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iClean - Loitering attack UAV
CDR
June 27th, 2012 Aerospace Faculty, Technion, Haifa
Moshe EtlisDaniel LevyMor Ram-OnMatan ZazonYa’ara KarnielMeiran HagbiOshri RozenheckYanina DashevskiNathaniel LelloucheMenahem Weinberger
Supervised by Dror Artzi
PDR Overview
Remarks from PDR
Airfoil and Propeller Selection
Geometry Improvements
Performance Calculations
Wing Detailed Design
Wings’ Folding Mechanism
System Installation Layout
Weight and Balance
Wind Tunnel Model Design
Wind Tunnel Test
Conclusions and Recommendations
Table of Contents
Operational capabilities: Suicide UAV. Endurance: 5 hr. Range: 400 NM (approx. 750 Km) Man in the loop. Launching System: Mobile Ground Launcher with as many as
possible UAV's ready to be launched.
Target definition and acquisition:
Target type: Static and mobile. Truck Target: detection range of 30 Km, recognition of 12 Km. Target acquisition: Day and Night Capabilities. Attack capabilities:
Warhead: Approx. 20 Kg. Attack capabilities: Any angle - vertical or horizontal. Low Cost UAV unit.
PDR Overview - Customer Specifications
Diving at 150 kt
BOOM!!
Launch
Climb to 5000 ft
Cruise at 5000 ft at approx. 80 kt
Loiter at 5000 ft at approx. 60 kt
Mission Profile
PDR Overview - Chosen Components
Sensor: Controp ESP 600C (27 lbs, X15 zoom lens, 0.7-22.6 degrees FOV).
Engine: 3W 275 XiB2 (26 HP, 15.5 lbs).
Launching Method: Booster rocket (Launched from a canister).
PDR Overview - 2 Configurations
AB
PDR Overview - Final Geometry for PDR
Remarks and Solutions
Airfoil Selection
NACA 0012Eppler 560
NACA 0012 Eppler 560 Improvement)%(
Max CL 0.972 1.827 88
Max L/D 40.63 60.08 48
Stall angle 7.5 14.5 93
NACA 0012 Eppler 560 NACA 4412 Improvement )%(
Max CL 0.972 1.827 1.507 21
Max L/D 40.63 60.08 57.209 5
Stall angle 7.5 14.5 6 142
Airfoil Selection
NACA 4412
Consulting the engine data and information .
The chosen engine 3W:275 XiB2 TS )from the PDR( .
Engine rotation speed : 1000-7000 RPMpower :26 horsepower =~ 19300 watts.Weight:15.5 lbs=~ 7 Kg.
two blade propeller : 26x16 or 26x14 (“) 3 blade propeller :of 22x14 or 24x14 (“).
Propeller Selection
)From our engine data:(
engine max RPM is 7000 RPM = 116.67 round per second.
Max speed at - 180 secft
550.34
11618m
roundinch
round
Propeller Selection - Calculations – Needed Pitch
3 5
1
Step angle
tan
p
D diameter
V velocity
n RPM value
Density
P power
Vadvanced ratio J
nDP
power coefficient Cn D
pDif is const thenr
Our propeller is a 2 bladed-back-folding propeller at the size of 25X18.
Propeller Selection - Calculations – Needed Diameter
Direction of flight
Stability Solution: Changing the Configuration
0 10 20 30 40 50 60 70 80 90-70
-60
-50
-40
-30
-20
-10
0
10
20
X: 82.8Y: 0.1136
Stability Margin as a function of the Wing Opening AngleS
tabi
lity
mar
gin
opening angle [deg]
Stability Margin [%]
Stability Margin [cm]
40%-60% Configuration’s Stability
Geometry as shown at PDR:
Geometry Improvements
The final geometry for CDR:
Old: The fuselage becomes thinner in the middle of it and
then expends
New: The guideline of the fuselage as much as monotonic
as possible
New: Wings’ hinges are
covered
New: 40-60 canard
Old: 25-75 canard
Old: Wings’ hinges were
exposed
Geometry Improvements
Geometry Improvements
Property Value
Airfoil (EPPLER 560)
Aspect ratio
Spans
Reference lift area
max ,0
zero liftline
11.83; 0.8; 4.1
14.5 0.37[ ], 6.5 0.11[ ]
l l l
stall
C C Crad
rad rad
9, 11, 0.5wing canard fuselageA A A
, 1.5[ ]9.8[ ], 8.9[ ] tailwing canard b ftb ft b ft
Weight 220[ ]W lbf
4.9[ ], 2.0[ ]0.4[ ], 0.1[ ]
w c
w c
x ft x ftz ft z ft
Aerodynamic center’s position
UAV’s Properties
2
2, ,
2, ,
17.6[ ]
10.5[ ], 1.2[ ], 0.9[ ]
7.1[ ], 0.9[ ], 0.7[ ]
ref
wing r wing t wing
canard r canard t canard
S ft
S ft C ft C ft
S ft C ft C ft
Fuselage 28.7[ ], 6.7[ ]fuselage fuselageS ft L ft
Vertical tail 2. , ,1.4[ ], 1.3[ ], 0.6[ ]v tail r tail t tailS ft C ft C ft
Assumptions:
The body as a lift generator componemt:
, constc w
Lift Coefficient’s Properties
1, 1c b
,
1C 2
radbl
0
C 0bL
, constLC
Property Value
,
16.086LC
rad
0.721 6.086LC
max2.11LC
0.228[ ] 13.1stall rad
0 0.118[ ] 6.8LC rad
Lift coefficient slope
Lift coefficient as a function of angle of attack
Minimal lift coefficient at height of 0ft and 5000ft
Maximal lift coefficient
Stall angle
Zero lift angle
min0
min5000
0.317
0.395height ft
height ft
L
L
C
C
Lift Coefficient’s Properties
Assumptions:
0
2
inducedform & skindragdrag
D D LC C KC
1
- Oswald efficiency number
KAe
e
Drag Coefficient’s Properties
0.775e
00.02DC
Property Value
Wing’s induced drag coefficient
Canard's induced drag coefficient
UAV’s drag
0.0456wK
0.0373cK
Fuselage's induced drag coefficient 0.821bK
UAV’s total induced drag coefficient
0.056planeK
20.02 0.056D LC C
Drag Coefficient’s Properties
Property Value
max15[ ] 53.5%T lbf T
min 14.7[ ]T lbf
max 28[ ]T lbf
Velocity
Property Value
0
5000
41.8[ ]
46.7[ ]height
hight ft
stall
stall
V knot
V knot
Maximum thrust
Minimum thrust
Thrust for cruise flight
Minimum velocity (stall) - height of 0ft and 5000ft.
Assumption:
Cruise flight:
T D
W L
Assumptions:
Cruise flight
Maximal velocity:
max 110[ ]V knots
Engine Thrust
Assumption & data:
Constant:
Range & Endurance
Property Result
Range for climb
Endurance for climb
Minimum range for cruise
Maximum endurance for cruise
climb 10.6[ ]R NM
climb 10 minE
cruise 371[ ]R NM
cruise 4.6E hr
Final results:
,V
3
0.75 0.2 10sec
lb lbSFC
hp hr hp
r
L
F
R
Booster Rocket Angle
Time of opening the wings:Velocity:Density :The mass :
:The lift coefficientThe acceleration of the booster:Area of wing that creates the lift:The force that booster applies :
The lift:
The total moment:
3
2
2
2
1.5 sec
117
0.07648 /
220
1.5
/ sec
s
F m a sin
.
N
0 5 l
l
t
v kts
lb ft
m lb
C
a kt
L
ft
v s C
s
20.5 lM L R F r v s C R m a r
Assumptions and data:Booster Rocket Angle
4
Booster Rocket Angle
Wing Detailed Design
The lift load distribution on a trapeze wing:
2
2 3
4
2 3 2
21
2 2 3 2
ROOT TIPTIP
ROOT TIPTIP
C CnW B BQ C bl bl
S B
C CnW B BM C bl bl
S B
Wing Detailed Design - Load Distribution
Wing Detailed Design - Load Distribution
2
125.71
ROOTROOT
ROOT
QQ bl Q
BQ kgf
2
01.5 925 1387.5 141.4
B
x blF SF Q SF Q N kgf
Assuming this lift load distribution the resultant force is:
2
0
2
0
0.5
B
xC B
x
F xX m
F
Wing Detailed Design - Web Thickness
2_ 6 /
1.5 125.710.54
16.16 364
100
wing
WEB
allow carbon fiber
WEB wingallow
Q
H t
Kg mm
Qt mm
H
Wing Detailed Design - Flange Area
1.5 LIMMA
H
bl [mm] 0 500 1000 1400
93643 39365 9220 348.5
33.52 14 3.28 0.12
Flanges
kgf mmM
2[ ]A mm
2[ ]A mm
bl [mm]
Wing Detailed Design - Skin Thickness
1 2
35.2BMPa
t
Thickness [mm]
0.25 140.8
0.5 70.4
1 35.2
1.5 23.5
2 16.25
2.5 13
1 2 MPa Material
Carbon Fibers 706.3
Aluminum 2024-T3
290.4
Aluminum 7075-T6
270.8
0.6y MPa FS
The selected method is the vertical pin for the Advantages below:
Structural simplicity Load paths determined with Confidence Minimum volume of hinge Simple actuator mechanism Very few moving parts Minimum weight
Wing Detailed Design - Joint Selection
A vertical pin through the pivot axis transfer the force-couple from the movable outer wing to the fixed center section
Wing Detailed Design - Joint Selection
Carbon fiber ±45°
Unidirectional
Wing's root
Wing Detailed Design - Final Formation
Wing Detailed Design - Strength Analysis
2 2
2 2
.
12.47 72
1.63 27.6
allowed
allowed
Carbonfibersfibers
hinge Al
kg kg
mm mm
kg kg
mm mm
1 2
70720
12772
root res c
root
M F x kg mm
MF F kg
Force and Moment Calculations:
Conclusion from Allowable and Actual Stresses Calculations:
Wing Detailed Design - Strength Analysis
Wing Detailed Design - Strength Analysis
Shear StressVon Mises
Material Max. Deformation]mm[
Aluminum6
Carbon 2.3
722.6
27.6allowed carbon fiber
allowed aluminum
Factor Calculation:
Wing Detailed Design - Strength Analysis
Pre Calculations:
the average velocity of
the UAV during launch time is:
Reference areas:
50 25.7 [ / sec]launchV knots m
2 2
2 2
10.471[ ] . ][
7.116 [ ] . [ ]
wing
canard
S ft m
S ft m
0 972
0 665
Wings’ Folding Mechanism
,
2 2
,
2 2
3.62 0.232 0.157[ ] 9
3.72 1 0.68 0.05 0.079[ ] 4.5
1 1L = 1.225 25.7 0.972 0.502 197 [ ]
2 21 1
L = 1.225 25.7 0.665 0.3232 2
w
c
sl laun
L w c w
L c w c
wing
canard canard
ch wing L
sl launch L
C i i i rad
C i i i rad
S C
S
N
C
V
V
87 [ ]N
The wing‘s lift during launch time
0 0
0
2
2 2
2 2
0.721
0.02 0.056
1 1= 1.225 25.7 0.972 0.049 19.26 [ ]
2 21 1
= 1.225 25.7 0.665 0.049 13.18[ ]2 2
L
wing ref
canard
D
ref
L
D
D
C
C C
D V N
D V N
S C
S C
The wing‘s drag during launch time:
Tension
Drag
Wings’ Folding Mechanism
Tension
Drag
Direction of flight
Drag
Wings’ Folding Mechanism
2
2
1= 19.26 [ ]
21
= 13.18[ ]2
wing ref
canard ref
D
D
D V N
D V
S C
NS C
The wing‘s drag during launch time:
Movement limiters Main spring
Connecting rods
Bearing
Wings’ Folding Mechanism - Other Related Parts Design
Booster for launch
Motor & Propeller
Integral fuel tank
WarheadEO Sensor
Wing & Canard Opening mechanism
Avionics
Internal Layout
S/N Part name Mass[gr] X[cm] My[N*cm]
1 Structure Fuselage 15000 115 16922.25
2 Wings 6740 145 9587.313
3 Mechanics –wing 3500 145 4978.575
4 Reinforcements- wing 2000 145 2844.9
5 Canard wings 4060 55.5 2210.4873
6 Mechanics-canard 2000 55.5 1088.91
7 Reinforcements- canard 1000 55.5 544.455
8 Tail 2000 185 3629.7
9 Fuel injection system 3000 130 3825.9
10 Engine Engine 6800 195 13008.06
11 Fuel 15000 107.48 15313.901
12 Fuel tank 1000 107.48 1020.9267
13 Oil 1000 110 1079.1
14 Warhead Warhead 20000 78 12556.8
15 Payload Sensor 8000 30.2 2370.096
16 Battery 5000 180 8829
17 Avionics Computer+Control system 2000 17.5 343.35
Total Mass: 98100
CGx[cm]: 107.48
Weight and Balance
Cg=107.5cm
Cp=111.2cm
Weight and Balance - C.G Location
10% chord stability margin
General instructions:
Max. length: 100 cm. Max. section area: 2-4% of cell’s section area. Wing tips should be away from the cell’s walls. Model shouldn’t be too small in order to get accurate results.
The model’s scale will be 1:7.
Wind Tunnel Model Design
Wind Tunnel Model Design
Steel reinforceme
nt
Hinge
Morse cone
CanardWing
Drawn nose
Steel holder
Wind Tunnel Model Design
Although in order to keep similarityrules, we had to use an air speed that is greater than 80 m/sec for the experiment, we wanted to avoid the situation in which model’s wings can’t handle the lift loads so we lowered the air speed to 45 m/sec .
Wind Tunnel Model Design
Experiment purpose:
Find the 2D lift coefficient slope, stall angle, pitch and yaw moments coefficients.Learn about UAV’s stability status.Learn about situations where a flow separation may occur.
No.Experiment
CodeConfiguration Plane
AOA Range [Degrees]
Air Speed [m/sec]
1 7369 Open Longitudinal -16-16 45
2 7373 Open Lateral -20-20 45
3 7376 Closed Longitudinal -20-20 45
4 7377 Closed Lateral -20-20 45
5 7381 Open Longitudinal -20-20 45
6 7383 Open wing only - With tufts Longitudinal -20-20 45
7 7385 Fuselage only Longitudinal -20-20 45
The different configurations:
Open
Closed
Open with tufts
Open, wings only with tufts
Fuselage only
Wind Tunnel Model Design
Wind Tunnel Model
Wing Tunnel Test ResultsLift Coefficient for Open Configuration
The graph above demonstrates the tunnel results of the total lift coefficient as function of angle of attack in comparison to the calculated theoretical lift coefficient.
-20 -15 -10 -5 0 5 10 15 20-1.5
-1
-0.5
0
0.5
1
1.5
2
2.5
3
X: 18.32Y: 2.667
[deg]
CL
CL for Open Configuration at 45 m/s
X: -6.891Y: -0.01198
Whole Aircraft - Tested
Whole Aircraft - Theoretical Calculations
Wing Tunnel Test ResultsLift Coefficient for Open Configuration
The graph above demonstrates the Lift coefficient as function of angle of attack of each lift-generator part of the UAV.
-20 -15 -10 -5 0 5 10 15 20-1
-0.5
0
0.5
1
1.5
2
2.5
3
X: 0.1681Y: 0.005301
[deg]
CL
CL for Open Configuration at 45 m/s
Whole Aircraft - TestedWings+Fuselage - TestedFuselage Only - TestedWings only - Est.Canards Only - Est.
Wing Tunnel Test ResultsMoment Coefficient for Open Configuration
The graph above demonstrates the tunnel results of the total lift coefficient as function of lift coefficient in comparison to the calculated theoretical lift coefficient.
-1.5 -1 -0.5 0 0.5 1 1.5 2 2.5 3-1
-0.8
-0.6
-0.4
-0.2
0
0.2
0.4
CL
CM
CM
as function of CL for Open Configuration at 45 m/s
Whole Aircraft - Tested
Whole Aircraft - Theoretical Calculations
Wing Tunnel Test ResultsMoment Coefficient for Open Configuration
The graph above demonstrates the Moment coefficient as function of angle of attack of each lift-generator part of the UAV.
-20 -15 -10 -5 0 5 10 15 20-2
-1.5
-1
-0.5
0
0.5
1
[deg]
CM
CM for Open Configuration at 45 m/s
Whole Aircraft - TestedWings+Fuselage - TestedFuselage Only - TestedWings only - Est.Canards Only - Est.
Wing Tunnel Test ResultsDrag Coefficient for Open Configuration
The graph above demonstrates the tunnel results of the drag coefficient as function of angle of attack in comparison to the calculated theoretical drag coefficient.
-20 -15 -10 -5 0 5 10 15 200
0.05
0.1
0.15
0.2
0.25
0.3
0.35
0.4
0.45
X: -5.592Y: 0.02025
[deg]
CD
CD as function of AOA
X: -5.592Y: 0.06562
TheoreticalExperimantal
Wing Tunnel Test ResultsLift Coefficient for Closed Configuration
The graph above demonstrates the tunnel results of the total lift coefficient as function of angle of attack in comparison to the calculated theoretical lift coefficient.
-20 -15 -10 -5 0 5 10 15 20-1.5
-1
-0.5
0
0.5
1
1.5
2
2.5
3
[deg]
CL
CL for closed Configuration at 45 m/s
Whole Aircraft - Tested
Whole Aircraft - Theoretical Calculations
Wing Tunnel Test ResultsMoment Coefficient for Closed Configuration
The graph above demonstrates the tunnel results of the total lift coefficient as function of angle of attack in comparison to the calculated theoretical lift coefficient.
-1.5 -1 -0.5 0 0.5 1 1.5 2 2.5 3-1
-0.5
0
0.5
1
1.5
CL
CM
CM
as function of CL for closed Configuration at 45 m/s
Whole Aircraft - Tested
Whole Aircraft - Theoretical Calculations
1. Improving geometry.
2. Performances estimation.
3. Wing design.
4. Wing & canard opening system.
5. Structural analysis.
6. Building a wind tunnel model or airplane model.
And more
Work Plan for Current Semester – as Seen on PDR
1. Strengthening the wind tunnel model.
2. Perform additional wind tunnel test on
the wings and canards in order to
evaluate their mutual affect on each
other.
Conclusions and Recommendations
Questions?