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How to design a fuel-powered quadcopter with 3D
printing
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Table of Contents
Executive Summary .................................................................................................................... 3
1. Introduction............................................................................................................................ 4
2. Rotor Design ......................................................................................................................... 5
3. Concept Design ..................................................................................................................10
4. Detailed Design ..................................................................................................................23
5. Prototype Results ...............................................................................................................31
6. Conclusion ...........................................................................................................................37
7. References ..........................................................................................................................39
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Executive Summary
The white paper will describe the process of designing a fuel-powered quadcopter with
the use of 3D printing. The document starts off with a short introduction and is followed
by four main parts of research, namely the rotor design, concept design, detailed design
and prototype results. The rotor design touches on the momentum blade element theory
and the beam theory where it is described how the design for the rotors was chosen and
why. Concept design talks about the control subsystem, structure subsystem and the
power generation subsystem. Detailed design is about choosing other components of
the quadcopter, for example the engine and why they were chosen. The prototype results
show data for the simulation and real test. Lastly, there is a conclusion along with
recommendations for further design.
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1. Introduction
This white paper intends to develop a design methodology to manufacture custom model
multi-rotorcraft rotors using 3D printing techniques available at Leapfrog 3D Printers. In
accordance with the above statement, the following requirements for the design
methodology will also be met. Firstly, the resulting design will be printable at Leapfrog
3D Printers. Secondly, the resulting designs will be able to withstand the operating
conditions of similar model rotorcraft rotors. Lastly, the design methodology will be easy
to adapt to rotors of different sizes. The white paper is divided into four main sections,
namely rotor design, concept design, detailed design and prototype results. The section
on rotor design explains how the rotors will be designed and why and it also touches on
the momentum blade element theory and the beam theory. The concept design breaks
down the following components of the quadcopter, namely the control subsystem,
structure subsystem and the power generation subsystem. It is explained why the final
designs have been chosen and how they influence the quadcopter as a whole. Detailed
design touches on other components of the quadcopter, namely the rotor hubs, rotor
blades, engine, transmission system and frame layout and why the final design for each
component has been chosen. Lastly, the prototype results section shows data about the
simulation and test results compared to a bigger quadcopter.
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2. Rotor Design
Momentum blade element theory
Beam theory
The methodology used to determine rotor’s characteristics is based on a simplified
version of momentum blade element theory by Ruijgrok (2009). In order to start the
design process, the wanted thrust, rotor radius, blade count, lift coefficient distribution,
airfoil shape and chord distribution have to be specified. Then, the momentum blade
element theory can be used to calculate the rotational rate and blade twist distribution
necessary to create the wanted thrust. Once these parameters are known, the
aerodynamic coefficient distributions over the rotor blades can be calculated. Since the
rotational rate is also known, it is possible to calculate the force distribution over the rotor.
This makes it possible to calculate the drag moment acting on the rotor blades and
calculating the necessary power to achieve the wanted thrust.
Furthermore, the beam theory by Hibbeler (2008) can be used to calculate the internal
forces acting on each rotor blade. Since the airfoil shape and twist angle distributions are
known, the area moment of the blade’s inertia can be calculated at every radius. By
integrating the aerodynamic and centripetal force distributions, an accurate estimation of
the maximum internal stresses on each radial segment of the rotor can be made. This in
turn makes it possible to check if the rotors will survive the required operating conditions.
Since the resulting airfoils have a curved and twisted shape, it is almost impossible to
print them using an FDM technique because they lack a flat side which can be printed
against the bottom plate of the printer. In order to solve this problem, we can transpose
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each blade element ever so slightly in order for the top and the leading edge of the rotor
blades to align over the entire radius of the blades.
Figure 1: Example blade element alignment side views. Left: before correction, right: after correction .
In order to test the design, several prototypes were made. The first prototype was
unsuccessful. We attempted to print the complete rotor at once on the largest scale
possible. A simple rotor hub was designed to connect the three rotor blades together and
the rotor was supposed to be printed on the flat upper side. This failed due to the rotor
warping upwards and detaching from the printer bed.
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Figure 2: Unfinished rotor prototype.
The second prototype was printed in multiple parts, since the first prototype failed. This
also made it possible to print the parts on the leading edge. To test what works, two
blades were made, one printed on the leading edge down and another one the same
way as the first prototype. The rotor printed on the leading edge provided far superior
results, which can also be seen in the figures below. This is due to the fact that the upside
down prototype requires support material in order to create the correct shape, while the
other component can be printed without any support material. A drawback; however, is
that it does result in a slightly rougher bottom side of the airfoil since the printing surface
is perpendicular to the blade surface.
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Figure 3: Rotor blade example, printed upside down.
Figure 4: Rotor blade example, printed standing on the leading edge.
For the final prototype a simple joint was designed to fit all blades together.
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Figure 5: Final prototype rotor example.
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3. Concept Design
Control subsystem
Structure subsystem
Power generation subsystem
In order to create the most optimal design, it is important to investigate multiple
technology directions and to do this, three different concepts will be investigated.
It is important to split up the design into several, mostly independent subsystems since
this will provide a good overview of the design space available for the concept
generation. For the model multi-rotorcraft, three different subsystems can be
distinguished, namely the control subsystem, the power subsystem and the structure
subsystem.
In terms of the control system, there are several possible options. Traditionally, the
control of the model rotorcraft is achieved by altering the lift generated by the different
rotors, which is done by changing their rotational rate. A more recent option is control by
pitch adjustment. More fine-tuned control of the rotor is possible by using more complex
variable pitch rotors (Cutler, 2010). The third option is controlling the rotorcraft by using
deflection vanes right below the rotors. This is less effective than the two previously
mentioned systems due to the limited torque generation. The other side of the control
system exists in the controller itself. Even though variation is possible here, almost all
commercial systems use the same approach. This uses a 6-Degree Of Freedom Inertial
Measurement Unit (6-DOF IMU) in combination with a Global Navigation Satellite System
(GNSS) like GPS and altitude measurement. By using simple controllers, a system like
this allows complete and accurate control of the rotorcraft.
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The structure subsystem controls the layout of the rotorcraft itself. With this system, there
is a lot of variation possible between different rotor configurations, materials, and
construction. The rotor configuration is one of the most defining aspects of a rotorcraft
design. The most common choices are the X and H frame configurations. In former
configuration, the rotors are located at the ends of several booms which extend from a
main central body. This configuration is one of the most versati le since it can handle any
amount of rotors. However, there are drawbacks in the limited space available in the
center of the rotorcraft. This configuration is more interesting for systems, which have
significant amounts of parts in the main body of the rotorcraft. While the center beam is
free for mounting miscellaneous components, the rotors are located on the tips of the
cross beams. The main drawback of this configuration is that it tends to be heavier than
comparable X frames due to the amount of material, which is needed to keep the center
frame sturdy.
The final subsystem of the rotorcraft is the power generation subsystem, which can be
further divided into energy storage, power generation and power transmission. Since
almost all model multi-rotorcrafts are electricity based, this system traditionally consists
of high-energy density batteries. These batteries can store and discharge energy very
efficiently; however, their main drawback is the relatively low amount of energy they can
store compared to systems, which rely on combustion. For endurance systems, this
component tends to be responsible for majority of the rotorcraft weight. The power
generation system is responsible for transforming stored energy into mechanical power.
With electricity-based systems this happens through the use of lightweight brushless DC
motors, which are controlled by Electronic Speed Controllers (ESCs). Another approach
is using a small combustion engine to generate mechanical power directly from the
combustion of the fuel.
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Finally, a hybrid solution exists by using a fuel-powered engine coupled with a generator
to power an electrical system. The transmission system is responsible for bringing this
mechanical power to the rotors of the rotorcraft. Many models have the rotors mounted
directly on top of the motors powering them; however, it can be an advantage to power
all rotors from one central power source. In this case, a combination of shafts and belts
can be used to connect the rotors directly to the motor. The layout of this system is closely
intertwined with the chosen structure layout.
Based on these possibilities, three different concept designs are chosen to be further
investigated. The choice in the power generation system is considered to be the most
important difference due to its impact on the rotorcraft’s endurance. The first concept
design is mainly a reference design, featuring a fully electrical drive system with one
motor per rotor. This design is the closest to the current commercially available
rotorcrafts, and is the simplest option to develop. The second concept design features a
single combustion engine as the power generation system. This system features a
mechanical transmission to power all the rotors, which means the control system cannot
rely on the speed control to control the rotorcraft. To compensate for this, pitch control
can be used to stabilize the rotorcraft. The final concept design features the hybrid power
generation system. This system combines the simplicity of commercial model rotorcraft
control systems with the endurance of fuel-based systems. Its major drawback; however,
is the weight required for complex power generation system.
In order to make an accurate comparison between different concept designs, accurate
theory for sizing different rotorcraft models must be established. In terms of the
electricity-based design, an estimation of the rotorcraft endurance can be made by
solving the equations for total mass and power consumption using data of motor
efficiency and battery energy density. Firstly, data collected from T-motor (2013) is
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analyzed by correcting the produced lift and efficiency of the motor weight and the
matching rotor. The result can be seen below.
Figure 6: Net motor lift compared to motor efficiency of several brushless DC motors for model rotorcraft. Rotors
used are the largest rotors described in the engine model datasheets .
From this graph, we can conclude that the highest currently possible efficiency by using
COTS technology is 14g/W. Further research shows that the maximum storage density
of commercially available batteries has a maximum of approximately 0.2 Wh/g. If we
assume there is constant power consumption in hover, we can use these two variables
to calculate the maximum possible quadcopter endurance by solving the following
equation:
It is important to note that in this equation, the thrust is units of mass because the thrust
data is provided in grams at sea level. By substituting the mass of the batteries and the
thrust by power and endurance, we can write the following equation:
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The final unknown is the mass of the frame and the payload over the mass of the
batteries. If we assume the weight of the frame is negligible, we can calculate a maximum
theoretical flight time of 2.8 hours independent of mass. This flight time quickly decreases
as the frame and payload mass fraction increases. A practical number can be derived
from the fact that the frame mass of model rotorcraft is often approximately equal to the
mass of the batteries, which yields a practical flight time of 1.4 hours.
The sizing of the second concept design is more difficult. Due to the motor and rotor
performance data not readily available in this sector, estimations of performance will
have to be made using mostly theoretical analysis of the rotorcraft.
Firstly, it is necessary to generate a performance estimation of the variable pitch rotors
this concept requires. By using the momentum blade element theory adapted from
Ruijgrok (2009), we can provide an estimation of the thrust and power coefficients of
rotors similar in size to the rotors used by the motors from the first concept. These rotors
have an estimated radius of 350mm, a constant chord of 33 mm and a NACA0009 airfoil
similar to the rotors of model helicopters (Cutler, 2010).
By using XFOIL (Drela, 2013), the aerodynamic characteristics of this airfoil were derived
at a tip Mach number of 0.3 and a Reynolds number of 200000, which were estimated
from operating conditions of similarly-sized helicopter rotors. By using the momentum
blade element theory and assuming a two-bladed rotor with a 50mm radius rotor hub,
this data is enough to calculate the values of the rotor power and thrust coefficients
depending on the blade pitch angle and assuming hover conditions. With performance
characteristics of the rotors being known, it is then possible to estimate the endurance of
the resulting rotorcraft design itself. By substituting the thrust required by the weight of
the vehicle as well as accounting for the presence of multiple rotors, we gain the following
relation:
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This relation indicates that for minimum power consumption, the ratio must be
maximized. For the used rotors, the ratio is equal to 1.48.
We can then calculate the endurance of the vehicle by solving the mass of the vehicle.
Because the mass change of the vehicle is caused by the fuel being burned and
accounting for the efficiency of the power conversion, we can write the following
equation:
This is the first order nonlinear differential equation and can easily be solved. The
boundary conditions of the equation have to be calculated first in order to solve the
system. The first boundary condition can be derived from the fact that at the end of the
flight, the mass of the system equals the empty mass of the system. The other boundary
condition follows from the maximum power that can be provided by the system. For
rotorcraft control, it is essential there is more power available than is being used since
control happens through thrust adjustment. For concept one, the motors are operating at
50% power, and to keep the comparison fair, the same useful power fraction will be used
for this concept. These boundary conditions can then be written down as the following
relation:
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The system can now be completely solved, which results in the following equation for
the endurance of the fuel-powered quadcopter:
As can be seen from the equation above, an estimation of the fuel-powered rotorcraft’s
performance requires knowledge about the empty mass of the rotorcraft, the maximum
power output of the engine, the efficiency of the system, and the amount of rotors present
on the rotorcraft. These factors will be estimated for the requirement of one kilogram of
payload. For simplicity of the transmission system, we will assume the rotorcraft has four
rotors. Research shows the smallest commercially available engines produce around 2
kW while weighing approximately 1 kilogram including peripherals (DLE Engines, 2015;
O.S. Engines, 2015).
To keep the comparison fair, we will use the empty mass of the first concept as an
indication of the empty mass of this concept. Since it will run at 50% power, it is able to
weigh 14 kg due to the efficiency of 14 g/W. It is assumed that half the weight of the
system are batteries, therefore, the result will be an empty weight of 7 kg including
payload.
The final, unknown, parameter is the engine efficiency. While gasoline engines can reach
an average efficiency of 20%, it is not expected this efficiency will be reached on this
rotorcraft because of the drawbacks of downscaling a gasoline engine. Fuel consumption
data on this engine size is not well documented, therefore, an assumption of efficiency
is made at 10% as a safety factor because the choice and the performance of the engine
are not yet known. As for fuel energy density, it is assumed gasoline has an energy
density of 44 MJ/kg.
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With these parameters, it is possible to make an estimation of the rotorcraft’s
performance, which results in an expected flight duration of 7.7 hours with a take-off
mass of 11.4 kg. The sizing of the third design involves a mixture of techniques from both
designs. The new components necessary for this design include a high power alternator
and rectifier. If we assume the engine size is the same as for concept two, commercial
components of this power rating would weigh at least 2 kg at a maximal 92% efficiency
(Sullivan Products, 2015). By using the mass differential equation from the second
concept and the rotor efficiency approach from the first concept, we can then derive a
new mass differential equation for this concept:
This equation has the following boundary conditions:
Solving these equations then yields:
Due to the craft’s mass being variable, it is impossible to always fly at top performance
of the fixed pitch rotors. This variation is rather small thus we can see that it is still
possible to fly with an average efficiency of 13.5 g/W. If we then assume the empty mass
is equal to the empty mass of the second concept with the added mass of the required
generator systems on top, we can calculate the total flight time of 4.9 hours. We can also
further calculate that the concept’s take-off mass is 12.42 kg.
Both concept two and three have the possibility of providing a much better performance
than what is possible with the electricity-based concept one. While both concepts have
a lower maximum take-off mass at the same power generation compared to the first
concept, they make up for this by being able to carry significantly more energy.
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Both these concepts have the potential to meet the flight time requirement, however, the
expected flight time of concept two provides a significantly higher design safety margin
than concept three. The improved efficiency due to fixed pitch rotors in concept three
seems to not make up for the mass and losses added by the generator subsystem.
With current technology, it is theoretically impossible for the first concept to meet the
required flight time so it will not be discussed any further. A big problem with concept two
is the system’s complexity. Both concept one and three have relatively small moving
parts; however, concept two requires a complicated transmission system in order to
power the rotors. This limits the freedom of structure’s design and requires a strong frame
to function well. Furthermore, it requires variable pitch rotors which bring significant
complexity to the rotor mounting itself. The development time for the rotorcraft is very
limited thus it could pose a risk for the project. The variable pitch rotors do; however,
have several advantages. The large rotors which are necessary to make concept three
efficient have significant inertia and can make control of the vehicle difficult because of
the delayed response. Meanwhile the response of the variable pitch rotors of concept
two can react faster than any fixed pitch design and can also provide negative lift.
Furthermore, no energy is wasted during maneuvering since it is not necessary to slow
down and accelerate the rotors. The variable pitch rotors also allow for a safety measure
to be in place, in case the engine of the rotorcraft runs out of fuel. In concept three, a
separate battery-powered system will have to be created in order to allow safe landing;
however, concept two has the innate possibility of using autorotation to safely descend.
Finally, concept two is a better fit for the scalability requirement. The fixed pitch rotors of
concept three cannot be scaled up efficiently because of their increasing inertia. This can
be solved by adding more rotors instead of scaling up rotors; however, this results in
reduced efficiency and an overall more complicated system. Meanwhile, the structure of
concept two can be scaled up while still having functional rotors.
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The first concept was quickly discarded because it did not meet the requirements. The
second concept easily meets the expected performance and the third concept can be
adapted to meet the goals. Since it was found out that commercial parts are easily
available to create the complicated transmission of concept two and this removed the
largest risk, it was decided to continue the preliminary design with concept two.
Once the concept design was settled on, we had to make a choice about the engine type.
For a rotorcraft of this scale, there are three major engine types, namely nitro engines,
gas engines and model turbine engines. Nitro engines are the smallest engines available
on the market. Their main feature is that they run on a special nitro fuel mixture, which
allows them to run at very high rotational speeds and this enables very lightweight engine
designs to still produce considerable power. Their main drawbacks exist in how
expensive their fuel is and the very oil-rich exhaust they produce which can cause
damage to the systems they are integrated in. The second option is a miniature two-
stroke gasoline engine. Even though it cannot deliver the extreme power to weight ratio
as a nitro engine can, they do provide significantly more efficiency and can operate on
normal gasoline. One drawback; however, is that they need a spark-plug based ignition
system, but recent developments have made this less of an issue because of the rise of
electronic ignition systems. The final option exists in model turbine engines. These
engines can run on almost any fuel; however, they require a special controller to regulate
the engine. While they can produce incredible power to weight ratio, they tend to be
inefficient (Salt, 2015). The choice was made to go with the two-stroke gasoline engine
because the project’s goal is to design a long-range quadcopter. Even though the engine
is slightly heavier, the motor’s efficiency easily weighs up to the slightly increased mass.
The largest issue in terms of the physical size is the sourcing of the variable pitch rotor
hubs required for the design to function. These components are very complex and to
produce them to order is impractical due to project’s deadlines, which means these parts
will have to be sourced commercially. The largest commercial availability of small scale
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variable pitch rotor hubs exists in the market of model helicopter’s tail rotors. Even though
the main rotors used by these helicopters are unsuitable because of their full coaxial
control, their tail rotors are an excellent fit in our case since the assemblies often already
include a right angle connection. Helicopter tail rotors normally support significantly
smaller blades thus it is important to find the largest commercially available tail rotor hub
since the efficiency of the rotorcraft increases with the size of the rotors.
This led to the Align T-rex 800E tail rotor hub being chosen as the rotor hub assembly
for this project. It is commercially available and can fit rotor blades as large as 325 mm
in length, like for example the Align T-rex 325 mm carbon fiber blades (van Natter, 2014).
Larger blades are more efficient thus this blade size was taken as the final rotor size for
the design. Next, is choosing the frame design. This choice will have a large impact on
the transmission design because it determines how axes and belts can be used to power
different rotors. The chosen rotor hub components allow an axis connection directly
perpendicular to the rotor hub so it was decided to use an H-frame with four rotors due
to the simplicity of the resulting transmission system. Furthermore, the rotors will be
located on top of the frame arms in order to prevent problems due to rotor flexibility and
will also simplify the landing gear.
With the last remaining core design parameters fixed, we can make a full estimation of
the rotorcraft design. The data below lists all data used for this design.
Airfoil (profile estimated based on
the thickness of the chosen rotor
blades)
NACA 0013
Amount of blades per rotor 2
Amount of rotors 4
Rotor radius [mm] 350
Rotor hub radius [mm] 50
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Rotor chord [mm] 33
Rotor shape Straight, no twist
Engine efficiency [%] 10
Fuel energy density [MJ/kg] 44.4
Engine maximum shaft power [kW] 2
Steady state maximum power usage
[%]
50
Engine efficiency [%] 10
Empty mass (no payload) [kg] 6
Payload mass [kg] 1
Figure 7: data used to estimate the performance of the rotorcraft during preliminary design. Rotor airfoil
characteristics were derived using XFOIL at M=0.3, Re=200000.
With this data, a relation between maximum flight time and payload mass can be
developed.
Total fuel mass [kg] 5.33
Maximum flight duration [h] 9.94
Rotor rotational rate at start of
flight [rot/min]
2677
Rotor rotational rate at end of
flight [rot/min]
2017
Rotor rotational rate at maximum
thrust [rot/min]
3373
Power usage at the end of flight
[W]
428
Figure 8: Calculated performance characteristics of the rotorcraft specified in table 7.
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Figure 9: Expected maximum flight time for variable payload mass of the rotorcra ft design specified in table 7.
Interestingly, it turns out the rotor design used in this performance estimation is
significantly better than the rotor design previously assumed. Furthermore, in the above
figure, it can be seen that the resulting rotorcraft is not only able to lift 1 kilogram of
payload for a significant amount of time, but it can also lift the payload close to the
rotorcraft’s weight for shorter amounts of time.
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4. Detailed Design
Firstly, it is important to investigate the most critical components of the design. Several
subsystems have significant impact on other components of the system. These
subsystems are the rotor hubs, the rotor blades, the engine, the transmission system
and the frame layout. While the rotor hub and blade subsystems have already been
decided on, the rest of the power systems still has to be designed.
The first choice to make is choosing the engine. The requirements for this component
come from the necessary rotational rate range, power range and the rotational range on
which the clutch engages. Due to the fact that gas helicopter engines require a clutch for
smooth start-up behavior, the availability of this component will be researched first.
Market research shows that all lightweight clutches for the wanted engine size are so-
called centrifugal clutches. They operate in a concentric fashion where the inner flywheel
expands because of the rotational rate of the flywheel until it locks against the clutch bell
at a set rotational rate. Due to the size and rotational rate of the wanted clutch, the only
commercially available models engage at 6000 RPM. This means a fast motor is
necessary, followed by a stepdown conversion in the transmission system. The only
engine available that meets these requirements is the OS GT15HZ engine model. It has
an effective range of 2000 up to 16000 RPM, a 4:5:1 stepdown in the transmission
system so it provides the rotational rates required to fly the rotors.
With the choice of the engine settled on, the transmission system can be designed. To
ease the design process, a clutch is selected that is known to be compatible with the
engine model. This clutch already comes with a 14 teeth pinion gear thus a matching
spur gear with 64 teeth is selected which results in a step down conversion of 4.57:1,
which meets the engine requirements. The frame layout is decided to be an H-frame so
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this gear will power an axis which extends through the central beam of the frame. Similar
axes will be mounted between the rotor hubs and by using miter gears where these axes
meet the central axis, the correct rotation of rotors can be ensured. In order to keep
friction in the transmission system low, it will be suspended on bearings rated for both
the rotational rate of the rotors and the axial loads placed on the axis due to the gear
joints. An overview of the transmission system can be seen below.
Figure 10: Layout of the complete transmission system including the engine.
With the critical subsystems designed, it is possible to continue with the general design
of the rotorcraft. Firstly, we will discuss the general frame layout and after that the
different subsystems. One of the core requirements is that parts of the frame will be 3D
printed. While the 3D printing technology does not offer many advantages for the central
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beam of the frame, the multitude of different interfaces present on the rotorcraft arms
make these components excellent for 3D printing.
A big problem with these arms is that they are significantly larger than the maximum size
that can be printed with the printers available at Leapfrog 3D Printers. Therefore, it was
decided to split the arm into multiple pieces that can be glued together.
Meanwhile, the frame of the center beam poses a different problem. It has to provide
support for the transmission system, and also house the components of the other
subsystems while remaining easy to produce as well as lightweight. Keeping the
requirement of upgradability in mind, it was decided to construct this component from
sparse folded aluminum sheets. This allows easy mounting of different components and
provides the strength required to house the transmission system while remaining
extremely lightweight. The engine, clutch and central transmission shafts can then be
supported through several milled pieces of aluminum mounted inside the structure.
Since the size of the central beam structure is known, it is time to focus on the joint
between the central beam and the arms supporting the rotors. Keeping maintainability in
mind, we decided this joint should be easy to unlock in order to inspect the gearing
system which will be located at this joint. A big issue; however, is accurately attaching
the manufactured aluminum structure to the rough 3D printed structure of the arm
frames. A solution to this problem is putting the 3D printed part in between the metal
parts of the joint and a small aluminum plate on the other side of the frame.
An iterative process is performed to design the 3D printed arm frames. Because of the
way 3D printed parts are manufactured, the stresses on the part will be almost completely
carried by the outside of the part, while the inside of the part contains a simple supporting
structure that prevents the outside skin from collapsing. Due to the distribution of the
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loads, a simple design is made featuring a tapered stretched cylinder design which will
not significantly interfere with the airflow of the rotors. Testing shows that an optimal
result is reached by combining a part wall thickness of 0.8 mm with a supporting structure
with 10% infill. The resulting part structure can be seen below:
Additional features include the presence of two extrusions at the arm tips which act as
landing gear, and a pocket in which servos can be mounted in order to control the
variable pitch rotors. Another interesting point is that the arm frame is asymmetrical. This
is caused by the asymmetry of the used rotor hubs, which means that the servos will
always be located on the same side of the outer rotor arm components. It also means
that both outer rotor arm parts are identical, which simplifies the production of the parts.
Figure 11: Unfinished test print of the central section of the arm frame.
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Figure 12: Overview of the resulting arm frame design.
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Figure 13: Overview of the center frame design.
The frame design is completed so the control subsystem can now be focused on. This
system can be divided into different subsystems, namely measurement, controller and
actuation.
The first system we will look at is the actuation subsystem. The rotorcraft features four
pitch-controlled motors, and one engine which can be throttle controlled. All of these
components receive their input through the movement of a mechanical lever and
therefore, a mechanical actuator is necessary. Due to the relatively high speed control
loop of the model multi-rotorcraft, these actuators are able to quickly apply feedback. In
order to achieve this, servo motors which update at 330 Hz were chosen for this system.
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The controller subsystem has to drive these actuators. In order to keep the development
time short, a COTS system was selected for the controller subsystem. This system can
be customized to fit the needs of the rotorcraft and has the capability to power all servos.
It also includes a pressure sensor and a six degrees of freedom inertial measurement
unit.
The measurement system is responsible for gathering data for the control subsystem. In
order to fulfil the requirements, several sensors are selected next to the sensors already
present in the control system. The first sensor is an ultrasonic distance sensor which can
be used to accurately measure the distance to the ground for soft landings. Furthermore,
a Hall Effect switch sensor will be included which can be used to measure the rotational
rate of the rotors due to a set of magnets mounted on the clutch. This is necessary since
the performance of the rotors significantly varies with this rotational rate, therefore, the
controller needs to be able to effectively regulate the engine. These systems run on
electrical power, the same as the electrical ignition used by the engine. While it is
possible to generate this power on board, it makes the design unnecessarily complex.
Due to this, it is decided to power the electrical systems of this prototype rotorcraft from
batteries. The current consumption of the electrical ignition is rated at 750 mA at 14000
RPM; however, the current consumption of the control system is not well known because
of the variable power draw of the servos. To compensate for this, the current sensor will
be integrated in the servo power supply to estimate the servo current draw. Until this
performance is tested, two 10 Ah two-cell Lithium polymer batteries will be used as the
energy source. The combined capacity of these batteries should be able to handle a
current draw of 4A for the required flight duration, providing a safe power budget of 0.65A
per servo. The current draw of the controller and measurement subsystems is negligible
compared to this and is therefore not included in the analysis.
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The final subsystem to be designed is the energy storage subsystem. This rotorcraft
design incorporates a gasoline engine, therefore this means a large amount of gasoline
has to be stored on board. If the full 5.3 kg of fuel are stored in one large tank, this can
cause issues due to the fuel sloshing around in the tank, making it impossible to stabilize
the rotorcraft. Another issue lies in the change of the center of gravity of the rotorcraft
due to the fuel being depleted over the duration of the light.
These issues are solved by storing the fuel in multiple small tanks located in the center
of the craft, at the sides of the center frame. Since the fuel cannot move freely around
large distances, the impact of sloshing on the stability of the rotorcraft is reduced.
Furthermore, the fuel is stored very closely to the center of gravity, which reduces the
influence of the fuel mass on the movement of the center of gravity. As an added benefit,
this also allows the rotorcraft to remain relatively flat since there is plenty of space not
covered by the rotors in this area.
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5. Prototype Results
The following estimation is made for the prototype:
Airfoil (characteristics derived
using XFOIL)
NACA 0009
Amount of blades per rotor 2
Amount of rotors 4
Rotor radius [mm] 350
Rotor hub radius [mm] 50
Rotor chord [mm] 33
Rotor shape Straight, no twist
Engine efficiency [%] 15
Fuel energy density [MJ/kg] 44.4
Engine maximum shaft power
[kW]
2
Steady state maximum power
usage [%]
50
Engine efficiency [%] 10
Engine mass [kg] 1
Rotor and rotor hub mass [kg] 0.25
Frame mass [kg] 3
Figure 14: Prototype estimation.
Based on this information, the simulation gives the following results:
Maximum fuel mass [kg] 7.3
Maximum flight duration
[h]
19.0
Rotor RPM when full
[rot/min]
2603
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Rotor RPM when empty
[rot/min]
1554
Figure 15: Simulation results.
This helps establish the following relation between the payload weight and flight duration:
Figure 16: Relation between the payload weight and flight duration.
Then, we performed these same tests on a larger 12 kW rotorcraft and these are the
results:
Airfoil (characteristics derived
using XFOIL)
NACA 0009
Amount of blades per rotor 2
Amount of rotors 4
Rotor radius [mm] 860
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Rotor hub radius [mm] 120
Rotor chord [mm] 81
Rotor shape Straight, no twist
Engine efficiency [%] 22
Fuel energy density [MJ/kg] 46
Engine maximum shaft power
[kW]
12
Steady state maximum power
usage [%]
50
Engine mass [kg] 10
Rotor and rotor hub mass [kg] 1.5
Frame mass [kg] 18
Figure 17: 12 kW rotorcraft test results.
The simulation gave the following results:
Maximum fuel mass [kg] 34.5
Maximum flight duration
[h]
26.9
Rotor RPM when full
[rot/min]
1058
Rotor RPM when empty
[rot/min]
745
Figure 18: 12 kW rotorcraft simulation results.
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This helps establish the following relation between the payload weight and flight duration:
Figure 19: Relation between the payload weight and flight duration.
These results show that quadcopters that are fuel-powered have the potential of having
a much longer flight time when compared to an electrical multi-rotorcraft since those are
limited to a flight time of maximum two hours. This is due to an increased amount of
energy that can be stored in the fuel as well as the sharp rise in flight time at low payload
mass implies there is another effect present. This rise can be made more clearly visible
by looking at the power consumption of the craft during the flight.
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Figure 20: Relation between the rotorcraft’s power use and flight time.
The above graph shows that the rotorcraft’s power use drops significantly as the flight
time increases. Power consumption of electrical multi-rotorcraft is, on the other hand,
constant. The explanation for this is that fuel-powered rotorcraft loses mass while flying,
while electrical rotorcraft has to carry the batteries even when they are empty. This,
therefore, doubles the efficiency of the rotorcraft near the end of its flight time, which
allows these rotorcrafts to have an extremely good endurance.
Another element to consider is the rotorcraft’s maximum range. However, due to the
aerodynamics of the multi-rotorcraft this is hard to estimate because of the complex
interactions between the multiple rotors and the rotorcraft body itself. It is possible to give
an estimation of the range at a low speed by comparing it to other quadcopters of the
same size; however, the effects of quadcopters’ forward flight on their power
consumption is not very well understood. Due to the variable pitch rotors, the rotorcraft
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should be able to maintain reasonable efficiency at higher forward velocities, however, it
is hard to estimate the effect it will have on the power consumption and the stability of
the rotorcraft. If looking at the performance of a normal fixed-pitch rotorcraft, it is definitely
possible to fly 50 kilometers per hour without experiencing major drawbacks in power
consumption. This makes it possible for the prototype design to have a range of nearly
800 kilometers while the large scale design could travel over 1000 kilometers. These
numbers would; however, decrease because of the payload in the same way as the
maximum flight time decreases.
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6. Conclusion
This white paper has proven it is possible to construct a rotorcraft that meets all the
requirements mentioned at the beginning of the report. The designed rotorcraft has the
potential to fly with one kilogram of payload for more than nine hours using power
provided by a gasoline engine.
To ensure the project results in an operational rotorcraft that meet the requirements, it is
essential the recommendations below are followed. These recommendations are based
on problems encountered during the construction of rotorcraft’s parts, as well as
concerns raised during the design of the rotorcraft.
In terms of structural integrity, it is important to analyze the expected deformation of the
frame under maximum load. Next, it is important to investigate the effects of continuous
operation on the 3D printed components of the frame. Furthermore, different components
of the transmission system have to be further researched.
When it comes to vibration, the possibilities of damping between the frame and the
controller have to be researched. It is also important to measure vibrations caused by
the engine and the rotors. Finally, a vibrational analysis on the frame has to be
performed.
In terms of performance, it is important to analyze the thrust losses due to downwash on
the frame, investigate the use of an expansion chamber on the engine, analyze the actual
efficiency of the engine and investigate the effects of forward flight on fuel usage.
With control, there has to be an investigation of different controller possibilities for
regulating engine performance and it is important to integrate GNSS or remote control
into the design.
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When it comes to enhancements, an aerodynamic cover for the center beam has to be
designed as well as an inbuilt starter instead of requiring the use of a manual starter for
the engine.
The aim of these recommendations is to aid future development based on this prototype.
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7. References
DLE Engines (2015). Small gasoline engines. http://www.dle-engines.com/index.html
Ger J. J. Ruijgrok (2009). Elements of airplane performance. VSSD
J. Salt (2015). Choosing the best RC helicopter engine.
http://www.rchelicopterfun.com/rc-helicopter-engine.html
M. Drela. (2013). Xfoil- an analysis and design system for low Reynolds number airfoils.
http://web.mit.edu/drela/Public/web/xfoil/
M. J. Cutler (2010). Design and Control of an Autonomous Variable-Pitch Quadrotor
Helicopter. Massachusetts Institute of Technology
O.S. Engines (2015). Heli engines. http://www.osengines.com/engines-heli/index.html
R. C. Hibbeler (2008). Mechanics of Materials. Prentice Hall
S. van Natter (2014). Gas Powered Single Engine Variable Pitch Quadcopter.
http://diydrones.com/profiles/blogs/gas-powered-single-engine-variable-pitch-
quadcopter
Sullivan Products (2015). Alternators and Regulators for Unmanned Vehicles.
http://www.sullivanuav.com/home.html
T-motor (2013). http://www.rctigermotor.com. Consulted on the 28th of May 2015
