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Analytical Fuselage and Wing Weight Estimation of Transport Aircraft

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NASA Technical Memorandum 110392 Analytical Fuselage and Wing Weight Estimation of Transport Aircraft Mark D. Ardema, Mark C. Chambers, Anthony P. Patron, Andrew S. Hahn, Hirokazu Miura, and Mark D. Moore May 1996 National Aeronautics and Space Administration Ames Research Center Moffett Field, California 94035-1000 NASA Technical Memorandum 110392 Analytical Fuselage and Wing Weight Estimation of Transport Aircraft Mark D. Ardema, Mark C. Chambers, and Anthony P. Patron, Santa Clara University, Santa Clara, California Andrew S. Hahn, Hirokazu Miura, and Mark D. Moore, Ames Research Center, Moffett Field, California May 1996 National Aeronautics and Space Administration Ames Research Center Moffett Field, California 94035-1000 2 Nomenclature KF1 frame stiffness coefficient, IAFF/ A fuselage cross-sectional area Kmg shell minimum gage factor AB fuselage surface area KP shell geometry factor for hoop stress AF frame cross-sectional area KS constant for shear stress in wing (AR) aspect ratio of wing Kth sandwich thickness parameter b wingspan; intercept of regression line lB fuselage length bs stiffener spacing lLE length from leading edge to structural box at theoretical root chord bS wing structural semispan, measured along quarter chord from fuselage lMG length from nose to fuselage mounted main gear bw stiffener depth lNG length from nose to nose gear CF Shanley’s constant lTE length from trailing edge to structural box at CP center of pressure theoretical root chord C root chord of wing at fuselage intersection R l1 length of nose portion of fuselage C′ theoretical root chord of wing R l2 length of tail portion of fuselage Cs1 portion of wing leading edge not used for lπ length from nose to breakpoint of fuselage structural box L lift Cs2 portion of wing trailing edge not used for structural box LT maximum vertical tail lift CSR structural root chord of wing m buckling equation exponent; slope of regression line CST structural tip chord of wing M longitudinal bending moment CT tip chord of wing n normal load factor d frame spacing nx longitudinal acceleration dW optimum web spacing of wing Nx axial stress resultant D maximum diameter of fuselage A Nx bending stress resultant e wing buckling exponent B NxP pressure stress resultant eC wing cover material factor + Nx tensile axial stress resultant E Young’s modulus of shell material − Nx compressive axial stress resultant EF Young’s modulus of frame material Ny hoop direction stress resultant Fcy compressive yield strength P perimeter FS shear strength Pg internal gage pressure Ftu ultimate tensile strength Ps perimeter of shell h thickness of sandwich shell P perimeter of walls th w hei step function for i engine on wing P exponent of power law of nose section of th 1 hlgi step function for i landing gear on wing fuselage I frame cross-sectional area moment of inertia F P2 exponent of power law of tail section of Iy area moment of inertia about the y-axis fuselage Iy′ Itys/ r radius of fuselage iii r(y) total wing chord as a function of position tSC shell thickness required to preclude along quarter chord compressive failure rs(y) structural wing chord as a function of position tSG shell thickness required to meet minimum along quarter chord gage constraint R correlation coefficient used for regression tST shell thickness required to preclude tensile failure Rfin fineness ratio tT smeared tension tie thickness RHT ratio of horizontal tail station to fuselage length tw smeared wall thickness RLE ratio of wing leading edge station at twG thickness of wall to meet minimum gage theoretical root chord to fuselage length constraint RMG ratio of length to main gear to fuselage length twT thickness of wall required to prevent tensile failure RNG ratio of length to nose gear to fuselage length T torque on wing carrythrough structure RP1 ratio of length to leading edge of fuselage mounted propulsion to fuselage length VB fuselage volume RP2 ratio of length to trailing edge of fuselage VW volume of wing structural box, including mounted propulsion to fuselage length structural components Rt(y) thickness ratio of wing as a function of V1 volume of nose section of fuselage position along quarter chord V2 volume of tail section of fuselage RTAP taper ratio of wing wC width of carrythrough structure of wing S plan area of the fuselage B W weight of aircraft structure S stroke of landing gear LG W′ weight of wing per unit span SP plan area of wing WB weight of fuselage structure and attached t(y) thickness of wing box as a function of position components along quarter chord WFT weight of fuel tc core thickness WI ideal fuselage structural weight tf face sheet thickness WNO weight of nonoptimum material tg material gage thickness, tKsmg/ WS vehicle longitudinal weight distribution tmg material minimum gage thickness WTO gross takeoff weight of aircraft t skin thickness s W/S shell structural weight per unit surface area t stiffener thickness w x longitudinal fuselage coordinate t total equivalent isotropic thickness of shell and x weight calculated by PDCYL frames calc x distance from nose to theoretical quarter chord t total equivalent isotropic thickness of fuselage HT B of horizontal tail structure x distance from nose to leading edge of wing at t smeared equivalent isotropic thickness of LE F theoretical root chord frames x distance from nose to leading edge of fuselage t equivalent isotropic thickness of shell P1 S mounted propulsion t shell thickness required to preclude buckling SB x distance from nose to trailing edge of fuselage failure P2 mounted propulsion iv y transverse fuselage coordinate; wing εW wing web structural efficiency coordinate measured along quarter chord Λ wing sweep y actual weight act µ wing loading y estimated weight after regression est ρ structural material density z vertical fuselage coordinate ρB gross fuselage density Z(y) total width of wing box as a function of ρ frame structural material density position along quarter chord F σS allowable shear stress for wing ZS(y) width of wing box structure as a function of position along quarter chord Σ sum over fuselage or wing length; solidity of wing δ frame deflection ψ truss core angle ε shell buckling efficiency εC wing cover structural efficiency v Analytical Fuselage and Wing Weight Estimation of Transport Aircraft MARK D. ARDEMA,* MARK C. CHAMBERS,* ANTHONY P. PATRON,* ANDREW S. HAHN, HIROKAZU MIURA, AND MARK D. MOORE Ames Research Center Summary because of the detailed weight information available, allowing the weights output from PDCYL to be compared A method of estimating the load-bearing fuselage weight to actual structural weights. The detailed weight state- and wing weight of transport aircraft based on funda- ments also allow nonoptimum factors to be computed mental structural principles has been developed. This which, when multiplied by the load-bearing structural method of weight estimation represents a compromise weights calculated by PDCYL, will give good representa- between the rapid assessment of component weight using tive total structure weight estimates. These nonoptimum empirical methods based on actual weights of existing factors will be computed through a regression analysis of aircraft, and detailed, but time-consuming, analysis using a group of eight transport aircraft. the finite element method. The method was applied to eight existing subsonic transports for validation and corre- PDCYL is able to model both skin-stringer-frame and lation. Integration of the resulting computer program, composite sandwich shell fuselage and wing box PDCYL, has been made into the weights-calculating constructions. Numerous modifications were made to module of the AirCraft SYNThesis (ACSYNT) computer PDCYL and its associated collection of subroutines. program. ACSYNT has traditionally used only empirical These modifications include the addition of detailed weight estimation methods; PDCYL adds to ACSYNT a fuselage shell geometry calculations; optional integration rapid, accurate means of assessing the fuselage and wing of a cylindrical fuselage midsection between the nose and weights of unconventional aircraft. PDCYL also allows tail sections; addition of landing and bump maneuvers to flexibility in the choice of structural concept, as well as a the load cases sizing the fuselage; ability to introduce an direct means of determining the impact of advanced elliptical spanwise lift load distribution on the wing; materials on structural weight. variation of wing thickness ratio from tip to root; ability to place landing gear on the wing to relieve spanwise Using statistical analysis techniques, relations between bending loads; distribution of propulsion system compo- the load-bearing fuselage and wing weights calculated by nents between wing and fuselage; and the determination PDCYL and corresponding actual weights were deter- of maximum wingtip deflection. mined. A User’s Manual and two sample outputs, one for a typical transport and another for an advanced concept vehicle, are given in the appendices. Brief Description of ACSYNT The Aircraft Synthesis Computer program, ACSYNT, Introduction is an integrated design tool used in the modeling of advanced aircraft for conceptual design studies (ref. 5). A methodology based on fundamental structural ACSYNT development began at NASA Ames Research principles has been developed to estimate the load- Center in the 1970s and continues to this day. The carrying weight of the fuselage and basic box weight of ACSYNT program is quite flexible and can model a wide the wing for aircraft, and has been incorporated into the range of aircraft configurations and sizes, from remotely AirCraft SYNThesis program (ACSYNT). This weight piloted high altitude craft to the largest transport. routine is also available to run independently of ACSYNT, and is a modification of a collection of pre- The ACSYNT program uses the following modules, not viously developed structural programs (refs. 1–4). The necessarily in this order: Geometry, Trajectory, Aero- main subroutine called by ACSYNT is PDCYL.
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