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CONCEPTUAL DESIGN OF A MODULAR 150 KG
VERTICAL TAKE-OFF AND LANDING UNMANNED AERIAL VEHICLE
F. Götten, D. F. Finger
FH-Aachen, Institute of Aircraft Engineering
Hohenstaufenallee 6, 52064 Aachen, Germany
Abstract
The authors present the conceptual design for a new unmanned surveillance aircraft. The aircraft is sized to
carry a 14.5 kg electro-optical payload, and must deliver the longest possible endurance, while being
constrained to a maximum take-off mass of 150 kg. This limit is set by the STANAG 4703 certification rules.
The unmanned aerial vehicle (UAV) is able to take off and land vertically. By using a modular propulsion
concept, the UAV can be converted from the vertical take-off and landing (VTOL) role, into a conventional
aircraft with ease. This is made possible by using lift-jets for the vertical take-off and landing phase, and a
piston engine for cruise flight. The determination of an aircraft's mass- and aerodynamic properties are critical
in the conceptual design phase. Therefore, these are the main focus points of this paper. To improve
confidence in the mass estimation, a first structural layout is presented and the main structural members are
sized. Results indicate that the modular approach to VTOL UAV design should be considered for further
development, because it offers great flexibility for medium altitude long endurance surveillance missions.
Nomenclature
AR = aspect ratio
C = coefficient
CD = coefficient of drag
CFRP = carbon fiber reinforced plastic
CL = coefficient of lift
CTOL = conventional take-off and landing
e = Oswald’s aircraft efficiency factor
FEM = finite element method
HT = horizontal tail
MALE = medium altitude long endurance
MTOM = maximum take-off mass
MZFM = maximum zero fuel mass
n = load factor
RCS = reaction control system
UAV = unmanned air vehicle
va = maneuver speed
vc = cruise speed
vD = dive speed
VT = vertical tail
VTOL = vertical take-off and landing
1. INTRODUCTION
At present, the UAV market is rapidly expanding.
Technological innovation and progress makes new
aircraft and mission concepts feasible, which would
be literally unable to take-off, employing conventional,
manned design approaches. Transitioning vertical
take-off and landing (VTOL) aircraft combine the best
of two worlds: A vehicle that merges a helicopter’s
ability to take-off and land almost anywhere, with the
speed, range, endurance and load carrying capability
of a fixed wing aircraft. Some sources call these
vehicles ‘convertiplanes’ or ‘hybrid aircraft’ and many
different configurations are used [1] [2].
A market study of unmanned aerial vehicles (UAVs)
shows the need for flexible long endurance VTOL
aircraft. By limiting the maximum take-off mass
(MTOM) to 150 kg such UAVs can be treated as light
unmanned aircraft systems. When compared to
heavier UAVs, light systems enjoy a simplified
certification process, according to NATO’s STANAG
4703 [3].
There are many missions where the VTOL aircraft is
superior to a conventional take-off and landing
(CTOL) solution. Ground infrastructure is minimal,
especially when small, unmanned VTOL aircraft are
considered. Without the need for launch and recovery
equipment, they can attain a mobility that is not
realizable for CTOL vehicles. VTOL aircraft can be
launched from any reasonably flat surface. Typically,
a clear area, the size of a helicopter landing-pad
(15 m x 15 m) is more than sufficient. However,
civilian and military users may not require VTOL
capabilities when circumstances allow starting and
landing from a runway. Then, VTOL UAVs suffer from
reduced flight performance compared to CTOL UAVs
due to their high propulsion system weight. It is
therefore desirable to develop hybrid solutions that
allow on-demand VTOL capabilities when needed,
but do not suffer from the weight penalties when used
in CTOL configuration.
Many recent developments focus on classic fixed-
wing UAVs and enabling vertical take-off and landing
by attaching multi-rotor technology to these aircraft
[4]. Multi-rotor technology requires electric motors to
drive the propellers in order to achieve the necessary
control speed. The cruise propulsion system,
however, often relies on fuel based energy storage to
achieve very long endurance. This propulsion
architecture has to either incorporate a generator with
batteries that act as a power buffer, or batteries that
are charged on the ground in order to provide
electrical power for the lift system. This increases
complexity and adds additional workload to the
ground crew.
Several civilian and military users therefore require
UAV systems to only rely on a single source of energy
and prefer heavy fuel solutions. In such cases, jet
engines can be advantageous in providing the
required thrust for vertical take-off and landing. They
combine high thrust with small weight and do not
require electrical energy like multirotor systems.
To satisfy the requirements of long endurance flights,
on-demand VTOL capabilities, and a heavy fuel
propulsion system, the conceptual design of a novel,
modular UAV is presented in this paper. It is called
JetFalcon. The JetFalcon design combines lift jet
engines with a heavy fuel two-stroke engine for
forward flight. It aims to use currently available
technology in order to keep development time and
costs at a minimum. Maximum take-off mass is
limited to 150 kg for certification according to
STANAG 4703. The paper focuses on the UAV’s
mass and aerodynamic properties, as these are
especially critical for VTOL designs.
This paper is structured the following way: Following
this introduction, an overview of the requirements is
provided and the conceptual design is presented in
section 2. In section 3, aerodynamic and performance
analyses are shown. Finally, section 4 gives a
comprehensive conclusion.
2. REQUIREMENTS AND CONCEPT DEFINITION
2.1. Requirements and Mission
JetFalcon is designed as a VTOL UAV for
surveillance missions with extended loiter
capabilities, and – for certification reasons – a
maximum take-off mass of 150 kg. Therefore, it falls
under the certification specifications of STANAG
4703 [3]. Additional guidance for certification was
provided by EASA’s CS-VLA [5].
Primary Requirements
• Payload: Day/Night HD Observation System for
Air Applications Controp SHAPO
• VTOL capability
• Mission: medium altitude long endurance
(MALE) surveillance
Secondary Requirements
• Stall speed less than 23 m/s
• Cruise speed of 35 m/s
• Maximum speed higher than 55 m/s
• High maneuverability is not required
Maneuvering beyond 2 g is unnecessary
• For easy transport, the UAV must disassemble
into parts with a maximum length of 2.5 m
2.2. JetFalcon Design Concept
JetFalcon (Figure 1 and Figure 16) is designed as a
low wing tractor UAV. Two lift jet engines are used to
give it VTOL capability. The small form factor of
turbojets with radial compressors allows the vertical
installation inside the fuselage. This uses significant
internal space, and requires careful inlet design.
However, these downsides are balanced by
JetFalcon’s modular design. The aircraft can be used
for VTOL missions, as well as for CTOL missions,
because the lift-jets are easily removable from the
airframe. If the VTOL capability is not required by a
certain mission, the jet engines’ space and mass can
be replaced with additional fuel or additional payload.
Thus, the airframe’s capabilities are very much
extended.
The payload is integrated ahead of the wing but
between the lift-jets. This is not a perfect solution, as
high heat loads must be expected in this area, while
the jets are operating. The challenge was overcome
by retracting the optical sensor when the lift-jets are
operating. During take-off and landing payload
operation is not required, anyway and this allows for
the added benefit of improving the aerodynamic
efficiency during cruise, if the payload is not needed
during the time to target.
A low wing configuration was chosen to allow for a
short landing gear. Additionally, this frees up space in
the upper fuselage, which allows for easy integration
of an emergency parachute.
VTOL Concept
In hover, the aircraft’s mass is supported by two lift-
jet engines. Both jets are equipped with 3D-thrust
vector controls that can be used to control the aircraft
in the x-y-plane. Roll control is maintained by
integrating small electric ducted fans at the wingtips.
This is necessary, because small jet engines do not
offer bleed-air offtake and thus a conventional
reaction control system (RCS) cannot be integrated.
Additionally, the ducting for a RCS is prohibitively
heavy for small aircraft. The control concept is shown
in Figure 1. The jet engines are angled back at 5°
against the vertical axis. This facilitates a nose-up
attitude during hover. The transition from hover to
cruise and back is easier this way, as the wing
generates more lift at higher angles of attack.
Additionally, during landing, the nose-up attitude
allows ground contact to be made on the main wheels
first, which keeps the forces on the nose gear down
and limits the bending reaction of the fuselage.
Unfortunately, there is no redundancy during hover. If
a single part of the lift propulsion system fails, it is
likely that the aircraft will be lost. This is, however, a
problem that is faced by many VTOL aircraft, and the
risk is typically accepted.
Figure 1 – VTOL control concept
There are a number of challenges that need to be
analyzed in depth in the later design stages. It is
possible that the jets’ 800 °C, 690 m/s exhaust gases
cause problems. The hot jet is exhausted 40 cm
above the ground. This can cause problems with
ground erosion. While the erosion of the ground might
be tolerable by itself, it can cause damage to the lifting
system and the aircraft itself. Eroded soil can form a
bucket (like a thrust reverser) and can project the
exhaust gas and debris upward into the vicinity of the
inlet. Foreign object damage can lead to the loss of
the complete aircraft.
Recirculation of exhaust gases into the jet engine’s
inlet must also be considered. Hot air ingestion will
then decrease engine thrust. This is depicted in
Figure 2.
Figure 2 – Recirculation and hot-has ingestion
(adapted from [6])
However, these problems are hard to quantify without
detailed simulations or tests. Therefore, they must be
revisited in the later design stages.
Fuselage
The fuselage is shown in Figure 3, and it is made of
two segments: the forward part is 2.15 m in length
and carries the propulsion section, payload, and the
safety parachute system. The aft fuselage is 1.5 m in
length and carries the empennage and the avionics
systems. Splitting the fuselage is necessary to satisfy
the maximum length requirements for easy transport.
The physical separation is of advantage when ease
of maintenance is considered, as the propulsion
section and the avionics can be maintained
simultaneously. This will allow for shorter turnaround
times.
Figure 3 – Fuselage sideview
Wing
The wing is sized by the stall speed constraint and the
maximum length constraint for transportation.
The stall speed constraint limits wing loading to
460 N/m², if a typical maximum lift coefficient for low
Reynolds number airfoils of CLmax = 1.4 is considered.
This is a relatively low wing loading, but appropriate
for light single piston engine aircraft [7]. Given the
fixed MTOM of 150 kg, this gives a wing area of
Sref = 3.2 m².
The maximum size constraint makes the wing design
challenging. If a two-part wing is considered,
maximum span is limited to 5.0 m, if a three-part wing
is considered, maximum span is limited to 7.5 m. The
former gives an aspect ratio of 7.81, the latter gives
an aspect ratio of 17.58. Because VTOL aircraft are
very mass sensitive, and every joint increases mass,
the authors decided on a two-part wing. This limits
aspect ratio, but gives a rather sturdy structure and
avoids the aeroelastic challenges of high aspect ratio
wings. Additionally, the increase in wing chord allows
easy integration of the electric ducted fans for roll
control. Therefore, only a moderate taper ratio of 0.68
is selected.
The wing’s control surfaces are sized using statistical
analysis from [8] and [9]. A planview of the wing is
shown in Figure 4.
Figure 4 – Wing planview
At the given cruise speed of 35 m/s, the wing flies at
CL values between 0.5 and 0.6. An airfoil analysis
shows that a NACA airfoil with 3% camber and 14%
thickness is a good fit. The NACA 3414 balances
aerodynamic performance and structural depth, to
allow lightweight construction. Selecting a four-digit
NACA airfoil is appropriate for the conceptual design
stage. Optimized airfoils can be selected in later
design phases to optimize aerodynamic properties.
Weight
Roll-
Impeller
Roll-
Impeller
Turbine
Turbine
x
y
z
To delay wingtip stall, the wing is linearly twisted
nose-down by 1° from root to tip. Maximum
performance is traded against good stall behavior,
which will help transition between hover and cruise
flight. Because the transition phase is highly critical,
compromises regarding absolute performance have
to be made.
Tailplane
The H-tail configuration is selected for the
empennage. While not ideal from a drag point of view,
this choice has several advantages that make up for
that: During assembly of the aircraft, the horizontal tail
can easily be mounted on top of the fuselage, then
the vertical tails can just be stuck on the side of the
horizontal tail, using simple bolts. The vertical tails
also act as endplates for the horizontal tail, thereby
reducing trim drag and improving control efficiency.
The conceptual design is carried out using volume
coefficients from [10] and [11]. However, the stabilizer
surfaces are oversized (CHT = 0.80, CVT = 0.075),
compared to the light aircraft category, to offer
adequate control power in the transition period
between hover and cruise flight.
The horizontal tail is designed with a rectangular
planform for easy construction. The vertical tail is
primarily designed from an aesthetics’ point of view.
Both horizontal and vertical tail use NACA 0010 airfoil
sections.
Landing Gear
The landing gear is designed as a fixed tricycle gear.
As shown by [12], it is much more important for MALE
UAVs to keep the mass of the landing gear low, than
to improve the aerodynamics. This is especially true
for high-drag configurations, like JetFalcon. In case
that the payload’s field of view is too restricted by the
front landing gear, a solution could be to retract only
the front gear. This principle is, for example,
employed by Rutan’s VariEze.
2.3. Component Selection
Cruise Propulsion System
Propulsion for wingborne flight is provided by a 3W-
370 HF two stroke engine. It uses heavy fuel and
produces a maximum power of 21.3 kW. This gives
the novel UAV a power loading of 7.04 kg/kW, which
is adequate for this class of aircraft [10].
Lift Propulsion System
Two AMT Nike jet engines are chosen to provide
thrust for vertical take-off due to its class-leading
thrust-to-weight ratio. At maximum rpm, 784 N of
regular thrust can be generated with one engine. If
necessary, short overloading periods are possible.
Sensor System
Controp’s SHAPO lightweight multi sensor payload
observation system is chosen to provide day and
night surveillance and reconnaissance performance.
The sensor unit is well suited for UAV’s due to its
compact dimensions.
Rescue parachute
The parachute rescue system “Finsterwalder-Charly
HFa192” is designed as an emergency system for
paragliding. It achieves a sink rate between 4.3 m/s
and 4.8 m/s for the given UAV mass. The system is
designed as an emergency system only and not
suited for regular landings. It will reduce JetFalcon’s
impact energy, should control be lost.
2.4. Mass Estimation
Mass estimation in early design stages is important
for any aircraft design project but it is especially
critical for VTOL designs. JetFalcon’s maximum take-
off mass is limited to 150 kg for regulatory reasons.
Experience shows that the aircraft must be designed
exactly at this limit to achieve maximum mission
performance. JetFalcon’s mass estimation therefore
follows in an inverse approach. It aims to estimate the
aircraft’s structural and system mass and assumes
the maximum take-off mass to be 150 kg. In such, the
difference between MTOM and system and structural
mass gives the aircraft’s maximum fuel mass. The
fuel mass will be used in the mission performance
evaluation to validate JetFalcon’s ability to achieve
the mission goals (longest possible endurance).
The aircraft is divided into systems and subsystems
for which the mass is estimated individually. An
overview about the aircraft’s main systems and their
estimated installed masses is given in Table 1. The
masses of systems and components that are
available from third-party manufacturers are taken
from the respective datasheets and can be provided
with high accuracy. Avionics and electrical system
mass are taken from a previous, similar design.
The structural mass is estimated using class-II weight
estimation methods found in [10], [13] and [14]. These
class-II weight estimation equations differ
significantly from one another and even give
unrealistic results in some cases. Naturally,
unrealistic results were not considered for this study.
In the absence of a more rational analysis, the results
of the different class-II weight estimation methods
were averaged for the individual components.
The propulsion system accounts for about a third of
the total aircraft mass, which is typical for VTOL
configurations [4]. Structural mass is estimated to be
about 20% of the total aircraft mass, which is
ambitious but not unrealistic [10]. According to the
initial mass estimation, JetFalcon can carry about
37 kg of fuel to keep the maximum take-off mass at
150 kg.
While system masses can be provided with high
accuracy, the structural mass estimation with weight
equations is subject to high uncertainties. However,
the structural mass estimation is critical as an
increase in structural mass directly decreases the
allowable fuel mass having and immense impact on
flight performance. In order to increase the accuracy,
the conceptual design of the main structural
components is carried out, according to the critical
load cases presented in the following section.
2.5. V-n Diagram and Load Cases
According to the requirements, JetFalcon is not
supposed to be a highly maneuverable aircraft, and
load factors beyond 2g are not expected in normal
operation. However, STANAG 4703 requires a
minimum sustainable g-load of +3.8g and -1.5g.
This load spectrum is expanded by gust loads.
Typically, positive and negative gusts of 15.2 m/s
must be considered at cruise speed and 7.6 m/s
gusts must be considered at dive speed. However,
ccording to [15], it is possible to reduce those values.
For a typical operating altitude of 1000 m, and while
staying clear of thunderstorms, a gust of more than
10 m/s can be expected only once every 1.6 million
kilometers flown. At a cruise speed of 35 m/s, this
would occur once every 12700 flight hours. If the
aircraft is designed for 10000 flight hours, and the
aircraft is forbidden from flying into thunderstorms,
the additional risk of reducing the gust loads is
acceptable. Therefore, for the gust load analysis, only
gusts of +/- 10 m/s are considered at cruise speed,
and gust of +/- 5 m/s are considered at dive speed.
Using the first mass estimation as a baseline, the v-n
diagram is constructed for MTOM (150.0 kg) and for
maximum zero fuel mass (MZFM) (112.8 kg). The
diagrams are shown in Figure 5 and Figure 6.
Figure 5 – v-n diagram 150.0 kg
Figure 6 – v-n diagram 112.8 kg
Table 1 – Aircraft systems and subsystems – initial mass estimation
System
Subsystem
Mass, kg
Mass, %
Propulsion system
3W-370 HF combustion engine
electrical power supply
2x AMT NIKE jet engines
2x roll control impellers
15.0
8.0
23.3
1.5
10.0
5.3
15.5
1.0
Avionics and electrical systems
flight controller
navigation system
mission computer
3.6
2.0
4.0
2.4
1.3
2.7
Payload
Controp SHAPO
video computer
13.0
1.5
8.7
1.0
Structure
wing
fuselage
tailplane
landing gear
11.8
10.0
3.1
5.5
7.9
6.7
2.1
3.6
Fuel system
tank
misc.
4.0
0.5
2.7
0.3
Rescue system and miscellaneous
Finsterwalder & Charly parachute
reserve mass
5.0
1.0
3.3
0.7
Empty mass
112.8
75.2
Fuel mass
37.2
24.8
MTOM
150.0
100.0
Table 2 gives and overview about the critical load
cases that are considered in the initial structural
sizing. The load cases are derived in accordance to
STANAG 4703 and partly derived from CS-VLA. The
aerodynamic forces for the specific load cases are
simulated with an in-house vortex-lattice code for
both MTOM and MZFM. Inertia forces are calculated
at discrete stations, while system masses are treated
as point masses and assigned to the nearest discrete
inertia station. The structural mass is distributed over
the discrete stations. An exemplary force distribution
of one wing is shown in Figure 7. As can be seen,
bending moment relief is provided by the fuel mass
and the landing gear mass.
Table 2 – Critical load cases
Load case EAS [m/s] load factor n [/] mass [kg]
Pos. gust v
C
51.5
+4.17
150.00
Pos. gust v
C
51.5
+4.28
112.75
Neg. gust v
C
51.5
-2.17
150.00
Neg. gust v
C
51.5
-2.28
112.75
Pull-up v
A
41.0
+3.80
150.00
Pull-up v
D
72.1
+3.80
150.00
Roll
41.0
+2.53
150.00
Yaw
41.0
+1.00
150.00
Landing
0.0
+8.41
150.00
Figure 7 – Force distribution over wing span, load
case pos. gust vc
2.6. Preliminary Structural Design
The conceptual structural design of wing, tail and
fuselage allows a much more accurate estimation of
the structural masses and, in turn, increases the
accuracy of the fuel mass estimation. The sizing
process is only conceptual and focuses on the main
structural components. The detailed structural design
is part of later development stages and not treated in
this publication.
The aircraft’s structure is composed of a classic semi-
monocoque design for lifting surfaces and a
framework design for the fuselage. A framework
design for UAV fuselages might be unusual but has
distinctive advantages for medium sized UAVs in the
150 kg class. Such UAVs are too big to be handled
like model aircraft that can be man-carried and easily
disassembled, but too small in order to provide easy
access space for system integration and
maintenance. Accessibility is therefore a critical issue
for the UAV’s fuselage. JetFalcon’s framework
fuselage design is advantageous as it allows easy
access from all sides. The fuselage skin does not
carry any loads and is completely detachable.
Wing
The structural sizing of JetFalcon’s lifting surfaces
follows classic handbook methods treating static
strength and static stability. Fatigue or aeroelastic
effects are not considered in the early design stage.
The calculations are spread-sheet based and too
extensive to be shown here. The process uses
conceptual design parameters for high-tension
carbon fibers found in [16] and respective
manufacturer data of foam core material.
JetFalcon’s wing utilizes a single spar design with an
additional stub spar to support landing gear loads.
The main spar is designed as an I-beam and supports
bending loads. The spar flanges consist of
unidirectional fibers while the web uses +-45° carbon
fabric supported by a foam core. The wing skin is
composed of a sandwich structure and designed to
support torsional loads. The conceptual structural
design of the wing is shown in Figure 8.
Figure 8 – Wing structural design
The wing’s pure structural weight according to the
conceptual sizing is 9.9 kg. A mass break-down of the
individual components is shown in Table 3.
Table 3 – Wing structural mass break-down
Part mass [g]
main spar (CFRP+foam) 2400
landing gear support spar (CFRP) 1220
wing skin (CFRP+foam) 5500
ribs (CFRP+foam) 750
Total
9870
-400.0
-200.0
0.0
200.0
400.0
600.0
800.0
0.0 0.5 1.0 1.5 2.0 2.5
Fz [N]
half span position [m]
Lift and inerta forces, load case pos. gust vc
lift
inertia
resulting force
The wing skin makes for a significant part of total wing
mass followed by the main spar. The landing gear
support spar is relatively heavy despite its small
length due to the enormous load requirements for a
hard landing (8.41 g shock load). The masses shown
in Table 3 do not include any systems or additional
installation masses. The mass for redundant flap and
aileron actuators (e.g. 6x VOLZ DA-26) will be
approximately 1.6 kg. Additional weight must be
reserved for the required resin and several structural
reinforcements for panels and additional
attachments. In order to account for this, the wing
mass is increased by an additional 1.6 kg that is
extrapolated from experience with another similar
UAV. The updated wing mass after initial structural
sizing is 13 kg and therefore about 10% higher than
estimated with a class-II method presented above.
Tailplane
The tailplane has the same structural design as the
wing and is shown in Figure 9. The sizing is carried
out analogue to the wing’s structural sizing.
Figure 9 – Tailplane structural design
Table 4 shows the structural mass of horizontal and
vertical stabilizer. For the sake of brevity, the masses
presented are not divided into individual structural
components.
Table 4 – Tailplane structural mass break-down
Part mass [g]
horizontal stabilizer 1650
vertical stabilizer 1050
Total
2700
Redundant rudder and elevator actuators (e.g. VOLZ
DA-22) will increase the mass by about 400 grams.
An additional 400 grams is reserved for resin and
attachments increasing the total tailplane mass to
3.5 kg. The updated tailplane mass is therefore about
13% larger than estimated with the class-II method.
Fuselage
The fuselage framework consists of off-the-shelf
wound carbon fiber tubes that carry all loads. In order
to select appropriate tubes and determine their size,
a fuselage framework is developed and analyzed with
a finite element method (FEM). The critical load case
for the fuselage is a tail-down hard landing with 8.4 g,
in accordance to STANAG 4703. The inertia forces of
the fuselage’s structure and internal systems are
treated as external point forces in the FEM model
(yellow arrows in Figure 10). The model is mounted
at the location where the shear bolts that support the
landing gear and main spar are attached to the
fuselage (red arrows in Figure 10). The required
material data for carbon fiber tubes is taken from
respective manufacturer data. The resulting stresses
of the fuselage framework are shown in Figure 11.
The maximum stress is 275 N/mm². This is tolerable
by a wide variety of small diameter tubes. However a
further analysis has shown that buckling is critical and
only tubes with a minimum inner diameter of 18 mm
and 1 mm thickness are able to provide an adequate
safety margin towards the buckling stress.
Manufacturer data provides a mass of 120 g/m for the
chosen carbon fiber tubes which gives a total tube
mass of 6.5 kg. Covering the fuselage with a glass
fiber sandwich skin of 4.25 m² will likely add about
2.5 kg depending on the chosen number of layers that
are required to provide adequate handling qualities.
This increases fuselage mass to 9 kg. Additional
masses include the framework connecters and
attachments for all internal components. Experience
with another UAV shows that these additional
components can increase the mass significantly. The
class-II weight estimation method gave a fuselage
mass of 10 kg, which is unlikely to be achieved
according to the results of the initial structural sizing.
To be conservative, the mass determined by
structural sizing is increased by 1.33 and assumed to
be 12 kg.
Figure 10 – Fuselage FEM model
Figure 11 – Framework stresses for hard landing
load case, deformation 50 times enlarged
2.7. Updated Mass Estimation
Table 5 shows the updated mass break-down of the
novel UAV. The numbers in brackets indicate the
previous results using class-II weight estimation
equations (compare also Table 1). On first glance, the
individual deviations between the structural masses
from conceptual sizing and weight estimation
equations seem small. However, the masses
determined by structural sizing are consistently
higher than estimated with weight equations, which
leads to an 11% reduction in available fuel mass.
Such a reduction might decrease loiter flight time by
nearly 45 minutes. This drastic impact shows the high
sensitivity of VTOL designs on accurate estimations
already in early design stages.
3. FLIGHT PERFORMANCE EVALUATION
3.1. Aerodynamic Performance
The aerodynamic efficiency of the design is estimated
with the help of drag-buildup methods. A significant
allowance is made for trim drag, cooling drag,
payload drag, drag of the lift-propulsion system, and
other undesirable drag. This brings the CDminDrag to
about 434 counts. This drag coefficient is achieved at
the lift coefficient for minimum drag CLminDrag = 0.05.
Oswald’s aircraft efficiency factor is chosen as e = 0.8
in a conservative manner. The aircraft drag polar is
described as shown in equations 1 and 2:
= + (−)²
∙ ∙ (1)
= 0.0434 + (−0.05)²
∙7.81 ∙0.8 (2)
The performance is showcased in Figure 12.
Figure 12 – Aerodynamic Performance
At the design lift coefficient of 0.6, an cL/cD of 10 is
achieved. This is acceptable for such a high-drag
configuration. Drag reduction measures can be taken
in the later design stages.
-0.2
0.0
0.2
0.4
0.6
0.8
1.0
1.2
1.4
-0.2
0.0
0.2
0.4
0.6
0.8
1.0
1.2
1.4
00.02 0.04 0.06 0.08 0.1 0.12
cL
cD
JetFalcon Drag Polar
0.0 0.2 0.4 0.6 0.8 1.0 1.2 1.4
0
2
4
6
8
10
12
14
0
2
4
6
8
10
12
14
0.0 0.2 0.4 0.6 0.8 1.0 1.2 1.4
cL/cD, cL^(3/2) / cD
cL
cL/cD
cL^(3/2) / cD
Table 5 – Aircraft systems and subsystems with updated structural masses
System
Subsystem
Mass, kg
Mass, %
Propulsion system
combustion engine 3W-370 HF
electrical power supply
2x AMT NIKE gas turbine
roll control impeller
15.0
8.0
23.3
1.5
10.0
5.3
15.5
1.0
Avionics and electrical systems
flight controller
navigation system
mission computer
3.6
2.0
4.0
2.4
1.3
2.7
Payload
Controp SHAPO
video computer
13.0
1.5
8.7
1.0
Structure
wing
fuselage
tailplane
landing gear
(11.8) 13.0
(10.0) 12.0
(3.1) 3.5
5.5
(7.9) 8.7
(6.7) 8.0
(2.1) 2.3
3.6
Fuel system
tank
misc.
4.0
0.5
2.7
0.3
Rescue system and miscellaneous
Finsterwalder & Charly parachute
reserve mass
5.0
1.0
3.3
0.7
Empty mass
(112.8) 116.4
(78.2) 77.6
Fuel mass
(37.4) 33.6
(24.8) 22.4
MTOM
150.0
100.0
3.2. Notional Missions
As the refined mass estimation shows, JetFalcon can
carry 33.6 kg of fuel. However, for trapped fuel and
reserves, 6% of the fuel mass are considered
unusable. Therefore, 31.6 kg of fuel are available for
the VTOL missions. If the lift-jets are replaced by fuel
tanks, the capacity might be increased to 51.9 kg
(54.9 kg minus 6% reserves + 2 kg for tanks and
plumbing) for conventional missions.
The first mission profile is shown in Figure 13.
Following a vertical take-off and climb to altitude, the
aircraft enters cruise. Following a descent, the
landing is executed vertically, too. This is a typical
VTOL surveillance mission.
Figure 13 – Mission 1 – VTOL
A traditional CTOL surveillance mission is shown in
Figure 14. Here, the lift propulsion system is
exchanged for extra fuel, which will increase
endurance. However, a runway is required for take-
off and landing.
Figure 14 – Mission 2 – CTOL
Figure 15 shows a possibility for a support mission.
Here take-off is commenced conventionally from a
runway to save fuel. Supplies are released in the field,
where the aircraft executes a vertical landing and
vertical take-off. The mission terminates at a base
with a conventional landing.
Figure 15 – Mission 3 – Mixed
3.3. Mission Performance Calculation
The three missions are analyzed using an in-house
flight performance tool. The results are given in Table
6. JetFalcon can offer a maximum operating radius of
roughly 200 km for VTOL missions, and a 450 km
operating radius for CTOL mission. This shows the
massive impact of requiring VTOL capability: flight
time and cruise distance are reduce by approximately
55% compared to the CTOL case.
An interesting observation can be made for mission
3. Because a supply drop of 10 kg is considered mid
mission, and a part of the fuel is exhausted during the
time to target, less fuel is required for the VTOL
phase. Thus, flight time and mission radius can be
slightly increased, when compared to mission 1.
Table 6 – Mission performance
Mission
Flight
Time
Cruise
Distance
Required
fuel
1 3 h 23 min
439 km 31.6 kg
Mission Radius ~200 km
2 7 h 21 min
939 km 51.8 kg
Mission Radius ~450 km
3 3 h 38 min 471 km 31.6 kg
Mission Radius ~215 km
The disadvantage of shortened endurance times for
VTOL aircraft can – at least partially – be made up for
by increased ground mobility. For a very simple
surveillance mission, a CTOL aircraft must be
launched at a base, cruise to the target area, loiter
over the target and cruise back to base.
Because of significantly reduced dependence on
ground infrastructure, VTOL aircraft can be launched
much closer to the target zone, eliminating two
extended cruise segments and getting a relative
improvement of time over the target.
Time on station might also be extended by employing
a “stakeout” approach. A VTOL aircraft can land close
to a target, shut down the engines and function as a
highly mobile ground post. Communications and
surveillance data can be transmitted while on the
ground, with minimal power consumption. If a theater
view from altitude is required, the vehicle can be
launched and commence loitering flight. This greatly
expands mission capabilities for VTOL UAVs in
relation to the traditional CTOL aircraft [1].
4. CONCLUSION
The study revealed that it is very challenging to meet
the requirements for a 150 kg VTOL MALE. However,
it is shown that it is possible to design an unmanned
aircraft in this mass category, which is able to take-off
and land vertically, and can carry a significant
payloads for an extended flight-time. A flight time of
3.5 hours is reached for VTOL missions and
7.3 hours endurance are achievable during
conventional missions.
Mission scenarios were discussed, where VTOL
aircraft are superior to a CTOL solution. This mission
performance is a benefit, which must be considered
Cruise
Vertical Take-off
Descent
Climb
Vertical Landing
Take-off + Climb
Descent + Landing
Cruise
Take-off + Climb
Cruise
Cruise
Descent + Landing
VTOL
Descent
Climb
against the mass and thus endurance impositions of
the VTOL requirement.
The conceptual design study of the JetFalcon UAV
(Figure 16) is only the first step towards a
development program. Future work must aim to
optimize the configuration, build confidence in the
design and assess the risks and benefits of this
design. Significant engineering work is still required.
Figure 16 – JetFalcon 3-view
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