Aircraft Design Studies Based on the ATR 72

Transcription

1 Aircraft Design Studies Based on the ATR 72 Mihaela Florentina Niţă Supervisor: Prof. Dr.-Ing. Dieter Scholz HAW Hamburg RRDPAE 2008 Recent Research and Design Progress in Aeronautical Engineering and its Influence on Education Brno University of Technology, Czech Republic, October 2008

2 Two Design Steps Preliminary sizing (Paper on RRDPAE CD) Conceptual Design (Master Thesis on WWW) Emphasis of this presentation

3 Overview Preliminary sizing Gives input parameters for the conceptual design:» Maximum take-off mass,» Fuel mass,» Maximum operating empty mass,» Wing area, m F m MTO m OE S W TTO PTO» Take-off thrust, or take-off power,

4 Overview Conceptual design

5 Fuselage The Fuselage Requirements:» Passengers comfort» Drag» Weight Cross section:» Given: Number of passengers n PAX = 70» Yields: Number of seats abreast n SA = 0.45 n = 4 PAX and number of aisles nsa 6 1Aisle (CS )

6 Fuselage» Interior diameter of the fuselage Empirical Equation dfi, = (2 Bench width + Aisle width) m m = 2.57m» Exterior diameter of the fuselage d = d d = 0.084m d d = 2.77m FO, FI, FI, FO,

7 Tail length Fuselage Cabin and fuselage:» Seat pitch: 31in = 0,78 m; k CABIN = 1m An average seat pitch including galleys,lavatories» Cabin Length npax lcabin = kcabin = 19.25m n SA» Fuselage length l = l d + 4m= 27.13m F CABIN F Cockpit length» Emergency exits: 2+2 type I and III

8 Fuselage Other parameters:» Slenderness parameter lf λ F = = d F Important parameter that determines drag and structural weight

9 Wing The Wing Design boundaries All interconnected!!!

10 Wing Design method

11 Wing Results SW M CR = = 61.3m A = ϕ 25 = 3 b= 27.13m Chosen airfoil: NACA 430xx Non linear regression ( t/ c) = k M cosϕ C k = t u v w t DD 25 L M ( t/ c) = 18% r (/ t c) = 13% t NACA NACA ε = 3 t 0 i w Abbot, Pankhurst C = + α0 0.4 ε = 4 C α LCR, 0 t L ν W = 0 0 λ ϕ 25 opt = 0.45 e or statistics λ = c / c = 0.59 t r

12 Wing c r 2b = = 2.6m A[(1 λη ) + λ+ λ] k i c = λc = 1.5m t r 1 1+ λ τ + λ τ Vtank = 0.54 S ( t/ c) = 9.3m > V = 4.5m A (1 + λ) W r 2 tank,nec where (/ t c) t τ = = 0.72 (/ t c) r From preliminary sizing

13 High Lift System The high lift system Design method Start C = 1.1 C = 2.2 L,max L,max, INITIAL SIZING C L,max type of flap : double slotted flap and slats Statistical reasearch Increase in Lift calculation Verify equation Iterative process no yes Stop c = k k k ( c ) = L,max, f L,max base S C = c K = W, f L,max, f L,max, f ϕ SW ' c L,max, s cl, δ,maxηmaxηδδ f c = = c SWs, CL,max, s = cl,max, s cosϕh, L = S W 0.95 CL,max, f + CL,max, s CL,max CL,max, clean DATCOM

14 Empennage Estimating the empennage area from statistics Horizontal tail, vertical tail Configuration: T-tail (engine location on a high wing) Surface area from statistical approach Tail volume Tail volume coefficient Lever arm l = l = 50% l = m V H F C H SH lh = = 1.05 S c W MAC Tail volume coefficient C V SV lv = = S b W S S H V CH SW cmac = = m lh CV SW b 2 = = m l V 2 Results

15 Empennage Other parameters Aspect ratio and taper ratio: A A H V = 0.5A = 6 = 1.6 w λh = 0.6 λ = 0.6 Dihedral and sweep: V V H V = 80 = ϕ ϕ V 25, H 25, V = 8 0 = 25 NACA Airfoil: for the vertical tailplane NACA 0009 for the horizontal tailplane

16 Mass and CG Mass estimation and CG location Estimation per each component using a Class II method (Torenbeek) Example calculation: wing mass 0.30 m b W ref 0.55 bs / t r = bs 1+ nult = 0.17 m MZF b s mmzf / SW m = 0.17 m = 3045kg W MZF The approximations are made by taking into account variations with specific parameters, as it is shown in the next table

17 Mass and CG Parameters used for the mass estimation Results [kg] Wing B ref /b s ; m MZF /S W ;n ult 3045 Fuselage S wet,f ; l H ; V D ; d F 2323 Horizontal Tailplane S H ; V D 124 Vertical Tailplane S V ; V D 179 Landing gear m MTO and coefficients 961 Engine nacelle T, respectively P,η,V 242 Installed engine n E ; m E 1533 Systems m MTO 3114 Supplemental mass n Seat ; n Pax 1050 Operating empty mass Sum of components 12834

18 Mass and CG CG position and position of the wing towards the fuselage - CG position of the wing x mi xi, = = m m WG LEMAC -CG position of the fuselage x FG mi xi = = m m i i Equilibrium of moments 25%C MAC Position of the wing: m x x x x x m ( ) WG LEMAC = FG CG, LEMAC + WG, LEMAC CG, LEMAC = 11 mfg TORENBEEK, E.: "Synthesis of Subsonic Airplane Design Delft University Press, 1988

19 Sizing the empennage according to stability and control requirements Horizontal Tail Sizing after control requirements SH / SW = a xcg AC + b CL CMW, + CME, a = = l b = = H CLH, ηh lh CLH, ηh cmac cmac Sizing after stability requirements SH / SW = a xcg AC CL, α, W a = = ε l H CL, α, H ηh 1 α cmac Intersection of requirements SH SH 9.701m S = = W Stability and Control 2 Following the introduction of the stability margin, according to the next graph

20 Stability and Control

21 Stability and Control Vertical Tail S V Sizing after control requirements NE + ND = = m 1 2 c L, δ ' ρvmc δf ( cl, δ ) theory K KΛ lv 2 ( cl, δ ) theory Sizing after stability requirements SV C C b S C l N, β N, β, F W = = W Y, β, V V S = 9.57m V Evaluation of the results If the area S H does not match Empennage results then: m H would need to be re-evaluated and wing position adjusted For the vertical tail the larger area of the two was chosen 2

22 Landing gear Landing Gear To prevent tail tipping Position: corelated with the CG aft position Turn over angle in the x direction: min. 15 Distance between wheels of the main LG Tail clearence: 11 x ytrack, LG, M = 10.77m LG N = 4.10m To prevent side tipping Lateral clearence: min. 7 required

23 Drag and Polar Drag estimation and polar Three major components: Zero lift drag it is being estimated for each component, according to the formula: CD,0 = Cf FFc Qc Swett / Sref Lift dependent drag Mach drag we neglect this from the beginning, as the aircraft flies at lower speed C f FF c Q c S wett /S ref C D,0 Fuselage Wing Horizontal Tailplane Vertical Tailplane Nacelle Total

24 Drag and Polar The polar is given by 2 L 2 CD = CD,0 + C CD = CL π A e» In the preliminary sizing calculation the value e = 0.85 was used» The resulting L/D is E = 15.8

25 DOC Estimation Design evaluation AEA method (Association of European Airliners) for estimating the direct operating costs (DOC) Parameters used for the estimation Depreciation Service life, residual value 0.99 Interest Average interest rate, total price of the aircraft Results [mil$/year] 0.73 Insurance % of aircrafts price 0.07 Fuel Maintenance Price and mass fuel, no. of flights per year Labor and material, inflation factor Crew No. of crew members 2.07 Fees: Landing Maximum take-off mass, no. of flights/year, inflation factor Navigation Maximum take-off mass, inflation factor Handling Maximum payload, inflation factor

26 DOC Estimation Total DOC = the sum of the costs of each of the following elements: CDOC = CDEP + CINT + CINS + CF + CM + CC + CFEE C DOC =10.5 mil US$/year

27 Summary Components Redesign Original Deviation Fuselage Length Diameter Cabin Length Wing Wing Span Wing Surface Wing Loading High Lift Device m 2.77 m m m 61.3 m kg/m 2 Double sloted flaps and slats m 2.57 m m m 61.0 m kg/m 2 Double sloted flaps 0.1% -2.0% 0.1% 0.3 % 0.5 % 0.0 % Power Plant Power Loading W/kg W/kg -0.1 % Horizontal Tail Surface 9.7 m m % Vertical Tail Surface 14.1 m m % Mass Maximum Take-Off Mass kg kg 0.5% Operating Empty Mass kg kg 0.9%

28 For more information please visit the digital library: and check the RRDPAE CD