aircraft design project
DESCRIPTION
important for Aerospace studentsTRANSCRIPT
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AIRCRAFT DESIGN PROJECT
PRABHJOT KAUR DHAWANREG. NO- 1191210067
AER0-B
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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.
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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
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RESULT:The required parameters have been estimated from the graphs and survey has been done for the Propeller aircraft.
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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)• • •
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CALCULATION:• R=1700 (from literature survey)• Ƞp = 0.8•
• • = (1-0.912)*(1.06) =0.092• • =2178.5kg
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• RESULT: The maximum take-off weight is calculated both theoretically and from literature survey:
Calculated theoretical value=2178.5kg From literature survey=1900kg
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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:
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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
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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
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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
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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
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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
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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
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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
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• 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
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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)•
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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)
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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
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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
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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
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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
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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
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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
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RESULT:Dimensions of fuselage has been calculated through various parameters.
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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
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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
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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)
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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 .
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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
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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
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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:
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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:
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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