FFD 2011 Best Design Report

Transcription

PAFA BURAQ
FUTURE FLIGHT DESIGN 2011
COLLEGE OF AERONAUTICAL ENGINEERING, PAKISTAN AIR FORCE ACADEMY
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PAFA BURAQ
FFD 2011
“BURAQ”
BURAQ”
COLLEGE OF AERONAUTICAL ENGINEERING
PAKISTAN AIR FORCE ACADEMY RISALPUR
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PAFA BURAQ
Table of Contents
1. Executive Summary..................................................................................................................... 5
1.1. Design Strategy..................................................................................................................... 5
1.2. Design Characteristics ......................................................................................................... 5
1.3. Aircraft Performance ............................................................................................................ 6
2. Management Summary................................................................................................................ 7
2.1. Team Introduction ................................................................................................................. 7
2.2. Workload Distribution ........................................................................................................... 7
2.3. Timeline ................................................................................................................................. 8
3. Conceptual Design ...................................................................................................................... 9
3.1. Mission Analysis ................................................................................................................... 9
3.2. Airfield Observations .......................................................................................................... 10
3.3. Aircraft Requirements ........................................................................................................ 10
3.4. Payload Requirements........................................................................................................ 10
3.5. Score Analysis .................................................................................................................... 11
3.6. Problem Statement ............................................................................................................. 11
3.7. Design Requirements ......................................................................................................... 11
3.8. Concept Collection ............................................................................................................. 13
3.9. Figures of Merit ................................................................................................................... 16
3.10. Concept Selection ............................................................................................................. 16
3.11. Final Concept .................................................................................................................... 19
4. Preliminary Design .................................................................................................................... 20
4.1. Mission Profile .................................................................................................................... 20
4.2. Design Point........................................................................................................................ 21
4.3. Weight Estimation ............................................................................................................... 21
4.4. Airfoil Selection................................................................................................................... 21
4.5. Geometric Sizing................................................................................................................. 23
4.6. Aerodynamic Analysis........................................................................................................ 23
4.7. Propulsion Analysis............................................................................................................ 25
4.8. Structural Loads ................................................................................................................. 26
4.9. Stability Analysis ................................................................................................................ 27
4.10. Aircraft Performance ........................................................................................................ 28
5. Detail Design .............................................................................................................................. 29
5.1. Dimensional Characteristics .............................................................................................. 29
5.2. Structural Systems and Capabilities .................................................................................. 30
5.3. Systems and Sub Systems ................................................................................................. 33
5.4. Weight and Balance ............................................................................................................ 35
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5.5. Final Aircraft and Mission Performance ............................................................................ 37
5.6. Drawing Package ................................................................................................................ 37
6. Manufacturing Plan ................................................................................................................... 43
6.1. Figures of Merit ................................................................................................................... 43
6.2. Fuselage .............................................................................................................................. 44
6.3. Wings................................................................................................................................... 45
6.4. Tails ..................................................................................................................................... 46
6.5. Manufacturing Timeline ...................................................................................................... 46
7. Testing Plan ............................................................................................................................... 47
7.1. Ground Testing Plan ........................................................................................................... 47
7.2. Check List ........................................................................................................................... 47
7.3. Flight Testing Plan .............................................................................................................. 48
8. Performance Results ................................................................................................................. 49
8.1. Ground Testing Results...................................................................................................... 49
8.2. Flight Testing Performance ................................................................................................ 51
9. Conclusion ................................................................................................................................. 52
9.1. References .......................................................................................................................... 53
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1. Executive Summary
The following report presents the design and development of a fire
extinguisher unmanned radio controlled aerial vehicle by team “BURAQ”
comprising of undergraduate students of College of Aeronautical Engineering,
Pakistan Air Force Academy, Risalpur. The team will compete at the Future
Flight Design Contest 2011 being held at Istanbul to celebrate the centennial
anniversary of the Turkish Air Force. The target will be to complete all three
designated missions in the most befitting manner and obtain maximum design and flight score at FFD.
1.1. Design Strategy
The first step was requirement identification and selection of target technical specifications along
with desirability values like minimum weight and cost. After a detailed analysis of FFD requirements
including payload, performance and sizing, a target technical specifications for the UAV was defined
and design process was initiated to satisfy these specifications with minimal weight and cost.
The design process started with the conceptual design which not only helped finalize the basic
configuration but also helped determine the initial weight and geometric sizing of the UAV. After
finalizing the initial dimensions of the UAV the design process entered into the preliminary stage
where high fidelity analysis was carried out using Computational Fluid Dynamics to determine the lift
and drag characteristics of the UAV. Using this data the stability parameters were calculated and the
structural loads were determined. Limited theoretical propulsion analyses were also carried out during
the same design phase. Depending upon any shortcomings felt during the preliminary design,
required changes were made and the external dimensions were frozen. The detailed design was
made using the CAD software. The major area of concern during the detailed design was to address
the integration and manufacturing issues with adequate strength to bear the aerodynamic loads.
Detailed parts, template and mould drawings were thus generated and manufacturing process was
initiated. The UAV was completely built by team members, without any external help. Fiberglass
moulds, composite reinforcements, precise balsa sheeting and many other techniques were applied
by the team members to construct the UAV parts. Final assembly was performed along with avionics
and propulsion integration. Structural testing was carried out by applying the predicted aerodynamic
loads and weight balancing was carried out in order to bring the centre of gravity close to the
aerodynamic centre for better stability features. Finally the flights started with each test flight
constituting of post flight and pre flight checks. The performance parameters of the actual UAV were
also determined.
1.2 Design Characteristics
Obtaining the best solution for the FFD requirements was a challenging task. A scoring matrix
was formulated which revealed that minimizing aircraft weight and flight time will yield the maximum
score. In addition to this UAV must satisfy the constraints to make a short takeoff within 40m distance
and carry payload balls that have to drop during the flight on specified targets. The maximum payload
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capability of the UAV will help in maximizing the last mission score. This variation of scoring criteria
made the solution of FFD UAV quite complex as the lighter aircraft will have better scores in first two
mission but will face problems during the third mission when payload drop will result in large C.G
change for a lighter aircraft. Therefore detailed analysis was carried out to choose the best solution
for all three missions.
The conventional design UAV featuring light rectangular wing mounted on a fuselage with
conventional tails was found to be the most appealing solution as it gave best results for achieving
good performance along with ease of manufacturing and being cost effective. The wing span of 5.8 ft
2
with an area of 7.8 ft was selected to compensate for the thrust limit imposed due to 40A current
requirement. The total length of the aircraft is 5.16 ft with an empty weight of 12 lb. Conventional tail
with horizontal tail span of 32 in and vertical tail height of 15.6 in was incorporated on a 29 in long
boom. The wings were made with traditional foam core method with carbon fiber spars. The two piece
carbon fiber fuselage was extracted from upper and lower female moulds. Tails were made using the
pink foam reinforced with carbon spars. The UAV was powered by AXI brushless motor using a 6 cell
Lithium Polymer (LiPo) battery.
1.3. Aircraft Performance
The UAV was designed to complete all the three contest missions and satisfy all the technical and
safety requirements. The performance parameters of the UAV with maximum takeoff weight are
documented as follows;
Empty Weight
12 lb
Maximum Takeoff Weight
20.6 lb
Maximum Payload Capacity
8.6 lb
Wing Loading
2.66 lb/ft2
CL max
1.6
Maximum Velocity
65 ft/s
Maximum Rate of Climb
240 ft/min
Minimum Takeoff Roll
90 ft
Stall Velocity
38 ft/s
Instantaneous Turn Rate
95 deg/sec
Max Thrust
7.5 lb
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2. Management Summary
2.1. Team Introduction
The team “BURAQ” consisted of four undergraduate students from PAF College of Aeronautical
Engineering. The team was supervised by team advisor Wing Commander Dr Messam Abbas Naqvi
of Aerospace Engineering Department. He teaches subjects of Aerospace Vehicle Design and Aircraft
Performance at CAE and holds a doctorate degree in Aircraft Design and Multi Disciplinary Design
Optimization from Georgia Tech, USA and has been involved in various national level projects of
UAVs. The team was selected on the basis of personal interests and expertise in the field of UAVs.
The brief introduction of team is as follows.
Wing Commander
Dr Messam Abbas Naqvi
Team Advisor
Squadron Leader
Dr Irtiza Ali Shah
Co Advisor
Pilot Officer Waqas Akram
Pilot Officer Syed Aoun
Pilot Officer Usman Umer
Pilot Officer Amir Abdullah
Team Leader/Pilot
Team Member
Team Member
Team Member
2.2. Workload Distribution
Team management is the key to any successful project. Therefore the team was placed in a
certain hierarchy where the team advisor kept an overall watch and took care of all the official
formalities. The team leader led the team assigning job to the rest of the team member during every
design phase and subsequently dividing the workload during the manufacturing process and report
writing. The work distribution among the team member during various design phases is as follows
Waqas
Usman
Abdullah
Aoun
Mission Analysis
Scoring
Mission Performance
Analysis
Propulsion Feasibility
Design Report
Requirements
Conceptual Design
Fuselage Design
Wing Design
Payloads
Tail Design
Preliminary Design
CFD
Structural Loads
Mission Profile
Optimization
Detailed Design
CAD
Structural Analysis
Weight Calculations
Drawing Package
Manufacturing
Fuselage
Wings
Payloads
Empenage
Ground Testing
Propulsion Testing
Wing Strenght Test
Paylolad Release System
CG Testing
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Every design segment was initiated after a meeting of team members and advisor in which the
work strategy was finalized and tasks were assigned to all the team members on the basis of their
aptitude and expertise and milestones were defined for task completion as per the timeline defined.
Based on the individual efforts and complexity of different tasks, compatibility issues were resolved
and where ever needed, either the human resource was beefed up for a complex task or more time
was allotted.
2.3. Timeline
A timeline was also defined with the initiation of the project keeping in view the limitations
imposed by the military training schedule of the trainee officers. All efforts were made to follow the
time line and it helped in the timely completion of the UAV design
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3. Conceptual Design
Conceptual design is technically the most important phase in aircraft design. Any wrong decision
in this design phase can lead to failures at later stage resulting in time and cost penalty. A thorough
research was carried out to ascertain the design of the UAV. All the important factors like cost, weight
penalty, stability, manufacturing and innovations where catered during the selection phase.
Mission
Analysis
Mission
Requirements
Target
Specifications
Concept
Survey
Concept
Selection
However the most important stage before actual commencing of the design phase is to analyze
the required specifications and the mission that the UAV is supposed to carry out. All the factors like
contest requirements, airfield issues and team limitations were studied and a set of target
specifications were chalked out as per the RFP (Request for Proposal) provided by the FFD 2011.
3.1. Mission Analysis
The UAV will be required to fly three different kinds of missions during the contest. The general
flight path of the mission was provided by the FFD 2011. Each mission has slightly different scoring
criteria. The detailed description of the three missions is as follows.
3.1.1. Mission One
In this mission the UAV will fly a two lap mission to determine the drop zone or fire area. The UAV
will not be flying with any kind of payload and the scoring will be done on the basis of total flight time
and the aircraft weight for that mission.
3.1.2 Mission Two
In this mission the UAV will fly a one lap mission with one fire extinguisher ball. The scoring will
be again done on the basis of flight time and weight. However the flight time for this mission will
include the loading time of the payload.
3.1.3. Mission Three
This mission will be flown with the maximum payload capability of the UAV and the balls will be
dropped on the respective fire areas designated by the red lines. The mission score will be judged by
the payload carried and the flight time. Mission three carries 50% more weightage than the other two
missions.
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3.1.4. Mission Requirements
Apart from the scoring factors for each mission there are certain requirements that are to be
fulfilled by the designed UAV which will otherwise result in score deduction. These requirements are
listed as;
(a) Takeoff distance should be less than 40m (130 ft)
(b) Aircraft should make a successful landing to be eligible for the mission score
(c) Aircraft should not sustain any significant damage during the landing
3.2. Airfield Observations
The contest will be held at Hezarfen Airport in Istanbul. The climate
conditions during May are appropriate for RC flying. Also the airfield is
quite vast and seems quite suitable for good flying. The UAVs will be
easily able to make straight in approaches.
3.3. Aircraft Requirements
Along with the mission requirements, there are certain aircraft requirements that the UAV should
satisfy in order to pass the technical inspection prior to the contest.
(a) Aircraft should fit in a box of 4.9 ft x 2.6 ft x 2.6 ft dimension.
(b) Aircraft should be powered by off the shelf Electric Motor and Propeller.
(c) The propulsion system cannot draw more than 40 Amperes of current.
(d) Aircraft should perform unassisted takeoff from onboard batteries.
3.4 Payload Requirements
The aircraft will be required to carry fire extinguisher balls as per the mission requirements. Two
types of balls will be provided by the contest authorities. Ball A has a diameter of 10 cm and a weight
of 1 kg whereas Ball B has diameter of 15 cm and weight of 1.3 kg. The aircraft will have to carry two
or three balls during the last mission which means four types of combinations will be available.
S. No
Ball Type
Number of Balls
Total Weight (kg)
1
A
2
2
2
A
3
3
3
B
2
2.6
4
B
3
3.9
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3.5 Score Analysis
3.5.1. First and Second Mission
The flight score for first mission will be based on relative basis with maximum score going to the
lightest and fastest aircraft. The maximum achievable score is 100 points per mission.
3.5.2. Third Mission
The flight score for the third mission will be based on maximum payload capacity and flight time.
As mentioned earlier the weightage is 50% more than mission 1 and 2. The maximum achievable
score is 438 points with the maximum payload of 3.9kg. This clearly means that carrying the
maximum payload will earn the team maximum points and can overcome the shortcomings in mission
1 and 2.
This also means that the total maximum achievable score is 638 and appropriate weight and time
targets will be selected keeping in mind the limitations and practicality of the team’s approach along
with the scores of highest scoring teams in previous similar competitions.
3.6. Problem Statement
In light of the mission requirements the team’s target was now set to design and develop a high
performance, easy to manufacture and a cost effective UAV that can carry the maximum payload
effectively and complete all three missions successfully satisfying all the other aircraft requirements.
The minimum targeted score was set to 85% of the maximum achievable score.
3.7. Design Requirements
The aircraft was designed keeping in mind the performance of the third mission as it comprised
maximum marks and payload. If the aircraft is well designed for the third mission its performance for
first and second mission will improve automatically.
3.7.1. Takeoff
The aircraft is supposed to takeoff within 130 ft distance. All wheels should be off the runway
before the max distance line. So the target of 100 ft was selected with a margin of 30 ft for any under
or over estimation in design.
3.7.2. Payloads
In order to obtain the maximum score as mentioned in problem statement, the aircraft has to carry the
maximum payload combination of 3 balls having a total weight of 3.9 kg. The payloads have to be
dropped during the flight at designated spots.
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3.7.3. Cruise Velocity
In order to obtain good flight score, the performance of former UAVs in similar configuration were
studied against their payload capabilities and it was decided that the UAV will be have a cruise
velocity of 65 ft/s. A higher target was selected as the actual performance often varies with the
theoretical calculations.
Payload Capability vs Max Velocity
75
VELOCITY
70
65
60
55
50
45
40
2.2
2.4
2.6
2.8
3
3.2
3.4
PAYLOAD
3.7.4. Weight Target
A weight target was set keeping in mind the payload release and weight & balance issues along
with the weights of former UAVs just like the cruise velocity selection. According to histoirical trends
the best empty weight for a payload of 8.6 lb is 9.4 lb. However keeping in mind the 85% maximum
score target and weight & balance issues, 12 lb target empty weight of the UAV was decided.
WEIGHT
Payload Capability vs Max Weight
8.5
8
7.5
7
6.5
6
5.5
5
4.5
4
10
12
14
16
18
PAYLOAD
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3.7.5. Design Score
In view of the above set design targets the expected score were calculated that the team is
expected to obtain during each mission. Since the payloads will be mounted internally so it will not
effect the cruise velocity to a great extent except the induced drag element due to higher lift. Also the
same aircraft with same battery package will be used for all the three missions. The scores were
calculated using the scoring formula provided in the RPF and the refrence weights and time were
taken from previous competitions.
Mission
Flight Score
Weight Score
Mission Score
1
0.88
0.78
68.64
2
0.86
0.75
64.5
3
0.95
2.95
420.375
Total Score out of 638
553.515
3.8. Concept Collection
In order to select a feasible and viable concept configuration, the first step is to make a population
space of a good collection of applicable concepts and ideas so that the best option can be selected.
In addition to conventional concepts, new innovative ideas are required to satisfy the stringent design
requirements. Different ideas from various books, journals and online databanks were obtained and
studied. All kinds of ideas were welcomed from all the members and finally a large pool of concepts
was created.
3.8.1 Aircraft Configurations
Different types of aircraft configurations being used in R/C flying and UAVs were studied. Each
concept was analyzed for fly worthiness in the required design space.
• Canards prevent aircraft
stall to an extent but
their controls are difficult
to design and they
present
stability
problems
• Favourtie wings of earlier
designs. They tend to give
a good amount of lift but
again the weight issues are
a hinderance.
• A great weight saver and
low drag profile. The only
problem remains is the
controlability and stability
issues.
Canards
Bi Plane
Flying Wing
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• This is the most simplest of
all configurations. Suited
for most of the roles in
conventional spectrum.
• Have a good lift to drag
ratio
but
complicated
manufacturing always is an
issue in this configuration.
• If designed carefully can
prouduce a lot of lift but
can be tricky due to weight
penalties
Conventional
Blended Wings
Tandem Wings
3.8.2. Wing Configuration
A lot of wing designs were collected but only a few of them were shortlisted based upon the
strength and lift requirements. Generally the UAVs are designed for long endurance flights however in
this case the aircraft is supposed to carry out a low endurance mission
• These wings are best
suited for short takeoffs
and high lift. However
they are very stable and
have pitch up tendency
which are sometimes not
desirable for pilot
• Conventionally the most
appropriate wing as it
gives the highest velocity
and has good lift and
stability characteristics
• Low wings have longer
takeoff
distances
however they are quite
maneuverable and have
higher
speeds
as
compared to high wings.
High Wing
Mid Wing
Low Wing
3.8.3. Tail Configuration
Tail design is very important for the stability of the aircraft. However, a stable tail design may
incorporate controllability concerns and other issues like the weight penalties or the complexities of
manufacturing. Tail design is also very much linked to the aesthetics of the aircraft which force
designers to select a particular tail design.
• Mostly used tail
design becuase of
its easy stability
characteristics and
design.
• Act as a weight
saver in some
cases but the V Tail
can sometimes be
trickier to handle
• As compared to the
conventional tail
the cruciform tail
has lesser
downwash effects.
Conventional
V-Tail
Cruciform
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• T Tail is best suited
for avoiding
downwash effects
on the tail, however
they are difficult to
manufacture
• A similiar design as
the V Tail. However
it has a slightly
different behaviour
but acts as a weight
saver
• To keep the vertical
tails out of propwash
the H tails are often
used to give better
controls
T-Tail
Inverted-V
H-Tail
3.8.4. Motor Location
The right motor location has an important impact on the aircraft performance and control
behavior. The motor locations can be decided on the basis of accessibility and thrust requirements.
Also increasing the number of motors always remains a valid option to increase the overall thrust of
the aircraft.
• The propeller gets
cleaner air and gives
more thrust. However
the propwash affects
the aircraft controls.
• The
pusher
configuration is good
for stability as it has no
propwash
effects
however due to lesser
clean air the thrust is
affected.
• A good choice when
using multiple motors.
Both the motors get
clean air. However the
calibration of these
motors
is
a
complicated job.
Tractor
Pusher
Wing Mounted
Motor
3.8.5. Landing Gears
Type of landing gears is important for takeoff and landing characteristics. Landing gear bear the
impact loads during the landing phase and give directional control during the takeoff phases when the
prop wash is significant because of low speed high thrust conditions. They are also important drag
adders. The use of adequately designed landing gears can give great comfort to the pilot.
• A very stable
configuration and
good for pilot
handling. However it
is a bit draggy.
• Good for short
takeoffs as the
aircraft is already at
an incidence angle
however the ground
controls are tricky
• Used for heavier
configurations
where landing gears
have to sustain a
heavy impact
• A great weight
saver configuration
but very difficult to
operate with.
Generally used with
HALE UAVs
Tricycle
Tail Dragger
Tandem
Single Main
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3.9. Figures of Merit
Certain major and minor factors that affect the decisions regarding the concept selection were
analyzed and short listed based on the impact on design configuration. However, the primary targets
remained the satisfaction of RFP requirements.
(a)
Weight is the most important factor in FOM as it affects all the other factors directly
or indirectly. The decreased weight means better score, smaller engine, higher
velocity, better controls and lesser cost. The weight also governs the wing loading
which is the most important aerodynamic parameter.
(b)
Portability is one of the important constraints in the RPF as any design no matter
how effective can be useless if it is not portable. Portability means that aircraft should
be able to fit into the required box and can easily be disassembled for this purpose.
(c)
Manufacturing plays a decisive role in any decision because the UAV has to be
manufactured by students themselves who had very little experience and expertise
with model building. Hence, any complex manufacturing could be a very difficult task.
(d)
RFP Performance means that the UAV should be able to satisfy all design
requirements set earlier. These include parameters like stall velocities, takeoff rolls
and maneuverability of the UAV.
(e)
Stability and Controllability is a very factor which defines the control and handling
of the UAV. No matter how capable the UAV may be if it’s difficult to fly then it can
lead to crashes because of poor control and handling characteristics. Since the
airfield analysis showed that airfield is not well suited for the model flying so this FOM
is very important.
(f)
Velocity is the main scoring factor in the flight missions. The aircraft which will
complete the mission in earliest possible time will be awarded the highest marks. So
the design should be able to give good cruise speeds for a faster mission completion.
(g)
Aesthetics have no technical value but from the designers point of view it has its
weightage. It’s a famous saying that “Anything that looks good, flies good.”
3.10. Concept Selection
After studying and short listing different concepts along with the figures of merit the selection
process started. Each FOM was given a certain percentage weightage. Every concept was given
marks out of 10 against each FOM. Subsequently the marks percentage was multiplied by the
weightage percentage. All the marks scored against each FOM were summed up and a total score
out of 100 was obtained. The final score of each concept was compared with the other concepts and
the best or highest scoring concept was selected for the UAV.
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RFP
Performance
Manufacturing 15%
Aesthetics Velocity
5%
10%
20%
Weight
20%
Portability
20%
Stability and
Controlability
10%
3.10.1. Aircraft Configuration
Conventional
Bi Plane
Tandem
Wings
Canard
Blended
Flying
Wing
Wing
Velocity
10
8
6
5
7
6
6
Weight
20
7
6
5
7
5
8
10
8
7
6
7
7
4
Portability
20
7
8
7
7
5
6
Manufacturing
20
8
6
5
8
4
5
RFP Performance
15
7
8
6
7
8
7
Aesthetics
5
6
7
6
8
8
9
100
73.5
68.5
57
72.5
57
63
Stability and
Controllability
Total
Portability and manufacturability proved to be the most decisive factor in overall configuration
decision. Advanced concepts like blended wings and flying wings were good options for payload
flights but their maneuverability and portability was not good. So conventional design and canard
design were two main competitors but conventional design took the edge because of ease of controls
and manufacturing.
3.10.2. Wing Configuration
High Wing
Mid Wing
Low Wing
Velocity
10
6
8
7
Weight
20
8
6
8
Stability and Controllability
10
8
7
5
Portability
20
7
6
7
Manufacturing
20
8
5
8
RFP Performance
15
8
8
6
Aesthetics
5
6
8
7
100
75
65
70.5
Total
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PAFA BURAQ
The only clear solution was the high wing. As its takeoff performance and stability are good as
compared to other options so it was an obvious decision. The mid wing could have been a good
option but it faced a good set back from portability and manufacturing point of view.
3.10.3. Tail Configuration
Conventional
V Tail
Inverted V
H Tail
T Tail
Cruciform
Velocity
10
7
8
7
6
7
7
Weight
20
7
8
8
6
5
6
10
8
7
7
8
8
7
Portability
20
7
8
6
5
6
6
Manufacturing
20
8
8
7
6
5
5
RFP Performance
15
8
7
7
8
7
7
Aesthetics
5
8
6
9
7
8
7
100
77.5
74
71
63.5
61.5
62
Stability and
Controllability
Total
Conventional tail is the most successful configuration for use in tail designs. They are light weight
and have very good stability and control characteristics. They are also very easy to design and
manufacture as vast techniques are available because of their abundant usage.
3.10.4. Motor Location
Tractor
Pusher
Wing Mounted
Velocity
10
8
6
7
Weight
20
7
7
5
10
7
8
4
Portability
20
8
7
5
Manufacturing
20
8
7
4
RFP Performance
15
8
8
6
Aesthetics
5
6
7
8
100
76
71.5
52
Total
The motor location was selected as tractor configuration. Since the pusher motor has thrust
losses due to unclean air. A clean air for higher thrust was required especially in the short takeoff
mode for which the tractor position was an obvious choice. The twin motor or wing mounted motors
were not a good option because of their complexities.
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Tricycle
Tail Dragger
Tandem
Single Main
Velocity
10
7
8
7
8
Weight
20
7
8
6
8
10
8
6
5
5
Portability
20
7
8
5
7
Manufacturing
20
8
7
5
7
RFP Performance
15
8
7
7
8
Aesthetics
5
8
7
6
6
100
75
74
57.5
72
Total
A tight competition existed between tricycle and tail dragger configurations. Even though tail
draggers are difficult to control during landings and takeoffs but their portability and weight savings
gave them an advantage. Here the aesthetics came in and the slight edge was given to the tricycle
landing gears based on designer choice.
3.10.6. Payload Location
The maximum payload will be three balls that will be
dropped during the flight one by one. This creates a
balance issue because when the payload will be dropped
it will disturb the C.G of the aircraft and the pilot will have
to trim the aircraft accordingly to adjust the balance of the
aircraft. Therefore the decision of the payload locations
and alignment was made by the pilot and it was decided that the payloads will be arranged along the
longitudinal axis and will be mounted internally.
3.11. Final Concept
As a result of decision matrices analyzed above the final
configuration of the UAV was finalized. The conventional
fuselage was modified by adding a tail boom due to portability
issues. The final design consisted of;
(a)
Conventional Fuselage with Payload Storage
(b)
Two Piece High Wing Design
(c)
Boom mounted Conventional Tail
(d)
Tricycle Landing Gears
(e)
Tractor Motor Configuration
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4. Preliminary Design
After finalizing the configuration of the UAV, next
step is the detailed sizing and aerodynamic analysis
of the aircraft. At the end of this phase dimensions
will be frozen. However, during this sizing phase,
iterative schemes are followed to optimize the
aircraft weight and size. The correct airfoil for UAV
will
be
selected
and
analyzed
along
with
determination of aircraft lift and drag characteristics.
This data will determine the stability, structural loads
and performance of UAV. Based on the results
necessary changes will be incorporated and desired
process will be iterated.
For the preliminary analysis the design methodology of “Daniel P Raymer” was used. Starting with
the design point which was evaluated on the basis of design specifications, initial weight estimation
was performed. Depending upon the weight the geometry was finalized and then aerodynamic
analysis of this geometry was carried out. Using the aerodynamic characteristics, the propulsion,
structural loads and stability of UAV were calculated. Finally the performance was predicted and
optimization was carried out.
4.1. Mission Profile
A mission profile was chalked out on the basis of design specifications. This is the exact mission
that the aircraft will fly on the final day. This mission profile served as a reference point for the further
analysis and optimization process
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4.2. Design Point
Design point means selection of the Wing Loading and Thrust to Weight ratio of the UAV. The
design point was calculated against the design specifications. Based on the all the mission segments
and other performance, the design point that is most appropriate for our design was selected. For this
purpose, wing loading and thrust to weight ratio for each requirement was calculated and then the
smallest values of T/W and W/S were selected. It was noted that the stall speed proved to be the
decisive factor in terms of wing loading and the takeoff helped decide the value of thrust to weight
ratio.
T/W
0.35
W/S
2.66
4.3. Weight Estimation
For all the initial calculations a weight estimation of aircraft was required. After some basic home
work it was found that weight target is previous sections was achievable so the total gross weight of
the aircraft for each mission was as follows.
Mission
Weight
1
12 lb
2
14.9 lb
3
20..6 lb
4.4. Airfoil Selection
Airfoil is no doubt heart of the aircraft. It’s just the magic of airfoil that makes dead weight reach
the skies. Selecting the correct airfoil is a very complicated decision because of the impact it can have
on the aircraft performance as well as weight and manufacturing process. An airfoil selection criterion
was defined and based on that the best airfoil was picked. The three main criteria were;
(a)
Maximum Lift Coefficient: The CL max of an airfoil directly affects the stall and
takeoff properties of the aircraft so high value of maximum CL max is desired.
(b)
Lift to Drag Ratio: Maximum lift to drag ratio can be termed as aerodynamic
efficiency of the aircraft. Higher the value of L/D)max, the better performance is
expected out of an aircraft
(c)
Maximum Thickness: The thickness doesn’t only define the stall behavior but also
adds to the weight of the wings; hence reasonably thick airfoil had to be selected.
Airfoil
CL Max
Best L/D
Thickness
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SD7062
1.59
50
14%
MH113
2.14
60
14.7%
MH114
1.98
45
13%
Clark YM-18
1.70
46
18%
Clark YM-15
1.59
52
15%
GOE 611
1.84
44
14%
4.4.1 Airfoil Analysis
Since the MH113 airfoil’s data was collected from internet sources, it was decided to analyze the
airfoil using numerical and computational techniques. Initial testing was carried out utilizing “Design
Foil” software that uses panel method (potential flow solver with viscosity) to evaluate the lift curve
slope of the airfoil. Since the viscous effects could not be defined in the panel method accurately,
validation of CL max of the airfoil was carried out by CFD analysis. A meshed grid was generated
around a 2D airfoil in GAMBIT 2.4 software. The pressure inlet and outlet boundary conditions were
given and the case file was exported to FLUENT 6.3. Here the airfoil was simulated at cruise speed of
15 m/s and sea level conditions. The drag polar from “Design Foil” and CFD contours from FLUENT
along with aerodynamic characteristics of MH113 airfoil are given below:
Max CL:
2.14
Max CL angle:
15.0
Max L/D:
Max L/D angle:
60.475
4.5
Max L/D CL:
1.453
Stall angle:
17
Zero-lift angle:
Thickness:
-8.0
14.7%
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4.5. Geometric Sizing
Traditionally, aircraft geometric sizing is done based on statistical equations, analytical formulae
and design heuristics, but the constraint of aircraft to be fitted in a predefined box took priority over all
others.
The payload capacity also dictated the wing size for adequate lift generation. A detailed
dimensional drawing of the UAV can be seen in the drawing package in the detailed design segment.
However the brief summary of the UAV sizing is as follows:
4.5.1 Fuselage
The fuselage has to house the internal payloads along with wing attachments and propulsion
package. So the sizing was done keeping in mind the payload and box constraints. The length of the
aircraft could be maximized by using a tail boom as it will provide required moment arm to the tails.
4.5.2. Wings
Based on the selected wing loading the reference area of the aircraft came out to be 7.8 ft2. The
span was selected as 70 in for a two piece wing. In order to maximize area, chord length was
increased, though it decreased aspect ratio which resulted in reduced lift but increased the stall angle
which was desirable. No dihedral was given as wing was already of high wing configuration. Also the
sweep angle was zero to increase the lift curve slope as stall will be delayed already by the lower
aspect ratio. Similarly there was no tapering to avoid the manufacturing complexities. Aileron sizing
was done with historical trends mentioned in Raymer’s Design Book
4.5.3. Tail
Initially the tails were sized using the traditional tail coefficients but later during the stability
module it was noted that due to downwash effects the tail size needed to be increased. Hence, the tail
sizes along with the tail boom length were revised in order to increase the static margin.
4.6. Aerodynamic Analysis
The complete aerodynamic characteristics of the UAV can be summed up to the drag polar
evaluation. The lift producing capability along with the amount of drag penalty produced. Analytical
methods purposed for such analysis were used to determine the lift curve slopes and maximum lift
coefficient of the aircraft. Validation of the analytical results was performed by a CFD analysis of one
configuration at one flight condition. This helped improve the parasite drag estimation of the aircraft as
the flat plate skin friction analysis underestimate the aircraft drag.
The analytical methods for drag estimation cannot be utilized with reasonable accuracy for actual
model building. Hence, a CFD analysis of the aircraft geometry was carried out to determine the zero
lift skin friction drag of the aircraft and validate/refine the analytical results. Fudge factors were added
to the analytical results based on the CFD to make them close to real values.
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For CFD a tri mesh was structured using the GAMBIT software after which the case file was
exported to FLUENT for analyses. Pressure based analysis were carried out by applying the pressure
far field, pressure inlet and pressure outlet boundary conditions. Analysis was carried out at cruise
conditions.
4.6.1 Lift Curve Slope
The lift produced by the aircraft was determined by CFD at different angles of attack. Later the
values were interpolated to obtain a straight line. The stall characteristics were estimated initially
based on the airfoil values and later validated with CFD.
CL
Lift Curve Slope
-20
-10
2
1.5
1
0.5
0
-0.5 0
-1
AOA
10
20
4.6.2. Total Drag
The total drag force experienced by the UAV during sea level cruise is plotted below. This is also
the minimum thrust required to keep the UAV airborne. It can be seen that design cruise point lies
exactly on the L/D max velocity
Drag (lb)
20
Drag
15
10
5
0
-5
20
45
70
95
120
Velocity
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4.6.3. Drag Polar
A drag polar is the essence of the complete aerodynamic analysis. It states the lift to drag
relationship of an aircraft. The best L/D ratio can be predicted from drag polar which yields best
performance parameters. The best L/D for this UAV is 10. UAVs generally have higher L/D ratios but
they are optimized for endurance missions where as in this case, it is a maneuverable aircraft which
will never fly endurance missions. The drag polar along with other aerodynamic parameters are;
Drag Polar
2
CL
1.5
1
0.5
0
0
0.1
0.2
0.3
CD
CLα
4.02 / rad
CDo
0.047
CL max
1.58
L/D (Max Weight)
10
K
0.078
Typical Re No
4 Million
e
0.91
4.7. Propulsion Analysis
A propulsion system must be able to account for the worst case drag penalties and at times is
also required to augment lift and handle weight component for better aircraft performance. All the
performance considerations in the air machinery are governed by the propulsion system. The more
powerful the propulsion system is the more easily the aircraft will be able to perform different mission
segments and achieve high velocities with higher accelerations. Installed motor performance is
always different from uninstalled motor performance due to the losses, additional drags and
interference effects.
The AXI-4130/20 brushless motor was selected due to its appropriate RPM/Volt capability and
good internal resistance. The motors with a smaller RPM/V required bigger propellers and were
heavier where as the motors with larger RPM/V were not able to rotate large props.
Motor
AXI-4130
AXI-5320
AXI-4120
RPM/V
305
206
512
Resistance(m ohm)
0.99
0.84
0.7
Efficiency(%)
88
93
86
Weight(g)
409
495
320
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The next step was to analyze different propellers and battery combinations to get the required
thrust using the minimum weight penalty. LiPo batteries were allowed in the contest so they were an
obvious choice over the Ni-Cad batteries due to their better efficiency. Propellers of different
diameters were selected and analyzed for required thrust within the allowable current limit.
Thus using the 6 cell 22.2 V LiPo battery and a 15 in diameter propeller the required thrust was
achievable. The analysis for the mentioned setup using the Blade Element Momentum theory is;
Velocity(m/s)
Current(A)
RPM
Power(W)
Thrust(lb)
0
30
6500
361
7.43
5
27.9
6577
340
6.06
10
25
6685
310
4.66
15
21.5
6830
268
3.35
20
17
7015
211
2.19
25
11.3
7238
137
1.2
4.8. Structural Loads
Before the actual structural members can be sized and analyzed, the loads applied during flight
must be determined. Aircraft’s loads estimation is a separate discipline of aerospace engineering. It
combines aerodynamics, structures, and weights categories for loads estimation. Loads on aircraft
can be of different categories like tensile, compressive, bending and torsional. As different types of
aerodynamic and inertial loads are applied at different parts of the aircraft during flight, different parts
are designed structurally to withhold these loads.
Maneuver load is the load generated when an aircraft performs high g maneuvers. Maneuver
loads are generally the highest loads sustained by the aircraft structure. This largest load which
aircraft experience is called the “Limit” or “Applied” load. Aircraft loading is expressed in terms of load
factor ‘n’. The aircraft structure is designed to have the limit structural strength of more than the limit
loads. The V-n diagram depicts the aircraft limit load factor as a function of the air speed. The aircraft
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can be stalled at a higher speed by trying to exceed the
available load factor, such as in a steep turn. The aircraft
maximum speed, or dive speed, at the right of V-n diagram
represents the maximum dynamic pressure q limit. The
point representing maximum q and maximum load factor is
clearly important for structural sizing. At this condition, the
aircraft is fairly at a low AOA because of high q, so the load
is approximately vertical in the body axis.
Category
Type
Load
Lift
Distributed
80 lb
Thrust
Point
25 lb
Landing
Impact
75 lb
Tail
Distributed
25 lb
Payloads
Distributed
8 lb
4.9. Stability Analysis
The basic concept of stability is simply that a stable aircraft, when disturbed, tends to return by
itself to its original state (pitch, yaw, roll, velocity etc). But Stability and Control are opposite to each
other, so the best combination is a good tradeoff between stability and controllability.
Static stability is present if the forces created by the disturbed state push in the correct direction to
return the aircraft to its original state. The requirement for good stability, control and handling
quantities are addressed through the use of tail volume coefficient method and through location of
aircraft centre of gravity at some percent of wing mean aerodynamic chord. The static stability of
aircraft has been computed for longitudinal, lateral and directional axes. The main equations
governing these stability factors are;
Tail Airfoil
NACA 0010
Cl α Horizontal Tail
0.06 / deg
Aspect Ratio Horizontal Tail
2.66
Aspect Ratio Vertical Tail
1.32
C.G Location
At Aerodynamic Centre
Cm α
-0.012 / deg
Cn β
0.055 / deg
Cl β
-0.018 / deg
Static Margin
0.23
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4.10. Aircraft Performance
After sorting out the weight estimations, lift, drag and propulsion characteristics of the UAV,
aircraft performance was evaluated. Standard analytical methods were used to evaluate different
performance parameters. These performance results are based on the estimated weights.
4.10.1. Rate of Climb
excess power that an aircraft has along with defining the
flight envelope and parameters like maximum and minimum
velocities. The following graph was obtained for ROC in
ft/sec with maximum payload.
ROC
Rate of climb of an aircraft is the measure of specific
5
4
3
2
1
0
0
20
40
Velocity
60
80
4.10.2. Turn Rate
UAV has to perform certain maneuvers in each lap of
the mission. In order to achieve these mission requirements,
a good turning performance is mandatory. Also due to small
drop zone the UAV will have to make tight turns. The
instantaneous turn rate along with the sustain turn rate was
plotted as shown.
4.10.3. Takeoff and Landing
The aircraft when loaded with full payloads will be able to takeoff within the 117 ft distance and
land within 200 ft distance without brakes
Mission
Takeoff Roll (ft)
1
55
2
80
3
117
4.10.4. Mission Performance
The performance of the UAV predicted during each mission is documented below. Mission 1 has
no payload but 2 laps where as the distance for mission 2 are shortest.
Mission
Aircraft Weight (kg)
Maximum Cruise Speed (ft/s)
Stall Velocity (ft/s)
1
5.5
74
28
2
6.8
69
32
3
9.4
65
37
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5. Detail Design
After the preliminary design is completed, the
detail design phase was started. All systems and
components were designed, selected and integrated.
The aircraft structural analyses were done on critical
areas and a complete aircraft sizing was carried out
in Solid Edge V19 CAD software.
Weight and
balance for both external and internal payloads was
carried out. Finally, flight and mission performance
parameters were calculated, along with the Rated Aircraft Cost of the whole UAV package.
5.1. Dimensional Characteristics
The dimension of main aircraft parts are documented below. The following dimensions were
finalized after integration test and were then strictly implemented during the manufacturing phase.
Fuselage
Horizontal Tail
Length
36 in
Span
32 in
Width
6.5 in
Area
2.66 ft2
Height
7.5 in
Chord
12 in
Aspect Ratio
2.66
Tail Boom
Length
29 in
Elevator Chord
2.5 in
Width
1 in
Elevator Deflection
30 deg
Incidence Angle
0
Wing
2
Vertical Tail
Area
7.8 ft
Span
70 in
Span
15.6 in
Chord
16 in
Area
1.33 ft2
Aspect Ratio
4.35
Sweep Angle
36 deg
Aileron Length
24 in
Rudder Chord
2.5 in
Aileron Chord
3 in
Dihedral
0
Rudder Deflection
30 deg
Main Landing Gears
Sweep Angle
0
Height
8 in
Width
16 in
Tire Diameter
3 in
Box
Length
40 in
Width
24 in
Height
24 in
Weight
Nose Gear
Height
8 in
Tire Diameter
3 in
Steering Angle
45 deg
Airframe
4 kg
Payloads
3.9 kg
Propulsion
1.1 kg
Diameter
15 in
Avionics
0.4 kg
Pitch
10 in
Total Maximum
9.4 kg
Propeller
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5.2. Structural Systems and Capabilities
All the structural designing and material selection was done keeping in mind the FOMs. The
structure was designed keeping in mind the assembling ease and portability.
5.2.1 Fuselage
Designing the fuselage was the most complicated job as it had to house the propulsion system,
control avionics, batteries and the most important “payloads”. The final product was a carbon fiber
fuselage. The section strength was provided by the plywood bulkhead and pink foam partitions for the
fuselage. The critical areas like wing attachment and landing gears were reinforced with extra layers
of composite and plywood for stiffness. A styro-foam piece was added at the tail attachment to give
additional grip to the tail boom. Lower portion of the fuselage comprised compartment doors for
payloads.
The plywood section consists of four parts. First one is the motor firewall that is made of double
plywood and will work as motor and nose gear attachment. The second section is the front bulk head
that will restraint the batteries and nose gear servo. The third is the rear bulk head that was the tail
boom attachment followed by a small firewall that will restraint the boom moment. In-between the
bulkheads pink foam was placed that will not only restraint the payloads but also provide sectional
strength to the fuselage. The wings were attached in the dowel holes from front and screws from rear.
The main landing gear was attached near the rear bulkhead.
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5.2.2. Wings
Wings are the main load bearing members. They were strengthened by carbon fiber spars that
will bear the main bending aerodynamic loads. The rest of the wing is made up of Styrofoam sheeted
with 1/32” balsa sheet. The leading edge and trailing edge is also made up of balsa. Wing has been
covered with monokot sheet for smooth surface finish. Wings also house servos for aileron controls.
The ailerons were made up balsa reinforced with fiber glass near the control horns. The wing
attachments are two wooden dowels in the front and screws at the back. The two piece wing is joined
using the hardwood joiner that also works as reinforcement to carbon fiber spars.
Since wing spar was the main load bearing member that will bear all the extreme loads, its
analysis was done in ANSYS 12 in which the carbon fiber spar was given distributed elliptical loads at
4g conditions and the bending behavior was studied. The critical areas were near the roots which
were reinforced with the hardwood joiner. The limited deflection was also restricted with the fiber glass
layup on roots and tips.
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5.2.3. Tail Assembly
The conventional tail was made using the pink foam wings mounted on the aluminum boom. The
horizontal and vertical tails were made of pink foam with the density of 33 kg/m3. The trailing edge
was made of balsa along with the rudder and elevator. The tails were reinforced with carbon fiber
spars. The tails were joined to the boom using bolts. Similar bolts were also used to restrain the boom
moment from fuselage-boom attachment.
The aluminum boom was analyzed in ANSYS and again it was found weak to sustain maximum
loads so it was also reinforced with wood near the boom attachment. This also gave strength for
impact loads that are quite common in RC flying
Carbon Fiber main landing gear was taken off the shelf with about 8 in height. The tires were also
taken off the shelf to maintain the required clearance and get a softer impact on landings. The mild
steel nose gear was also taken off the shelf. The Landing gear analysis were also carried out for
impact loads in ANSYS 12.
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5.2.5. Aircraft Box
Aircraft Box was made up of fiber glass reinforced with wooden
rods. The size of the box was minimized to avoid the weight penalty.
The weight of the box is about 4 kg. Aircraft disassembles to fit into
the box. The two piece wing detaches along with the tail assembly.
The fuselage remains intact to save the 5 min assembling time limit.
The box house the whole package including the whole aircraft, radio
transmitter and small tools required for assembling of the aircraft.
5.3. Systems and Sub Systems
The aircraft airframe was integrated with propulsion system that comprised of electric motor,
batteries and speed controller. Also the control system consisted of servos, receiver and battery. The
payload release system was also designed so that payloads can be dropped during the flight.
5.3.1. Propulsion System
The propulsion system was selected on the basis of results in the preliminary design. The
equipment was chosen on the basis of weight and quality. The electric motor was mounted on the
front firewall where as the battery and electric speed controller was placed just behind the firewall in
the fuselage. The fuse was mounted on the fuselage behind the motor.
Motor
AXI-4130/20 Brushless Motor
ESC
Castle Creations Phoenix ICE 60
Battery
Thunder Tiger 6 cell 22.2V 20C Lithium
Polymer Battery
Propeller
15 x 10 APC E
Fuse
40 Amp Blade Style Fuse
5.3.2. Control System
Conventional control mechanisms were adopted because of their reliability observed during the
aero modeling activities. The selection of servos was based on the loads and weight. For this purpose
standard servos were installed on the wings for aileron controls. The mini ball bearing servos were
installed on the tail surfaces. The metal gear mini servos were installed for the payload release
system. The transmitter selected was a 9 channel radio which supported different kind of control
mixes and flight conditions that were used for payload release.
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Transmitter
Hi-Tech Aurora 9
Receiver
Hi-Tech Optima 9
Battery
Chameleon 2100mah 4.8V Battery
3 x HS-325HB Standard Servo
Servos
2 x HS-225BB Ball Bearing Mini Servo
3 x HS-225MG Metal Gear Mini Servo
Apart from standard controls certain control mixing was done for better flight characteristics and
balance during payload release.
Control Mix
Description
Flaperons
To decrease the takeoff roll during third mission
Nose Gear + Rudder
For conventional ground steering
Flight Condition 1
Elevator normal for no/full payload flight
Flight Condition 2
Elevator up and front payload released
Flight Condition 3
Elevator normal and rear payload release
Flight Condition 4
Elevator normal and center payload release
5.3.3. Payload Release System
A very simple payload release system was developed. The payload compartment door was
restricted by the servo heads as shown in the picture below. The servo was given 90 deg rotations
upon which the doors are unlocked and the ball falls out due to gravitational force. After the release
the door closes with the assistance of air flow and attached spring mechanism. The assembly is same
for all compartments.
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5.4. Weight and Balance
It is very important to determine the weight and center of gravity of the aircraft. These calculations
can refine the stability and performance parameters. For this purpose the Solid Edge V19 CAD
software was used to estimate the overall weight along with the CG position. The densities of every
component was measured manually and then inserted in Solid Edge PP module.
5.4.1. Center of Gravity
As shown below the green dot indicates the CG of the aircraft which is almost near to the quarter
chord. The payloads are placed at the CG positions so there is apparently no change in the CG of the
aircraft because this was an intentional design feature.
From the above snap shots from CAD it could be seen that aircraft when fully loaded will have its
C.G exactly on the quarter chord of the wing that is the design C.G point. When the front payload is
released the C.G shift about 2.5 inch rearwards which remain within the static margin and the aircraft
is stabilized by trimming the elevator. After the rear payload is released the C.G return to the original
position. However when unloaded the C.G is negligibly ahead of the designed C.G which is adjusted
before the flight by readjusting the RX battery. The following data was obtained from the CAD
software.
Mass Moment of Inertia
Ixx
0.424417 kg-m^2
Iyy
0.760202 kg-m^2
Izz
1.105789 kg-m^2
Ixy
-1.105789 kg-m^2
Ixz
0.051613 kg-m^2
Iyz
-0.000922 kg-m^2
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5.4.2. Detailed Weight Estimation
This is a very important aspect of aircraft design. We have to get exact weight estimations for
determining the performance of our system. Also this weight is used to get the cost estimate of the
aircraft.
5.4.2.1. Aircraft Empty Weight
Wings
1500
Wing Joiner
100
Fuselage
1200
Main Gear
150
Nose Gear
50
Boom
210
Tail Assembly
360
Motor
400
Propeller
100
Battery
600
Servos (8)
310
RX Battery
100
Control Attachments
100
Epoxies and Attachments
400
ESC
110
Total
5690 g
5.4.2.2. Package Weight
Airframe
4070
Propulsion & Avionics
1620
Transmitter
1100
Box
4400
Tools
300
Total
11490 g
Tools
3%
Airframe
35%
Box
38%
Propulsion
14%
Transmitter
10%
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5.5. Final Aircraft and Mission Performance
Mission
1
L/D
2
3
10
Weight (lb)
12.34
15.2
20.94
Payload (lb)
0
2.86
8.58
T/W
0.6
0.49
0.35
W/S (lb/ft2)
1.58
1.95
2.68
Takeoff Distance (ft)
47
68
114
Maximum Cruise Speed (ft/s)
70
68
65
Stall Speed (ft/s)
28.5
31.8
37.4
Rate of Climb (ft/s)
6.4
5.5
3.6
Loading Time (s)
0
15
45
Assembling Time (s)
4
4
4
150
60
80
Mission Time (s)
5.6. Drawing Package
The drawing package extracted from CAD consists of following parts;
(a) 3 View Drawing
(b) Structural Component Layout
(c) Propulsion and Avionics System Integration
(d) Payload Release System
(e) Compact Package Drawing
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6. Manufacturing Plan
After all the designing comes the actual fabrication. Manufacturing was carried out in the AE
Design Lab of CAE and Aero modeling Club at PAF Academy Risalpur. All the fabrication was done
by the students with cooperation of different departments and labs located within the premises of
CAE. Manufacturing is a very important phase as it requires a lot of planning because any kind of
error can now cost time and money. Also wrong techniques can result in weight penalty and
compatibility issues. Certain decision were made on the following figures of merit.
6.1. Figures of Merit
Strength
20%
Ease of Availablity
15%
Cost
10%
(a)
Ease of
Manufacturing
35%
Weight Penalty
20%
Ease of Manufacturing is an important factor as the team members are building a
UAV for first time so a complicated process can give a lot of issues so this was given
higher priority so that UAV can be manufactured easily and accurately
(b)
Cost is also very important factor as the funds available are limited and the aircraft
cost have to be kept below the prescribed limit.
(c)
Weight Penalty can lead to a lot of problems specially the mission performance so
any technique that has drastic effect on weight will not be considered
(d)
Ease of Availability is a practical issue that can cause unnecessary time delays
because our institution is far away from main cities so this factor was always kept in
mind
(e)
Structural Strength is very important for any aircraft to get airborne. The
manufacturing strategy should ensure adequate strength for the UAV
Based on the weightage given to each FOM the manufacturing technique and the material were
selected by weighting each concept against all the FOM in a decision matrix. Following is the
description of the manufacturing process of the UAV.
43
PAFA BURAQ
6.2. Fuselage
For the manufacturing of the fuselage certain techniques were discussed and analyzed. These
techniques were seen in the list of FOMs discussed earlier.
Technique
Male Mould/Glass
Fiber
Female Mould/
Carbon Fiber
Plywood/Balsa
Composite
Reinforced Wood
Ease of
Manufacturing
Weight
Cost
Ease of
Availability
Strength
Total
7
6
7
8
7
69.5
7
8
5
7
8
72
6
5
7
8
6
62
5
7
6
5
7
59
Making a fiber glass fuselage from foam male mould was easiest but the weight control over
epoxy was difficult and more glass layers were required for strength. The traditional balsa and
plywood fuselage used in RC models was also considered but again the weight penalty was the
hurdle. Reinforcing the same wooden fuselage with composites was considered but it was a very
complicated technique. Thus it was decided to use carbon fiber for fuselage.
First of all the construction of mould started. For this purpose the styrofoam was cut into the
desired shape of the fuselage. After that it was layered with fiber glass layers for adequate strength.
The foam was dissolved and the fuselage was cut into two upper and lower halves. Then the wooden
reinforcements were attached for strength and resist deformation of the fuselage shape. For finish of
the mould first it was finished with sand paper and then the gaps were covered with appropriate filler.
Subsequently the primer and gel coat were applied for the finish and the mold was ready. The
vacuuming bagging bag was formed and attached to the mould.
2
The carbon fiber cloth of 250 g/m density was applied then to the mould with LY-556 epoxy
hardened with HY-951 hardener. The vacuum bagging was done to extract any unwanted and extra
epoxy. The composite was cured at 60C temperature for best strength. After 24 hrs the upper and
lower parts of the fuselage were obtained. The plywood reinforcements were inserted and then the
upper and lower fuselage was joined by the fiber glass patch. The finally the access panels and
44
PAFA BURAQ
payload door were cut out and the fuselage was finished with grinder and sand paper before applying
the monocot for smooth surface.
6.3. Wings
Wings were supposed to be the most accurate and rugged part of the UAV as they have to
generate all the lift and then sustain these lift loads effectively. Foam core wings with aluminum spars
were selected as they were very easy to manufacture accurately
Technique
Manufacturing
Weight
Cost
Availability
Strength
Total
Balsa Sheeted Foam
8
8
8
8
7
64
Reinforced Pink Foam
7
5
5
6
8
48.5
Wooden Buildup
6
7
7
8
7
54
The Styrofoam was hot wire cut by the help of airfoil
templates that were made by drawings from plotter. Then
the carbon fiber spar was inserted and joined with epoxy
glue. After that the foam was fined with sand papers and
balsa sheet of 1/32” were sheeted on the foam with the
help of lattice glue. Solid balsa leading edge was glued
with epoxy glue and then shaped by using the planar.
Similarly the balsa trailing edge was also attached. The
lighting holes were made by removing the unwanted foam
and then the wing was reinforced with glass fiber cloth of
6K 330g/m2 density. The wings were then again fined and
covered with monokot sheets. Finally the balsa ailerons
were attached with the help of plastic pin hinges to the
trailing edge. The cylindrical joiner for the wings was made
with the help of lathe machine.
45
PAFA BURAQ
6.4. Tails
Some extra weight saving effort was required in the case of tails as they are located at a
considerable moment arm so any weight penalty can cause C.G balance problems. The following
techniques were considered;
Foam Core
Ease of
Manufacturing
7
Weight
Penalty
6
Pink Foam
Balsa Build Up
8
6
8
7
Technique
7
Ease of
Availability
8
6
8
7
8
Cost
Strength
Total
8
71.5
7
8
74.5
71
The pink foam tails were cut using the hot wire apparatus.
The tail wings were then reinforced with the carbon fiber rods of
4mm. The rods were glued into the slots created in the tails with
hot wire which were later covered with a foam piece. The tails
were then fined using the sand paper and finally a fiber glass
patch was added to the roots for additional strength especially
for the attachments. Elevator and rudders were made with solid
balsa with the help of a planar. Aluminum boom was purchased
off the shelf and modified as per requirements.
6.5. Manufacturing Timeline
The manufacturing plan was made to ensure in time fabrication of the UAV. The days were
allotted keeping in mind the military training commitments of the students. The black lines shows the
actual progress against the planned one.
46
PAFA BURAQ
7. Testing Plan
After the manufacturing was complete the ground testing of the aircraft structure was carried out
and then subsequently the test flights were started. These tests were designed to find out any
possible short coming in the design so that it can be rectified before the actual contest.
7.1. Ground Testing Plan
The following ground tests are planned to check the strength and integrity of the aircraft structure
and avionics system. The propulsion system will also be tested accordingly.
Wing Strength Test
Landing Gears Test
Point loads equivalent to the flight loads will be placed on the wings and
the strength of the wing for 4g loads will be tested.
The impact loads will be simulated on the landing gear by throwing the
aircraft from 3 ft height.
Motor Power Test
Motor thrust will be measured to verify it with the theoretical results
Motor Current Test
The current drawn by the motor will be measured
Payload Release
The payload release mechanism will be tested and improved
CG Test
Avionics System
Radio Range Test
Tail Boom Test
The CG of the fully loaded and unloaded aircraft will be checked and will
be corrected if deviated from theoretical value
The radio programming will be checked and it will be ensured that all the
controls are working the way they should
Radio will be tested for range by moving the transmitter away from the
aircraft and also the Fail Safe mode will be checked
The strength of the tail boom will be tested by applying the impact and
flight loads and also checking the strength of the tail boom attachment.
7.2. Check List
The following checklist will be used for every flight test as well as at the competition.
Check C.G at design location
Check all controls move in correct directions
Check wing attachment integrity
Check tail boom attachment integrity
Check tail attachment integrity
Check propeller integrity
Check battery voltage
Check RX/TX battery voltage
Check aileron throws
Check elevator throws
Check Rudder Throw
Check payload doors closed
Check batteries are secured
Check nose gear steering
Check flight conditions preset mode
Check fail safe settings
Check motor for full power
Check fuse installed properly
Radio Range Check
Check Transmitter Antenna
47
PAFA BURAQ
7.3. Flight Testing Plan
The following flight testing plan will be carried out before proceeding to the competition. A total of
15 flights are planned as part of testing. More flights can be carried out in case of improvements are
needed or more practice is required.
Flight
Payload
1st
Nil
2nd
Nil
3rd
4th
Nil
Nil
5th
1 Ball
6th
1 Ball
7th
8th
1 Ball
1 Ball
9th
3 Balls
10th
3 Balls
11th
12th
13th
14th
15th
3 Balls
3 Balls
Nil
1 Ball
3 Balls
Observations
High Speed Taxi
Takeoff Attitude
Stability Behavior
Glide Behavior
Landing Speed
Takeoff Distance
Maneuvering Capability
Maximum Speed
Stall Speed
Landing Roll Distance
2 Lap Flight Time
Battery Endurance
Takeoff Distance
Maximum Velocity
Stability Behavior
1 Lap Flight Time
Glide Performance
Stall Behavior
Landing Distance
1 Lap Flight Time
In Flight Payload Drop
Takeoff Distance
Maximum Velocity
Stability Behavior
1 Lap Flight Time
Glide Performance
Stall Behavior
Landing Distance
1 Lap Flight Time
In Flight Payload Drop
First Mission Rehearsal
Second Mission Rehearsal
Third Mission Rehearsal
48
PAFA BURAQ
8. Performance Results
8.1. Ground Testing Results
The testing plan decided earlier was implemented to check the performance and other parameter
of design.
8.1.1. Wing Strength Test
The equivalent flight loads were applied on the wing to check the strength. The distributed loads
of 10kg, 20 kg, 30 kg and 40 kg were applied for 1-4 g load conditions. There was very less deflection
noted that too because of styrofoam compression at the tips.
8.1.2. Landing Gear Test
The landing gears were initially tested by throwing the aircraft from 3 ft height and then they were
subjected to 80 lb equivalent impact load in the impact testing machine. The carbon fiber landing gear
cleared all the test will maximum 0.3 inch deflection.
49
PAFA BURAQ
8.1.3. Motor Power and Current Test
The motor was placed in the CAE Wind Tunnel and the thrust was measured with the help of
spring balance attached to the motor. The Castle creation data logging software was used to measure
the current and the thrust, power and current plots were obtained for the 15x10 APC E Prop.
Velocity
Forces (lbs)
Moments (lbs)
RPM
Power
Current
THRUST
mph
LF
DF
SF
PM
YM
RM
%
Watts
Amps
lbf
0
0
-0.78
-0.07
-0.02
0.08
-0.03
25
31
1.55
-0.52501
0
-0.1
-1.93
-0.11
-0.15
0.07
-0.26
50
101
5.04
2.011213
0
-0.2
-3.74
-0.11
-0.46
0.03
-0.37
75
208
13.35
5.776199
0
-0.3
-5.55
-0.13
-0.722
0.01
-0.47
100
497
25
9.189735
13.6
-0.1
-0.51
-0.12
0.14
0.06
-0.11
25
35
1.72
-0.87325
13.6
-0.2
-1.58
-0.11
0
0.06
-0.17
50
106
5.28
1.498258
13.6
-0.2
-3.38
-0.12
-0.29
0.03
-0.35
75
274
13.76
4.793674
13.6
-0.3
-5.08
-0.15
-0.59
0
-0.55
100
498
25.16
8.357225
27.3
-0.2
-0.18
0.03
0.24
0.06
-0.13
25
26
1.32
-2.22724
27.3
-0.1
-0.76
0.06
0.09
0.06
-0.28
50
81
4.06
-0.37626
27.3
-0.1
-2.82
0.03
-0.25
0.02
-0.48
75
287
15.05
5.06324
27.3
-0.2
-4.18
0.01
-0.5
-0.01
-0.52
100
486
24.7
6.676444
54.5
-0.1
0.42
-0.08
0.12
0.06
-0.14
25
18
0.9
-3.67524
54.5
-0.1
0.26
0.04
0.14
0.07
-0.03
50
41
2.08
-3.32272
54.5
-0.1
-0.15
0.03
0.01
0.05
-0.31
75
90
4.52
-2.65392
54.5
-0.1
-1.77
0.08
-0.13
0.03
-0.51
100
346
17.34
2.716378
8.1.4. Payload Release and C.G
The integrity of payloads and C.G shift was measured. After the front ball is dropped there is a
C.G shift of about 2 in. The releasing mechanism was tested and as a result the servo heads of the
servos were adjusted and payload doors were redesigned. The payload moment during the flight was
restricted by adding more foam in the fuselage.
50
PAFA BURAQ
8.1.5. Tail Boom
The tail boom is subjected to air as well as impact loads. The calculated equivalent loads were
applied to the tail boom. The impact testing of the boom failed so the hard wood reinforcement was
added to the boom.
8.1.6. Radio and Control Test
All the control links were checked for integrity and the
radio programming was carried out. All the four flight
conditions were set and checked on the ground. The aileron
throws were adjusted by varying the clevis positions. Similar
adjustment was made to elevator and rudder. The battery
link and radio range was tested and found to be about 2 km.
The endurance of RX batteries was also tested and found to
be 45 min of flight time.
8.2. Flight Testing Performance
As per the time line the flight testing will start in the start of April, so the actual performance of the
aircraft couldn’t be documented in the design report. Eagle Tree telemetry system will be used to
check the flight parameters of the UAV. The flight testing will be carried out as per the tesing plan in
earlier section.
51
PAFA BURAQ
9. Conclusion
The design and development of BURAQ UAV was a very unique experience for the students
where lot was learned and is a project for the future. The future of Aerospace Industry does not
belong to manned aircraft, but flying robots of all sizes and ranges to fulfill the entire range of future
missions. FFD is a very positive initiative to encourage young students to focus on the design and
development of UAV. The way the UAVs are progressing, it is not far off that designers will be
producing UAV not only to suffice their defense industry needs, but to meet all the requirements of
Private Sector Organizations. Media soon would have their UAV and so will the courier services.
These UAV have the advantage of penetrating into dirty, dull and dangerous environments in the
current arena of terrorism. The fabrication of BURAQ is almost complete and the UAV has been
designed to meet all the competition requirements. This is just the first step, may be in the next step,
similar efforts can be extrapolated to include semi-autonomous and autonomous UAVs. Another
lesson learnt was lot of stuff which is studied in analytical equations and formulations, does not apply
to ground realities. Practical experiences learnt during the development of BURAQ will be very useful
for all such future endeavors.
The design work of BURAQ is fully complete and so is the manufacturing. The testing is currently
being performed on the ground as the report is being submitted. This would be followed by flight
testing and improvements needed if any. BURAQ team is very motivated and excited to have a good
showing at the contest and bring laurels to the PAF Academy, Risalpur.
52
PAFA BURAQ
9.1. References
Reports and Books
(a)
Aircraft Design; A Conceptual Approach by Raymer
(b)
Aircraft Performance and Design by John D Anderson
(c)
Design Report TuAFA “Cheveri” DBF 2010
(d)
Design Report CAE “Zafir-II” DBFC 2008
Software
(a)
Solid Edge V19
(b)
Gambit 2.4
(c)
Fluent 6.3
(d)
ANSYS 12
(e)
MS Excel
(f)
Design Foil V6
Websites
(a)
www.worldofkrauss.com
(b)
www.rcpak.com
(c)
www.castlecreations.com
(d)
www.hitecrcd.com
(e)
www.javaprop.com
53

FFD 2011 Best Design Report

Transcription

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