break; they are much more rugged than the other
types of panels. However, problems may arise Figure 14.
with how to bond this technology to the aircraft
skin and keeping it bonded throughout the extreme
conditions associated with flight. With all of this
comes extra maintenance costs and other issues,
such as wiring the system, that were not taken into
consideration. Even with these matters aside,
overall with the high cost, minimal power
generation, and additional weight makes the use of
solar panel technology infeasible for aeronautical
applications at this time. In the future, once the
efficiency can be raised substantially, this would be
a key technology to look into. Solar paint is much
more practical than any type of solar panel. Its
ability to use light energy from the infrared makes
solar paint a very ideal choice when it comes to
solar energy, especially for planes. Since it can
produce energy on cloudy days and even at night,
the added weight won’t be going to waste during
Figure 14.
7
Weight Added by Solar Options
Cost of Solar Options
night flights as it would be for solar panels. Also, since planes are painted anyway, it adds very little weight, and if
technology advances enough to remove the layering process needed to make it generate electricity, the added weight
could potential reach zero. Solar paint works just like regular paint so durability, high pressure and altitude
characteristics, performance, and maintenance issues remain unchanged. The only down side is that dirt/soot build
up can decrease the efficiency of the paint. Overall, solar paint seems to be the way to go in the future of renewable
energy for aircraft despite the fact that current technology cannot efficiently produce the paint for commercial use.
V. Thermo-Electric Generator
The Thermo-Electric Generator’s (TEG) function is to capture the energy being wasted in the exhaust gases and
temperature. By lining the exhaust manifold with TEGs, the difference in temperature between the exhaust and the
ambient atmosphere creates flows of electrons and creates a voltage difference, as shown in Figure 16. This
happens based on the Seebeck effect, which provides coefficients for many different materials that can be used to
find the approximate voltage that will be created. This is illustrated in the following diagram, the heat source is the
hot exhaust gases in contact with the muffler and the cool source is the ambient air at cruise and takeoff. The
Seebeck coefficients and the voltage difference are the key factors in the productions of energy using a TEG.
Equation 1 is used to calculate the output voltage of a TEG:
V  (S B  S A ) * (T2  T1 )
(1)
The exhaust temperature and the Seebeck coefficients must be known to solve the equation. The program
AEDsys is used to determine what the exhaust temperature is for the engine that will
be used in the design of the aircraft. For an altitude of 20,000 feet, a cruising speed
of Mach=.48, the cruising temperatures were found. T9/T0 (exhaust temperature
over ambient temperature) is given in the output as 2.5437. Using the ambient
temperature at 20,000 feet the temperature of the exhaust at cruise is found to be
632 degrees Kelvin and the ambient temperature at this height is 248 degrees
Kelvin. The same process is done for the temperature at takeoff. The temperatures
that were found for takeoff were 658 degrees Kelvin in the exhaust and 288 degrees
Kelvin for the ambient air. The difference in temperature at cruise is 394 degrees
Kelvin and the temperature difference at takeoff is 370 degrees Kelvin.
The design of a good TEG is largely a materials problem. The correct materials
have to be found for the application at hand. The materials must be able to
withstand the high temperature differential that they are being applied to. They Figure 16. TEG Diagram
must have a high melting point and be able to withstand a lot of stress. Two
materials were chosen to have a high difference in their Seebeck
coefficients. The materials that were chosen are Silicon and Carbon, and
their Seebeck difference is 437 µV/C. This means that for every degree
difference, 437 µV is produced with these materials. The approximate
size of one unit was found to be 0.19x.09 inches using the total length as
0.75 inches. These dimensions were then used to determine how many
could fit around the engine exhaust. The circumference of the exhaust is
estimated to be 3.2 feet in diameter. Using these dimensions, it was
found that 404 units could be placed around the exhaust. For one loop
around the muffler it is possible to create a maximum of 65 Volts at
takeoff and 67 volts maximum at cruise. The resistance of the whole
system and the lithium-ion batteries that were used has a very low
Figure 17. Example of a TEG [9]
resistance (around 0.2 for the batteries and a few ohms for the system).
Testing will need to be conducted to determine the heat loss in a system such as this and the realistic power output.
A resistance value of 5 Ohms was used for calculations. The energy extracted from the exhaust will be used to help
charge the batteries and ultimately provide power to the electric motor.
VI. Final Selection
This section will discuss overall system requirements and integrating the various components to come up with a
workable propulsion system. Calculations will be shown to evaluate the aircraft’s performance characteristics and
8
ensure a realistic system has been achieved. There is a distinct gap between currently available technology and
future technology. For this reason, two versions of the hybrid system will be evaluated.
A. System Requirements
The propulsion system of the Pilatus PC-12 is required to have a total output of 875 kW available for take-off
and climb and 750 kW available for cruise. Therefore, the combination of gas turbine engine and electric motor
output must not be less than either of these. The useful load of the aircraft is approximately 1800 lbs. This is
normally used for passengers and baggage, but some will be used for the installation of additional hybrid
components. Consideration should be taken to ensure there is adequate weight capacity for passengers. The range is
1700 miles and the fuel weight is 2783 lbs.
Component details from previous sections of this report are
summarized in Table 4.
B. The Hybrid System with Current Technology
The Symmetron P-2 AC motor is currently available for production, and has a power to weight ratio that is
powerful enough to meet the requirements while staying fairly light. Symmetron also makes a lightweight controller
for its motors, which includes all electrical system requirements, such as a power inverter. We considered various
sized motors and their corresponding system requirements, resulting in a useful load approximation. The 100 kW
was chosen to be the best choice, because it results in the best allowable useful load, 1275 lbs., which allows enough
room for the aircraft to carry 6 people, including the pilot. The amount of battery required was calculated based on
take-off and climb alone. No alternative power sources were considered for this version of the aircraft, since TEG’s
have not come about yet, and solar technology is too heavy to be feasible at this time. This means that the batteries
would be completely drained after climb, rendering the electric motor unusable during cruise. Therefore the only
fuel savings comes from the take-off and climb portion of flight, and it can be calculated by multiplying the motor
output power by the engines SFC and the time duration. The fuel saved is calculated to be 18 lb; including this fuel
savings in the weight addition brings the useful load to 1294 lbs. Estimating a company to use the aircraft for 2
flights per day, 5 days per week, 52 weeks out of the year with Jet A at $3.50/gal yields an annual savings of about
$5450 in fuel. The cost of adding these components is about $150,000 for the batteries, and an estimated $10,000
for the motor and controller. This means it would take almost 30 years for the technology to pay for itself in fuel
savings. It is concluded that this current technology version of the hybrid aircraft is infeasible until alternative
power sources can be taken advantage of.
C. The Hybrid System with Future Technology
The future technology version allows the use of the Table 3. Component Details
lightweight DHARMA DC motor.
Again, we
considered various size motors and their corresponding
Component Details
system requirements, resulting in a useful load Parameter
Value
Units
approximation. Again, 100 kW of motor power was
AC Motor Power/Weight
0.43 kW/lb
chosen as the optimum value, with current technology
0.81 kW/lb
allowing 1400 lbs of useful load, enough for about 6 DC Motoer Power/Weight
91 W-hr/lb
passengers. Adding in the alternative power sources Lithium-Ion Power Storage/Weight
15 kW
yields an additional 61.5 kW to the electrical system, TEG Power
which can be used to run the motor and/or charge the Photovoltaic Paint
46.5 kW
battery. With this additional power available to the
motor, not as much battery power is needed. Assuming
only half the battery calculated previously still leaves an Table 4. Hybrid Component Additional Weight and
additional 11.5 kW (50+61.5-100) of power, which is Cost
acceptable since some of the power sources were
Additional
estimates. The motor is assumed to operate at full
capacity, 100 kW, during take-off and climb, and at Component
Weight (lb) Cost (USD)
about 60%, or 60 kW, during cruise. This is a DHARMA DC Motor
123.5
20000
reasonable assumption, since operating a motor at peak
Engine (keep same engine)
0
0
power periods can cause overheating and premature
137.5
75000
damage. The alternative power sources can then be used Battery (Li Ion)
Photovoltaic
Paint
272.4
5000
for both running the motor and recharging the battery in
flight. A fuel savings calculation can be done similar to TEG
25
20000
the one earlier, this time accounting for both take- Fuel Savings
-233.4
0
Total
9
325
120000
off/climb and cruise. A summary of the components additional weights is found in Table4. The fuel saving is
calculated to be 233.4 lb.
This total added weight, including fuel savings, gives a final useful load of 1450 lbs, enough for 5-6 passengers
and a pilot. With a similar estimate of 2 flights per day, 3 days a week, 50 weeks per year and Jet-A at $3.50 /gal, an
annual fuel cost savings can be estimated at $41,000. If Jet-A increased over $5.00 per gallon, fuel savings could
easily reach upwards of $60,000 per year. This means that the hybrid modifications to the PC-12 could pay for
themselves in fuel savings within 3-4 years. Also, offsetting 11,000 gallons of fuel each year, per aircraft, could
have significant, positive environmental impact.
VII. Conclusion
An optimum hybrid-electric system was designed for a turbo-prop aircraft. In terms of cost, weight, and fuel
efficiency, the hybrid system is comparable to the traditional system based on Jet-A fuel, The hybrid system,
however, has less pollution effects In the near future, it is expected that the initial and running cost of the hybrid
system will go down while that of the fossil fuel will go up. Therefore, it is believed that the hybrid system will soon
be advantageous over the traditional system.
REFRENCES
[1]http://en.wikipedia.org/wiki/Series_hybrid#Series_hybrid
[2]http://www.efunda.com/DesignStandards/sensors/thermocouples/thmcple_theory.cfm
[3] www.rasertech.com
[4]afefoundation.org/v2/.../alan.cocconi.optimized.electric.drive.systems.pdf, Alan Cocconi, April 2008,
[5]http://upload.wikimedia.org/wikipedia/commons/4/4e/Turboprop_operation.png (picture of the engine)
[6] www.janes.com
[7]https://engineering.purdue.edu/AAE/Research/Propulsion/Info/comb/jets/tjets/images/jets/tprops/pt6a.gif
[8] www.power-talk.net
[9]http://en.wikipedia.org/wiki/File:Thermoelectric_Generator_Diagram.svg
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