Conceptual Design Of 180-Seater Passenger Aircraft

November 2017 IJIRT Volume 4 Issue 6 ISSN: 2349-60026.0 POWER-PLANT SELECTIONAgain, from the historical and comparative data,the engine is selected. The aircraft will require thethrust about 220 kN. To achieve the desired thrust,two high-bypass turbofan engines are chosen, eachproducing a minimum thrust of 110 kN. Eachengine is mounted below the wing by the help ofpylons, at 6.78 m from the aircraft centerline. Thetype of engine which would meet the designstandards will be IAE-V2527-A5.Table 2: ENGINE DATATypeIAE V-2527-A5Thrust118.32 kNBypass ratio4.8:1Compression ratio32.8:1Fan diameter1.613 mLength of engine3.2 mWeight2,359 kg7.0 WING GEOMETRYThe geometry of the wing is mainly described byits planform shape, its aspect ratio, the type ofairfoil at wing root and tip, the thickness of wingalong the span, wing sweep angle, taper ratio, andgeometric twist. The taper ratio is obtained fromreference-1, and it is found to be 0.24. Othergeometric aspects like the wingspan, wing area,and aspect ratio are found from the analyticalresults of historical data, and they are found to be34.09 m, 122.6 m2 , and 9.47, respectively.From these parameters, wing root chord andwingtip chord is calculated by using Eqn-IX. Also,length of mean aerodynamic chord and its relativedistance from the aircraft centerline is calculated byusing Eqns-X, and XII, respectively.Table 3: WING GEOMETRIC DATARoot chord length5.8 mTip chord length1.392 mMean aerodynamic chord length4.046 mDistance of mean aerodynamic chord 6.78 mfrom aircraft center lineinterference drag and gives the best stability if thewing is dihedral. Whereas the low wing aircraft aremore stable than mid-wing aircraft, but as stable ashigh-wing aircraft. But, low-wing configuration ismostly employed because of its structuraladvantage and ease of landing gear retraction intothe wing box.Selection of airfoil is the key phase in the winggeometry. The airfoil should not only reduce theform drag but also should provide enough pressuredistribution around its surface to contribute in theA software quality is defined based on the study ofexternal and internal features of the software. Theexternal quality is defined based on how softwareperforms in real time circumstances in operationalmode and how it is useful for its users. The internalquality focuses on the essential aspects that arereliant on the quality of the code which isdeveloped. The user concentrates more on thesoftware how it works at the external level, but thequality at external level can be maintained only ifthe coder has written a meaningful and goodquality code. The quality system encompassesdifferent activities like Staff development ofpersonnel employed within the quality area. Thedevelopment of standards procedure andguidelines.generation of lift. The airfoil which would matchthe design requirements is NACA 66(4)-221. Theadvantage of choosing NACA 6-digit airfoil is thatit maintains laminar flow over the large part of thechord and they maintain CD(min) compared with 4 or5-digit airfoils.Figure 4: NACA 66(4)-221 airfoil.8.0 FUSELAGE SIZINGFigure 3: Wing geometryThe location of the wing plays a key role in thestability of the airplane. Aircraft having high-wingare comparatively more stable in lateral and rollingmotion. While the aircraft with mid-wing have lessIJIRT 144888The role of the fuselage is to hold the other parts ofaircraft. Its major function is to provide overallstructural integrity and house for the payload. Thefuselage is also expected to maintain thepressurization inside the cabin and the cockpit. Thefuselage is divided into sections, such as, nose,cockpit, cabin, and tail fuselage. The cockpit iswhere all the controllers, navigational systems, etc.,are placed. The cabin houses the payload. Themajor geometric parameters that decide the cabinare cabin diameter and its length. These in turn aredecided by more specific details like number ofseats, the width of each seat, the arrangement ofINTERNATIONAL JO URNAL OF INNOVATIVE RESEARCH IN TECHNOLOGY76

November 2017 IJIRT Volume 4 Issue 6 ISSN: 2349-6002seats, pitch of the seat, aisle width, and lastly, thenumber of aisles. The typical seating abreast is 6with one aisle. The number of seats across will fixthe number if rows in the cabin, and thereby thecabin length.Figure 5: Fuselage cross sectionThe length of the fuselage is determined by usingthe Eqn-XII. Also, the geometric parameters of thefuselage, as stated above are calculated using theEqn XII, XIV, XV, and XVI. The profile of the rearfuselage is chosen to provide a smooth (lesser drag)shape which also supports the horizontal stabilizersand tail. It houses the Auxiliary Power Unit (APU).The lower profile of the rear fuselage must provideadequate clearance for aircraft during take-off. Thelength of the tail is determined by using the EqnXVII.Table 4: FUSELAGE SIZING DATALength of fuselage, Lfus34.18 mLength of nose, Lnose1.024 mLength of cabin, Lcabin30.74 mInternal diameter of the fuselage, 3.38 md fus(internal)External diameter of the fuselage, 3.56 md fus(external)Structural thickness, tstructural0.093 mLength of the tail, Ltail8.188 mFigure 6: Fuselage sizing9.0 AERODYNAMIC CHARACTERISTICS9.1 LIFTLift is mainly due to the pressure distribution onthe surface of the wing. The amount of lift dependson the planform area, air density, velocity ofaircraft, and lift coefficient. This component ofaerodynamic force (lift) is generated on aircraft,IJIRT 144888perpendicular to the direction of flight. The liftcoefficient, CL , is a measure of lift effectivenessand mainly depends upon section share, the angleof attack, compressibility effects (Mach number),planform geometry, and viscous effects (Reynold‟snumber). The role of CL plays vital during lowspeeds, especially during landing and takeoff.During these maneuvers, high lift is desired, whichis obtained by increasing the CL . High lift devicesare installed on the surface of the wing to achievehigh CL . They change the net angle of attack, andincrease the surface area, thereby obtaining highlift.The value of lift keeps varying along the missionphases. Lift is high during takeoff and landing,comparatively lesser during climb and descendphases, and lift is equal to the weight of the aircraftduring the cruise, as the aircraft will neither gainnor lose altitude.The lift is estimated using the Eqn-XVIII,considering the changes in flight velocity, density,and CL , as the aircraft gains or loses altitude.Table 5: LIFT ESTIMATION DATALift DuringCLLift value 581383.619.2 DRAGThe component of aerodynamic force which actsopposite to the direction of flight, which reducesthe overall performance of the airplane, is calleddrag. There are many factors due to which the dragwould arise. Of those factors, the main contributorsare, skin friction drag- which arises due to shearstresses produced in the boundary layer, form dragwhich arises due to the shape of the body, and dueto the static pressure distribution around the surfaceof the body, and wave drag- which arises due to thepresence of shockwaves on wings and fuselage,propeller blade tips moving at transonic andsupersonic speeds.The amount of drag depends on air density,planform area, aircraft‟s velocity, and dragcoefficient. The value of drag can be estimated byusing Eqn-XIX. But, the value of drag is not onlydependent on above-mentioned parameters, it alsodepends on the alignment of aircraft with therelative head wing, cleanliness and surfaceroughness of aircraft skin, the maneuvering ofaircraft, the effect of angle of incidence, etc. Hence,the drag value is a variable quantity, and it can onlybe determined if all the dependent variables areknown.INTERNATIONAL JO URNAL OF INNOVATIVE RESEARCH IN TECHNOLOGY77

November 2017 IJIRT Volume 4 Issue 6 ISSN: 2349-600210.0 LANDING GEAR CONFIGURATIONThe basic type of landing gear used are- tandemlanding gear, tail-wheel type landing gear(conventional gear), and tricycle-type landing gear.In tandem landing gear, the main gear and tail gearare aligned to the longitudinal axis of the airplane.The aircraft which use this type of landing gear ismilitary bombers B-47 and the B-52, and fewgliders are configured with tandem landing gear. Intail wheel-type landing gear arrangement, the maingear is located forward of the c.g, causing the tail torequire support from the third wheel assembly.Maule MX-235 Super Rocket and McDonnellDouglas DC-3 use this type of landing geararrangement. The landing gear configuration whichfinds more application in civil aircraft is tricycletype landing gear. The advantage of using this typeof arrangement in civil aircraft is that it preventsground-looping of the aircraft, and it allows moreforceful application of the brakes without noisingover when braking, which is advantageous duringhigher landing speeds.The landing gear arrangement which meets thedesign standards is tri-cycle type arrangement, with4 wheels in the main landing gear (two on eachaxle) and 2 on the nose landing gear.The geometry of the wheels is calculated by usingEqn-20, by substituting the values of A and B asgiven in Table-6. The weight acting on each mainlanding wheel is 18,426 kg and that on noselanding wheel is 16,152 kg.Table 6: LANDING GEAR SIZING DATAABWheel diameter1.63 inch0.315 inchWheel width0.1043 inch0.48 inchTo determine the position of landing gear,aerodynamic and geometric parameters of the wingis considered. The static margin is assumed to be18% and the position of the neutral point (XNP ) isdetermined using Eqn-XXI. The horizontal tailvolume ratio and lift slope ratio of tail to wing istaken from the historical and analytical data, and itis found to be 0.253 and 4.545, respectively. Thelocation of the main landing gear is at the center ofthe wing.Table 7: LANDING GEAR WHEEL DATAMain landing Nose landinggeargearWheel0.91 m0.876 mdiameterWheel width0.29 m0.277 mIJIRT 144888Table 8: LANDING GEAR POSITION DATAPosition of neutral point from aircraft 19.14nose, XNPmPosition of aerodynamic center of wing, 17.98X(ac)wingmPosition of main landing gear from 19.27aircraft nosemPosition of nose landing gear from 3.95 maircraft nose11.0 RANGE AND ENDURANCE11.1 RANGEThe range is the total distance covered by theaircraft during its mission on one load fuel. Therange of an aircraft depends on the weight of fuel itis carrying, the engine‟s thrust specific fuelconsumption (TSFC), cruising altitude, wingplanform area, andratio. The flightconditions for maximum range for a jet-propelledairplane are- the aircraft should fly at a highaltitude where the free stream air density is small;the engine should have lowest possible TSFC; theaircraft should carry a lot of fuel, and the aircraftshould fly with a maximu m.The range is calculated using the Eqn-II, and it isfound to be 6025.10 km.11.2 ENDURANCEEndurance is the total time that an airplane can stayin the air on one load of fuel. The range of anaircraft depends on the weight of fuel, the engine‟sTSFC, and L/D ratio. For an aircraft to havemaximum endurance, it should carry a lot of fuel;the engine should have lowest possible TSFC, andthe aircraft should fly at maximum (L/D). Therange is calculated by using the Eqn-II, and it isfound to be 16.7 hours. But, practically, the aircraftwill not have such endurance. The aircraftbelonging to the design requirement categorypossess endurance of 5 to 7 hours. Since thecalculation is done considering the maximumdependent parameters, such a huge value ofendurance is achieved.12.0 TAKEOFF PERFORMANCEThe takeoff distance for an airplane is the totaldistance covered by the airplane, with respect to theground, from the point of release of brakes(mission point-2) to the point where the airplanecleared the obstacle height. The takeoff distance isprimarily divided into two phases - ground rolldistance, and distance while airborne to clear anobstacle. The obstacle height for the militaryairplane is 50 feet, and for civil aircraft is 35 feet.INTERNATIONAL JO URNAL OF INNOVATIVE RESEARCH IN TECHNOLOGY78

November 2017 IJIRT Volume 4 Issue 6 ISSN: 2349-6002The ground roll distance is the total distancecovered by the aircraft from the point of release ofbrakes to the point where the aircraft becomesairborne. During takeoff, the airplane would requirea greater lift, which is obtained by increasing thecoefficient of lift with the help of high lift devices.The ground roll is calculated by using the Eqn XXIII. During this phase, the flaps are extendedand kept at takeoff position of 20º, so that themaximum coefficient of lift will be increased to2.508.Figure 6: Illustration of ground roll (s g ), andairborne distance (sa )The flight path after liftoff is essentially pull-upmaneuver in the radius „r‟. The turn radius iscalculated using the Eqn-XXIV. The angle inducedby the flight path between the point of takeoff andthat for clearing the obstacle height is ϴOB. Duringthis phase, the airplane will cover a distance, S a,which is calculated by using the Eqn-XXV.Figure 7: Illustration of obstacle airborne distance.Table 9: TAKEOFF DISTANCE DATAElevation6.7 mDensity1.225 kg/m3Ground roll distance603.35 mAirborne distance336.27 mTotal takeoff distance939.62 mIJIRT 14488813.0 LANDING PERFORMANCEThe landing distance begins when the airplaneclears an obstacle. The landing distance is majorlyclassified into three phases - approach distance,flare distance, and ground roll. The distancemeasured along the ground from obstacle to thepoint of initiation of flare is called approachdistance. It can be calculated by using the EqnXXVIII. The airplane then begins to flare, which isthe transition from the straight approach path to thehorizontal ground roll. The transition path is a pullup maneuver in the turn radius „rflare‟, which can becalculated from the Eqn-XXVI. The flare isinitiated from a height „h f‟, measured from theground, and it can be calculated by using the Eqn XXVII. The angle induced by the flight pathbetween the point of initiation of flare to the pointof touchdown is called flare angle, ϴf. The flareangle will be almost equal to the approach angle,which is about 3º. The flare distance can becalculated by using the XXIX.After the point of touchdown, the aircraft willfreely roll down the runway for about 3 seconds,then the pilot applies thrust reversers, releases thespoilers, and other braking systems to bring theaircraft to halt. The distance covered by theairplane from the point of touchdown to the pointits velocity becomes zero is called the ground roll.It can be calculated by using the Eqn-XXX.Figure 8: Illustration of landing distance.The flare velocity is typically taken as 1.23 timesthe stall velocity, and the touchdown velocity (VTD )is assumed as 1.15 times the stall velocity.Table 10: LANDING DISTANCE DATAElevation6.7 mDensity1.225 kg/m3Approach distance112 mFlare distance358 mGround roll distance458.29 mTotal landing distance928.29 mINTERNATIONAL JO URNAL OF INNOVATIVE RESEARCH IN TECHNOLOGY79

August 2017 IJIRT Volume 4 Issue 3 ISSN: 2349-60023D ILLUSTRATIONFigure 9: Isometric view of the designed airplane.Figure 10: Front view of the designed airplane.Figure 11: Top view of the designed airplane.Figure 12: Side view of designed airplane.DESIGN POINTSFigure 13: Cruise speed vs takeoff weightFigure 14: Cruise speed vs Empty weightIJIRT 144769INTERNATIONAL JO URNAL OF INNOVATIVE RESEARCH IN TECHNOLOGY80

November 2017 IJIRT Volume 4 Issue 6 ISSN: 2349-6002Figure 15: Cruise speed vs serviceFigure 16: Cruise speed vs wing spanFigure 17: Cruise speed vs aspectFigure 18: Cruise speed vs length of aircraftIJIRT 144888ceilingCONCLUSIONThe aircraft design methodology is presented in thispaper. The methodology gives a brief designframework in which weight estimation; selection ofpower-plant; design criteria for wings, fuselage, andlanding gear; and performance evaluation is studiedin a coherent fashion. The design and calculation of180-seater passenger aircraft have been illustrated,and the skin-model is created by using CATIAV5R20. The obtained results are closely associatedwith the actual data of the aircraft of the twinturbofan medium-range category.REFERENCES[1] Dr. Jan Roskam, “Airplane Design”, 3rd edition,Roskam Aviation and Engineering Cooperation,Kansas, 1989.[2] John D Anderson, “Introduction to Flight”, 2ndedition, McGraw-Hill.[3] Daniel P Raymer, “Aircraft Design: AConceptual Approach”, 4th edition, AIAAEducation Series, American Institute ofAeronautics and Astronautics, Inc., Washington,DC, 1982.[4] John D Anderson, “Aircraft Performance andDesign”, 2nd edition, McGraw-Hill, 1998.[5] Mohammed Sadraey, “Wing Design”, Section5.4.6, Daniel Webster College.[6] Llyod R Jenkinson and James F Marchman III,“Aircraft Design Projects”, 2003.[7] “Theory of Wing Section”, report NACA-824,NASA TN-D-7428.[8] Egbert Torenbeek, “Advanced Aircraft Design,ConceptualDesign,TechnologyandOptimization of Subsonic Civil Airplanes”,1988.[9] A P Hettema, „Vertical Tail Design‟, DelftUniversity of Technology.[10] Perkins C and Hage R, “Aircraft PerformanceStability and Control”, Wiley, New York, 1963.[11] Nelson R C, “Flight Stability and AutomaticControl”, McGraw-Hill, 1998.ratioINTERNATIONAL JO URNAL OF INNOVATIVE RESEARCH IN TECHNOLOGY81

The design process is majorly broken into three phases- conceptual design, preliminary design, and detail design. 1.1 CONCEPTUAL DESIGN Conceptual design is the most important stage in the production and development of an aircraft. The primary components like wings, f