TURKISH AIR FORCE ACADEMY

Transcription

F
FFD 2011
F
D
1.0 E XE CUTIVE SUMMA RY .............................................................................................. 3
2.0 MANAGEME NT SUMMA RY ......................................................................................... 4
2.1 Organization Scheme ............................................................................................... 4
2.2 Scheduling............................................................................................................... 5
3.0 CONCEPTUAL DESIGN .............................................................................................. 5
3.1 Mission Requirements .............................................................................................. 5
3.2 Payloads ................................................................................................................. 7
3.3 Scoring A nalysis ...................................................................................................... 8
3.4 Competitive Design Requirements ...........................................................................13
3.5 Conceptual Design Selection Process ......................................................................14
4.0 PRELIMINA RY DES IGN .............................................................................................20
4.1 Critical Design Parameters ......................................................................................20
4.2 Mission Model .........................................................................................................23
4.3 Optimization Methodology........................................................................................23
4.4 Aerodynamic Optimization .......................................................................................24
4.5 Stability and Control ................................................................................................29
4.6 Structure.................................................................................................................31
4.7 Propulsion System ..................................................................................................32
4.8 Preliminary Lift and Drag Estimation .........................................................................34
4.9 Performance Parameters .........................................................................................38
5.0 DE TA IL DES IGN ........................................................................................................38
5.1 Dimensional Parameters..........................................................................................38
5.2 Structural characteristics and capabilities ..................................................................39
5.3. Systems and Sub-Systems Selection and Int egration ...............................................42
5.4 Flight Performance Parameters ................................................................................43
5.5 Weight and Balance ................................................................................................43
5.6 Rated Aircraft Cost ..................................................................................................44
6.0 Manufacturing Process ................................................................................................45
6.1 Manufacturing Figures of Merit .................................................................................46
6.2 Investigating the Manifacturing Tec hniques ...............................................................46
TURK ISH AIR FORCE ACADEMY
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6.3 Components Manufacturing Methods .......................................................................47
7.0 TES TING ...................................................................................................................49
7.1 Structural Testing ....................................................................................................50
7.2 Payload Mechanism Testing ....................................................................................50
7.3 Propulsion System Testing ......................................................................................50
7.4 Ground Testing .......................................................................................................50
7.5 Flight Testing ..........................................................................................................51
7.6 Testing Schedule ....................................................................................................52
8.0 Performance Results ...................................................................................................52
8.1 Key Subsystem Performance ...................................................................................52
8.2 Complete System Performance ................................................................................54
9.0 References .................................................................................................................56
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1.0 EXECUTIVE SUMMARY
This report documents the design, testing, and manufacturing efforts of Turkish Air Force
Academy Fesa Team to produce an unmanned aerial vehicle system that will be competitive in the
2011 Future Flight Design competition. Our primary objective is to design and control an aerial vehicle
to reach a fire area in the flight arena and leave fire extinguisher balls over the area covered with a red
canvas and return. The competition is placed into three categories: Category I, Category II and
Category III. Completely original radio-controlled aerial vehicles are designed and constructed for
Category I. Previously manufactured radio-controlled aerial vehicles are modified for Category II.
Original and autonomous air vehicles are designed and constructed for Category III.
Category I is the target of our team to compete in. All aspects of the design maximize the
scoring function. To win the competition, the team is aimed to get the maximum score in overall
missions. The overall team score is made up of both a written report and a total flight score. Total flight
score is determined by the system performance in three different missions: a two-lap ferry flight, a
one-lap flight loaded with a single ball to one of the fire areas and leave the ball over the red
canvassed area, and a one-lap flight loaded with two or three balls whose weights are 1 kg or 1.3 kg
to different fire areas and leave each ball respectively. Score analysis results showed us that if our
aircraft could carry three heavier balls, the team would have the maximum score. That’s why all the
aircraft design requirements were made carrying three 1.3-kg balls.
One of the essential parameters to success in the second and third missions was forming the
best possible mechanical system to leave the extinguisher balls and still providing stability of the
aircraft. Because of this fact, the team really focused on forming this system successfully. Also,
Mission 3 time recorded for our team was a vital parameter affecting the flight score and Mission 3
weight recorded for our team had no effect to the flight score. For this reason, Mission 3 flight time was
tried to be least, so Mission 3 cruise speed was tried to be high as much as possible. In order to
reduce the Mission 3 flight time (in order to speed up the Mission 3 cruise velocity), some parameters
such as bigger motor and larger control surfaces than how big they were required for Mission 1 and 2
were optimized. But we were aware of that Mission 1 and 2 must be successfully completed to get
score from the Mission 3. The team was also goaled to achieve Mission 1 and 2. Thus, ground
handling during taking-off and landing, stall characteristics, landing speeds, flying with asymmetric
load and stability were additionally paid attention to.
In order to improve the aircraft‘s performance in all the missions, great effort was made to
minimize the turning radius which affects the flight time. Analysis showed that turning flight comprises
a significant portion of the mission course and the flying pattern has eight 180 degree turnings whose
radius may change variously. The aircraft structure was designed to take advantage for turning time
and is capable of withstanding high g-loads for rapid turns. The structure of the vehicle was also
designed to minimize the weight. This was accomplished not only through the use of lightweight
materials, but also by designing existing structural elements to perform multiple roles.
Figure of Merits (FOM’s) are utilized to select the aircraft configuration and components. So a
mono plane configuration with a single tractor propeller, a single middle-mounted wing, tricycle landing
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gear and conventional tail were selected by the directions of FOM’s results. We have used a
fibercarbon build for a lightweight and robust aerial vehicle. The aircraft’s storage box was constructed
from a fiberglass-over-foam support frame and covered with a thin nylon. With a 2.17 m wingspan, a
2
total wing planform area of 0.66 m , the plane flew with a Hacker C50-14 L Acro 6,7:1 competition
motor. Fifteen GP 2200 LiPo batteries and a 20x12 propeller were used for all missions.
The predicted performance capabilities of Missions 1 / 2 / 3 are as follows: takeoff distances
are 12 / 18 / 30 m, thrust-to-weight ratios are 0.75 / 0.62 / 0.46 kg/kg, wing loadings are 9.24 / 11.2 /
2
15 kg/ m , stall speeds are 10 / 11 / 13 m/s, and cruise speeds are 19.4 / 21.5 / 25 m/s. These
predicted performance and scoring parameters couple to result in a highly competitive mission oriented vehicle.
2.0 MANAGEMENT SUMMARY
We formed our team of eight undergraduate students. During the conceptual design phase,
we worked together and decided the basics. In order to complete the design details and for the
production we organized the groups as shown in figure 2.1 below.
2.1 Organization Scheme
We formed our groups according to a classical scheme. A group leader who is responsible for
general subjects, a coordinator who is responsible for the coordination and communication of four
technical groups: aerodynamics, propulsion, structures and CAD designs. The objective of the
aerodynamics group is to optimize the aircraft configuration for maximum mission performance,
ensure stability and control, and predict flight testing performance. The propulsion group is responsible
for the optimization of the propulsion system, the analysis of power requirements for each mission,
and the selection and testing of the propulsion system components. The structure group is responsible
for the detailed design and construction of the integrated aircraft system. This includes the aircraft, box
and payload deployment mechanisms. The CAD designer is responsible for compiling system
component designs into the master CAD drawing for manufacturing and publication.
Team
organization chart is shown in figure 2.1.
Tolga PATNOS
Team Leader
Hüseyin TORUN
Eyüp ACAR
Nail SARIALTIN
Ertuğrul C. SUNGUR
Uğur ZİYAL
Coordinator
Aerodynamics Lead
Propulsion Lead
Structures Lead
CAD Designer
Figure 2.1: Team organization chart
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2.2 Scheduling
In order to succeed in the competition, first thing to do is a good planning. We planned our
program according to our school’s yearly plan. In our schedule we had 21 clear weeks to prepare our
aircraft for competition. We tried to organize our spare times and use the best of it. So we managed to
complete the schedule shown in table 2.1.
Table 2.1: Milestone schedule
3.0 CONCEPTUAL DESIGN
Conceptual design begins with spotting the mission requirements and design limitations.
Parameters, which affect stability, fertility, and performance more, must be found by analyzing the
score. To be able to reach the maximum score in the competition, some unimportant parameters can
be ignored instead of crucial parameters which affect score more. So the best design parameters
which respond to mission requirements have been selected.
Figure 3.1: Conceptual design methodology
3.1 Mission Requirements
There are 3 flight missions and total flight score will be obtained by adding the scores that are
acquired in each mission. Total competition score will be formed by the total flight score and written
report point. Missions have to be completed successfully and the aircraft has to be undamaged after
landing in order to score and pass another mission. The general constraints and flight pattern are
expressed below;

40 A maximum current is allowed for the battery package.

The battery type may be NiMh, NiCad or LiPo.
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
Maximum weight of battery package may be 3 kg.

All parts and hardware of aircraft have to be fitted into a box has 80x80x150 cm dimensions.

The assembly of aircraft and checkout must be completed in less than 5 minutes.

It is allowed a maximum of 5 flight attempts or 4 successful scoring flights whichever comes
first.
Figure 3.2: Flight pattern
Total flight score and total competition score are shown in equations 3.1 and 3.2.
(3.1)
(3.2)
Mission 1: 2 Lap Ferry Flight
It will be 2 laps ferry flight without payload. Flight time starts with the first movement of aerial
vehicle and ends with the conclusion of the second lap. Although landing duration is not added to flight
time, it is required to be able to land succes sfully to score and to deserve to compete in the next
mission. When the route of flight is scrutinized, distance may be assumed approximately 2600 m.
The scoring of Mission 1 is shown in equation 3.3.
(3.3)
Mission 2: One Lap Payload Flight
One lap payload flight which has to be loaded with a single fire ball (weight depends on the
teams’ preferences) and it has to be left 2
nd
flight segment. Mission 2 score is not dependent on flight
time but loading time of fireball to the aircraft has a vital role in scoring. Loading time is not only
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dependent on the easiness of loading mechanism of aircraft but also dependent on the team crew’s
physical performance. After loading there will be no interaction to change flight conditions of aircraft.
To be able to score and pass the next mission a successful landing is compulsory. Distance may be
assumed approximately 1300 m. The scoring of Mission 2 is shown in equation 3.4.
(3.4)
Mission 3: A single lap payload flight
It is formed of a single lap payload flight. Aircraft may be loaded 2 or 3 balls in different
weights depending on the teams’ preferences. If the team selects 2-Ball mission, the balls should be
released on flight segment 1 and 3, respectively. If the team selects 3-Ball mission, the balls should be
released on flight segment 1, 2, and 3, respectively. Scoring is not effected by loading time of balls but
it should be completed in 5 minutes. The flight time starts with advancing the throttle and ends with the
completion of the lap in the air. Successful landing is compulsory to be able to score. Lap distance
may be assumed approximately 1300 m. Mission 3 scoring formulas are seen in equations 3.5 and
3.6.
(3.5)
(3.6)
3.2 Payloads
The sizes and weights of payloads are optionally selectable. Although number of fire balls may
rd
change in 3 mission, there are two standard types of balls have to be used. 1-kg fire ball is limited by
50 mm radius or 1.3-kg fireball is limited by 75 mm radius. It is not allowed to use different types of ball
at Mission 3. Types of fire balls and their dimensions are seen in figure 3.3.
Figure 3.3: Fire ball dimensions
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3.3 Scoring Analysis
Firstly it must be analyzed that how our options affect maximum scores by comparing our
results with some optional selections. So the number and type of the balls should be determined. For
this reason, relation between 3
rd
mission and the other missions, and also the importance percentage
rd
of each mission must be analyzed. Related to the options, maximum scores of 3 mission are shown
in table 3.1.
OPTIONS
WEIGHT (kg) BALL NUMBER
MAXIMUM M3 SCORE
1
2
150
1,3
2
292,5
1
3
225
1,3
3
438,75
Table 3.1: Options and scores
When the importance percentage among the missions analyzed, we have reached the results
which are shown in table 3.2.
Mission
Scoring
Maximum Score Percentage of Influence
Report
100×0,5
50
7,83
Mission 1
50×1
50
7,83
Mission 2
100×1
100
15,66
Mission 3
438,75×1
438,75
68,68
Table 3.2: Influence of missions to the score
It is seen in the table that Mission 3 has a vital influence percentage in competition. As a
natural result of this case, our aircraft is aimed to get the highest score and it may be succeeded by
carrying three 1.3-kg fire balls at Mission 3. It may be thought that carrying the least weight (2 kg) at
Mission 3 ensures to take the weight and time advantage in Mission 1 and Mission 2. In this case, it is
seen from equation 3.3 and 3.4 that maximum 150 point score advantage may be provided at first two
missions. But also it has to be thought that 288.75 point will be lost in Mission 3 according to the
heaviest (3.9 kg) loaded aircraft. Therefore we have preferred to carry 3 balls each one weighs 1.3 kg
and according to this choice we have shaped our design parameters. But we must have kept eye on
one last subject. It will not be allowed to compete in the next mission if the previous mission could not
be succeeded. The influence percentage of Mission 3 was calculated without paying attention to this
case. If we are aimed to take maximum score, it is not an extraordinary case assuming to perform the
previous missions without failure.
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How can the aerial vehicles which are designed to carry lighter balls prevent the disadvantage
rd
of 3
mission regarding provided weight and time advantages of 1
st
nd
and 2
missions? In order to
answer this question, it should be estimated that our design’s weight and time results will have the
lowest scores and should be compared with the other teams’ possible best scores.
3.3.1 Weight Analysis
Total system weight has no vital role in 3
rd
mission. For this reason, only the balls’ weights are
st
analyzed to be able to reach a result. Take-off weights in 1
results are shown in table 3.3. In 3
rd
nd
and 2
missions are compared and the
mission, there are four different possibilities for payload which
depends on the teams’ preferences. Our aircraft is designed in order to carry three balls that each one
weighs 1.3 kg. According to the historical experiences, it may be assumed the lightest aerial vehicle
may be obtained empty weight to payload ratio 1/1. Therefore, empty weight to payload ratio may be
taken 3/2 for the worst possibility. Thus, our empty weight was estimated 6 kg.
PAYLOAD
NUMBER WEIGHT(kg)
WPL(kg)
WE(kg)
W PL/W E
MISSION 1 MISSION 2
WT/O(kg)
W T/O(kg)
2
1
2
2
1
2
3
2
1,3
2,6
2,6
1
2,6
3,9
3
1
3
3
1
3
4
3
1,3
3,9
6
3/2
6
7,3
Table 3.3: Take-off weight estimations
The influence of weight ratio on scoring of first two missions is sown in figure 3.4. It is seen
that take of weight of 2
nd
st
mission is more important on scoring than 1 mission.
80
S
C
70
60
50
O
40
Mission 1
R
30
Mission 2
E
20
10
0
0,30
0,35
0,40
0,45
0,50
0,55
0,60
0,65
0,70
0,75
W ref / W team
Figure 3.4: Weight analysis of Mission 1 and 2
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3.3.2 Time Analysis
If flight pattern is examined, range is obtained 2600 m. for Mission 1 and 1300 m. for Mission
3. An average speed of 20 m/s, flight pattern is completed in 130 sec. at Mission 1 and 85 sec. at
Mission 3. If Landing, take - off and accelerations are considered, 20 sec. extra time may be added to
the missions. For the worst possibility, if it is obtained that our aerial vehicle achieves 20 sec. later at
Mission 1 and Mission 3, the results are seen in table 3.4. It may be considered to have best loading
time at Mission 2. Although loading mechanism easiness may be same with the other teams’, we can
assume that our team’s physical performance will be enough to provide the reference time. Therefore,
the loading time was not compared or analyzed for Mission 2.
Mission 1
T1 (sec)
Mission 3
Possibilities
T3 (sec)
150
2×1
85
150
2×1,3
85
150
3×1
85
170
3×1,3
105
Table 3.4: Estimated flight times for mission 1 and 3
The lightest (2 kg) aircraft’s weight was chosen for W ref in table 3.3. Our aircraft’s weight (6 kg)
is also shown in table 3.3. Weight ratio was taken constant at Mission 1 while analysing the influence
of time on score which is shown in figure 3.5.
Mission 1
18
S
C
O
R
E
16
14
12
10
8
6
4
2
0
150
155
160
165
170
175
180
185
190
195
TIME
Figure 3.5: Time-dependent score analysing of Mission 1
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The influence of time on score is analyzed in figure 3.6 for different payload options. It is seen
that if 3×1.3 kg payload option is selected, flight time changes affect the score more than the other
options. Tref may be assumed 85 sec.
Tref = 85 sec.
500
450
S
C
O
R
E
400
2x1,3
350
2x1
300
250
3x1
200
3x1,3
150
100
50
0
85
90
95
TIME
100 105 110 115 120 125 130
Figure 3.6: Mission 3 score analyzing for different options.
3.3.3 Preferred Option Score Analyzing
SELECTION
SCORE
Number × Weight
M1
M2
M3
∑M
2 × 1 (kg)
50
100
150
300
Wteam
14,7
41
355,17
410,87
3 × 1 (kg)
50
100
225
375
Wteam
22,05
57,14
355,17
434,36
2 × 1,3 (kg)
50
100
292,5
442,5
Wteam
19,11
53,42
355,17
427,7
Table 3.5: Available scores for selected options
It can be easily realized that from table 3.5, in the 3
rd
mission, the team who prefers to carry
two 1-kg balls may take less points than our team, even if both of teams would successfully complete
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all the missions. In this case, it is figured out that according to our aerial vehicle it will have less
rd
chance to win the first place in the competition. Likewise, in the 3
mission, the aircraft which was
designed to carry three 1-kg balls will be scored less than our team, accomplishing all the missions
even if we would encounter cases like pilotage control disabilities or longer loading time in 2
nd
mission.
Therefore, it is estimated that this aircraft has also little chance to win the first place.
But it is predicted that the aircraft carries two 1.3-kg balls may get higher score than our team.
Even though it seems to be risky, in 3
rd
mission, the case of lowering our 20 seconds later flight time to
15 seconds makes it feasible to get higher score. Because the safety factor that was selected for 1
st
mission and worst possibility weight estimations of other missions help us to take this risk.
T
kı
Tref
85
85
90
95
100
105
110
115
120
125
130
439
414,4
393
372,9
355
339
324
311
298
287
438,8
416
394,9
376
359
343
329
316
304
439
416,8
397
379
362
347
333
321
438,8
418
399
382
366
351
338
439
419
401
384
369
354
439
420
402
386
371
439
420
404
388
90
95
100
105
110
115
Table 3.6: Mission 3 time-dependent score analysing
Red coloured area which is seen in table 3.6 shows the scores that may be taken in the
competition if we would finish Mission 3 with the best flight time. Orange coloured area shows the
scores that we aim to reach while the mission would be completed the best flight time by another
aerial vehicle which prefers to carry same weight. In this case, it will be required to score more at
previous missions to be able to win the first place. Pink coloured area shows our objective minimum
score values that if the best flight time scored aircraft carries less ball and weight according to us.
In consequence of analysing Mission 1 and 2, available score changes with both time and
weight parameters are seen in figure 3.7.
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Wref / Wteam
T ref / T team
Figure 3.7: Score analysing for Mission 1 and 2
Consequently, it can be said that to be able to win the first place in this competition, three 1,3kg balls option preferred team which has a same preference with us will be an obstacle for us to get
over. The score difference which will be formed between us will be dependent on the flight time factor
at last mission and weight and flight time factors at previous two missions. Therefore, our aerial
vehicle should be optimized many times to be able to design and manufacture the lightest and fastest
aircraft. Surely, Mission 3 flight time which affects score more than the other missions’ flight time and
weight values should be minimized. In conclusion, instead of choosing lighter battery or motor types to
get higher scores at first two missions, it should be designed the best level controllable and fastest
aircraft by giving up the idea of ensuring high efficiency.
3.4 Competitive Design Requirements
After evaluating all the competition rules and our score analysis for the missions, we
determined the competitive design requirements as follows:
a. RAC: RAC is the total system weight including the aircraft, assembly tools and storage box. Since
RAC has a vital role at first two mission, it should be designed the lightest aircraft and storage box in
order to get the highest score.
b. Flight Time: Flight time should be minimized in order to get the highest scores at Mission 1 and 3.
It is aimed to obtain 170 sec. flight time for Mission 1 and 85 sec. for Mission 3.
c. Loading Time: At Mission 2, score is affected from loading time directly. Our team aimed to reach
the best loading time and it can be assumed to limit between 8-10 seconds.
d. Types and Number of Balls: After creating a mission score analysis, the types of balls to carry in
Mission 3 should be three 1.3-kg balls in order to get the best score.
e. Efficiency: All configuration or components (motor, battery etc.) have to be used efficiently in order
to maximize the score.
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3.5 Conceptual Design Selection Process
3.5.1 Figures of Merit
RAC: Overall weight of the aircraft has to be considered at Mission 1 and 2 in consequence of
affecting mission score directly. Therefore, reducing system weight is very crucial to be able to get the
highest score at first two missions.
Loading and Releasing Convenience: Aircraft body configuration selection has to minimize loading
nd
time in 2
mission to reach the best score. At the same time, releasing system should not be blocked
by preferred configuration.
Drag: It is the effect each configuration has on the total drag of the aircraft. Drag is taken into
consideration because it affects aircraft performance and battery weight. Because, more drag
generated means more energy is required. That means we need a heavier battery, which we do not
appreciate.
Stability & Control: It is very important that the aircraft is stable and controllable both in the air and
on the runway so that the pilot can easily control and navigate the plane during take-off, cruise and
landing.
Ground-Handling: This feature is needed for an effective take-off and landing.Because, successful
take-off and landing is compulsary to receive score and to deserve performing the next mission.
Take-off Distance: It should be paid attention to determine aircraft configuration because of 40 m.
limitation of takeoff distance for all missions.
Manufacturing: In some cases, various parameters may be changed in the name of convenience in
production. Furthermore, we should consider the skills, materials, and time needed to properly
construct the design.
3.5.2 W empty and W 0 Estimation
Take-off weight parameter is one of the important parameters required for the start of a
design. Therefore, the maximum weight of the aircraft should be estimated and the required
parameters should be considered based on it. This estimated weight may not be matched the outcome
of the manufacturing. But, it will not create a problem because safety margin and the heaviest mission
are taken into consideration during design process. In order to make a decision about payload to
empty ratio, W pay load/W empty ratios of successfull aircraft of AIAA DBF contest are analyzed and reached
an average value 1/1 for W pay load/W empty ratio. We have decided to carry 3,9 kg payload at score
analysing section so it can be said that our empty weight will be approximately 3,9 kg. But, we have
also included the safety margin and considered 10 kg maximum weight for the worst possibility.
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3.5.3 W/S and S Estimation
We have known that the higher score is dependent on the shorter flight time (higher velocity)
at Mission 3 and at the same time Mission 3 influences score the most. For this reason, minimizing
wing area will reduce drag, increase cruise velocity and also provide advantage for RAC. There are
two important points have to be paid attention. Since lift coefficients are limited for an airfoil, wing area
is the only parameter left in order to reduce stall speed. Therefore, it should not be decreased more
than required to provide controllable landing speed (1,2 V stall). Controllable landing speed was
determined 13 m/s with the experiences from AIAA DBF contests. In this case, it was needed the
2
necessity of larger wing area than 0,6 m during landing (W e=6kg, CLmax=1,5). Another important
point is W/S parameter that directly affects wing structural strength. This parameter may change
variously depending on the construction materials of the wing. Therefore, after analyzing similar
3
designs made from balsa (our construction material), wing loading is agreed 15 kg/m . Furthermore,
2
wing area was emerged 0,66 m in consequence of 10 kg maximum take-off weight estimation. We
2
have agreed this wing area which was larger than 0,6 m provides stall and landing speed.
3.5.4 Payload Configuration
Internal or External
As mentioned before, it was decided to carry three 1,3-kg balls at Mission 3 to be able to
reach best score. For this reason, body should be considered for 3 balls carrying mission while taking
into consideration loading time at Mission 2. It is required t o make a decision between external and
internal configuration. Even though it can be considered external configuration allows smaller body, it
is required a stronger and larger body that wings, landing gear and loading-releasing mechanisms can
be mounted. Also, external payloads and mechanisms will be more complex and create added weight
and drag forces. Furthermore, external payload configuration will have disadvantages about loading
time according to the internal type. For this reason, internal payload c onfiguration was preferred to not
cause too much weight drawbacks , to minimize drag forces and to provide shorter loading time. There
are several internal payload configurations shown in figure 3.8.
Figure of Merit
Loading & Releasing
Capability
Stability & Control
Weight
3 Vertical Row
1 Front 2 Back
2 Front 1 Back
3 Horizontal Row
1 Up 2 Down
2 Up 1 Down
3
-1
0
0
1
-1
-1
3
1
0
0
-1
0
0
Drag
2
-1
0
-1
1
-1
-1
Manufacturing
1
1
0
0
1
0
0
TOTAL
9
-1
0
-2
3
-5
-5
Figure 3.8: Payload configuration decision matrix
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Even though our preference (3 horizontal raw configuration) will be affected by asymmetric
flight releasing balls respectively, we have considered resolving c.g. shift with adequate horizontal tail
control surfaces.
3.5.5 Aircraft Configuration
Flying wing and blended types are eliminated in consequence of deciding to carry payloads
inside the body of aircraft. Configuration selection as a FOM, loading time, RAC and S&C capability
should be considered. Additionally, ease of construction should be included among these criteria.
Bi-plane: It generates less lift and more drag according to conventional type at a same reference wing
area. Considering the 2
nd
mission, loading ball from the top is not suitable for this type. Furthermore,
the additional wing and carry-through structure causes an increase in RAC.
Canard: Although, it is a quite convenient option in terms of stall characteristics, there is a lack of
stability and control. In addition, it is more demanding in terms of production according to the
conventional type.
Conventional: This type provides for the best RAC while still attaining good lift and drag
characteristics. Also, it is more suitable in terms of stability and control while allowing loading balls
from the top at Mission 2. As a result, conventional type which best meet mission requirements was
chosen for the final design.
Figure of Merit
Weight
Stability & Control
2
RAC
2
Loading & Releasing
3
capability
Manufacturing
1
TOTAL
8
Delta
1
1
Conventional
1
1
Canard
0
1
-1
1
1
0
1
1
7
-1
4
Figure 3.9: Aircraft configuration comparison
3.5.6 Tail Configuration
Stability and control capability is the most important factor for tail configuration selection. RAC
and drag should be minimized while ensuring adequate stability and control. Convenience is also
taken into consideration in terms of manufacturing. Accordingly, various tail configurations were
investigated and the decision matrix shown in Figure 3.10 below.
Page 16 of 56
Figure of Merit
Stability & Control
RAC
Manufacturing
Drag
TOTAL
Weight
V-tail
Conventional
T-tail
Cruciform
3
0
1
1
1
2
1
0
-1
-1
2
0
1
-1
-1
1
1
-1
1
0
8
3
4
0
-1
Figure 3.10: Tail configuration decision matrix
T-tail: The horizontal stabilizer acts as an endplate for the vertical stabilizer reducing drag; however,
the added weight that the vertical stabilizer needs to support the aerodynamic load from the horizontal
stabilizer negates the tail as the best option.
V-tail: This type has two stabilization surfaces that work together to provide both elevator and rudder
responses, which theoretically reduces interference drag and weight of the tail; however for adequate
stability and control, the stabilization surfaces must be enlarged therefore counteracting any drag and
weight benefits. Also it cannot be controlled separately for vertical and horizontal tail and this makes it
more complicated to use.
Cruciform: It is a combination of T-tail and conventional tail. It is heavier than conventional tail and
generates more drag than T-tail.
Conventional: It is the most commonly used type which provides adequate stability and control while
minimizing weight. Also it has a simple manufacturing process. These provided conveniences enabled
us to choose the conventional tail.
3.5.7 Landing Gear Configuration
The most important features of landing gear are ground handling capability and take-off
distance efficiency. Also RAC should be taken into consideration to be able to minimize. In addition,
there is an important point to decide about the placement of the landing gear. The take-off angle is
limited if the landing gear is placed too far forward, but if the landing gear is placed too far aft; it is
difficult to create a big enough moment to take-off.
Tail-Dragger: In consequence of airflow, pitch up moment increases while moving for take-off in the
runway. So, it has poor ground handling abilitiy. This type may be considered lighter (shorter strut and
smaller wheel) than the other types.
Bicycle: Rear-wheel is far aft from the c.g. so that it increases the take-off distance. It can be said that
the increase in the number of wheels will increase weight.
Page 17 of 56
Single Main: This configuration has one main landing gear on the nose and smaller wheels under the
wing tips and empennage. It may be considered lighter because of having only one main landing
gear. Although it provides shorter take-off distance, it causes ground control problems and suffers
from weight problems.
Tricycle: Even though landing gear configuration heavier, it provides required level ground-handling
capability which played an influential role in our choice.
Figure of Merit
Ground-handling
Take-off distance
RAC
TOTAL
Weight
Tricycle
Bicycle
Tail-Dragger
Single Main
3
1
0
-1
-1
1
0
-1
1
1
2
-1
0
1
1
6
1
-1
0
0
Figure 3.11: Landing gear decision matrix
3.5.8 Propulsion System and Location
Both the number and placement location of the engine has to be decided. Although increasing
the number of motors allows us the election of smaller motors which will cause weight increase .
Tractor type aircraft body is affected by the propeller based airflow disturbances and accordingly,
efficiency decreases. At this time, pusher type aircraft propeller is affected by the body based airflow
disturbances and efficiency also decreases. Furthermore, pusher type motor requires higher aft
landing gear strut. This will cause weight increase which we do not appreciate. Doubling motors will
create mounting problem in manufacturing. Also motors mounted to the wings will cause weight
increase due to the consolidation of the junction points.
Double
Tractor
Double
Pusher
Tractor &
Pusher
0
-1
-1
-1
1
0
1
0
1
1
0
0
1
1
1
-1
1
-1
0
2
1
1
-1
-1
-1
8
6
3
-4
-5
-4
Figure of Merit
Weight
Single Tractor Single Pusher
RAC
Drag
Stability & Control
Efficiency
Manifacturing
TOTAL
3
1
1
-1
1
Figure 3.12: Propulsion system and location decision matrix
Page 18 of 56
3.5.8 Wing Vertical Placement
Low-wing: It is required longer take-off distance because of ground effect airflow disturbances. Due to
the risk of hitting ground, longer landing gear necessity should be taken into consideration.
High-wing: Provides shorter take-off distance in consequence of ground effect reduction. Weight gain
is ensured as it allows for the use of a shorter landing gear. Despite all these conveniences, options
are limited in terms of payload loading.
Mid-wing: Generates the least drag forces among these configurations. Contribution to stability may
be considered neutral. Beside these features It allows the best payload loading time and releasing
options which helped us to choose this type.
Figure of Merit
Loading & Releasing
capability
Stability & Control
RAC
Take-off Distance
Drag
TOTAL
Weight
Low
Mid
High
5
-1
1
-1
4
3
2
1
15
-1
0
-1
0
-11
0
-1
0
1
3
1
0
1
0
1
Figure 3.13: Wing vertical placement comparison
FINAL CONCEPTUAL DESIGN-SIDE VIEWS
Page 19 of 56
FINAL CONCEPTUAL DESIGN-MAIN VIEW
4.0 PRELIMINARY DESIGN
After completing conceptual design and determining the aircraft configuration, it is passed to
preliminary design section for the size and shape optimization.
4.1 Critical Design Parameters
As mentioned previously, during the conceptual design process some of the requirements
were renounced in order to accomplish more important ones. (During the optimization process, we
have paid less attention to the less effective parameters on scoring).For example, since Misssion 3
flight time (T3) has the most effective role on scoring between other missions parameters, we have
taken into consideration to get lower scores at first two missions in order to find out how to minimize
T3. It can be understood better from figure 4.1 represents this process.
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Figure 4.1: Optimization Process
Page 21 of 56
4.1.1 Aerodynamic Design Parameters
Wing Loading and Wing Area: Wing loading is a critical parameter that affects the size of the wing.
Even though higher wing loading provides smaller wing areas it is required more structural durability.
In this case, it is emerged the importance of wing construction materials. High wing loading is
preferable if the wings are made from composite. But, the wings made from balsa may not provide
sufficient strength if high wing loading is selected. Also it should be considered that higher wing area
provides more lift and shorter take-off distance while it generates more drag and requires more
energy.
Wing Span: Wing span has to be limited in order to provide enough convenience to fit the box. It
should be selected higher wing span while taking into consideration box size. Because, wing span is a
parameter affects the lift coefficient and induced drag.
Airfoil: It is one of the important parameters that directly affect lift, drag and aircraft weight. The airfoil
should be determined to meet required performance in terms of all flight conditions.
Fuselage Length and Empennage Size: While determining the dimensions of the fuselage and
landing gears it should be targeted that criteria which had been determined before must be done with
minimum weight. Because of the decision carrying payload in the fuselage, obligatory of having
dimensions minimum 45cm for fuselage can be given as an example for c riteria which had been
determined before. Also the necessity of tail performance for minimum maneuver radius or minimum
take-off distance can be given as example for criteria which has to been provided by aircraft.
4.1.2 Propulsion Design Parameters
Motor Selection: Power, weight and efficiency are the crucial parameters which must be analysed
carefully for motor selection. Even if brushless is more expensive and comlex when we compare with
brush motor, brushless motor has been preferred due to its long-lived, high efficiency and high torque.
Also one of the parameters which affect score is RAC. So thrust/weight ratio of motor which has been
selected should be maximized.
Battery Selection: LiPo batteries are lighter than NiMH and Nicd batteries. LiPo batteries can also
supply high current capacity more. They are long-lived and they can be manifactured in every shape
and size. It is the only disadvantage that they require special attention in charging and usage. So LiPo
batteries have been selected. But for battery selection, flight time or instantaneous battery current
drawn should be thought. So an optimization should be done for battery capacity, resistance and
weight.
Speed Controller Selection: For ESC selection, the most important criterion is the value of peak
current but also the efficiency and capability of being programmed is an important criterion. It should
be well-matched with engine and battery. If it is possible, even trademarks should be same.
Propeller Selection: Static thrust, efficiency and cruise velocity are the specialities which affect
propeller selection. It has to have enough static thrust which can achieve take-off in the limits and pitch
angle which can give the opportunity to increase the cruise velocity as much as possible.Having a
Page 22 of 56
longer diametered propeller than we need, causes drawing more current and exceeding 40 A limit. But
also having a shorter diametered propeller than we need, causes not ensuring enough static thrust
and exceeding 100 feet take-off distance limit.
4.2 Mission Model
1 Take-off: The take-off distance should not exceed 40 metres in mission-3.
2 Climb: Full throttle
3, 5, 7, 9 Turn: Full throttle to prevent the loss of altitude and velocity for the Mission 1 and Mission 3.
4, 6, 8 Cruise and payload releasing: Weight will be decreased by payload releasing.In order to
provide constant altitude it should be preferred trimming instead of reducing velocity and protected
maximum velocity. For the Mission 2, velocity may be reduced because the flight time is not important.
10 Because of the flight time will be stopped while the aircraft passing the finish line in the air,
maximum speed should be kept but after that successful landing should be achieved.
Figure 4.2: Mission model
4.3 Optimization Methodology
Aerodynamics: In this part, airfoil have been selected and the lift and drag coefficients have been
calculated. For 2-D airfoil calculations Profili and Design Foil programmes have been used. After that,
for the 3-D airfoil values, 2-D results are transformed by common analytical methods.
Stability and Control: In this part, dynamic and static stability have been calculated. The pitching
moment that will be emerged after releasing the balls, have been calculated and then control surfaces
are sized while taking into consideration the required trim.
Propulsion: Aircraft is planned to fly as fast as it could, because of the importance of the cruise
velocity at Mission 3.So we have selected a powerful motor which can overcome the drag that
emerged by the high velocity and the trims. Then, propeller selection and the bataries optimized
according to these features.
Page 23 of 56
Structure: Since it was decided to carry payload inside the aircraft we have preferred to provide
adequate structural strength to fuselage. Also, wings were made from balsa in order to ensure weight
gain. It was mounted a boom between body and empennage to provide required moment.
Performance: The performance of the aircraft have been calculated simultaneously with the other
components. Because since we could not meet the requirements (40m take-off distance, system
weigth, cruise speed, required energy, turn radius) that we tried to reach, some changes have been
made to achieve them.
4.4 Aerodynamic Optimization
4.4.1 Airfoil Optimization
Airfoil is chosen accordingly to some parameters such as aircraft’s cruise speed, stall
characteristic, take-off and landing distances. This is why it’s an important criteria.
Cambered or Uncambered: Camber is directly effective over the lift coefficient of the airfoil.
Cambered airfoils have higher maximum lift coefficients than uncambered ones. Maximum lift
coefficient affects stall and landing speeds. In order to perform a successful landing, maximum lift
coefficient value should be as high as possible. Because of the fact that we don’t plan to use flaps, we
decided to use a cambered airfoil in order to get a high lift coefficient value.
Thickness: Thickness is important for the stall characteristic, weight and drag value of the wing.
Statistical equations for wing weight show that the wing structural weight varies approximately
inversely with the square root of the thickness ratio. Therefore the thinner airfoil we choose the lighter
wing emerges and provide advantage about RAC. But this advantage won’t be so much because
these aircraft designs are smale-scale. Also higher thickness ratio increases the drag of the wing. But
the most suitable airfoil in terms of stall characteristics will be the airfoil has higher (14% and more)
thickness ratio. Because stall starts from the trailing edge and gradually proceed through the leading
edge and it does not occur any sudden changes. If we consider that a successful landing is important
for having the top score, we can see how stall characteristic is important. This is why we decided to
choose a profile with a thickness ratio of around 14%.
During the optimization of the airfoil, in generally maximum L/D is considered. If L/D ratio is
maximized, it provides also an increase in the efficiency. But we want to maximize t he performance
criteria. For example maximum endurance time is not important on scoring at all missions. So we need
to focus on maximizing the speed. Therefore we should find the optimum cruise lift coefficient that will
provide to not decrease our aircraft’s speed to protect our altitude. But also it should be paid attention
to increase CLmax in order to reduce landing and stall speed.
Page 24 of 56
Figure 4.3: Aerodynamics coefficients comparison for different airfoils
We can see in the figure 4.3 that MH114 has the best L/D ratio. With a +3° incidence angle,
this airfoil can be used efficiently during the cruise flight. But we want to make our aircraft faster. So in
the figure 4.3, optimum airfoil should be preferred best meet our requirements (cruise lift coefficient) to
protect level flight without trimming.Required level flight lift coefficient is calculated in table 4.1.
Table 4.1: Level flight lift coefficient calculation
Page 25 of 56
For the last mission, level flight lift coefficient value is 0.39 according to our calculations. We
should look for the airfoils that meet this value and at the same time provide higher C Lmax. But it
should be paid attention to one more essential point. This level flight lift coefficient value (0.39) is
emerged in consequence of 2-D airfoil calculations. In order to calculate aircraft lift coefficient value it
shoul be considered the lift contribution of wing and fuselage. 2-D airfoil results were transformed to 3D aircraft calculations using Prandtl equations and it is learned that in order to provide 0.39 CL value
for our aircraft, 0.5 Cl value is required for a 2-D airfoil.
As it mentioned before we should keep the lift coefficient as high as possible. Eppler 420,
Eppler 422 and MH114 airfoils can meet 0,5 Cl value. But in order to provide 0.5 Cl value, Eppler 420
must have a -5° incidence angle. Because of the fact that this airfoil is highly cambered, this angle
would make our drag increase as we can see in the figure 4.3. For this reason we discarded Eppler
420. MH114 has a thickness of 13.04% which is less than Eppler 420 (13.99%). It makes our wing
lighter. But Eppler 422 has a better stall characteristic. And its maximum lift coefficient value is slightly
higher than MH114. For this reasons we have chosen Eppler 422.
CFD Analysis of the Eppler 422 Airfoil
After we got the aerodynamic coefficients of the Eppler 422 by the profili and designfoil
programmes, we decided to make CFD analysis by fluent programme to check the results we had
before. So, to specify the airfoil geometry which was shown in figure 4.4, we imported a file containing
a list of vertices along the surface and had Gambit join these vertices to create two edges,
corresponding to the upper and lower surfaces of the airfoil.
Figure 4.4: Eppler 422 airfoil
After we had the airfoil, we defined a farfield boundary and meshed the region between the
airfoil geometry and the farfield boundary. The farther we were from the airfoil, the less effect it had on
the flow and so more accurate was the farfield boundary condition. Since we used the ambient
conditions to define the boundary conditions at the farfield, we placed the farfield boundary well away
from the airfoil and meshed them as it was shown in figure 4.5 and figure 4.6.
Page 26 of 56
Figure 4.5: Meshed region
Figure 4.6: Farfield boundary condition
We had 13680 cells and 13969 nodes by meshing the geometry. After that, we exported the
geometry for the Fluent programme and set up problem in Fluent. We defined materials, models and
boundary conditions to solve. After results, we compared the values got by profili and designfoil
programmes. Finally, because of the similarity between the results, we decided that we could use the
values got before.
We could not get the pressure coefficient of the airfoil at cruise flight conditions. We got it by
fluent and plot the graphic of pressure coefficient depending on the airfoil’s chord in figure 4.7.
Figure 4.7: Cp vs. chord position
Since we have plotted the graphic for the -2 deg. angle of attack condition, it seems that the
upper surface of the airfoil has more pressure than the lower surface by the 20 percent of the chord.
And also it can be said the stagnation point is in the somewhere on the upper surface.
Page 27 of 56
4.4.2 Wing Geometry Optimization
Wingspan and AR: If we consider that we can part our wings into two parts and our box sizes are
limited we can have a maximum wingspan of 300cm. W e calculated our wing area as 0.66 m². In order
to decide a wingspan it is better to use AR value. As the AR value increases, aircraft weight, lift
coefficient and induced drag values also change. If we increase the wingspan we should make wingfuselage connections and spars stronger. And this increases our total system weight. But as we
increase the AR we reduce the induced drag. In the graphics below, AIAA DBF competition’s previous
attendants’ AR and angle of attack values are given. In order to make it as close as possible to the 2 -D
airfoil, we can choose AR value as 8. As we can see in the graphic below, an approximate value of 6-8
seems logical. But wings’ structural durability and materials are effective on AR value so we looked for
the aircrafts which is made of balsa wood and decided the AR value as 7.
In consequence of
2
determining AR as 7, wingspan is emerged 2.15 m because of having 0.66 m wing referance area.
12-14
18%
10-12
22%
4-6
17%
8-10
17%
6-8
26%
Figure 4.8: AR comparison results
Taper Ratio: We decided to use tapered wings in order to reduce the induced drag of the aircraft and
reach the elliptical wing loading distribution. But as the taper ratio reduces, stall begins from wing tips.
This means our control surfaces will be stalled earlier which we don’t appreciate. A 0,45 tapered wing
causes only 1% more induced drag than an elliptical wing. Even though it seems logical to choose
0,45 taper ratio, we have wanted our spar to be straight and decided to use 0,6 tapered wing.
Incidence Angle: Incidence angle is usually used to reach the most efficient angle of attack for the
wings while the body flying with an AOA which provides minimum drag for the body.
We determine incidence angle to provide the required lift coefficient. We have designed our
fuselage symmetrically. So body drag will be minimum at 0° angle of attack. Because of that, during
our cruise flight we will try to keep this angle.Since we have wanted our airfoil to have a lift coefficient
of 0,5 and chosen Eppler 422 airfoil -2° incidince angle is preferred. In the figure 4.3 we can see that
for -2° angle of attack Eppler 422’s lift coefficient is 0,5.
Even though sweep angle,
dihedral,
aerodynamic
and geometric twist may provide
advantages in some characteristics, when we have taken into consideration disadvantages in terms of
construction, we have made a decision to not use these features.
Page 28 of 56
4.4.3 Fuselage Sizing
The most important parameter to size the fuselage is payload.Therefore, aircraft configuration
is emerged as shown in figure 4.9 in consequence of placing the payload inside of the body,
considering the ribs between them and planning the spar mounted in. The aft part is designed
symmetrically and slightly narrowing in order to reduce the drag force. And as releasing the fire
extinguisher balls, asymmetric flight conditions will occur. So we have planned to use a boom to
increase the effectiveness of the control surfaces and empennage stabilizers. Boom length was
decided in conclusion of stability and control calculations.
Figure 4.9: Aircraft side view
Main landing gear is designed to place under the reinforced area between 2
Also spar is plannned to be monted in the gap between 2
nd
nd
rd
and 3
balls.
rd
and 3 balls.
4.4.4 Empennage Sizing
The horizontal tail must provide enough moment to rotate for take off and provide longitudinal
stability. Initial horizontal stabilizer sizing was performed using empennage volumes from previous
DBF aircrafts that flew similar missions successfully. Extending the boom length makes empennage
smaller and provides weight advantage. But if we extend the boom more than required empennage
may hit the ground during take-off. For this reason we optimized the boom length and got the result of
40 cm. Horizontal stabilizer volume coefficient was determined as 0.35. Therefore horizontal stabilizer
area was calculated as 1050 cm². Taper ratio was choosen the same with the wings (0.6). In this case
horizontal stabilizer’s wingspan, root chord and tip chord was emerged as 72/18/11 cm respectively.
Angle of attack for the empennage was calculated as -2° during cruise flight to provide stability.
Volume coefficient for the vertical stabilizer was determined as 0,025 and 18 cm root chord
was adjusted in order to have the same chord with the horizontal stabilizers. So, 36 cm vertical
stabilizer length and 11 cm tip cord were figured out.
4.5 Stability and Control
Our aircraft fuselage has a symmetrical structure according to the central axis. As a result of
this, we have assumed that body angle of attack changes have not influenced yawning and rolling
Page 29 of 56
moment. Therefore, we have divided calculations into two parts as longitudinal (pitch) and lateral (roll
and yaw). Also, in order to provide stability, it was calculated neutral point ( ̅ np =0.505 in.) with the
equation 4.1 by assuming pitching moment and thrust changes constant.
(4.1)
In order to calculate aircraft static margin using equation 4.2, cruise flight stability values were
examined in table 4.1 for all missions.
(4.2)
Factor
Mission 1 Mission 2 Mission 3
Static Margin
7,50%
19,90%
12,60%
Cmα
-0,3797
-0,9627
-0,6098
Table 4.1: Stability results
After the static stability analysis, we made the dynamic stability analys is which is done to
see the aircraft's ability to fly straight and to do necessary maneuvers in the missions. Stability
derivatives for the dynamic stability calculations were found by using methods from Nelson and
Raymer are seen in table 4.2.
Stability and Control Derivative Coefficients
Cl,δr
0,013
Cl,p
-0,881
Cl,r
0,161
Cl,β
-0,01
Cl,δa
0,695
Cn,δr
-0,084
Cn,p
-0,072
Cn,r
-0,275
Cn,β
0,145
Cl,δa
-0,103
Cy,δr
0,103
Cy,p
-0,001
Cy,r
0,385
Cy,β
-0,405
Cm,α
-0,391
Cm,δe
-0,857
Cm,α
-0,215
Table 4.2: Stability and Control Derivative Coefficients
Control Surface Sizing
The control surfaces were sized using previous aircrafts as a benchmark, with general
sizing for a number of aircrafts being considered. Since we have a conventional tail, there are three
empannage parts which include two horizontal stabilizer parts and one vertical stabilizer part.
Elevator control surfaces have a size of 40% of the horizontal tail area and a ±40° maximum
Page 30 of 56
deflection. Ruddder control surface has a size of 20% of the vertical tail and ±20° maximum
deflection. It can be easily seen that elevator surfaces play more influential role than rudder surface
for radio-controlled aerial vehicles.
The ailerons were sized using Raymer historical trend to provide enough roll control during
turning and take-off. The ailerons were sized to a total area of 17% wing area with +20°/ 10°maximum deflections.
4.6 Structure
Catia V5r19, a multi-disciplinary 3-D drawing program which analyzes stresses and
displacements, was used in the structural analysis of the aircraft. The program was used to analyze
the wing, tail boom, fuselage and landing gear. The program lets us specify the material type,
constraints, and loads to be applied.
Figure 4.10: Catia load analysis of the components
As our aircraft's landing weight is maximum 4,5 kg (there is no mission required landing with
payload) and it is subjected to a 2,5 g load while landing, we determined the load of fuselage, landing
gear and wing as maximum 15 kg distributed force. Since the tail boom carries the tail, we set the
load as 1 kg distributed force. After the program was applied to the wing which is made of balsa wood,
we saw that the maximum stress occurs at the wing root and the first picture on the top left shows the
detail stress analysis of that region. As it is seen from the figure 4.10, the spar (made of carbon fiber)
length provide enough strength to our wing. Then the program was applied to the aluminium tail boom
which is shown as the second Picture on the top. We decided to change its material. Because it was
unable to carry the load. We set carbon fiber as the material and get the results we wanted to see.
Page 31 of 56
When the program was applied to the fuselage, we saw that it needs to be reinforced. So, we decided
to strengthen it several rips. Finally, we applied the program to the landing gear. It can be seen that
the highest stress concentrations occur along the curvature of the landing gear.
4.7 Propulsion System
4.7.1 Motor Selection
Motor should both provide sufficient static thrust in order to take-off from a distance of 40
meter and should be as light as possible. Static thrust obtained from motor should both overcome
rolling friction, aerodynamic drag and ensure the necessary acceleration for take-off.
Frequired
Fmotor - Frolling friction - Faerody namic drag
⁄
Faerody namic drag
Frolling f riction
Frequired
2
[k(W-L)]
SCd
(Increases while advancing in the runway)
(Lift
increases
while
advancing
in
the
runway)
ma
2
For Mission 3 it was calculated V stall=12,5m/s and V take-of f =15m/s (W=10 kg; c lmax = 1,5; S=0,66 m )
2
0
2
2
+ 2ax
ax
2
15
0 + 2ax
112,5
Take- off distance is limited 40 meter. We have thought the worst situation and taken into
consideration tail wind. So it is planned to take-off 30 meter distance. In this case;
30a
112,5
Frequired
a
ma
10.3,75
3,75
37,5 N
Rolling friction and aerodynamic drag were considered affecting the aircraft constantly with the
highest value. (L = 0, Vtake-of f = constant)
Faerody namic drag
Frolling f riction
Fmotor
2
0,5.1,225.15 .0,8721. (0,023 )
0,005.10.9,81
2,8 N
4,9 N
37,5+2,8+4,9 45,2 N
4,6 g
It should be preferred the motor generates minimum 5 kg static thrust in order to take into
consideration safety margin. İt is known that engine thrust is decreased with increas ing air velocity.
Therefore, it should be found the required thrust and engine power for Mission 3 which is planned to
have the maximum cruise flight among all missions .
2
D
0,5.1,225. 25 .0,8721.0,023
P
T
7,7.25
7,7 N
192W
Assuming the motor and propeller efficiencies 70%; Prequired
192.
⁄ .
⁄
391 W
Page 32 of 56
Motor selection FOMs were considered as expressed below;
Power/Weight: Since it is aimed to provide high power with low weight, power/weight ratio should be
selected as high as possible.
Efficiency: It is allowed maximum 40 A current drawn. Therefore, engine most efficient operating
range should be less than 40 A. Furthermore, engine most efficient current value should be applied
cruise flight conditions.
Resistance: It should be as low as possible.
RPM: It should be considered as high as possible in order to provide required stability unless the noise
limit exceeded. If it is possible, propeller speed should be adjusted by using gearbox.
We have taken into consideration these criteria and chosen brushless motor. Because
brushless motors provide higher efficiency and torque compared with brush motors. Also it was
considered to make a choice among the inrunner types because of easier cooling systems than
outrunner types. Furthermore, outrunner types efficiencies
are lower and draw more current to
generate same torque.In consequence of these criteria, it is examined several motors and compatible
propellers in terms of 40 A current limit and 5 kg static thrust requirement. As it can be seen from here
we have tried to determine prop size simultaneously. Consequently, we have chosen Hacker C50-14L
Acro 6,7:1 competition model motor shown in figure 4.11.
Figure 4.11: Hacker C50-14L Acro 6,7;1 competition model motor
4.7.2 Propeller Diameter Calculation
Aircraft stationary state propeller tip velocity is expressed with the equation 4.3. Considering
forward movement of the aircraft, propeller helical velocity can be calculated with the equation 4.4. If
propeller velocity is approached sound speed, it may cause excessive noise and structural damage.
Vtip static = π.D.n
(4.3)
2
Vtip helical = Vtip static + Vaircraf t
2
(4.4)
There is a risk of structural damage to occur, if the speed exceeds 260 m/sec. for wooden
propellers and 290 m/sec. for metal propellers. If excessive noise is not desired, 213 m/sec. helical
speed must not be exceeded.
If it is considered our design maximum velocity is 25 m/sec D ≤
211,5
3,14.n
is emerged.
Page 33 of 56
D = 16 inch  n ≤ 9944 RPM
D = 17 inch  n ≤ 9360 RPM
D = 18 inch  n ≤ 8820 RPM
D = 19 inch  n ≤ 8340 RPM
D = 20 inch  n ≤ 7920 RPM
It was taken into consideration that as the propeller diameter grows the current drawn under
the same voltage increases, as the propeller diameter becomes smaller static thrust diminishes.
In order to provide required static thrust and not exceed 40 A current limits were the basic
constraints. Propellers which do not exceed RPM limits above were tested and 20 x 10 inch propeller
was selected. Pitch angle value was calculated by the help of RPM value provided 25 m/sec. c ruise
velocity at Mission 3. Drive calculator and prop c alculator package programs were used during
calculations.
Figure 4.12: Hacker C50-14L AC 6,7:1 performance results
4.8 Preliminary Lift and Drag Estimation
4.8.1 Lift Estimation
First of all we calculated the lift coefficient of our airfoil by using software packs (Profili,
Designfoil). Then 2-D airfoil lift coefficient values were transformed into 3-D wing values by the help of
Prandtl equations. And we applied the same method to tail surfaces. We considered the fuselage and
tail effects while calculating total aircraft lift coefficient. But we can say that our fuselage will not
generate any lift at 0 deg AOA because of its symmetrical shape. Aerodynamic coefficient values of
tail and wing are given in table 4.3 and 4.4.
Page 34 of 56
NACA 0009 - Re = 300000
EPPLER 422 - Re = 600000
/
α
/
α
-3.0
0.3495
0.0125
27.9600
-0.1244
-3.0 -0.3863 0.0083 -46.5422 0.0085
-2.0
0.4591
0.0113
40.6283
-0.1230
-2.0 -0.2910 0.0074 -39.3243 0.0119
-1.0
0.5659
0.0105
53.8952
-0.1214
-1.0 -0.1822 0.0072 -25.3056 0.0136
0.0
0.6752
0.0107
63.1028
-0.1205
0.0
0.0000 0.0074 0.0000
1.0
0.7844
0.0107
73.3084
-0.1195
1.0
0.1823 0.0072 25.3194 -0.0135
2.0
0.8908
0.0108
82.4815
-0.1181
2.0
0.2910 0.0074 39.3243 -0.0118
3.0
1.0246
0.0104
98.5192
-0.1232
3.0
0.3863 0.0083 46.5422 -0.0084
4.0
1.1257
0.0107
105.2056 -0.1210
4.0
0.4786 0.0100 47.8600 -0.0053
5.0
1.2248
0.0112
109.3571 -0.1185
5.0
0.5683 0.0124 45.8306 -0.0019
6.0
1.3206
0.0118
111.9153 -0.1156
6.0
0.6586 0.0150 43.9067 0.0019
7.0
1.4160
0.0128
110.6250 -0.1131
7.0
0.7467 0.0185 40.3622 0.0059
8.0
1.5057
0.0132
114.0682 -0.1094
8.0
0.8332 0.0239 34.8619 0.0097
9.0
1.5836
0.0139
113.9281 -0.1037
9.0
0.9124 0.0330 27.6485 0.0129
10.0
1.6532
0.0149
110.9530 -0.0969
0.0000
10.0 0.9685 0.0444 21.8131 0.0177
Table 4.3: Wing aerodynamic
Table 4.4: Tail aerodynamic
coefficient values vs. AOA
coefficient values vs. AOA
values of tail
We assume that maximum
A
aerodynamic drag will occur during the last mission because of the
high velocity. This drag value will be an important factor when choosing our motor. For this reason we
calculated the cruise flight lift coefficient values which affect induced drag. These values are 0.3542 for
the wing and -0.13 for the tail. We assume that lift generated by fuselage is zero because of the
symmetrical shape. In this case total aircraft lift coefficient can be calc ulated by using the equation 4.5.
CL
CL + CLtail × (Stail / Swing )
(4.5)
Maximum lift coefficient was calculated as 1.8 by using Profili and 1.6 by using Designfoil. For
2-D airfoil, 1.6 value was chosen as the worst possibility. 3-D maximum lift coefficient for the wing was
calculated by the help of equation 4.6.
CLmax
Cl × 0,9 × Cos(
0,25)
(4.6)
Tail lift coefficient was calculated using the same method and total aircraft lift coefficient was
emerged as 1.5.
Page 35 of 56
4.8.2 Drag Estimation
Component build-up method was applied with the help of equation 4.7 in order to calculate the
drag values. In order to apply this method reference wetted areas were calculated.
(CDo)subsonic
∑ (C fc FF c Qc Swet c )
Sref
+ (CD )misc + (CD ) l
(4.7)
p
Top and side views of fuselage are the same because of the symmetrical shape and given in
the figure 4.13.
Figure 4.13: Sref and S wet areas of components
We calculated reference area for the wing as 0,666 m² in the conceptual design phase. When we
2
consider root cord is 39 cm and fuselage width is 17 cm, Sexp is emerged as 6003 cm . And equation
4.8 can be used to calculate wing wetted area.
Swet 2×(1+0,2 × t/c)×S exp
(4.8)
12342 cm²
Same method can be applied to tail and reference area 1312 cm² wetted area 2671 cm²
values can be found.
Carbon-fiber boom has 30 cm length and 1.5 cm radius. Reference and wetted area are
calculated as 75 cm² and 235 cm², respectively. As a result;
(
)
8721 cm²
⁄
(
)
=2,18
20415 cm²
Turbulence is not affected by only Re number but also affected by Mach number and surface
roughness. For our design, there will be probably laminer flow up to a point of the wing chord and then
turbulence will begin. But we have assumed the worst possibility that the flow was completely turbulent
around the wing. Herewith, maximum possible drag was kept in mind.
Page 36 of 56
Equation 4.9 was used for Re number and 4.10 for
Re
Cf
.
1,053
38,21×(L / k)
(4.9)
0,074
(4.10)
1/5
Re
For the landing gear D/q table was used from Raymer. The total drag was increased for the
leakage-protuberance and scrubbing drag.
Table 4.5: Aircraft components drag results
Figure 4.14: Drag percentage of each component
Drag force for all missions is the same because we don’t carry any external payloads.
alues in the
table represent drag values of the last mission while flying with maximum velocity.
Page 37 of 56
4.9 Performance Parameters
Mission
Mission 1
Mission 2
Mission 3
1,5
1,5
1,5
0,39
0,39
0,39
0,0187
0,0187
0,0187
Thrust (take-off) (kg)
4,6
4,6
4,6
Take-off Weight (kg)
6,1
7,4
10
W/S (kg/m )
9,24
11,2
15
Take-off Distance (m)
12
28
30
Total Flight Time (s)
130
85
85
Cruise Speed (m/s)
19,4
21,5
25
Stall Speed (m/s)
10
11
13
Parameter
2
Table 4.6: Aircraft performance results
5.0 DETAIL DESIGN
After the preliminary design was completed, we consantrated on the detail design phase.
During detail design, all systems and components were selected and integrated. The aircraft structural
characteristics and capabilities were finalized and a weight and balance summary compiled for each
mission. Finally, flight and mission performance parameters are documented.
5.1 Dimensional Parameters
The table 5.1 below shows the final lengths, widths, and diameters of the fuselage, tail, and
wing in addition to other important aircraft parameters.
FUSELAGE
Length(cm)
BOX
Length(m)
1,2
Width(m)
0,7
96
Width(cm)
17
Height(m)
0,3
Height(cm)
17
Weight(kg)
0,8
Page 38 of 56
COMPONENTS
WING
HORIZONTAL TAIL
VERTICAL TAIL
Airfoil
Eppler 422
NACA 0009
NACA 0009
Span(cm)
217
72
35
Root Chord(cm)
39
18
18
Tip Chord(cm)
23
11
11
Incidence Angle(deg)
-2
-2
Height
36
(cm)
Taper
0,6
0,6
0,6
Area(m )
0,66
0,10
0,05
AR
7
5
5
Volume coefficient
-
0,35
0,025
2
Control Surfaces
-
Elevator
Area
0,0648 m
2
Rudder
0,014 m
Area
Aircraft Weight
Mission 1
Mission 2
Mission 3
W empty(kg)
6,1
6,1
6,1
W payload (kg)
-
1,3
3,9
6,1
7,4
10
Maximum Gross Weight
(kg)
2
Table 5.1: Aircraft dimensional parameters
5.2 Structural characteristics and capabilities
Fuselage System
Our fuselage system was designed around the best fireball placement, minimum drag,
minimum weight and minimum loading time. We have made a decision to manufacture a structural
strength body in order to provide robust mounting points for both wings, landing gears and other
aircraft basic components. Therefore, as a result of analysis carbon-fiber body was preferred in
cosequense of providing higher strength/weight ratio. We have aimed to carry three 1,3-kg balls inside
the body at last mission. In order to provide a convenience releasing system, it was opened three
hatch under the body. Also considering the easiness of loading time at Mission 2, one more hatch is
opened above the body. To solve this loss of endurance, body was strengthened by three rips which
were replaced each spaces between three balls. Fuselage hatches opened view can be seen in figure
5.1.
Page 39 of 56
Figure 5.1: Aircraft fuselage
Loading and releasing mechanisms
It is planned to place hatch systems separately under each balls in consequence of releasing
them respectively at Mission 3. Hatch systems will be kept by servo controlled pins and if the pin is
pulled by servo, hatch will be opened by ball’s weight and will be closed by spring connected to the
body. Since it is not required shorter loading time at Mission 3, balls are planned to load from releasing
hatches by reversing the aircraft. Therefore, we don’t need to place hatches or mechanisms which will
cause extra weight increase on upper side. But loading time has a vital role on scoring at Mission 2
and one ball should be loaded as soon as possible. For this reason, one more hatch mechanism
(which can be easily opened and closed) was placed on upper side. Loading and releasing
mechanisms are seen in figure 5.2.
Figure 5.2: Ball loading mechanism for Mission 2
Page 40 of 56
Figure 5.3: Ball releasing mechanism for mission 3
Wing Assembly
In consequence of making a decision to mount wing middle of the body, our wings were
planned as two parts. As shown in figure 5.4 wing is mounted by inserting the spar inside the larger
diametered pipe which has the same length with body width. Wings were made from balsa to provide
weight gain. In order to ensure adequate strength to the spar mounting place, airfoil rips were
thickened towards the body connection. It was used tapered wing to approach elliptical wing loading.
Figure 5.4: Wing mounting system
Storage Box
The storage box was designed around two indispensable objectives; RAC and the assembly
time. In order to meet these requirements all structural parts and elements of the box must be
lightweight. Although the assembly time is not scored, it must be done in 5 minutes. Also it must have
an easy opening mechanism in order to enable us a quick taking and loading time of the softball since
the loading time is scored in Mission 2.
Page 41 of 56
We determined the size and shape of the box according to the overall configuration of the
aircraft but it is limited with 80x80x150 cm dimensions in the competition rules. If the aircraft is
seperated all its components storage box size does not create an obstale for us.
We used carbon tubes to create the structural frame of the box. Also, the box structure has
crosswise truss carbon tubes to ensure an easy opening mechanism. The carbonfiber sheeting covers
the structural frame because of its lightweight. Also, we use bals a truss elements to place the
components of the aircraft into the box safely as shown in the figure 5.5 below.
Figure 5.5: Total system storage box
5.3. Systems and Sub-Systems Selection and Integration
Propulsion System Selection
The propulsion system components were selected after optimization studies done in the
preliminary design phase. In order to keep center of gravity in front of the aerodynamic center, the
batteries were placed on the top of the first ball, on the ceiling of the fuselage. Since that the maximum
current that can be used in the competition is 40 amps, the peak value of the ESC was determined
according to this. The size of the propeller was optimized in order to maintain the static balance of the
last mission’s heaviest conditions. So, it can be said that the static balance of the other missions can
be maintained. Pitch angle was chosen as big as possible, because the flight time is important in
scoring.
Propulsion System Selection
Motor
ESC
Hacker C50-14L Aero 6,7:1
Futaba R6014HS 14-Channel
2.4GHz FASST Receiver
Battery
8S2P
Propeller
APC 20 X 12
Table 5.2: Propulsion system elements
Page 42 of 56
Control Sub-System Selection
The Control Sub-System Selection includes the selection and placement of servos, receiver,
receiver battery and transmitter. The rudder servos are located on the pressure side of rudder. The
aileron servos are located on the pressure side of the wing. The last servo is located on the front
landing gear for ground handling. The servos are selected based upon the minimum weight with
sufficient torque values. Receiver and receiver battery are located at the back and front side of the
fuselage. We considered weight for the receiver, enough capacity along with the weight for the
receiver battery.
Control Subsystems
Subsystem
Model
Weight (gr)
Selection Justification
225
Lightweight, Sufficient Torque
20
Lightweight
Sanyo 1000 Mah
100
Lightweight, Adequate Capacity
Futaba R6014FS 2.Ghz
1100
Pilot Preference
Servos (9)
Receiver
Hitec HS-125MG
Futaba R6014FS 10-Channel
Receiver
Battery
Transmitter
Table 5.3: Control elements
5.4 Flight Performance Parameters
Flight Performance Parameters
Mission 1
Mission 2
CL_max
1,5
CL_cruise
0,39
CD0
0,0187
Mission 3
Cruise speed (m/s)
19,32
21,31
24,94
Stall Speed (m/s)
9,85
10,86
12,72
Max. Turn Rate (deg/s)
90
55
35
Max. Rate of climb (m/min)
325
205
120
Table 5.4: Flight performance parameters
5.5 Weight and Balance
Since missions need different payload placements, it affects the c.g. location of the aircraft
between missions. Besides, since that the location of the c.g. will differ while releasing the balls, we
aim to keep this change at minimum. But, after releasing all the balls in the third mission, the c.g. isn’t
behind the a.c. For this reason, all components are located in order to keep the c.g. in front of the a.c.
in ferry flight. In last mission, even though the center of gravity is little bit behind the aerodynamic
center, sufficient moment has been provided by using larger elevator.
Page 43 of 56
Component
Weight (gr)
Arm (cm)
Moment (gr.cm)
Fuselage
2385
50,5
120442,5
Wing
771
56
43176
Horizontal Tail
300
137
41100
Vertical Tail
150
137
20550
Battery Pack
537
12
6444
Motor
423
12
5076
Boom
70
116
8120
Esc
22
40
880
Propeller
135
1
135
Receiver
20
116
2320
Av. Battery
72,5
10
725
Nose Landing Gear
80
2
160
Main Landing Gear
226
51
11526
Ball-1
1300
23
29900
Ball-2
1300
39,5
51350
Ball-3
1300
46
59800
Missions
Total Weight (gr) Total Moment (gr.cm) Center of Gravity (cm)
Mission-1
5191,5
260654,5
50,20793605
Mission-2
6491,5
312004,5
48,06354464
Mission-3
9091,5
401704,5
44,184623
Aerodynamic center of the wing (cm)
50,5
Table 5.5: Weight and balance calculations
5.6 Rated Aircraft Cost
RAC is one of the most important qualities that affect the flight score directly. Rated aircraft
cost was determined as 6 kg in the competitive design requirements phase. The table below shows
the founded weights for every component of our final aircraft design.
Page 44 of 56
Component
Weight (gr)
Percent (%)
Fuselage
2385
34,63297756
Wing
771
11,1958179
Empennange
450
6,534524069
Motor
423
6,142452625
Esc
22
0,319465621
Batteries
537
7,797865389
Landing Gear
306
4,443476367
Box
800
11,61693168
Transmitter
1100
15,97328106
Receiver
20
0,290423292
Av. Battery
72,5
1,052784433
Total
6886,5
100
Table 5.6: RAC Breakdown
Fuselage
Wing
Empennange
Motor
Esc
Batteries
Landing Gear
Box
Transmitter
Receiver
Av. Battery
Figure 5.6: RAC Pie Chart
6.0 Manufacturing Process
From the begining of the design period, our manufacturing team worked paralelly with the
design group in order to find out the optimum way of manufacturing. Because manufacturing method
and materials affects the weight of the aircraft dramatically. So, many different types of construction
methods and materials were researched and analyzed. We preapared a figure of merit and scheduling
approach to develop a reasonable manufacturing plan.
Page 45 of 56
6.1 Manufacturing Figures of Merit
A list of foam was prepared for the manufacturing plan. The factors used in FOMs were used
to compare the manufacturing options and provide a basis for the optimum selection off materials and
processes. Below, there are the basic selection parameters which were considered to have
importance mostly.
Weight: Total weight is one of most important figures that directly affects the flight score. So, weight
reduction is our main target and must be a serious factor in choosing the manifacturing process and
the materials. But, strength loss should be considered while reducing weight.
Ease of manifacturing: Manifacturing process should provide required easiness in order to complete
components in scheduled time.
Reparability: The repairing and replacement convenience plays a vital role in cas e of an emergency.
Strength: Even though all components should provide a certain extent strength, fuselage also must
ensure higher structural durability in consequence of holding aircraft parts together and carrying
payload.
6.2 Investigating the Manifacturing Techniques
The following techniques were considered while analysing.
Balsa build up: The structure consists primarily of a wooden frame. Parts can be laser cut for
accurate dimensioning. It includes creating the balsa frames and covering them with a monocoque
skin. Frames define the shape and plywood frames connect loaded parts like wings and payload.
Foam core: Foam was cut at the shape of the components. Then fiberglass and epoxy are laid on top
of it. Since it is essential to ensure a smooth surface, it was emeried in order to have a rigid structure.
Lost foam: This method is the following step of the foam core method. After covering the foam with
composite materials, the internal foam is removed. This makes the components lighter in weight. But
the difficulty of this method is a disadvantage.
Molded composite: A foam male mold is constructed in a similar fashion to the foam core composite
method. Female molds for the upper and lower wing surface are built around a male mold using thick
fiberglass cloth. A release agent allows the molds to separate, from which multiple identical wings can
be fabricated.
Page 46 of 56
6.3 Components Manufacturing Methods
6.3.1 Wings and Empennage
In order to determine the best techniques for both the wing and the empennage, we create
decision matrixs shown below.
OPTIONS
Figure of Merit
Weight
Balsa build up
Lost foam
Molded composite
Foam core
Weight
4
1
0
-1
0
Ease of
Manufacturing
3
1
0
0
0
Reparability
1
0
1
1
1
Strength
2
-1
0
1
1
TOTAL
10
5
1
-1
3
Figure 6.1: Wing manifacturing technique decision matrix
As shown in figure 6.1 we decided that balsa build up is the most convenient technique for the
wing. Firstly individual components were made and then assembled together. After the balsa skeleton
was constructed, the surface was coated in a film. The film was used instead of fiberglass, which
decreases the weight of the wing without removing all structural support. As it can be seen here that
there is an easy manufacturing process.
OPTIONS
Figure of Merit
Weight
Balsa build up
Lost foam
Molded composite
Foam core
Weight
3
1
0
-1
0
Ease of
Manufacturing
2
1
0
0
0
Reparability
1
0
1
1
1
Strength
4
-1
0
1
1
TOTAL
10
1
1
2
5
Figure 6.2: Empannage manifacturing technique decision matrix
As a result of FOM seen in figure 6.2 it is tended to a different manufacturing method (Foam
core) for the empennage. Because the influence of the strength was increased to reach higher
strength for the empennage. In consequence of asymmetric flight it was required to overcome more
moments. In order to provide higher moments without damage it was preferred to strengthen
horizontal, vertical stabilizers and control surfaces.
Page 47 of 56
6.3.2 Fuselage
The techniques were investigated and molded carbon technique was selected as a result of
Catia load structural analysis. Composite materials provided higher durability, flexibility and weight
gain.
Figure 6.3: Fuselage manufacturing process
First of all, we have cut the foam by using a CNC foam cutter considering the sizes of the
fuselage. So it was created the male and female molds. But we have not used the male one. It used
epoxy, wax, polish and sandpaper to get a smooth inner surface for the mold. The inner surface was
covered with carbon fiber. We have used a compressor to remove the air from the inside of the molds
to vacuum. One day later, when the carbon fiber hardened, we have taken the fuselage out of the
molds.
6.3.3 Landing Gear
The main landing gear manufacturing materials can be selected from aluminium, steel or
composite materials. All of these options were considered and analysis were made to select the best
one. After the analysis, we have chosen composite materials considering their durabili ty, weight and
flexibility.
It was created a mold at the beginning of the process. The inner surface of the mold was
covered with first carbon fiber then with kevlar and finally with foam and the remaining spaces were
covered with carbon fiber. One day later, it hardened and we got the landing gear.
Page 48 of 56
6.4 Manufacturing Milestone Chart
Manufacturing progress can be seen is seen in figure 6.4.
Figure 6.4: Manufacturing milestone chart
7.0 TESTING
In order to finalize the optimization process of our aircraft, we have carried out how to perform
required tests. The main objective of all the testing was to verify the predictions about aircraft
performance characteristics. Additionally, the test program was used in the development and
improvement of the airplane and its components. Finally the following tests were handled: structural,
payload, propulsion, ground and flight testing. Our team has also tested the aircraft performance for
all the missions. This section summarizes the main testing objectives, plans, and methodology that
drove the final design presented in the previous sections. It is shown in table 7.1 our objectives during
testing processes.
Turkish Air Force Academy Fesa Team Testing Checklist
Test
Wings
Landing gear
Loading
Releasing
Battery/Motor
Propeller
Ground
Flight
Objectives
Are they strong enough to meet flight dynamics?
What is the height at which gear fails at 10 kg?
What is the weight at which gear fails 22 cm?
What should be done to make it enough?
What is the time for loading fireballs?
Are there any problems while loading them?
Do the releasing mechanisms work correctly?
Are they optimum preference or not?
Is it compatible with the system?
Is it checked ease of taxiing straight, turning radius?
Are the parameters adequate for flight conditions?
Response
Table 7.1: Testing checklist
Page 49 of 56
7.1 Structural Testing
This section includes the wing and the landing gear tests. First, the body mounted spar
strength was tested by putting weight on the tip side of the wings. Although the wings were exposed
to the weight heavier than fully loaded time, they have provided the required strength. Our aircraft
landing gears must meet 2.5 g sudden force during landing. Different landing gears were tested, the
only difference among them was their materials. At the end, we found that the landing gear made from
kevlar and carbon was the most suitable option and provided lightweight, too.
7.2 Payload Mechanism Testing
We have tested our loading mechanism and loading time for Mission 2. Payload releasing
mechanism and their external integrity ports were tested to prevent any unexpected situation. We
didn’t meet any trouble about loading-releasing mechanisms or placement spaces of payload inside
the body.
7.3 Propulsion System Testing
This test has provided approximate results about aircraft performance. An original system
was established in order to measure the power and thrust of the motor mechanically to see whether it
will be enough for the aircraft or not. Different battery packs and propellers were tested to get the best
results. Hacker C50-14L Acro 6,7;1 competition model motor with 8S2P 3800 mah LiPo batteries
along with 20x12 inch propeller was selected.
Figure 7.1: Propulsion System Test
7.4 Ground Testing
Proper operation of electrical subsystems and the propulsion system is verified upon
completion of the aircraft. Firstly, we put all the components in our storage box with all other
necessary things to see the best positioning of the parts. After that, we assembled our aircraft and
test the ground handling both laoded and unloaded cases. Also we measured our assembly time in
order to check whether it exceeds 5 minutes or not.
Page 50 of 56
7.5 Flight Testing
Flight testings were made in order to see the aircraft flight performance, assess dynamic
modes, and ensure proper completion of all mission profiles. A radar gun was used to measure the
speed of the aircraft during takeoff and cruise. Estimated values and real values were compared.
These flight testings were made according to a flight test checklist as shown in table 7.2 below.
Turkish Air Force Academy Fesa Team Test Flight Checklist
Flight
Time
Temperature
Date
Wind
Pilot
Objective
Before Flight Controls
In Flight Controls
After Flight Controls
CG Control
Takeoff Speed
Batteries
Secure
Cruise Speed
Radio
Batteries
Stall Speed
Receiver
Propeller
Takeoff Contrability
Speed Controller
Speed controller
Sink Rate
Propeller
Fuse
Thrust Characteristics
Control Surfaces
Stability and Control
Payload Mechanisms
Roll
Landing Gear
Yaw
Taxi Handling
Pitch
Radio
Receiver
Fail Safe
Comments
Comments
Comments
Table 7.2: Flight Checklist
Page 51 of 56
7.6 Testing Schedule
The milestone chart for the testing progress is shown in table 7.3 below.
Figure 7.2: Milestone Chart
8.0 Performance Results
After testing, we have reviewed the results and component designs were corrected if it was
inadequate. The performance details of critical components and changes to the design taking into
account the test results are explained in the section below.
8.1 Key Subsystem Performance
8.1.1 Propulsion
We have predicted that 20x10 propeller would be suitable for all missions after our
optimization studies. However the demonstrated performance of our aircraft showed that the batteries
were enough but the flight time was not what we expected. Considering takeoff distance and battery
consumption are not an issue, we decided to select a higher pitch propeller which will yield a higher
score because it allows for more thrust and therefore a greater maximum velocity. So, we changed
our propeller dimensions to 20x12 inch for Mission 1 and 2. This time, it was reached the results we
wanted. For Missions 2, since efficiency is more important than top speed, the 20x10 inch propeller
was tried and the results were positive, so we decided not to change the propeller for this mission.
After aircraft performances, we faced a problem with the overheating of the electronic speed
controller. In order to reduce this heat, we decided to place the ESC outside the aircraft and solved
the problem.
Page 52 of 56
8.1.2 Wing
We made two tests to see the performance of the wing; spar strength and wing loading tests.
Our aim was to see whether these are enough to stand the maximum loaded weight (8.25 kg)
required by the missions. Root side of the wing was put on the table by the help of spar and the tip
side hand-helded. Thus, it is amied to see both the spar and wing loading strength. Although we have
2
2
2
designed our wing endure 15 kg/m wing loading, it was applied 22 kg/m (7kg weight to 0,33 m wing
area) without failure. In the same way spar was designed to endure 267 kg-cm moment. Buth it was
applied 376 kg-cm moment without failure. It was emerged 1.41 safety factor which can be
acceptable enough for both spar and wing loading strength.
Figure 8.1: Wing spar
Figure 8.2: 7kg distributed load applied wing
8.1.3 Tail boom
The torque on the tail boom from the ruddervator was the major factor tested in the design
and fabrication of the boom. We calculated that the tail boom should provide 180 kg-cm. The boom
was able to carry 253 kg-cm without failure. Test showed us that the tail boom's demonstrated
performance is sufficient for our mission requirements.
Figure 8.3: Tail boom performance test
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8.1.4 Landing Gear
Considering that there is no loaded landing we have tried to test with our emty weight. Our
emty weight is 4,25 kg and it is subjected 2.5 g sudden force while landing. With a 1.41 safety factor it
was applied 15 kg net weight without failure.
Figure 8.4: Landing gear performance test
8.2 Complete System Performance
The aircraft had to be consistent in both loading and flight. Therefore, extensive testing was
done to ensure a reliable and quick laoding time. Flight testing was also performed to practice flying
the missions in the optimized mission plan and ensure the aircraft was capable of completing each
mission.
8.2.1 Loading Performance
The prototype aircraft has been used to demonstrate, practice and evaluate ground handling
performance and loading time. Ground crew try-outs were performed to select the fastest set of team
members that will perform the loading task. The preliminary rounds of tests show that loading time for
the Mission 2 is about 10 seconds and after some performances, we reached our goal which is 6
seconds shown in figure 8.5. Also we have reached 3,9 minutes assembly time after some trials seen
in figure 8.6.
Page 54 of 56
Loading Time Trials
Assembly Time Trials
5
Assembly time(min)
Loading time(sec)
12
10
8
6
4
2
0
4,5
4
3,5
3
0
5
10
15
20
25
30
0
2
Trial number
4
6
8
10
12
Trial number
Figure 8.5: Loading time tests
Figure 8.6: Assembly time tests
8.2.2 Flight Testing Performance
Firstly, we have tested our aircraft taxi performance. After that, for the first flight, we just
aimed to fly and do some basic maneuvers with empty weight. Then, we started to fly with the real
mission conditions on the progress.
Figure 8.3: Taxi and Flight Tests
Take-off distance has not created any problem for any mission because it never passed 15
metres.
st
At 1
mission our flight time has changed between 150 and 170 seconds as we have
expected.
In Mission 2, since the flight time is not scored, our aim was just to improve loading time of a
one ball. We have started 10 seconds loading time at first performance. But we have learned that
hatch closure mechanism was slowing us. When we changed it with magnetic mechanism we
reduced loading time 6 seconds after some trials. We have also improved our physical conditions to
perform it quicker.
For Mission 3, We had to overcome c.g. shift and asymmetric flight conditions. At first
performance we have understood that our elevator area could not enough to prevent c.g. shift with
trimming. Therefore, we increased our elevator area to provide enough pitch moment when it is
created c.g. shift during releasing balls. After several trials our aircraft has released the three balls and
Page 55 of 56
performed the Mission 3 successfully. Considering our estimated flight as 85-100 seconds, our pilot
will try harder to perform flight pattern better and reduce our flight time 85 seconds until the date of the
contest.
9.0 References
[1] Raymer,Daniel P., Aircraft Design : A Conceptual Approach.
[2] Kroo,Ilan,Ph.D.,Aircraft Design : Synthesis and Analysis.
[3] Jenkinson,Lloyd R./ Marchman III,James F., Aircraft Design Projects for engineering students.
[4] Nelson, R. C. Flight Stability and Automatic Control. 2nd Ed. McGraw-Hill Higher Education. 1998.
[5] Anderson Jr,John D., Aircraft Performance and Design.
[6] Turkish Airforce Academy D/B/F Competition Databas e.
Page 56 of 56

TURKISH AIR FORCE ACADEMY

Transcription

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