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NASA NTRS1992

Composite fuselage shell structures research at NASA Langley Research Center

Fuselage structures for transport aircraft represent a large portion of both the weight and cost of these aircraft primary structures. Composite materials offer the potential for reducing both the weight and cost of these structures, but only limited studies of the response and failure of composite fuselage structures have been conducted. The research is described which is being conducted and planned at NASA Langley to understand the critical behavior of composite fuselage structures and to validate the structural mechanics methodology being developed for stiffened composite fuselage shell structure subjected to combined internal pressure and mechanical loads. Stiffened shell and curved stiffened panel designs are currently being developed and analyzed, and these designs will be fabricated and then tested to study critical fuselage behavior and to validate structural analysis and design methodology. The research includes studies of the effects of combined internal pressure and mechanical loads on nonlinear stiffened panel and shell behavior, the effects of cutouts and other gradient producing discontinuities on composite shell response, and the effects of local damage on pressure containment and residual strength. Scaling laws are being developed that relate full scale and subscale behavior of composite fuselage shells.

Starnes, James H., Jr.↗

NASA NTRS2000

Finite Element Model Development and Validation for Aircraft Fuselage Structures

The ability to extend the valid frequency range for finite element based structural dynamic predictions using detailed models of the structural components and attachment interfaces is examined for several stiffened aircraft fuselage structures. This extended dynamic prediction capability is needed for the integration of mid-frequency noise control technology. Beam, plate and solid element models of the stiffener components are evaluated. Attachment models between the stiffener and panel skin range from a line along the rivets of the physical structure to a constraint over the entire contact surface. The finite element models are validated using experimental modal analysis results. The increased frequency range results in a corresponding increase in the number of modes, modal density and spatial resolution requirements. In this study, conventional modal tests using accelerometers are complemented with Scanning Laser Doppler Velocimetry and Electro-Optic Holography measurements to further resolve the spatial response characteristics. Whenever possible, component and subassembly modal tests are used to validate the finite element models at lower levels of assembly. Normal mode predictions for different finite element representations of components and assemblies are compared with experimental results to assess the most accurate techniques for modeling aircraft fuselage type structures.

Buehrle, Ralph D.↗

NASA NTRS2000

Finite Element Model Development For Aircraft Fuselage Structures

Buehrle, Ralph D.↗

NASA NTRS1984

Noise-reduction measurements of stiffened and unstiffened cylindrical models of an airplane fuselage

Noise-reduction measurements are presented for a stiffened and an unstiffened model of an airplane fuselage. The cylindrical models were tested in a reverberant-field noise environment over a frequency range from 20 Hz to 6 kHz. An unstiffened metal fuselage provided more noise reduction than a fuselage having the same sidewall weight divided between skin and stiffening stringers and ring frames. The addition of acoustic insulation to the models tended to smooth out the interior-noise spectrum by reducing or masking the noise associated with the structural response at some of the resonant frequencies.

Willis, C. M.↗

NASA NTRS1984

Crashworthiness of light aircraft fuselage structures: A numerical and experimental investigation

The dynamic behavior of aircraft fuselage structures subject to various impact conditions was investigated. An analytical model was developed based on a self-consistent finite element (CFE) formulation utilizing shell, curved beam, and stringer type elements. Equations of motion were formulated and linearized (i.e., for small displacements), although material nonlinearity was retained to treat local plastic deformation. The equations were solved using the implicit Newmark-Beta method with a frontal solver routine. Stiffened aluminum fuselage models were also tested in free flight using the UTIAS pendulum crash test facility. Data were obtained on dynamic strains, g-loads, and transient deformations (using high speed photography in the latter case) during the impact process. Correlations between tests and predicted results are presented, together with computer graphics, based on the CFE model. These results include level and oblique angle impacts as well as the free-flight crash test. Comparisons with a hybrid, lumped mass finite element computer model demonstrate that the CFE formulation provides the test overall agreement with impact test data for comparable computing costs.

Nanyaro, A. P.↗

NASA NTRS1942

Stability of elastically supported columns

A criterion is developed for the stiffness required of elastic lateral supports at the ends of a compression member to provide stability. A method based on this criterion is then developed for checking the stability of a continuous beam-column. A related method is also developed for checking the stability of a member of a pin-jointed truss against rotation in the plane of the truss.

COMPRESSION MEMBERS-STABILITY↗

NASA NTRS1999

Experiment Verification of the Analytical Methodology to Predict the Residual Strength of Metallic Shell Structure

Experimental and analysis results for a curved, stiffened aluminum fuselage panel tested in a combined loads test machine with combined internal pressure, axial compression, and torsional shear loads are described. The experimental and analytical strain results for the panel with and without discrete source damage are presented. The effect of notch tip geometry on crack growth predictions is addressed. The crack growth trajectory predictions for the panel are presented for the applied loading conditions at failure.

Starnes, James H., Jr.↗

NASA NTRS1997

Advanced Technology Composite Fuselage - Materials and Processes

The goal of Boeing's Advanced Technology Composite Aircraft Structures (ATCAS) program was to develop the technology required for cost and weight efficient use of composite materials in transport fuselage structure. This contractor report describes results of material and process selection, development, and characterization activities. Carbon fiber reinforced epoxy was chosen for fuselage skins and stiffening elements and for passenger and cargo floor structures. The automated fiber placement (AFP) process was selected for fabrication of monolithic and sandwich skin panels. Circumferential frames and window frames were braided and resin transfer molded (RTM'd). Pultrusion was selected for fabrication of floor beams and constant section stiffening elements. Drape forming was chosen for stringers and other stiffening elements. Significant development efforts were expended on the AFP, braiding, and RTM processes. Sandwich core materials and core edge close-out design concepts were evaluated. Autoclave cure processes were developed for stiffened skin and sandwich structures. The stiffness, strength, notch sensitivity, and bearing/bypass properties of fiber-placed skin materials and braided/RTM'd circumferential frame materials were characterized. The strength and durability of cocured and cobonded joints were evaluated. Impact damage resistance of stiffened skin and sandwich structures typical of fuselage panels was investigated. Fluid penetration and migration mechanisms for sandwich panels were studied.

Scholz, D. B.↗

NASA NTRS1994

Residual life and strength estimates of aircraft structural components with MSD/MED

Economic and safe operation of the flight vehicles flying beyond their initial design life calls for an in-depth structural integrity evaluation of all components with potential for catastrophic damages. Fuselage panels with cracked skin and/or stiffening elements is one such example. A three level analytical approach is developed to analyze the pressurized fuselage stiffened shell panels with damaged skin or stiffening elements. A global finite element analysis is first carried out to obtain the load flow pattern through the damaged panel. As an intermediate step, the damaged zone is treated as a spatially three-dimensional structure modeled by plate and shell finite elements, with all the neighboring elements that can alter the stress state at the crack tip. This is followed by the Schwartz-Neumann alternating method for local analysis to obtain the relevant crack tip parameters that govern the onset of fracture and the crack growth. The methodology developed is generic in nature and aims at handling a large fraction of problem areas identified by the Industry Committee on Wide-Spread Fatigue Damage.

Singh, Ripudaman↗

NASA NTRS1995

Load transfer in the stiffener-to-skin joints of a pressurized fuselage

Structural analyses are developed to determine the linear elastic and the geometrically nonlinear elastic response of an internally pressurized, orthogonally stiffened, composite material cylindrical shell. The configuration is a long circular cylindrical shell stiffened on the inside by a regular arrangement of identical stringers and identical rings. Periodicity permits the analysis of a unit cell model consisting of a portion of the shell wall centered over one stringer-ring joint. The stringer-ring-shell joint is modeled in an idealized manner; the stiffeners are mathematically permitted to pass through one another without contact, but do interact indirectly through their mutual contact with the shell at the joint. Discrete beams models of the stiffeners include a stringer with a symmetrical cross section and a ring with either a symmetrical or an asymmetrical open section. Mathematical formulations presented for the linear response include the effect of transverse shear deformations and the effect of warping of the ring's cross section due to torsion. These effects are important when the ring has an asymmetrical cross section because the loss of symmetry in the problem results in torsion and out-of-plane bending of the ring, and a concomitant rotation of the joint at the stiffener intersection about the circumferential axis. Data from a composite material crown panel typical of a large transport fuselage structure are used for two numerical examples. Although the inclusion of geometric nonlinearity reduces the 'pillowing' of the shell, it is found that bending is localized to a narrow region near the stiffener. Including warping deformation of the ring into the analysis changes the sense of the joint rotation. Transverse shear deformation models result in increased joint flexibility.

Johnson, Eric R.↗

NASA NTRS1993

Test results from large wing and fuselage panels

This paper presents the first results in an assessment of the strength, stiffness, and damage tolerance of stiffened wing and fuselage subcomponents. Under this NASA funded program, 10 large wing and fuselage panels, variously fabricated by automated tow placement and dry-stitched preform/resin transfer molding, are to be tested. The first test of an automated tow placement six-longeron fuselage panel under shear load was completed successfully. Using NASTRAN finite-element analysis the stiffness of the panel in the linear range prior to buckling was predicted within 3.5 percent. A nonlinear analysis predicted the buckling load within 10 percent and final failure load within 6 percent. The first test of a resin transfer molding six-stringer wing panel under compression was also completed. The panel failed unexpectedly in buckling because of inadequate supporting structure. The average strain was 0.43 percent with a line load of 20.3 kips per inch of width. This strain still exceeds the design allowable strains. Also, the stringers did not debond before failure, which is in contrast to the general behavior of unstitched panels.

Madan, Ram C.↗

NASA NTRS1986

A comparison of the structureborne and airborne paths for propfan interior noise

A comparison is made between the relative levels of aircraft interior noise related to structureborne and airborne paths for the same propeller source. A simple, but physically meaningful, model of the structure treats the fuselage interior as a rectangular cavity with five rigid walls. The sixth wall, the fuselage sidewall, is a stiffened panel. The wing is modeled as a simple beam carried into the fuselage by a large discrete stiffener representing the carry-through structure. The fuselage interior is represented by analytically-derived acoustic cavity modes and the entire structure is represented by structural modes derived from a finite element model. The noise source for structureborne noise is the unsteady lift generation on the wing due to the rotating trailing vortex system of the propeller. The airborne noise source is the acoustic field created by a propeller model consistent with the vortex representation. Comparisons are made on the basis of interior noise over a range of propeller rotational frequencies at a fixed thrust.

Eversman, W.↗

NASA NTRSDate not supplied

Development and Characterization of the Integrally Stiffened Cylinder (ISC) Process for Launch Vehicles and Aircraft Fuselage Structures

Over the past decade, NASA Langley has led development of the Integrally Stiffened Cylinder (ISC) Process, a near-net shape, flow-forming technology. This innovative process is being evaluated for launch vehicle and commercial aircraft manufacturing. A thick-walled, cylindrical preform is flow-formed in a single operation into a thin-walled barrel with longitudinal stiffeners. The one-piece, stiffened barrels offer a direct replacement for conventional multi-piece welded or riveted structures. A cost-benefit analysis for launch vehicles estimated that the ISC process offers up to a 50% reduction in manufacturing costs and a 10% reduction in mass. NASA, the European Space Agency (ESA), and industry partners have teamed to manufacture 10-ft. diameter aluminum ISCs. The intent is to demonstrate the commercial viability of the process at a scale relevant to the commercial launch and aircraft industries. Development activities will be presented, including process scale-up, formability assessment, microstructural analysis, and mechanical property testing.

ISC, flow forming, near net shape↗

NASA NTRSDate not supplied

Development and Characterization of the Integrally Stiffened Cylinder (ISC) Process for Launch Vehicles and Aircraft Fuselage Structures

Over the past decade, NASA Langley has led development of the Integrally Stiffened Cylinder (ISC) Process, a near-net-shape, flow-forming technology. This innovative process is being evaluated for launch vehicle and commercial aircraft manufacturing. A thick-walled, cylindrical preform is flow-formed using a single operation into a thin-walled barrel with integral longitudinal stiffeners. The one-piece, stiffened barrels offer a direct replacement for conventional multi-piece, welded or riveted structures. A cost-benefit analysis for launch vehicle cryogenic propellant tanks estimated that the ISC process offers up to a 50% reduction in manufacturing costs and a 10% reduction in mass. NASA, the European Space Agency (ESA), and industry partners have teamed to manufacture 3-m(10-ft.) diameter aluminum ISCs. The intent is to demonstrate the commercial viability of the process at a scale relevant to the commercial launch and aircraft industries. Development activities will be presented, including process scale-up, formability assessment, and mechanical property testing.

ISC, flow forming, near net shape↗

NASA NTRS1986

A comparison of the structureborne and airborne paths for propfan interior noise

In the present comparative study of aircraft interior noise levels due to airborne and structureborne paths for a given propeller source, a structure model is used that treats the fuselage interior as a rectangular cavity with five rigid walls; the sixth, or fuselage sidewall, is a stiffened panel. The wing is modeled as a simple beam carried into the fuselage by a large, discrete stiffener representing the carry-through structure. Comparisons are made on the basis of interior noise over a range of propeller rotational frequencies at a fixed thrust. It is found that the relative importance of the structural and airborne paths is highly dependent on the structural parameters.

Eversman, W.↗

NASA NTRS1981

Structural synthesis - Its genesis and development

An historical account is given of the development, from its conception in 1960, of the structural synthesis method. While synthesis techniques lag behind analytical ones in both sophistication and application, the structural design procedures created by combining finite element analysis and mathematical programming algorithms have progressed to the point of maturity. As in the case of finite element analysis, use and acceptance of structural synthesis methodology depends on the development and distribution of easily used and well-documented, production-quality computer programs. Attention is given such elementary applications of synthesis methods as the three-bar truss, an integrally stiffened waffle plate, a stiffened cylindrical shell, aircraft fuselage window panels, the structural efficiency of graphite-epoxy hat-stiffened panels, and an idealized delta wing.

Schmit, L. A.↗

NASA NTRS1981

A study of methods of prediction and measurement of the transmission sound through the walls of light aircraft

Several aspects were studied. The SEA theory was used to develop a theoretical model to predict the transmission loss through an aircraft window. This work mainly consisted of the writing of two computer programs. One program predicts the sound transmission through a plexiglass window (the case of a single partition). The other program applies to the case of a plexiglass window window with a window shade added (the case of a double partition with an air gap). The sound transmission through a structure was measured in experimental studies using several different methods in order that the accuracy and complexity of all the methods could be compared. Also, the measurements were conducted on the simple model of a fuselage (a cylindrical shell), on a real aircraft fuselage, and on stiffened panels.

Forssen, B.↗

NASA NTRS1997

Advanced Technology Composite Fuselage - Manufacturing

The goal of Boeing's Advanced Technology Composite Aircraft Structures (ATCAS) program is to develop the technology required for cost-and weight-efficient use of composite materials in transport fuselage structure. Carbon fiber reinforced epoxy was chosen for fuselage skins and stiffening elements, and for passenger and cargo floor structures. The automated fiber placement (AFP) process was selected for fabrication of stringer-stiffened and sandwich skin panels. Circumferential and window frames were braided and resin transfer molded (RTM'd). Pultrusion was selected for fabrication of floor beams and constant-section stiffening elements. Drape forming was chosen for stringers and other stiffening elements cocured to skin structures. Significant process development efforts included AFP, braiding, RTM, autoclave cure, and core blanket fabrication for both sandwich and stiffened-skin structure. Outer-mold-line and inner-mold-line tooling was developed for sandwich structures and stiffened-skin structure. The effect of design details, process control and tool design on repeatable, dimensionally stable, structure for low cost barrel assembly was assessed. Subcomponent panels representative of crown, keel, and side quadrant panels were fabricated to assess scale-up effects and manufacturing anomalies for full-scale structures. Manufacturing database including time studies, part quality, and manufacturing plans were generated to support the development of designs and analytical models to access cost, structural performance, and dimensional tolerance.

Wilden, K. S.↗