The document describes a conceptual design for a take-off and landing system (TLS) for the Insitu Integrator unmanned aircraft system. It provides background details on the aircraft configuration, structural scheme, and current TLS. Five concepts for a new TLS were generated and two were selected for further analysis based on a trade study. The feasibility of integrating the selected concepts with the aircraft was then evaluated through analytical calculations and simulations. A failure modes and effects analysis was also conducted to assess the safety of the proposed design. The results were discussed to evaluate the feasibility and suitability of the new TLS concept for the Integrator.

1. CARLETON
UNIVERSITY
AEROSPACE
ENGINEERING
AERO 4003
Conceptual Design of TLS for Insitu
Integrator
Team 2
DATE: 8 December 2014
Page 1 OF 20
AERO 4003: Aerospace Systems Design
Conceptual Design of Take-off and Landing System for
Insitu Integrator
by
Team 2:
Dustin Jee - 100847594
Boon Teh - 100866301
Kane Abbis-Mills - 100821006
Alex Lister - 100848225
Brian Sanders - 100864778
December 8, 2014
Carleton University

2. CARLETON
UNIVERSITY
AEROSPACE
ENGINEERING
AERO 4003
Conceptual Design of TLS for Insitu
Integrator
Team 2
DATE: 8 December 2014
Page 2 OF 20
1.0 INTRODUCTION ........................................................................................................................3
2.0 OBJECTIVES...............................................................................................................................3
3.0 BACKGROUND RESEARCH......................................................................................................3
3.1 Definition of Aircraft Configuration ...........................................................................................3
3.2 Definition of Aircraft Structural Scheme.....................................................................................4
3.3 Definition of Current Aircraft TLS, Dimensions and Performance Parameters...............................4
4.0 SYSTEM REQUIREMENTS........................................................................................................5
5.0 CONCEPTS AND SELECTION....................................................................................................6
5.1 CONCEPT 1.............................................................................................................................6
5.2 CONCEPT 2.............................................................................................................................6
5.3 CONCEPT 3.............................................................................................................................7
5.4 CONCEPT 4.............................................................................................................................7
5.5 CONCEPT 5.............................................................................................................................7
5.6 Concept Selection .....................................................................................................................8
6.0 System Overview and Breakdown of Selected Design Concepts.......................................................8
7.0 CONCEPTUAL DESIGN..............................................................................................................9
7.1 Landing Gear Interface..............................................................................................................9
7.2 Wing and Roof Rack Interface...................................................................................................9
7.3 Wiring....................................................................................................................................10
7.4 Feasibility Analysis .....................................................................................................................10
7.4.1 Weight Estimation................................................................................................................10
7.4.2 Aerodynamic Effects During Takeoff ....................................................................................10
7.4.3 Landing impact loads on landing gear....................................................................................11
7.4.4 Sizing of landing gear motor .................................................................................................11
7.4.5 Effect of TLS Implementation on UAS Performance ..............................................................12
7.4.6 Stressing of Critical Components...........................................................................................12
Locking Pin..................................................................................................................................13
Wheel Strut..................................................................................................................................13
8. Failure Modes and Effects Analysis ...............................................................................................14
11. Discussion and Feasibility Assessment.........................................................................................17
11.1 Weight Consideration and Aerodynamic Effects of TLS..........................................................17
11.2 Effect of TLS on UAS Performance .......................................................................................17
11.3 Stressing of Critical Components ...........................................................................................17
12.0 Conclusion................................................................................................................................18
13.0 References................................................................................................................................18

3. CARLETON
UNIVERSITY
AEROSPACE
ENGINEERING
AERO 4003
Conceptual Design of TLS for Insitu
Integrator
Team 2
DATE: 8 December 2014
Page 3 OF 20
1.0 INTRODUCTION
This report outlines the design process and implementation of a takeoff and landing (TLS)
system on the Boeing Insitu Integrator. This report also contains background research about the
UAS followed by detailed system requirements imposed on the designed TLS. Five concepts
were generated based on the system requirements from which two were selected based on a
trade-off study performed on all the concepts. The feasibility of the selected concepts was then
evaluated through analytical calculations and numerical simulations. The integration between the
new TLS and the Integrator was explored, analyzed and solidified. A failure modes and effect
analysis (FMEA) of the proposed design was then conducted. Lastly, the results of the
aforementioned analyses are discussed and the report is concluded with an assessment of its
feasibility.
2.0 OBJECTIVES
Currently, the Integrator is launched from a pneumatic catapult launcher called Mark IV which is
a large complex machine not suitable for civilian use. It requires highly trained personnel for safe
operation, and also weighs 997 kg [1]. For landing, a hook recovery system called SkyHook is
used. Similar to the launcher, it is far too heavy and large for civilian applications. Therefore, the
objective of this project was to design a TLS for the Integrator that is light, cost effective, and
easy to operate, allowing the Integrator to be utilized for civilian applications.
3.0 BACKGROUND RESEARCH
Prior to the design process of the TLS, research was done on the aircraft and its current TLS to
have a clear understanding of the structure of the aircraft. Information on the performance of the
aircraft was also obtained to be used in the later stages of the design process when feasibility of
the design is analyzed.
3.1 Definition of Aircraft Configuration
The Boeing Insitu Integrator shown in
Figure 1 is a single engine pusher-
propeller aircraft with two longitudinal
booms fixed to its main wing on either
side of its centre line from extended
nacelle-like bodies. These twin booms
provide mounting points for its horizontal
and vertical tail surfaces. The Integrator’s
design incorporates winglets at the
wingtips of its 4.8m (16ft) wide high
aspect ratio wings which yields a high
aerodynamic efficiency in terms of lift-to-
drag (L/D) ratio and endurance. The
aircraft also features six configurable
payload bays each with their own power
and Ethernet connections. These include a
nose bay compartment, a centre of gravity bay, and a wing and winglet bay per wing [2].
Figure 1: Aircraft layout [2]

4. CARLETON
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ENGINEERING
AERO 4003
Conceptual Design of TLS for Insitu
Integrator
Team 2
DATE: 8 December 2014
Page 4 OF 20
3.2 Definition of Aircraft Structural Scheme
The Integrator is currently launched using a pneumatic wedge catapult launcher, and is retrieved
using a hook and cable recovery system. This unique system requires that the aircraft and each of
its components be capable of
withstanding the forces caused by the
accelerations associated with launch
and recovery. Consequently, it is
suspected that the entire airframe be
reinforced appropriately, especially
along the span of the aircraft’s wings.
The reinforcement devices likely
adopt the form of lateral wing spars,
which is also shown by the five
circular elements shown in the wing
section (Figure 2 Magnified View A)
indicative of wing spars or a rib.
The connection points on either main
wing to the twin booms also serve as
an aircraft hardpoint. The detail shown
in the component view (Figure 2 Magnified View B) suggests that the central fuselage is
connected to the nose and engine via multiple connectors, possibly bolt and nut fasteners. The
arrangement and position of the fasteners could imply the presence of the equivalent of forward
and aft bulkheads. It is also predicted that the aircraft have additional hardpoints directly under
the wing fuselage interface that could be utilized with the alternative take-off and landing
systems integration. The location of the aircraft’s hardpoints are important as they influence the
integration of an alternative TLS.
3.3 Definition of Current Aircraft TLS, Dimensions and Performance Parameters
Table 1: Aircraft description [1]
Dimensions Performance Parameters
Length: 8.2 ft {2.5 m}
Wingspan: 16 ft {4.8 m}
Empty weight: 80 lb {34 kg}
Maximum take-off weight: 135 lb {61.2 kg}
Maximum payload weight: 40 lb {18 kg}
Endurance: 24 hours
Ceiling: 19500 ft {5944 m}
Maximum speed: 90 knots {46.3 m/s}
Cruise speed: 55 knots {28.3 m/s}
Powerplant: Electronic Fuel Injection (EFI)
Fuel: Jet Propellant 5 (JP-5), JP-8
Figure 2: Component View of the Integrator. Modified by Brian
Sanders [2]

5. CARLETON
UNIVERSITY
AEROSPACE
ENGINEERING
AERO 4003
Conceptual Design of TLS for Insitu
Integrator
Team 2
DATE: 8 December 2014
Page 5 OF 20
4.0 SYSTEM REQUIREMENTS
The following table list the requirements the alternative take-off and landing systems shall be
governed by. Each requirement is defined and complemented with the source of the requirement,
the justification for said requirement and the proposed verification method.
Table 2. Take-off and Landing Systems Requirements
Source Requirement Justification Verification Method
GeoSurv II
Requirements
Document
Section 2.1 [3]
The launch and recovery of the
UAV shall be achievable within
a flat area clear of obstructions
meeting one of the following
definitions:
• A square measuring not more
than 50 m on each side;
• A circle measuring not more
than 55 m diameter;
Take-off and landing
should be possible in
limited spaces to
accommodate civilian
applications.
Measurement of takeoff
and landing distances on
multiple takeoff and
landing tests.
CARs Part V
Subchapter
523-
VLA.1309 [4]
When performing its intended
function, the TLS shall not
adversely affect the response,
operation or accuracy of any
equipment essential for safe
operation
The aircraft must function
safely with the addition of
the TLS.
1) Aerodynamic analysis
for stable flight.
2) Drawings for TLS
integration
3) Fly-by during operation
to verify aircraft
components are
operational
GeoSurv II
Requirements
Document
Section 2.1 [3]
The system shall be designed to
operate in various geographical
areas,which may include
remote and underdeveloped
areas.
Ideal airfields may not
always be present during
aircraft operation.
Simulate and/or test
landing/take-off on
different surfaces with
varying impact loads and
obstacle sizes.
Performance
Requirement
The TLS shall be able to
withstand impulse loads during
takeoff, landing, catapult or
recovery.
The TLS should not be
damaged during normal
operation.
Structural analysis of
TLS structure at
maximum loading
condition
GeoSurv II
Requirements
Document
Section 3.6-4
– Modified [3]
The TLS shall be able to land
with crosswinds up to 0.6Vstall
and withstand any associated
loads.
The GeoSurv II
requirements specify flight
capability during
crosswinds of 0.6Vstall.
1) Aerodynamic analysis
for flight stability.
2) FEA analysis for TLS
integrity.
Performance
Requirement
The cycles to fatigue failure of
the TLS must be comparable to
the aircraft’s fatigue ability.
Aircraft and TLS must
remain economically viable
compared with competing
UAS.
Simulate takeoff and
landing cycles for desired
aircraft lifespan on
prototypes through fatigue
tests.
CARs Part V
Subchapter
523-
VLA.1309 [2]
The TLS shall be designed to
minimize hazards to the aircraft
in the event of a probable
malfunction or failure.
Failure of the TLS should
not cause further damage to
the UAS.
Failure Modes and Effects
Analysis will be
performed to determine
the modes of failure.

6. CARLETON
UNIVERSITY
AEROSPACE
ENGINEERING
AERO 4003
Conceptual Design of TLS for Insitu
Integrator
Team 2
DATE: 8 December 2014
Page 6 OF 20
5.0 CONCEPTS AND SELECTION
This section contains five concepts of a new TLS for the Institu Integrator. Each concept is
presented with a basic sketch (illustrated by Brian Sanders) and a brief description of the concept.
These concepts were then evaluated using a weighted trade study. The selected concept will be
presented in section 6.0.
5.1 CONCEPT 1
This concept uses a bicycle configuration landing gear that retracts and extends via an electric
motor. The propeller blades fold back when the motor is idling to ensure the propeller doesn’t
strike the ground upon landing. Wheels are installed on the tips of the wings to prevent damage if
they contact the ground. The UAS uses a winch take-off to become airborne without the power
of the motor. This design is illustrated in the figure below.
A similar design is used on full sized gliders and is proven to work on both paved and
unprepared runways. The fully retractable gear, small wing tip wheels and folding propeller will
not drastically affect the drag of the aircraft. Also the simplicity of the design will keep the cost
of this TLS at a minimum.
5.2 CONCEPT 2
This concept employs a
vehicle take-off system and a
belly landing. The UAS is
attached to a vehicle using a
launch roof rack. To land, a
skid pad installed on the
bottom of the fuselage is used.
A folding propeller design is
implemented to prevent the
propeller from hitting the
ground during landing.
Figure 3: Concept 1
Figure 4: Concept 2

7. CARLETON
UNIVERSITY
AEROSPACE
ENGINEERING
AERO 4003
Conceptual Design of TLS for Insitu
Integrator
Team 2
DATE: 8 December 2014
Page 7 OF 20
This design will cause a minimal increase in drag, as the belly skid plate does not add a large
amount of cross sectional area. It would also be light in weight compared to conventional landing
gear systems. The launching mechanism is based on proven technology. Instead of the launching
frame being attached to a catapult, it is simply attached to a vehicle using a roof rack. The skid
plate however is not ideal for absorbing impact energy upon landing.
5.3 CONCEPT 3
This concept is a simple non-retractable
tricycle landing gear attached to the
fuselage of the UAS. This is illustrated
in Figure 5.
This design will be economical to
implement because of its simplicity.
However, the narrow wheelbase will
make the UAS susceptible to tip-overs
during take-off, landing and taxiing.
Also, because the landing gear cannot
retract and has a large frontal area, the
drag increase will have an adverse effect
on the performance of the aircraft.
5.4 CONCEPT 4
This concept again is a simple
tricycle landing gear but with a
wider wheelbase. The main landing
gears are attached to the existing
hard points located near the root of
the wing. This concept is
illustrated in Figure 6.
This concept is again simple,
however, the large cross sectional
area will increse drag. Also the
main gear struts are very long and
will require more material to
achieve the desired stiffness. This
in turn will add more weight to the
design.
5.5 CONCEPT 5
This design uses the same car launch take off method as concept 2, but for landing, a parachute
design is implemented as illustrated in Figure 7.
Figure 5: Concept 3
Figure 6: Concept 4

8. CARLETON
UNIVERSITY
AEROSPACE
ENGINEERING
AERO 4003
Conceptual Design of TLS for Insitu
Integrator
Team 2
DATE: 8 December 2014
Page 8 OF 20
To land the UAS with this design,
the parachute is deployed by opening
the main payload bay. This parachute
is attached to a ring that swivels
around the fuselage. The UAS then
floats down under the canopy,
landing on its belly. The
disadvantage of this design is that it
requires a massive parachute to
decelerate the aircraft to a safe
velocity (similar to the size of a
parachute used by skydivers). This
also adds weight and occupy
valuable space in the payload bay.
5.6 Concept Selection
A trade-off study was performed to choose the best concept, and concepts 1 and 2 were
combined as the final concept. The landing system from concept 1 - the bicycle landing gear
with folding propeller - and the vehicle launch system from concept 2 were utilized in finalizing
the selected concept. A small change was incorporated into the landing system from concept 1.
Instead of placing small wheels at the wing tips, small consumable skid plates were used instead
for structural simplicity and ease of maintenance.
For the trade-off study, parameters such as cost, proven technology, structural complexity,
integrality, and maintenance were considered. The following figure shows the results of the
trade-off study.
6.0 System Overview and Breakdown of SelectedDesignConcepts
The selected TLS design has a bicycle landing gear similar to landing gears commonly found on
gliders. Both the front and rear wheels are fully retractable via a metal-geared servo, and only
used upon landing. When extended, the wheels push open a spring loaded door that is held open
0
10
20
30
40
50
60
CONCEPT 1 CONCEPT 2 CONCEPT 3 CONCEPT 4 CONCEPT 5 CHOSEN
CONCEPT
Figure 9: Trade-off study results
Figure 8: Concept 5

9. CARLETON
UNIVERSITY
AEROSPACE
ENGINEERING
AERO 4003
Conceptual Design of TLS for Insitu
Integrator
Team 2
DATE: 8 December 2014
Page 9 OF 20
by the extended wheel. This door is closed by a spring when the gear is retracted and is no longer
holding it open. The skid pads placed at the wingtips prevent the wingtips from striking the
ground as the aircraft comes to rest after landing. Also, to prevent the propeller from striking the
ground upon landing, a folding propeller is used.
For take-off, the aircraft is mounted to the roof of a vehicle via a roof rack launching mechanism.
As the vehicle approaches the aircraft’s take-off speed, the aircraft will be released from the
mount. Detailed drawings of the systems can be seen on pages 19 and 20.
7.0 CONCEPTUAL DESIGN
This section outlines the interfaces between the TLS and the aircraft. The feasibility of the design
was evaluated through analytical calculations and numerical simulations.
7.1 Landing Gear Interface
The interface between the landing gear unit and the aircraft is a separate entity that fits into the
payload bay area of the aircraft. The frame of the landing gear unit is locked in place via nuts and
bolts. These units are designed in a way that allows simple replacement of the landing gear in
case it is damaged and needs to be replaced. A more detailed drawing showing the integration of
the landing gear can be found in the Figure 10.
7.2 Wing and Roof Rack
Interface
The interface between the wing
and roof rack consists of a
mechanism which allows a 1
degree of freedom sliding joint.
The wing surface has the male
side of the joint to minimize the
drag effects in flight. The male
side is a protrusion on the
underside of the wing which
slides into the female side located
Figure 10: Landing gear inteface
Figure 11: Roof rack - wing interface

10. CARLETON
UNIVERSITY
AEROSPACE
ENGINEERING
AERO 4003
Conceptual Design of TLS for Insitu
Integrator
Team 2
DATE: 8 December 2014
Page 10 OF 20
on the top surface of the roof rack. When the vehicle reaches the take-off velocity of the aircraft,
the actuator opens up the wing clamp, and the joint allows the aircraft to slide forward and
detach from the roof rack allowing the aircraft to climb with extra thrust from the engine.
7.3 Wiring
As seen in Figure 1, the power supply to
the actuator systems comes from the
internal aircraft power supply. The 75W
DC motors require 50V which is stepped
down by a voltage regulator positioned
between the aircraft power supply and the
motors. The onboard avionic system
supplies a signal to the motors to generate
the desired action. The avionic control unit
(ACU) is powered directly from the
internal power supply. Figure 1 shows only
the system power requirements for the
TLS. The ACU controls and other systems
onboard the aircraft however, was not
included in the schematic shown in Figure
12.
7.4 Feasibility Analysis
7.4.1 Weight Estimation
The volume of the components of the TLS was obtained using Creo Parametric 2.0. The volume
of the roof rack was found to be 1.435× 10−2
m3. The chosen material for the roof rack is carbon
fiber composite which has a density of 1.2 g/cc [5]. This yielded a mass of 17.22 kg.
The volume of the landing gear frame was found to be 1.067× 10−3
m3. With Aluminum 6061
T-6 as the material which has a density of 2.7 g/cc [6], this landing gear frame has a mass of 2.88
kg. Since the roof rack is only used for take-off, the total added weight to the aircraft is 5.76 kg
reducing the maximum payload weight to 12.24 kg.
7.4.2 Aerodynamic Effects During Takeoff
The drag force generated by the aircraft during take-off (zero-lift drag) was estimated by using
the wetted-area method. The calculations are shown below:
Swet of aircraft = 6.96 m2 , ARwet = 3.31
Sref = 1.8 m2, ARwing = 12.8
For most propeller-driven UAS ,
𝐶 𝑓
𝑒
= 0.005 [ref],
𝐿
𝐷
= √
𝑝𝑖∗𝐴𝑅 𝑤𝑒𝑡
4∗
𝐶 𝑓
𝑒
= 22.8
CDo =
𝑝𝑖∗𝐴𝑅 𝑤𝑒𝑡∗𝑒
4∗(
𝐿
𝐷
)2
= 0.0116
At take-off condition: D = 0.5ρ*V2*Sref*CDo = 1.637 N,
Figure 12: Wiring Diagram

11. CARLETON
UNIVERSITY
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ENGINEERING
AERO 4003
Conceptual Design of TLS for Insitu
Integrator
Team 2
DATE: 8 December 2014
Page 11 OF 20
where ρ is the freestream density of 1.225 kg/m3 and V is the take-off speed of 11.32 m/s
(1.2Vstall)
Therefore, the moment due to drag on the roof of the launch vehicle is as following:
M = 1.637N * 1.06 m (height of roof rack) = 1.735 Nm
Most cars are capable of transporting objects on their roof with similar wetted areas as the Insitu
Integrator (6.96 m2) and weight at much faster speeds (up to 120 km/hr). Therefore, it can be
safely assumed that the aircraft could be launched from most cars.
7.4.3 Landing impact loads on landing gear
This TLS has a maximum vertical decent speed of 10.3 m/s (glide scope of 6 degrees). From
research, it was found that the duration of the impact between the landing gear and the runway is
0.25 s . Bicycle landing gear configurations require that the rear wheel take all of the impact load
upon touchdown. Assuming no lift at landing and a margin of safety of 0.5 for unmanned aerial
vehicles, the impact load each gear must withstand can be calculated as follows.
𝐹 = 𝑚
Δ𝑉
Δ𝑇
= 1.5(62.971 𝐾𝑔)
(10.3
m
s
− 0)
0.25 s
= 3372𝑁
7.4.4 Sizing of landing gear motor
Sizing of the actuating systems for the proposed TLS began with an aerodynamic load
determination. The size of all components was first determined and the appropriate size of the
UAS’s tires was then estimated. The diameter of the tires, D (in inches), was found by the
equation given below [7].
log 𝐷 = log 𝐴 + 𝐵𝑙𝑜𝑔𝑊
For UAS’s up to 1000 kg, A=1.51, B=0.349. The maximum takeoff weight of the Integrator is
61.2kg. This weight is divided by the number of wheels.
Figure 13: Car roof take-off system for Penguin B [9]

12. CARLETON
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ENGINEERING
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Conceptual Design of TLS for Insitu
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DATE: 8 December 2014
Page 12 OF 20
log 𝐷 = log1.51 + 0.349𝑙𝑜𝑔(
61.2
2
)
𝐷 = 4.983 𝑖𝑛 = 0.127𝑚
Using a similar equation quoted in the sizing method the width of the tires, T, was found to be
0.0528 m. With all components sized, it was found that the peak aerodynamic load that the
landing gear would be subjected to results from combination of the drag induced by the exposed
tires and landing structure at full extension. A force-moment analysis was performed using the
peak aerodynamic load of 6.97N and the calculated gear ratio of 1.57.
𝑇 = 𝐹 ∙ 𝑑 = 6.97𝑁 ∙ 0.0508 𝑚 = 0.354 𝑁 ∙ 𝑚
𝐹 =
𝑇1
𝑑1
=
𝑇3
𝑑3
, 𝑇3 = 𝑇1
𝑑3
𝑑1
= (0.354 𝑁 ∙ 𝑚)
2.44𝑐𝑚
1.55𝑐𝑚
= 0.556 𝑁 ∙ 𝑚
Therefore, the selected motor must be capable of supplying a torque exceeding 0.556 N∙m. The
motor chosen for this task is the Beckhoff AS1030 stepper motor, which is capable of producing
0.6 N∙m of torque when supplied with 50 V and 1.5A (total maximum draw = 150W) [8]. This is
within the 350 W of available onboard payload power.
7.4.5 Effect of TLS Implementation on UAS Performance
The new TLS will affect the range and endurance of the UAS because of the added weight. Since
the landing gear is fully retractable it will not cause an increase in drag. The skid pads located at
the wingtips are small and streamline, thus the drag generated by them are neglected. This means
the SFC, cruise velocity and L/D will remain the same as they are independent of the weight.
The decrease in endurance can be found by using the Breguet endurance equation for the old and
new UAS configuration. Assuming the added landing gear does not affect the fuel capacity and
there is no significant change in the SFC, the decrease in endurance can be estimated by the
formula shown below.
Δ𝐸𝑛𝑑𝑢𝑟𝑎𝑛𝑐𝑒 = 1 −
(
1
√ 𝑊1
−
1
√ 𝑊0
)
𝑛𝑒𝑤
(
1
√ 𝑊1
−
1
√ 𝑊0
)
𝑜𝑙𝑑
= 1 −
(
1
√40
−
1
√61.2
)
𝑛𝑒𝑤
(
1
√34
−
1
√61.2
)
𝑜𝑙𝑑
= 30.1%
7.4.6 Stressing of Critical Components
In order to determine the structural integrity of this TLS
design, two simulations were performed on each individual
part that forms the landing gear. This will mainly serve to
determine critical loads and potential failure modes that
each part will encounter and the results shall be compared
to analytical results calculated. Note that all simulations
performed on all individual pars were done using the
ANSYS® Workbench 15 software. The location of each
Figure 14: Landing gear

13. CARLETON
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ENGINEERING
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Conceptual Design of TLS for Insitu
Integrator
Team 2
DATE: 8 December 2014
Page 13 OF 20
individual part which make up the landing gear module is shown in Figure 14.
Locking Pin
The locking pin serves to prevent the landing gear from being forced into retraction due to static
and dynamic loads encountered during landing.
In the simulation, a force of 3372 N (calculated analytically) at two locations along the pin were
imposed and the contact point between the locking pin and the gear frame was modeled as a
frictionless contact. In essence, the locking pin is experiencing a four point bend during landing.
The boundary conditions for the locking pin along with the deformation and resulting von Mises
stresses are shown below.
From the von Mises stress distribution of the locking pin, the discontinuity of the locking pin
served as a stress concentration during bending which means that this is the most probable area
of failure. However, the maximum von Mises stress is approximately 204 MPa and since this is
lower than the yield strength of Aluminum 6061 T-6 which is approximately 276 MPa, the pin
will be able to withstand the calculated landing impact load based on the von Mises yield
criterion. Also, note that the deformation of the locking pin conforms to the deformation of a
horizontal beam subjected to 4-point bending and the maximum deformation of the locking pin is
approximately 0.5 mm and the analytical deformation was calculated to be approximately 0.4
mm.
Wheel Strut
The wheel strut connects the wheel of the landing gear to the frame of the landing gear. A
simulation was performed on this part to determine the critical load for buckling of this part. The
boundary conditions and deformation result of the wheel strut is as follows.
Unit: MPa
Type: von Mises
stress
Figure 15: Stressing of locking pin

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Conceptual Design of TLS for Insitu
Integrator
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DATE: 8 December 2014
Page 14 OF 20
‘
Based on the results of the simulation, the critical load requried for buckling to occur in the
wheel strut is approximately 114 kN. For purposes of comparison, an analytical solution was
obtained by dividing the load equally between the two forks, and the critical load was calculated
for one of them. The analytical critical load for buckling to occur is calculated using the
following formula.
The calculated critical load for the entire wheel strut is therefore 77.5 kN. Note that the wheel
strut is assumed to be a pin-pin connection where K=1. Since the landing load was calculated to
be 3372N, it can be safely assumed that the wheel strut will not buckle due to landing impact
since the landing load is significantly lower than both the analytical and numerical critical load.
8. Failure Modes and Effects Analysis
This section presents the failure mode and effects analysis (FMEA) for the landing system of the
Insitu Integrator. The landing system is divided into subassemblies and parts, and the potential
failure mode of each part and their effects on the overall landing system are identified. Once
potential design and process failures are identified, design changes that are necessary to prevent
such failure will be determined.
8.1 System Breakdown and Categorization
The TLS of the aircraft was divided into two subsystems, take-off system and landing system.
The two subsystems were then broken down to their respective subassemblies and components
that can fail independently and cause the whole assembly to fail. These components were then
further divided into individual parts that are needed to construct these components which make
up the whole subsystem. FMEA was performed on critical individual parts to identify potential
process and design failures thereby minimizing the risk of failure during flight by implementing
any necessary design changes and inspection methods. The breakdown of the system is shown in
Figure 18.
Figure 16: Stressing of wheel strut

15. CARLETON
UNIVERSITY
AEROSPACE
ENGINEERING
AERO 4003
Conceptual Design of TLS for Insitu
Integrator
Team 2
DATE: 8 December 2014
Page 15 OF 20
Figure 17: System break-down

16. CARLETON
UNIVERSITY
AEROSPACE
ENGINEERING
AERO 4003
Conceptual Design of TLS for Insitu
Integrator
Team 2
DATE: 8 December 2014
Page 16 OF 20
Table 2: FMEA Summary
Ident. No. Item/Functional
Identification
Failure Mode Failure Cause Failure Event Target Action Required
A.1.2 Wingtip
protection
Delamination High peel stress Wingtip damage D Frequent inspection
A.2.1.1 Retracts and
deploys wheel
Stripped gear Excessive torque Landing gear does not
deploy
A Inspection on ground before take-off
or back-up actuator
A.2.2.1 Attach gear frame
to fuselage
Loose bolt Vibration from cyclic
loading
Frame detachment from
aircraft
A Implement lock wire for bolts
A.2.2.2 Landing load
distribution and
landing gear
placement module
Column
buckling
Compressive stress from
landing impact
Fracture of gear frame
structure
A Increase frame structure thickness or
use of stronger material
A.2.2.3 Mount servo
motor to gear
frame
Loose bolt Vibration from repeated
landing impact
Landing gear does not
deploy
A Implement lock wire for bolts
A.2.3.1 Secures landing
gear position
during
deployment
Shear failure Shear stress due to
landing impact
Landing gear collapse upon
touchdown
A Increase margin of safety or thickness
of pin
A.2.3.2 Attachment of
wheel to gear
frame
Buckling Compressive stress from
landing load
Sudden collapse of wheel
strut upon touchdown
A Increase strut thickness
A.2.3.2 Absorb landing
impact
Tire puncture Presence of foreign
object debris (FOD) on
landing runway
Loss of control during
ground roll
D Selection of tire with increased
thickness and inspection for FOD on
landing area.
A.3.1 Folds propeller
upon landing
Propeller does
not fold
Increased hinge joint
friction due to corrosion
Propeller strike during
landing
D Frequent lubrication
B.1.1.2 keeps roof rack
attached to car
Shear failure Shear stress from
inertial and drag forces
of roof rack and aircraft
Roof rack gets detached
from car
A, D,
P
Increase margin of safety for bolts
For the take-off subsystem, only B.1.1.2 was considered as the failure of all the other parts simply results in aborted take-off (stopping
of the vehicle) and no further damage on the aircraft or the operator,
Legend: A - Aircraft D - Downtime P - Personnel

17. CARLETON
UNIVERSITY
AEROSPACE
ENGINEERING
AERO 4003
Conceptual Design of TLS for Insitu
Integrator
Team 2
DATE: 8 December 2014
Page 17 OF 20
11. Discussion and Feasibility Assessment
11.1 Weight Consideration and Aerodynamic Effects of TLS
The estimated weight of the takeoff system is approximately 17 kg. Note that the tak-eoff system
is a separate unit from the UAS and shall be mounted on the top of a car. The zero lift drag on
the UAS was estimated to be 1.64 N and the resulting moment about the roof rack caused by this
drag during the takeoff run was 1.74 N∙m. A major assumption in this analysis is that the
moment about the roof rack will only be caused by the drag from the UAS. Based on research,
most cars are capable of transporting objects on their roof with similar wetted areas as the Insitu
Integrator (6.96 m2) at speeds up to 120 km/hr and a benchmark example considered is the
Penguin B roof rack launcher shown in Figure 13.
For the landing system, implementation of two landing gear units into the payload bay will
decrease the payload capacity of the UAS from 18 kg to 12 kg. The landing gear units are
implemented into the payload bay of the UAS to minimize structural changes to the UAS and to
maintain its structural integrity. This was also done for ease of access and maintenance as the
units are also designed to be easily replaceable. With this implementation of this landing gear,
the maximum payload decreased by 6 kg.
11.2 Effect of TLS on UAS Performance
The effect of implementing the landing gear unit into the payload bay of the UAS on its
performance was evaluated by means of comparing the endurance of the UAS with and without
the landing gear unit. It was found that implementation of the landing gear unit will decrease the
endurance of the UAS from 24 hours to approximately 16.8 hours assuming that the SFC, cruise
velocity and the L/D remains unchanged. The significant decrease in endurance can be attributed
to the fact that addition of two landing gear units increases the operating weight of the UAS.
Also, part of the thrust provided by the engine during flight will be used to sustain lift for the
additional weight. However, it should be noted that even with the landing gear unit implemented,
the Insitu Integrator still has a higher endurance than most UAS’s in the current market with
similar missions such as Barnard Microsystems’ InView which has an endurance of only 7 hours
[8]. Therefore, based on the aforementioned reasons, the decrease in endurance is justified and
implementation of the landing gear unit is feasible unless a higher endurance is desired. Also, the
decrease in endurance of the Integrator can be improved by selecting lighter materials such as
carbon fiber composite which is widely used in remote controlled (RC) aircraft.
11.3 Stressing of Critical Components
In order to determine the feasibility of the landing gear unit in terms of structural strength, two of
the most critical parts of the unit, the locking pin and the wheel strut, were analyzed and finite
element analysis (FEA) was performed. For the locking pin, a 4-point bending test was simulated
to approximate its loading condition upon touchdown. Based on the results of the simulation, the
pin will only experience a deflection of 0.4 mm and an average von Mises stress of about 100
MPa which is below the yield strength of Al6061 T-6. Therefore, based on the von Mises failure
criterion, the part will not fail due to landing load during impact. Note that von Mises criterion is
adapted as it predicts yielding more accurately compared to the Tresca theory which is more
conservative which can result in an overdesigned part that is heavier. Another factor supporting

18. CARLETON
UNIVERSITY
AEROSPACE
ENGINEERING
AERO 4003
Conceptual Design of TLS for Insitu
Integrator
Team 2
DATE: 8 December 2014
Page 18 OF 20
the structural feasibility of the locking pin is the fact that the maximum von Mises stress is only
200 MPa which is still below the yield strength of AL 606 T-6 and that this is located at the
geometrical discontinuity of the pin which is s source of stress concentration.
For the wheel strut, a simulation was performed to determine the critical buckling load and
results have shown that the part will buckle at a critical load of 114 kN which is significantly
higher compared to the analytical load of 77.5 kN. This discrepancy may be attributed to the
simplification of the wheel strut structure for the analytical calculation. Since the landing load of
3.37 kN is significantly lower than the predicted critical loads, it can be assumed that the wheel
strut will not buckle due to landing loads.
12.0 Conclusion
The conceptual design process of a take-off and landing system for Insitu Integrator was
performed. The selected concept uses two separate systems for take-off and landing: bicycle
configuration landing gear and folding propeller are used for the landing procedure, and a car
roof launch system is used for the take-off procedure. A detailed aerodynamic and structural
analysis was done to ensure the feasibility of the selected concept. A failure mode and effect
analysis was also performed for critical components of the system to evaluate the reliability of
the system and to come up with preventive measures for possible failures. The detailed drawings
of the system can be seen in the following two pages.
13.0 References
[1] "Launchers," Insitu, [Online]. Available: http://www.insitu.com/systems/launch-and-
recovery/launchers. [Accessed Nov 2014].
[2] "Integrator System," Insitu , 2013. [Online]. Available:
http://www.insitu.com/systems/integrator. [Accessed October 2014].
[3] T. James, ""GeoSurv II Systems Requirements Document," Carleton University, Ottawa,
2006.
[4] "Part V: Airworthiness Manual Chapter 523," Transport Canada, 2009.
[5] Performance Composites Ltd, "Mechanical Properties of Carbon Fibre Composite Materials,"
Performance Composites Ltd, July 2009. [Online]. Available: http://www.performance-
composites.com/carbonfibre/mechanicalproperties_2.asp. [Accessed 30 October 2014].
[6] "Aluminum 6061-T6; 6061-T651," Materials Web, 2014. [Online]. Available:
www.matweb.com/search/datasheet_print.aspx?matguid=1b8c06d0ca7c456694c7777d9e10be
5b. [Accessed 25 November 2014].
[7] A. Jha, "Landing gear layout deisgn for Unmanned Aerial Vehicle," in 14th National
Conference on Machines and Mechanisms, Durgapur, India, 2009.
[8] "Barnard Microsystems," 2014. [Online]. Available: http://www.barnardmicrosystems.com/.
[Accessed 25 November 2014].
[9] "7 High Tech Drones For Sale", [Online]. Available: http://www.thecoolist.com/7-high-tech-
drones-for-sale-today/