MI -5 M.KARTHICK

J. SHIVAKUMAR

MANISH KUMAR SINGH

NEEL HARSH PATEL

NAGENDRA KUMAR

AERODYNAMIC DESIGN – AS5210GROUP – M5

Historical survey

RC Aircraft Model No.

Wing span (mm)

Flying weight (gm)

Fuselage length (mm)

Motor(Brushless)

Catalina 1800 2700 1097 35-XX outrunner

EP-Cessna 206 1080 1190 980 1750 KVF.K. 10 950 1100 760 3536/1000 KVCessna 182 (EPO) 1410 1100 1030 3647/700 KVP-40 War hawk 1700 3500 1500 50-60 out runner/250 KV

V- tail V35 1280 1348 1000 3720/980 kVHobby 2280 4650 2015 70 CC gasoline engine

P-47Thunderbolt

1600 3000 1416 4258/400 KV

J3 Cub 1400 700 950 36-48Outrunner/700 KV

MI5 1000-1500 2000-3000 1000-1500 To be decided

Our Aircraft Specification

Range(Km) 0.8-1.0Thrust(N) To be determined

Endurance(min) 15-20Propeller Twin blade propeller

First Weight Estimation

W 0=W PL

1−(W PP

W 0)−(W E

W 0)

WPP = The weight of the power plant

WPL = The weight of the payload

WE = The empty weight

0 1000 2000 3000 4000 50000

0.050.1

0.150.2

0.250.3

0.350.4

0.450.5 Wpp/Wo vs Wo

Wo

Wpp

/Wo

500 1000150020002500300035004000450050000

0.1

0.2

0.3

0.4

0.5

0.6 Ws/Wo vs Wo

Wo

Ws/

Wo

Wo(gm) Wpp/Wo Ws/Wo Wo(gm) Calculated

2500(first initial

guess)0.324 0.445 1298.7

1298.7 0.300 0.444 1171.8

1171.8 0.300 0.444 1170

Airfoil, Wing Design and Second Weight Estimation

• Selection of airfoil is the most important design aspect of an airplane.• MI5 should fly at low altitudes and low velocity.• Low value for the stall speed and high maximum lift coefficient (high value).

5 10 15 20 25 30 350

102030405060708090

f(x) = 0.938473403135135 x + 42.528354687674R² = 0.328554708948436

W/S vs Wo

Wo (N)

W/S

(N/m

2 )

for Wo = 1170gm = 11.495 N

W/S = 53.3 N/m2

• The flying altitude is chosen : 25m• Designed cruising speed : 10

m/sec• Aspect ratio (AR) : 6.5

Span =

Airfoil Zero lift Angle of Attack

CLmax αstall Leading edge radius (%c)

NACA 0014 0 1.36 14.6 47 2.16NACA 2410 -2.1 1.46 14.6 42 1.10NACA 2412 -2.1 1.41 14.7 40.4 1.59NACA 6412 -5.5 1.51 14.8 36.8 1.59NACA 6414 -5.8 1.44 14.7 27.8 2.16

stalling angle range of 13 - 15

-0.04-0.02

00.020.040.060.08

1.0

Airfoil, Wing Design and Second Weight Estimation cont…

: 1.41

: 14.6

Aerodynamic centre : 0.265 c

Zero Lift angle : 2.1

() : 10 m/s.

NACA 2410

-10 -5 0 5 10 15 20

-0.5

0

0.5

1

1.5

2

Cl vs

Lift

Coe

ffici

ent

0.01 0.015 0.02 0.025 0.03 0.035 0.04 0.045

-0.5

0

0.5

1

1.5

2

Cl vs Cd

Drag Coefficient Cd

Lift

Coe

ffici

ent C

l

Airfoil, Wing Design and Second Weight Estimation cont…

𝐶𝑑=𝐶𝑑𝑜+𝑘𝐶𝑙2=¿0.0163+0.06121𝐶𝑙

2 k=1

π (e=Oswald ′ seff factor=0.8) AR=0.06121

𝐶𝑑𝑜=𝑘𝐶𝑙2 Cd=0.0326L/ Dmax=15.83

Wing CharacteristicsAspect ratio : 6.5Wing span : 1.184 mChord : 18.21 cmTaper : No taperLeading edge sweep : No sweep Wing incidence : 1

Performance Characteristics: 7.615 m/s : 10 m/s: 10 m/s

Climb angle : 10Climb rate : 1.75 m/sRange : 1 KmEndurance : 15-20 minAltitude : 25 mCeiling : 100m

Psteady = Tsteady × V = = 0.8286 × 10 = 8.286 Watt

Pclimb = Tclimb × V = = 2.573 10 = 25.73 Watt

Dmin=W

L /Dmax

=0.577 N

Airfoil, Wing Design and Second Weight Estimation cont…

Battery specificationKingMax3000mah 35C 22.2VSize : 115mm x 35.5mm x 52.8mm Net Weight : 428g Avionic 8g Servo (AV8A)Wire length : 18cm Weight : 8 gmMotor specification:Odin 2730 KV1300rpmWeight : 25 gmThrust : 440 gm Propeller specification:Prop - 11 x 6 cm

0 1000 2000 3000 4000 50000

0.1

0.2

0.3

0.4

0.5

0.6

f(x) = 1.36557075444183E-06 x + 0.44387231772415R² = 0.000831252632343071

Structural Weight / Gross Weight (Ws/Wo) – vs –

Gross Weight (Wo)

Wo in gm

Ws/

Wo

Wo(gm) Ws/Wo Wo(gm) Calculated

1170 0.444 1429.81429.8 0.434 1466.43

1466.43 0.444 14941494 0.444 1494

Wing Loading And Thrust to Weight Ratio

Approach (Sa) Flare (Sf) Ground Roll (Sg) = 30.114 m

50 m

R =

= 2

9.80

9 m

Vf = 1.23 V

stall

n = load factor = 1.3

a = 30

hf = R (1-cos a ) = 0.04085 m

~

1 m

Sf = R sin Sa = =18.326 m

Sg = j N +

Sg = 31.15 +

(W/S)Landing= 37.47 N/m2

Landing

Wing Loading And Thrust to Weight Ratio cont…

= = 0.51.222102 = 31.53 N/m2

= 36.23 N/m23 CD 0=k CL

2

= = 34.55 N/m2

Bank Angle, = 60o

Load Factor, n = sec = 2n = ( ) ( )max.( ) = 0.1133.

= 31.53 N/m2

Wing Loading And Thrust to Weight Ratio cont…

( TW )

cruise

= 1

( LD )

max

= 115.83

=𝟎 .𝟎𝟔𝟑𝟏𝟕

Take off(Sa)Ground Run (Sg) = 15.98 m

25 m

R = (6.96 × Vstall2)/g = 41.14 m

hOB = 1m

θOB = cos-1(1 – hOB/R) = 12.65o

Sa = R sin(θOB) = 9.015 m

Sg=

( TW )

Takeoff

=0.1408

= ) = +

= = 0.1579

T = 3.47 N

Initial Sizing and LayoutSpar

Wing

Y

Z

X

Z

Spar made of Balsa Wood (160 Kg/m3)

b/2 = 0.870m C = 0.267m

h =

26.7

mm

Lift DistributionLoad (L) = W/S * C = 31.45 * 0.267 = 8.405 N/m

Spar

t (thickness of Spar)

σ allowable=σ compressiveYield Strength of Balsa Wood

Factor of safety ¿literature ¿=

12.1 Mpamedium density

3=4.03 Mpa

I xx=M root × Zmax

σ allow

=3181 N −mm× 13.35mm4.03 N /mm2 =10537.56mm4= t × h3

12

Thickness of the Spar (t )=I xx ×12

h3 =10537.56 ×1226.73 =6.64 mm≈ 7 mm

Weight of Spar (mspar) = 52.003 gm

Initial Sizing and Layout cont…Ribs

Wing

267 mm

26.7mmX

ZArea = 4887.571mm2

Balsa Wood Number of Ribs = 20

Thickness of the Rib = 10mm = 0.01m

Weight of 20 Ribs (mRib) = 156.4 gm

Skin

Area = 944021mm2

1739mm

267mm

Foam (41 Kg/m3)

Thickness = 1mm

Weight of the Skin (mwing skin) = 37.76 gm

mWing= 245.80 gm

Initial Sizing and Layout cont…

Component Thumb Rule

Horizontal Tail 25 – 30 % Wing Surface Area

Vertical Tail 35 % of Stabilizer Area

Fuselage Length 75 % of Wing Span

Nose Length 25 % of the fuselage length

Wing trailing edge to stabilizer 40 % of the fuselage length

Horizontal tailmskin, hor tail= 11.61 gm

mRibs, hor tail= 10 nos X 3.91 = 39.1 gm

NACA 0010 134mm

mspar, hor tail= 15.60 gm

Spar X – n Area = 93.45 mm2

mhor tail= 66.31 g.Vertical tailmskin, Ver tail= 4.06 gm

mRibs, Ver tail= 5 nos X 3.91 = 19.55 gm

mspar, Ver tail= 5.50 gm

mver tail= 29.11 gm

Fuselage

60mm

60m

m

5mm

Initial Sizing and Layout cont…

mmain fuselage= 229.5gms

Fuselage Length = 1304 mm

35mm

35m

m

5mm

Nose Fuselage Length = 326.1 mm

mNose fuselage= 31.31gms

mfuselage = 260.8 gm

Structural components Weight(g)Fuselage 260.8Wing 245.8Horizontal tail 66.31Vertical tail 29.11Power plant 453Payload 100Servo controller + receiver 16Structural attachments ,glue etc. 30Landing gear 23

mstructure= 602.2gm.

Initial Aircraft Modelling

Initial Aircraft Modelling cont…

Initial Aircraft Modelling cont…

Initial Aircraft Modelling cont…

Initial Aircraft Modelling cont…

Propeller Design and Initial sizing APC 10x7 Thin Electric Propeller

Diameter = 22*(HP)1/4

P= 43.16.16 wattD=22*(43.16/746)1/4 D=10.78 inches Pitch ‘p’ = 7 inches

CONDITIONS VELOCITY (M/S)

Advance Ratio

J =V/(n*D)

THRUSTT= Ct × ρ × n2

× D4

(N)

POWERP = Cp × ρ ×

n3 × D5 (WATTS)

EFFICIENCYηp =JxCt/Cp

SPEED COEFFICIENT

Cs =J/(Cp)0.2

TAKE OFF 8.757 0.479 2.139 25.78 72.89 0.957

CLIMB 10 0.547 1.956 24.66 79.56 1.02

CRUISE 10 0.547 1.956 24.66 79.56 1.02

TURNING 14.14 0.774 1.196 19.06 88.78 1.522

Drag Polar Estimation

CD 0=∑C f ,(component Skin fric c oeff )× F orm F actorc omponent × Scomponent wet tesd area

Sref (wing area)

= d eq=√ (4 Amax ( X− n))

π =

= =

371999 mm2

899293.3 mm 2

275371.1 mm 2

103932.5 mm2CD 0 , parasite=0.00958

TotalC DO= (CDO )PARA+(C DO )MISC+(C DO )L∧P

= 0

CD=0.01276+0.0612 CL2

Drag Polar Estimation cont…

Thrust adequacy

-2 -1.5 -1 -0.5 0 0.5 1 1.5 20

0.050.1

0.150.2

0.250.3

Drag Polar

CL

CD

CD=0.01276+0.0612 CL2

Maneuver

Thrust required (N) Power required (W)

As perFirst T/W estimate

New ValuesAs per

First T/W estimate

New Values

Cruise 0.925 0.806 9.25 8.06Climb (θ = 10 ) 3.47 3.365 34.7 33.65Turning (n = ) 2.314 1.876 32.725 26.53

Take-off 2.58 2.22 22.59 19.44

CENTER OF GRAVITY LOCATION

Xcg =

Zcg =

Ycg =

Components Weight(gms)

X C.G(mm)

Y C.G(mm)

Z C.G(mm)

Wing 284 425.95 0 76.73Fuselage 289 592.64 0 53.76Hor Tail 96 1253.10 0 82.45Ver Tail 15 1243.04 0 264.94Propeller 22 33.97 0 20.36Landing Gear 23 652.69 0 89.80Battery 427 357.50 0 31.40Aircraft 1156 518.68 0 56.35

X

Y

Z

XC.G = 518.68 mmYC.G = 5.4e-5 mmZC.G = 56.35 mm

Stability and Trim Analysis

X np=CLα X acw− Cmαfus+ηh

Sh

Sw

CLαh

𝜕 αh

𝜕αX ach

CLα+ηh

Sh

Sw

CLαh

𝜕α h

𝜕 α

=4.7818 per radian

= 5.0134 per radian

=

acw = 0.25

906.9343mm

66.75mm

ηh=ηhT=0 ×(1+T cruise

q × A p)=0.9×(1+ 0.719

61 ×0.0508 )=1.1

X np=0.85787

Stability and Trim Analysis cont…

Cmcg=C Lα α ( Xcg − X acw )+Cmw

+Cm fus+Cmwδf

δ f − ηh

Sh

Sw

CLh( X ach − X cg )

• is elevator deflection

• For static trim condition the total pitching moment must be equal to zero.

Cmcg=− 2.0604α − 4.411δ f − 0.053 CL ,Total=6.436α +1.654 δ f

-4 -2 0 2 4 6 8 10 12

-1.5

-1

-0.5

0

0.5

1Chart Title

Wing angle of attack (α)

Cm-c

g

δ f =− 10o

δ f =− 5o

δ f =0o

δ f =5 o

δ f =10 o

Zero lift drag of landing gear

CD 0lg=∑

i=1

n= 3

CD lg( Slg i

Swing)Landing gear weight = 23 gm

Component CDo

Front Wheel 0.00529

Front Wheel Strut 0.00129

Rear Wheels 6.706 x 10-5

Rear Wheels Strut 3.876 x 10 -5

Total 0.00668

CD=

CD = 0.01276+0.0612CL2

V-n Diagram V st all=√ 2 ×Wρ × s×CLmax

=5.6 m /s N/

-6 -1 4 9 14 19 24

-1.5

-1

-0.5

0

0.5

1

1.5

2

2.5

3

3.5 V – n diagram

V (m/s)

n

n = (Clmax×0.5×ρ×V2×S) / W = 0.03136×V2

n = - 0.00132×V2

V B=2× ( nmax=3 ) ×W

( ρ ×CL , max × S )0.5=9.8 m /s

A

BC

D

VC = 21.42 m/s

P= Cd0×0.5×ρ×S×Vmax3+K×W2/(0.5×ρ×S×Vmax)

VC = 1.3×Vmax = 21.42 m/s