Student paper on the hybrid propulsion system for UAV

1. 50th AIAA Aerospace Sciences Meeting including the New Horizons Forum and Aerospace Exposition AIAA 2012-0147
09 - 12 January 2012, Nashville, Tennessee
SOLSTICE: Standalone-electric Optimized Lifting System,
Transitional Internal Combustion Engine
M. Cui, T. Drake, A. Kreuter, G. Kutil, B. Miller, C. Packard, M. Rahimpour and G. Soin 1
Aerospace Engineering Sciences, University of Colorado, Boulder, Colorado, 80309
This paper will focus on the innovative design architecture of the SOLSTICE hybrid
aircraft propulsion system. The goal of this project is to design, manufacture, integrate and
test a hybrid combustion/electric propulsion system to be integrated with a blended wing-
body aircraft currently being designed by Hyperion, an international team led by graduate
students at the University of Colorado. The concepts of operation will be discussed as well in
order to elaborate on the development of a transitionable power control system, allowing the
coaxial combustion engine and electric motor to operate in tandem or separately. Based on
requirements provided by the Hyperion graduate team, the propulsion system is designed in
part around power and thermal requirements. Analytical modeling of the engine system
with respect to these facets has been performed and multiple testing scenarios will be run in
order to validate these models. Verification of the hybrid propulsion system will be
performed on the ground, setting flight testing on the Hyperion aircraft as a stretch goal.
The SOLSTICE team plans to provide its customer, Dr. Jean Koster, with design
recommendations for future hybrid aircraft propulsion endeavors, moving aircraft
propulsion into a greener, fuel economy-friendlier realm.
Nomenclature
DAQ = data acquisition system
EM = electric motor
ESC = electronic speed controller
HPS = hybrid propulsion system
ICE = internal combustion engine
RPM = revolutions per minute
I. Introduction
T HE Standalone-electric Optimized Lifting System, Transitional Internal Combustion Engine, or SOLSTICE,
represents a turning point in how aircraft
will be flown in the future. The goal of this
project is to design, manufacture, integrate and
test a hybrid combustion/electric propulsion
system specifically for small aircraft. Currently,
hybrid propulsion systems for aircraft are nearly
nonexistent. However, the possibilities for their
uses in multiple fields of aviation would allow
for many performance benefits. An aircraft that
is able to utilize a standard combustion system
during normal flight and then transition to a
much quieter electric motor during landing
would greatly alleviate noise pollution, rampant
at and around today’s airports. A transitional
hybrid engine would also allow for greater fuel Figure 1. Hyperion Blended Wing-Body Aircraft Design.
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Undergraduate Students, Aerospace Engineering Sciences Dept., Engineering Center 1111 Engineering Dr, AIAA
Student Members
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American Institute of Aeronautics and Astronautics
Copyright © 2012 by the American Institute of Aeronautics and Astronautics, Inc. All rights reserved.

2. efficiency (similar to a hybrid car) by decreasing fuel expenditure, allowing the combustion engine to operate more
efficiently while the electric motor provides any remaining power necessary for flight.
The SOLSTICE HPS is one facet of a larger program known as Hyperion. The Hyperion team is an international
amalgamation, comprised of students from the University of Stuttgart in Germany, from the University of Sydney in
Australia, and from the University of Colorado, represented by both undergraduate and graduate teams. Under the
leadership of the CU graduate team, Hyperion is designing a model blended wing-body aircraft, seen in Fig. 1, with
the intent of developing a cleaner, greener and quieter aircraft system. The SOLSTICE undergraduate team is
responsible for the propulsion system of this new aircraft concept.
The Hyperion team plans to fly the modeled aircraft in multiple flight modes, including cruise, quiet and dash,
through the utilization of the SOLSTICE hybrid engine design. To achieve these different operational modes, the
SOLSTICE engine controls two motors within the hybrid system either individually or in tandem: dash mode
utilizes both the internal combustion engine and the electric motor while quiet mode uses the electric motor only.
Cruise uses the combustion engine alone. SOLSTICE is designing the control system to operate these flight modes
along with the physical HPS.
II. System Configuration
The SOLSTICE hybrid propulsion system makes use of two separate engines, one electric motor (EM) and one
internal combustion engine (ICE) to drive the propeller. A patent-pending gearbox converts the input torques of
these two engines into one output torque, turning the propeller shaft. Power sources for the EM and ICE consist of
batteries and fuel, respectively. By utilizing the EM and ICE either concurrently or independently the various flight
modes can be achieved such as EM only for quiet operation or EM and ICE concurrently for maximum power. The
entire system, excluding power
supplies, is placed on a single
base plate to allow for ease of
integration and structural
rigidity. The overall system is
constrained in mass and volume
as it is necessary for the HPS to
fit inside the Hyperion aircraft.
During the fall semester design
phase, multiple analytical models
were produced (thermal,
structural stress and strain, and
power) in order to ensure that the
system will provide the
necessary power to successfully
fly the Hyperion aircraft. Figure
2 illustrates the propulsion Figure 2. SOLSTICE Hybrid Propulsion System Configuration.
system configuration to be
manufactured in the spring semester, 2011.
This configuration was designed iteratively over the course of the fall semester utilizing a systems engineering
approach, working from broad to specific requirements. Initially the team worked with advisors and clients to
develop a preliminary goal and set of objectives. As is common practice, top level requirements were devised based
on these objectives and on certain parameters such as the overall aircraft mass restrictions. During the Preliminary
Design phase, these requirements were broken down to a subsystem level where enough detail was defined in order
to narrow down selection of components. Trade studies were performed and the design continued with those
selected. The overall system architecture was continually observed, allowing for a constant systems engineering
perspective. The results of these studies and their impact on the design were presented to a panel of aerospace
engineering faculty in a Preliminary Design Review (PDR). In the following Critical Design phase, parts were
selected and prototyping of major risks in order to find mitigating solutions was performed. During the Critical
Design Review (CDR) modifications to the preliminary system architecture were presented by the team as a part of
the complete system, its components and the analysis behind their design/selection.
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3. This educational process goes hand-in-hand with aerospace industry practices. Project definition is the first and
foremost step to the design process even with the existence of a multitude of unknowns at the time. This process-
and the process to follow in the spring semester- has invaluably prepared the SOLSTICE team for any future
engineering endeavors.
III. System Modeling
Modeling physical and conceptual systems is a necessary process within engineering. The ability to predict
performance increases the probability of successfully completing a project. Prediction through modeling can
produce a more thorough understanding of a system’s characteristics and interdependencies. In most situations,
modeling, combined with testing, generates this understanding earlier and more efficiently than testing alone.
Increasing efficiency not only allows a project to be completed in a shorter period of time, but also decreases the
amount of funds necessary for project completion.
A. Power Modeling
SOLSTICE’s main requirement from Hyperion, derived from the desired multiple flight modes, is that either
the electric motor or the internal combustion engine must supply the aircraft with 2 horsepower. During the design
phase of the project, a simple, yet
effective, power model was
developed to ensure SOLSTICE’s
HPS could meet Hyperion’s
requirement. The model takes
into account all major functional
inefficiencies to predict the
propulsion system’s power
Figure 3. Power Model Component Efficiency. output. The major functional
inefficiencies within the system
include friction losses from gearing and losses within the electric motor linked to converting electrical power to
mechanical power. Figure 3 represents the structure of the model visually. It is predicted through analysis that the
gearing is 75% efficient with a confidence of
70%. The confidence is based on the possible
range of coefficients of friction for meshing
gears. The electric motor was assigned an
efficiency of 80% based on manufacturer
specifications and lessons learned from
previous experience.
One of the novelties of the SOLSTICE Figure 4. Aircraft Flight Modes in Concept of Operations.
engine is the capability of running on
different modes. This ultimately allows the aircraft to also fly at different flight modes and can be used to optimize
and limit fuel consumption. An example of how the variable flight operations can be utilized in an aircraft is
shown in Fig. 4.
Figure 4 shows how an aircraft can take advantage of the engine to achieve periods of maximum velocity,
minimal fuel consumption, and regular flight operations. This concept was the foundation of the project: to
conceive and build an engine capable of performing each of these flight modes. Each flight mode has a power
output necessary to fly the
Table 1. Required Propeller Output for Flight Modes
aircraft as desired. The
Flight Mode Power (W) Velocity (m/s) Ideal Prop RPM
power output needed for
Takeoff 1500 >7.6 1795.3
each flight mode is
Cruise 750 24.3 5740.2
summarized in Table 1.
Quiet 410 13.6 3212.6 The output power is the
Dash 1500 32.4 7653.5 power required from the
Landing 460 9.9 2338.6 propeller.
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4. B. Thermal Modeling
Of major concern in integration of the hybrid propulsion system with the aircraft is the generation of heat by the
system. Because the propulsion system is located within the fuselage of
the aircraft, the heat generated by the system will heat the surrounding
aircraft structure and skin. The skin of the aircraft is fabricated from a
fiberglass-epoxy material that must remain below 60oC in order to avoid
material softening. To determine the thermodynamic characteristics of
the system, two assumptions were made. The first assumption is that all
power lost from inefficiencies in the system is fully converted to heat.
The second assumption is no force convection over the system, only free
convection. These two assumptions provide a worst case scenario of the
propulsion system during operation. Through lumped analysis, it was
determined that the system would produce 700W of heat via the three
mechanisms of heat transfer. The heat transfer present within the engine
cavity is shown in Fig. 5.
In order to mitigate the risk of weakening the fiberglass epoxy’s
integrity, forced convection over the system will be required. This shall
be accomplished by implementing a duct into the nose of the aircraft.
Further analysis of the system showed that a relation between the
necessary forced convection properties, such as the convection
Figure 5. Heat Transfer Model. coefficient and the flow speed, could be determined as a function of the
total heat transfer. The relation is shown in Fig. 6 which gives the
convection coefficient and flow speed required to keep fiberglass skin
under 60oC.
The model proved that as the heat transfer decreases, the required airflow over the system decreases. It was
determined that a heat transfer of 700W is the worst case scenario thus the design team is confident that passive air
cooling from the Hyperion aircraft will be sufficient for operational temperature.
150 Forced Convection Model
17
Convection Coefficient, h [W/m 2 K]
Airflow Speed, V [m/s] 16
15
100
Mass Flow Rate (g/s)
14
13
50 12
11
10
0
0 100 200 300 400 500 600 700 9
270 275 280 285 290 295 300
Heat Transfer Q [W] Anticipated Ambient Air Temperature (K)
Figure 6. Forced Convection Required. Figure 7. Necessary System Cooling.
Further modeling has provided a mass flow rate necessary for operation, based on expected atmospheric
conditions and the required interior temperature limit of 60 oC. The model predicts a mass flow rate range of 9.5 to
16.5 g/s. Currently a mass flow rate equal to or greater than 34 g/s is expected to account for the worse case
scenario and a factor of safety of 2. Figure 7 displays the mass flow rate’s dependence on the anticipated
freestream temperature range. The model assumes steady state flow, the air is perfectly mixed within the system’s
cavity, and heat transfer is limited to convection within the cavity and conduction through the aircraft’s skin.
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5. IV. System Verification
The next step following system modeling is performing tests with real hardware and software. The data
obtained from system and subsystem tests can then be analyzed and used to verify the different system models. In
order for these tests to represent the models accurately, the hardware including electronics and software used
during the tests should be the actual flight hardware and not prototypes. System verification tests are the final step
before the overall system design can be validated and approved for flight. As such, it is essential that these tests be
performed following detailed laid out procedures as well as be repeated multiple times to obtain accurate data.
A. Power Testing
One of the primary requirements that the SOLSTICE HPS has to supply a minimum of 2 horsepower from its
Electric Motor and Internal Combustion engine. These power
requirements are essential for the Hyperion aircraft to achieve flight
and also to verify their Concept of Operations. The SOLSTICE team
performed tests on its HPS and used the data to verify the power
output from the system. The primary apparatus used for these tests was
a reaction force dynamometer which utilizes a force transducer to
measure the reaction torque from the HPS. This is coupled with a
voltage expander and an RPM sensor which together measure the input
power to the system. The data obtained from these two sensors are
then analyzed and compared in order to obtain the efficiency curves
for the different components of the HPS. The setup for this apparatus
can be seen in Fig. 8.
100
95
SOLSTICE EM Efficiency [%]
Figure 8. Dynamometer with EM 90
Attached.
Figures 9, 10 and 11 represent the efficiency 85
and power curves for the most critical
components of SOLSTICE's HPS namely the 80
Electric Motor and the propeller. Using the
dynamometer, SOLSTICE has been able to 75
quantify the efficiency of both these EM Efficiency
components. Polyfit
70
Figure 9 shows the efficiency curve for the 4000 4500 5000 5500 6000 6500 7000 7500
SOLSTICE Electric Motor plotted against its RPM
RPM. The test was performed between a range Figure 9. EM Efficiency vs. RPM.
of 5 to 20 Amps at intervals of 5 Amps in order
to replicate the desired flight conditions for the EM. The design specifications for the EM provided by the
manufacturer rate the efficiency to be approximately 80%. From the EM dynamometer test, this specification was
verified and the efficiency of the motor was approximated to be 85% under the desired flight operation conditions.
In order to verify that the SOLSTICE Electric Motor has the required power to satisfy the given power
requirement, another dynamometer test was performed to quantify the power output of the motor. This was again
plotted against RPM and is shown in Fig. 10. Due to structural limitations of the SOLSTICE Dynamometer, the test
could not be performed at the optimum range of input current for the EM. However, the test performed provided the
relation through which the SOLSTICE team could then extrapolate the power output at higher current values.
Through these analysis, it was determined that at an RPM of 8750, the SOLSTICE EM will output 1500 Watts of
power which includes the 85% EM efficiency determined from the previous test. This power output is equivalent to
the 2 horsepower requirement. Hence, with the help of this test and its analysis, SOLSTICE has been able to
successfully verify its power requirements.
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6. 800
Although, the power requirements
give to SOLSTICE by the Hyperion
700 team accounted for a propeller
efficiency of 50%, SOLSTICE did
SOLSTICE EM Output Power
600 another dynamometer test to verify the
efficiency of the actual flight propeller.
500 This was done in order to be thorough
and obtain accurate power curves for
400 the HPS.
Figure 11 shows the efficiency
300 curve of the flight propeller plotted
against RPM along with a least squares
200 regression curve denoted as "Polyfit".
The profile changes because
100
SOLSTICE is using a single speed
4000 4500 5000 5500 6000 6500 7000 7500 propeller which is designed to have a
RPM
Figure 10. EM Output Power in Watts(Including EM Efficiency) maximum efficiency at a single point.
vs. RPM.
The approximate propeller efficiency 70
determined from this test is 55% to 65%
which is higher than what the Hyperion 60
requirements accounted for. Thus, the analysis
for this test further strengthens the verification 50
of the SOLSTICE power requirements.
Efficiency [%]
40
30
20
10
Propeller Efficiency
Polyfit
0
1500 2000 2500 3000 3500 4000 4500 5000 5500 6000
RPM
Figure 11. Propeller Efficiency vs. RPM.
B. Thermal Testing
To verify the thermodynamics and heat transfer of the hybrid propulsion system, tests are to be conducted to
verify the requirements set forth by the structural integrity of the aircraft. The tests include measuring the thermal
output from each component in the system. The major
components that must be considered from a
thermodynamic perspective are the internal combustion
engine, electric motor, electronic speed controller,
gearbox, and electronic circuits. Each component has
inefficiencies where power losses generate heat. To
experimentally determine the heat transfer of the
system components, a test box is utilized that has the
same geometry as the engine cavity within the aircraft.
The test box also utilizes a polycarbonate material that
has similar thermal properties as the fiberglass skin of
the aircraft. A representation of the test box is shown in
Fig. 12. In addition, the location of the temperature
Figure 12. Thermal Test Bed Setup. sensors placed throughout the box to measure the
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7. change in ambient temperature is shown. For the tests, seven LM 34 temperature sensors are to be used that
measure the ambient temperature in Fahrenheit. Using the Fahrenheit sensors allows for a better signal to noise
ratio than that of the Celsius LM35 sensors.
The first thermal test performed gave insight into how hot the engine cavity would get during the taxi portion of
the flight. To simulate this, the internal combustion engine is to run on idle with the engine completely sealed
within the box. During taxi, the propeller wash is negligible and is considered zero to test a worst case scenario.
The test results, shown in Fig. 13, proved that over a span of 17 minutes, the temperature throughout the engine
cavity does increase with the engine in idle. The maximum temperature measured during the 17 minute idle was
40oC located directly above the idling engine.
This is still below the maximum 60oC required be 45
Sensor 1
the aircraft structure. When the aircraft takes to Sensor 2
40
the air, the propeller will generate prop wash that Sensor 3
will flow over the system and provide passive air Sensor 4
35 Sensor 5
cooling.
Ambient Temperature [C]
Sensor 6
Testing of the electric motor and electronic Sensor 7
30
speed controller are still to be performed however
from preliminary dynamometer testing, the team
25
is confident that these component will not
generate any substantial heat and any forced
20
convection will be sufficient for the required
cooling. In addition, the gearbox will use
lubrication that has the appropriate viscosity that 15
will ensure minimal heat generation. Finally, a
full system test is to be performed several times 10
0 2 4 6 8 10 12 14 16 18 20
that will follow the aircraft mission and concept Time [min]
of operations. This test will be performed a Figure 13. ICE Idle Results, No Forced Convection.
multitude of times to ensure reliability in the
thermodynamics and heat transfer of the hybrid propulsion system to the aircraft.
V. System Applications
The advantage of the SOLSTICE system comes from the ability to change between fundamentally different
types of engines. The ICE has the advantage of burning fuel, reducing the overall weight of the aircraft over time.
The EM has the advantages of quiet operation relative to the ICE and it has the ability to run at higher altitudes.
A typical use of the SOLSTICE engine will cycle the ICE and EM based on the flight regimes. The flight
regimes include a dash mode which utilizes the power of both engines to reach a maximum speed for the aircraft.
This engine mode also has the ability to provide significant thrust for takeoff through running the two engines
together. This allows the selected engines to be smaller and tailored more towards the maximum thrust needed in
cruise conditions. The next flight regime utilizes the EM only. This flight regime can be considered a quiet mode
and can be used in urban areas where noise levels are an issue. This mode can be used for unobtrusive surveillance
of targets as well. It has also been proposed that this flight regime be used as a high altitude mode, allowing the
aircraft to fly at higher altitudes than those at which an ICE is capable of operating. The ICE-only flight regime can
be thought of as a general purpose mode. By burning the majority of the stored fuel during takeoff and climb the
thrust needed in cruise would be minimized for a higher percent of the mission duration. In the future the ICE may
be made to use a bio-fuel such as biodiesel to decrease its environmental impact. This could potentially lower the
fuel cost of operations which would be further offset by the use of the electric motor.
The basic design of the SOLSTICE engine is not limited only to a propulsion system. A further application
would be as a portable generator for remote operations such as Antarctica or disaster relief. This would involve
replacing the EM with a generator and the propeller with a turbine blade optimized for power generation. The
system provides near-limitless power from the wind. This is augmented with the ICE as a traditional gas generator.
By relying on the gas generator only when there is no wind the operation would need to carry less fuel for their
power needs. In windy places the ICE could be seen as an emergency option. The combination of the two sources
would reduce the total weight of the system and make it more portable than the two systems individually.
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8. VI. Conclusion
Although the application hybrid technology in aircraft has only recently become a topic of serious interest in the
aerospace community, it is a necessary technology in ensuring that the aircraft of the future are both fuel efficient
and environmentally friendly. The SOLSTICE project provides further development in the field of hybrid aircraft
propulsion through testing and modeling to verify the validity of such a system in aircraft. Current applications of
the SOLSTICE project are limited to UAV markets due to current battery technologies. However, with an ever
increasing interest in both UAV application and lessening carbon foot-prints, the SOLSTICE system is relevant in
today’s aircraft industry. The current state of testing verifies that the EM meets the power requirements and that the
propeller exceeds expected efficiency. Through the design process of the HPS, many valuable lessons have been
learned including the importance of communication within the team and with the customer. Since the project
represents a global effort between students from various countries it is necessary that communication is clear and
concise to ensure that any problems that may arise are mitigated diligently and efficiently. By developing a system
with the potential to rely less on fossil fuels for energy this project represents a global interest in ensuring that the
future of aircraft propulsion is cleaner and safer.
Acknowledgments
The authors thank Dr. Jean Koster and Dr. Donna Gerren for their continued advice on this project. The team
also appreciates the strong support and guidance received from Cody Humbargar and the entire Hyperion team
(Scott Balaban, Derek Nasso, Andrew Brewer, Julie Price, Chelsea Goodman, Eric Serani, Derek Hillery, Alec
Velazco, Mark Johnson, Richard Zhao, and Tom Wiley). We also thank the international partner teams from the
University of Stuttgart in Germany and from the University of Sydney in Australia who helped make this project a
fascinating and unique learning experience. Funding for this project is provided in part by The Boeing Company,
eSpace Inc., and NASA under grant NNX09AF65G.
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