AIRCRAFT DESIGN PROJECT

PRABHJOT KAUR DHAWANREG. NO- 1191210067

AER0-B

REPORT-1

AIM:. The data of existing 2-10 seater propeller aircraft are to be collected and values are to be obtained in report.

MEAN VALUES: The mean values which are obtained using different parameters are.

MEAN VALUES

Parameter Average

Seating capacity 4

Length 9.2mWingspan 2.8m

Chord length 11.8m

Cruising altitude 1.5m

Service ceiling 2800m

Range 5000m

Maximum speed 290km/hrPower 300 HP

Max. Take-off weight 1900kgEmpty weight 1200kg

Payload 760kgMaximum fuel capacity 0.42m^3

RESULT:The required parameters have been estimated from the graphs and survey has been done for the Propeller aircraft.

REPORT-2

AIM: The aim of this report is to find accurate estimation of weight of aircraft required to design it.FORMULA:• Wo= maximum take-off weight• We= empty weight• R=1700 (from literature survey)• • •

CALCULATION:• R=1700 (from literature survey)• Ƞp = 0.8•

• • = (1-0.912)*(1.06) =0.092• • =2178.5kg

• RESULT: The maximum take-off weight is calculated both theoretically and from literature survey:

Calculated theoretical value=2178.5kg From literature survey=1900kg

REPORT-3

AIM: To study the nomenclature of airfoil and to find the value of Cd & Cl max for the airfoil at fixed Reynolds number.

FORMULA USED:

CALCULATION:• =21072.9N• Vcr= 280 km/hr = 77.77m/sec• b= 11.8m• s= • c= = 1.4m• Cruising altitude= 2800m• ρ=0.9• µ=• CL= 0.44/9 = 0.5• • 0.6• 1.6

RESULT: Airfoil selected at mean position is NACA 2414• Airfoil selected at root position is NACA 2415• Airfoil selected at tip position is NACA 2413

Root Mean TipAirfoil selected NACA 2415 NACA 2414 NACA 2413

Cl 0.6 0.6 0.6Clmax 1.6 1.6 1.6

REPORT-4AIM:To estimate drag and lift for a given airfoil at different conditions.FORMULA USED:At cruise:• Cd=Cdo+kCl2

• Cdo=n x cfe

At Take off:• (Clmax)T.O=(Clmax)cr + ( Δ Clmax)flap

• (Cd)T.O=Cdo + k(Clmax)T.O

• Vsea=0.7x1.2xVstall

At Landing:• (Clmax)land=(Clmax)cr + ( Δ Clmax)land

• (Cd)LAND=Cdo + k(Clmax)LAND

• Vsea=0.7x1.3xVstall

CALCULATION:At cruise:• Lcr = W2 = 2148.5kg = 2148.5*9.81 = 21.37kN• CLcr = 0.5

• k= 0.047• = 3.3*10-3

• CD0 = 4*3.3*10-3 = 0.0132• CD= 0.0132+0.047+ (0.5)2

• =0.025• D=1.3kN

At take off:• (Clmax)T.O= (Clmax) cr + ( Δ Clmax)flap

• = (0.9*1.6) +0.7• =2.14• = 41.34m/sec• Vsea=34.72m/sec• = 3.5*106

• = 34.72/(1.4*287*288.15)^0.5 = 0.105• Cfe=3.57*10-3

• CDO= 4*3.57*10-3 = 0.229• D= 0.5*1.225*(34.72)2*19.32*0.229 = 3.2kN• L=0.5*1.225*(34.72)2*19.32*2.14 = 30.52kN

At landing:• (Clmax)land=(Clmax)cr + ( Δ Clmax)land

• = 0.9*0.6+0.9 = 2.34• Lland= 0.5*1.225*(37.61)2*2.34 = 39.168kN• Vsea=0.7*.3*41.34 = 37.61m/sec• =1.225*37.61*1.4/(1.7*10-5) = 3.7*106

• = 37.61/(1.4*287*288.15)^0.5 = 0.11• Cfe= 0.455/128.5 = 3.54*10-3

• CDO= 4*3.54*10-3 = 0.0141• CD= 0.0141+0.047(2.34)2

• = 0.271• D= 0.5*1.225*(37.61)2*19.32*0.271 = 4.5kN

• RESULT:

TAKEOFF CRUISE LANDING

Cl 0.5 2.14 2.34

Cd 0.025 0.229 0.271

L 21.37kN 30.52kN 39.168kN

D 1.3kN 3.2kN 4.5kN

REPORT-5AIM:To design a wing, horizontal tail, vertical tail and control surface by using mean parameters from literature survey.FORMULA USED:Fuselage wing:

Vertical tail:

• CtVT= λvtx (CrVT)•

Horizontal tail:• • CtHT= λHT* (CrHT)

• YHT=bHT/6*(1+2 λHT)/(1+λHT) • bHT = [SHT*(A.R)HT]1/2

Control surfaces:• AILERON: ba=80%bw and Ca=20%Cw• RUDDER: bru=80%bVT and Cru=20%CVT

• ELEVATOR: be=80%bHT and Ce=20%CHT • VVT= (LVT*SVT)/ (bw*Sw) • VHT= (LHT*SHT)/ (bw*Sw)

CALCULATION:Wing:• Sw=19.32• bw=11.8• cw= 1.4• A.R=8.4• λw=0.5• • = = 2.18m• Ct= λwxCr = 2.18x0.5 = 1.091m• = 11.8/6[(1+2x0.5)/(1+0.5)] = 2.61m

Vertical tail:• = 1.652m• bvt= (1.87x1.4)0.5 = 1.61m• CrVT= = 1.522m• Y= = = 0.268x1.259 = 0.337• CVT = = = 1.106m• Ctip VT = λVT x CrVT = 0.35 x 1.522 = 0.537m

Horizontal tail:• λht=0.3• A.R=4• Lf=9.2m• LHT=0.6xLf= 0.6x9.2 =5.52m• Cw=1.4• SHT=3.43m^2• A.RHT=4• bHT=(SHTxA.RHT)0.5= (3.43x4)0.5=3.70m• CrHT=1.426m• CtHT= λhtxCrHT = 0.3x1.426 = 0.4278m• Y= = = 0.758• CHT = = = 1.016m

Control surfaces:• AILERON: ba=50% of bw = 50% of 11.8=5.9m and Ca=15% of

Cw=15% of 1.4 m=0.21 m• RUDDER: bru=50% of bVT =50% of 1.61 =0.805 m and Cru=25% of

CVT =25% of 1.106 =0.276m• ELEVATOR: be=50% of bHT =50% of 3.70=1.85 m and Ce=25% of CHT

=25% of 1.016=0.254 m

RESULT:

Parameter Wing(m)

H.T(m)

V.T(m)

Cr 2.18 1.426 1.522Ct 1.091 0.4278 0.537Y 2.61 0.758 0.337S 19.32 3.43 1.652b 11.8 3.70 1.61

Control Surface b(m)

C(m)

Aileron 5.9 0.21Rudder 0.805 0.276Elevator 1.85 0.254

REPORT-6AIM:To design a fuselage for aircraftCALCULATION:• Lf= 9.2m• 6=9.2/6 = 1.53m• 3= (Lf-c)/df

• 3x1.53 = Lf-c

• 4.59m = Lf-c

• Θfc = 7 degree• Seat height = 3.5ft = 1.06m• Height= 1.524m• Aisle width=2=0.6096m• Seat pitch= 2.5 = 0.762m• Seat width= 1.5 = 0.4572m• Lcb=4x Seat pitch = 4x0.762 = 3.04m

RESULT:Dimensions of fuselage has been calculated through various parameters.

REPORT-7AIM:: To select an engine by calculating various parameters given.FORMULA USED:

• Pa= ȠPR x P•

• 2 blade,d = 22 x (H.P)1/4

CALCULATION:• hOb=15.2m• STO= 650m• SG= 391.5m• = = 0.071• TTO=154.69kg • TTO= 1517.5N• = 1.57 x 34.72/0.8 = 66.17 HP• Cruise:• = 1/ 16.43 = 0.06• Tcr = 0.06 x W2 = 0.06 x 1952.86 x 9.81 = 1.4kN• 1.14 x 77.77/ 0.8 = 110.8kW = 148.4HP • 2 blade,d = = 1.95m• Vtip = Πx n x d = 3.14 x 2500/60 x 1.95 = 255.125m/sec• Vcr= [(255.125)2+(77.77)2]0.5 = 266.715m/sec

RESULT:• : From the measured parameters Lycoming O-320 engine has been

selected.• Specifications:• Type: Four cylinder air- cooled horizontally opposite engine• Box: 5.125in (130.18)• Stroke: 3.875 (98.43)• Power output: 150HP• Compression ratio: 7:1• Power to weight ratio: 1.63lb/hp (0.99kW/kg)

REPORT-8AIM: To find parameters/ dimensions of landing gear.FORMULA USED:• At nose:• Diameter of nose (d) =Ad (Ww) Bd

.

• Where, Ww=0.1xWo/(no. of wheels)• width=Awi(Ww)Bwi .• At main landing gear:• Diameter of landing gear (d) =Ad (Ww) Bd

• Where, Ww=0.9xWo/2• width=Awi(Ww)Bwi .

CALCULATION:• Main tyre(Ww) = 2160.6lbs• Nose tyre (Ww)= 0.1 x Wo x 2.204 = 0.1 x 2178.5 x 2.204 = 480.14 lbs• Dia: Main tyre = A[Ww]B = 1.59[2160.6]0.349 = 23.18 inches• Nose tyre = A[Ww]B = 1.59[480.14]0.349 = 13.71 inches • Width: Main tyre= A[Ww]B = 0.7150[2160.6]0.312 = 7.84 inches • Nose tyre = A[Ww]B = = 0.7150[480.14]0.312 = 4.9 inches

RESULT: The parameters of tricycle type landing gear are measured as follows:

Main tyre Nose tyre

Ww 2160.6 lbs 480.14 lbs

Diameter 23.18 inches 13.71 inches

Width 7.84 inches 4.9 inches

REPORT-9AIM:To determine the performance characteristics and parameters of given propeller aircraft.FORMULA USED:• Take off:• • Vto= 1.2 x VStall = 1.2 x 41.34 = 49.608m/sec• Climbing:

• Level turn:• • • Turn radius:

• Gliding:

• Landing:

CALCULATION:• Take off:• = 927.4m• SG = 927.4m• Sto = 1.6 SG = 1539.4m• Vto= 1.2 x VStall = 1.2 x 41.34 = 49.608m/sec• Climbing:• = = 38.8m/sec =(R/C)max

• Level turn:• • = (77.77)2/tan(20)x 9.81 = 1691.11m• Turn radius:• θ= 20 degree• = 1.42• • • Gliding:• degree is glide angle• Landing:

RESULT:Performance characteristics are determined for propeller aircraft.• Take off: Sto = 927.4m V to= 49.608m• Climbing: R/C = 38.8m/sec• Level turn: R = 478m• Turn radius: R = 172.8m w= 0.23rad/sec• Gliding: Φ = 3.43 degree• Landing: Sto= 1147m