iclean - loitering attack uav cdr june 27 th , 2012 aerospace faculty, technion, haifa

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:


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


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:



Shear StressVon Mises

Material Max. Deformation]mm[

Aluminum6

Carbon 2.3

722.6

27.6allowed carbon fiber

allowed aluminum

Factor Calculation:


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


,

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


Tension

Drag

Direction of flight

Drag


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.



Steel reinforceme

nt

Hinge

Morse cone

CanardWing

Drawn nose

Steel holder


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 .


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

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



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


-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



Wing Tunnel Test ResultsMoment Coefficient for Closed Configuration


-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



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?

iclean - loitering attack uav cdr june 27 th , 2012 aerospace faculty, technion, haifa

Documents