Pressurized Fuselage Research Articles

Semi-analytical modelling of mode jumping phenomena in pressurized fuselage panels

In this paper, a novel semi-analytical model is developed to capture the mode jumping phenomenon in composite and metallic pressurized fuselage panels. The analytical formulation is derived using the Ritz-based approach and the resulting system of nonlinear equations are solved numerically by homotopy methods. Panels with two different aspect ratios and three different types of boundary conditions are investigated. The panels are loaded in two sequential steps. In the first step the panel is loaded by pressure resulting in a primary mode shape. In the second step, the panel is loaded by an in-plane axial compressive load which causes a sudden change in the primary mode shape leading to a mode jumping event. For verification purposes, finite element models are developed and analysed using the commercial software Abaqus. It is demonstrated that the semi-analytical model is able to successfully capture the mode jumping phenomenon and predict the corresponding load. New findings concerning the mode jumping problem in pressurized fuselage panels are analysed and interpreted.

АНАЛІЗ КОНСТРУКТИВНО-ТЕХНОЛОГІЧНИХ ОСОБЛИВОСТЕЙ ЗАБЕЗПЕЧЕННЯ РЕСУРСУ СЕРЕДНЬОЇ ЧАСТИНИ ФЮЗЕЛЯЖУ В ЗОНІ ВИРІЗУ ПІД ДВЕРІ ЛІТАКІВ ТРАНСПОРТНОЇ КАТЕГОРІЇ

This article is devoted to the analysis of the design and technological features of ensuring the resource of cutouts in the middle part of the pressurized fuselage of transport category aircraft. Aircraft fuselages, due to their functional features, have a large number of different cuts and irregularities. During flight, the fuselage structure perceives loads in the form of transverse and longitudinal forces, bending and torsional moments, which are caused by the general bending of the structure due to the action of aerodynamic, inertial and dynamic loads, as well as overpressure, which is formed in the middle of the pressurized part of the fuselage during operation of the air conditioning system.The main cutouts can be: cutouts for windows, doors, luggage, maintenance and emergency hatches, cutouts for the cargo hatch. Particular attention must be paid to the cutouts that are in the sealed area of the fuselage. The presence of a cutout in the structure of the fuselage shell leads to a weakening of the cross section of the fuselage in the cutout area. As an object of study, a cutout was chosen, which is located in the middle part of the fuselage under the door (1922x1012 mm). The presence of such a cutout in the fuselage shell creates a sufficiently high concentration of stresses, namely in the area of the cutout corner rounding radii. There, in addition to the concentration due to the corners of the cutout, which are rounded, there is also a concentration around the holes for fasteners, which are located near the rounding radii. It is in this zone that there is the highest probability of the formation and development of fatigue cracks.To achieve the desired durability indicators in the zone of significant stress concentration, a number of design and technological solutions are used. The use of constructive and technological solutions provides the requirements of static strength and durability in the areas of cutouts in the pressurized fuselage and puts forward the problem of obtaining a structure with a minimum mass

Fail-Safe Design Analysis of an Aircraft Fuselage with Crack Stopper Strap

This paper reveals one of the design philosophies, fail-safe is used to arrest the crack propagation of airframe. The two important cracks considered for the design of pressurized fuselage are circumferential and longitudinal. The fuselage cabin pressurization and depressurization cycle create a fatigue load on the airplane structure and produce cracks where the tensile stress is maximum on the structural components. If the crack propagation is not controlled, they propagate during the flight and eventually land on the complete loss of the airplane structure. The design feature introduced into the airframe bulkheads to stop the crack propagation during the flight is crack tear straps. The approach utilizes a finite element modeling and analysis to assess the damage in the bulkheads and stringers with crack stopper straps and without crack stoppers. The design and analysis tools used in the in this work are Catia V5 and ANSYS. The analysis is carried out under fixed support conditions of the fuselage structure at the maximum design load conditions. The fatigue life of the stiffened panel with tear strap is more than the panel without tear strap. The deformation under cyclic loads for the panel with strap is less than the fuselage stiffened panel without crack strap

Full-scale experiments on the reduction of propeller-induced aircraft interior noise with active trim panels

Results are presented of full-scale experiments on the active reduction of rotor-induced passenger cabin noise in a Dornier Do728 aircraft. Two active sidewall panels (smart linings) are used to reduce the sound pressure in a control region in front of the linings. The fuselage pressure distribution associated with the first five harmonics of a generic counter-rotating open rotor (CROR) engine is emulated with a loudspeaker array. Each smart lining is equipped with two inertial force actuators. It uses up to eight error microphones and implements an adaptive feedforward controller. The two smart linings are driven in parallel. A maximum SPL reduction of 11.3 dB is achieved in the controlled area. The mean SPL reduction over 18 microphones is 6.8 dB. At the most critical second frequency, a mean SPL reduction of 9.3 dB is achieved by the two smart linings working in parallel. It is observed that the SPLs near the lining (at the window seat positions) are significantly higher than the SPLs near the aisle. This leads to the conclusion that the major part of the sound energy is transmitted through the linings, which is seen as an argument for the suitability of the smart lining concept.

Validation of Higher-Order Interactional Aerodynamics Simulations on Full Helicopter Configurations

Time-accurate numerical predictions of the interactional aerodynamics between NASA's generic ROBIN fuselage and its four-bladed rotor were performed using the recently developed Reynolds-averaged Navier–Stokes solver HAMSTR. Two stencil-based reconstruction schemes (MUSCL, WENO), a second-order temporal accuracy, and the Spalart–Allmaras turbulence model were used. Three-dimensional volume meshes were created in a robust manner from two-dimensional unstructured surface grids using Hamiltonian paths and strands on nearbody domains. Grid connectivity was established between nearbody and background domains in an overset fashion. Two previously researched operational conditions were reproduced, i. e., a near-hover case and a medium-speed forward flight case at an advance ratio of 0.151. The results were compared with various experimental and numerical references and were found to be in good agreement with both. The comparison included the analysis of the rotor wake structure, tip-vortex trajectories, steady and dynamic fuselage pressure distributions in longitudinal and lateral directions, and rotor inflow predictions.

Manufacturing Aspects of Creating Low-Curvature Panels for Prospective Civil Aircraft

For this study, structural and manufacturing schemes for low-curvature pressurized fuselage panels were proposed, making it possible to provide high weight efficiency for the airframes of prospective civil blended wing-body (BWB) aircraft. The manufacturing scheme for low-curvature panels helped to achieve high strength characteristics of the composite details as well as decreased the labor input necessary for manufacturing and assembling. The beneficial features of the proposed structure are that the panels have a low weight, incur low manufacturing costs, and satisfy the demands of repairability.

A theoretical model of fuselage pressure levels due to fan tones radiated from the intake of an installed turbofan aero-engine.

An existing theoretical model to predict the pressure levels on an aircraft's fuselage is improved by incorporating a more physically realistic method to predict fan tone radiation from the intake of an installed turbofan aero-engine. Such a model can be used as part of a method to assess cabin noise. Fan tone radiation from a turbofan intake is modelled using the exact solution for the radiated pressure from a spinning mode exiting a semi-infinite cylindrical duct immersed in a uniform flow. This approach for a spinning duct mode incorporates scattering/diffraction by the intake lip, enabling predictions of the radiated pressure valid in both the forward and aft directions. The aircraft's fuselage is represented by an infinitely long, rigid cylinder. There is uniform flow aligned with the cylinder, except close to the cylinder's surface where there is a constant-thickness boundary layer. In addition to single mode calculations it is shown how the model may be used to rapidly calculate a multi-mode incoherent radiation from the engine intake. Illustrative results are presented which demonstrate the relative importance of boundary-layer shielding both upstream and downstream of the source, as well as examples of the fuselage pressure levels due to a multi-mode tonal source at high Helmholtz number.

Effects of loading modes on reliability of cracked structures using FORM and MCS

Studying the reliability of cracked structures under different types of failure modes such as a mixture of Mode I and Mode II loadings has been of great importance in many applications. The purpose of this study is to evaluate the effects of different modes of loading, i.e. pure Mode I, pure Mode II and Mixed mode I/II loadings on the reliability of cracked structures via three typical case studies including 1st: through crack in a very wide plate, 2nd: cracked pressurized fuselage and 3rd: cracked fastener holes under actual geometries and loading conditions using the FORM, the fracture toughness and the equivalent SIF for Mixed mode I/II loading which is extracted by the Energy release rate model 1. The MCS was used to check the accuracy of the FORM. It was found that the failure probability decreases significantly as the loading mode changes from the pure Mode I to pure Mode II. This means that the reliability in the Mixed mode I/II loading is higher than that of pure Mode I and smaller than for pure Mode II. Also, it was concluded that the failure probability decreases with increasing of the angle of crack for the Mixed mode I/II loading conditions. The results indicated that the angle of crack and the applied stress are the two variables which have the most significant influence on the reliability.

Finite element simulation of gaseous detonation-driven fracture in thin aluminum tube using cohesive element

Detonation-driven fracture problems in tube under dynamic load have received plenty of attention because of various ranges of applications, such as oil and gas pipeline systems, new rocket engine such as pulse detonation engine, and pressurized aircraft fuselages. This paper reports the crack growth modeling in a thin aluminum tube under gaseous detonation load. Because of three-dimensional fracture dynamics with gas dynamics coupled phenomena, analytical modeling is complicated. Thus, a finite element method was applied. The finite element modeling and simulation of the tube under detonation moving load were performed using commercial code Abaqus. This simulation leads to obtain structural response of the tube to detonation load. The simulations were compared with experimental and analytical results from the literature for elasto-dynamic response of cylindrical shells with finite length under internal detonation loading. Cohesive element with traction–separation law was used for crack growth modeling along with crack tip opening displacement value obtained from experimental–numerical analysis from previous research. The final section of the paper is dedicated to investigating differences and comparisons between the numerical crack propagation simulations and experimental results reported in the literature. It has been demonstrated that using cohesive elements with some modifications can improve the numerical accuracy. The obtained results are more similar to the experimental results than numerical results available in literature.

Experimental and numerical investigation of the power-on effect for a propeller-driven UAV

The power-on effect contributes to the performance and the stability of a propeller-driven aircraft. In the present study, the power-on effect for a propeller-driven unmanned aerial vehicle with a pusher propeller has been investigated through the wind tunnel test and the computational fluid dynamics simulation. The positive power-on effect is to increase the maximum lift and delay the stall of the aircraft. However, the most common power-on effect is to produce the slipstream drag and the nose-up pitching moment of the aircraft. The slipstream interaction with the wing and the fuselage increases the fuselage pressure drag leading to an overall increase in aircraft drag. Moreover, the slipstream interaction increases the downwash angle and dynamic pressure in the vicinity of the horizontal tail leads to an overall increase in aircraft nose-up pitching moment. Therefore, the power-on effects tend to reduce the performance and longitudinal stability of the propeller-driven aircraft. The power-on effect is pronounced at slower speeds and higher thrust levels. It is expected that the analysis results presented and discussed in this paper will contribute to the development of the propeller-driven aircraft with the pusher propeller.

Unstructured Overset Mesh Adaptation with Turbulence Modeling for Unsteady Aerodynamic Interactions

Schemes for anisotropic grid adaptation for dynamic overset simulations are presented. These approaches permit adaptation over a periodic time window in a dynamic flowfield so that an accurate evolution of the unsteady wake may be obtained, as demonstrated on an unstructured flow solver. Unlike prior adaptive schemes, this approach permits grid adaptation to occur seamlessly across any number of grids that are overset, excluding only the boundary layer to avoid surface manipulations. A demonstration on a rotor/fuselage-interaction configuration includes correlations with time-averaged and instantaneous fuselage pressures, and wake trajectories. Additionally, the effects of modeling the flow as inviscid and turbulent are reported. The ability of the methodology to improve these predictions is confirmed, including a vortex/fuselage-impingement phenomenon that has before now not been captured by computational simulations. The adapted solutions exhibit dependency based on the choice of the feature to form the adaptation indicator, indicating that there is no single best practice for feature-based adaptation across the spectrum of rotorcraft applications.

Blast Response of a Pressurized Aircraft Fuselage Measured Using High-Speed Image Correlation

As part of a project examining the blast response of a pressurized aircraft fuselage, high-speed digital image correlation was used to measure the change in surface strain and out-of-plane displacement during an explosive event. Two explosive charges were used for this test, one located in the aft cargo compartment and one located in the forward cargo compartment. Digital image correlation was used to monitor the charge in the aft compartment with the second charge in the forward compartment being detonated 100 ms after the first. Digital image correlation was able to monitor the change in out-of-plane displacement as a result of the high-temperature gas generated during the explosion. Strain measurements made during the explosion were used to calculate the Von Mises stresses and confirmed that the aluminum skin did not plastically deform. The second explosion in the forward baggage compartment did result in fracture of the fuselage skin and post-test analysis of the fracture surfaces showed that tensile stresses in the rivets likely led to failure.

Analysis of Residual Stress Effect on Fatigue Crack Propagation in Bonded Aeronautical Stiffened Panels

The application of adhesively bonded straps made of high-static strength materials on aeronautical stiffened panels to retard the fatigue skin crack growth is currently a topical research subject. The detrimental effect of the residual stress fields induced as a consequence of the dissimilar coefficients of thermal expansion of the skin and strap materials on the fatigue skin crack propagation was investigated. The residual stresses induced in a stiffened panel representative of a pressurized fuselage shell with titanium doublers in the middle of the stringer bays is numerically quantified for two likely operational temperatures. Their effect on the fatigue crack propagation is analyzed by means of a linear elastic fracture mechanics approach. The results show that adhesively bonded straps on the cracked surface can significantly retard the fatigue crack propagation but, in order to achieve reliable and conservative predictions on their performances, the effect of the residual stress fields they introduce must be taken into account.

Certification programme of airframe primary structure composite part with environmental simulation

The paper presents certification programme of an airframe primary structure composite part (pressurized bulkhead). Within the test the mechanical (static and fatigue) and environmental loading was simulated. The environmental loading is represented by 70 °C temperature and 90% humidity. The proposal and realization of the whole certification programme in the Aeronautical Research and Test Institute (VZLU) is presented. The designed programme fulfilled the conditions of the FAR-23 (or CS-23) regulations requirements. The composite bulkhead was designed and manufactured during the development of a new small Czech air carrier with pressurized fuselage. The bulkhead is honeycomb structure with carbon/epoxy skin. The technical and operational testing conditions, test facility possibilities and follow-up details related to the structure certification are stated. Results from residual strength test and results from advanced non-destructive testing are presented (Shearography method).

On the finite element modeling of fatigue crack growth in pressurized cylindrical shells

A methodology for the computational modeling of the fatigue crack growth in pressurized shell structures, based on the finite element method and concepts of Linear Elastic Fracture Mechanics, is presented. This methodology is based on that developed by Potyondy [Potyondy D, Wawrzynek PA, Ingraffea, AR. Discrete crack growth analysis methodology for through crack in pressurized fuselage structures. Int J Numer Methods Eng 1995;38:1633–1644], which consists of using four stress intensity factors, computed from the modified crack integral method, to predict the fatigue propagation life as well as the crack trajectory, which is computed as part of the numerical simulation. Some issues not presented in the study of Potyondy are investigated herein such as the influence of the crack increment size and the number of nodes per element (4 or 9 nodes) on the simulation results by means of a fatigue crack propagation simulation of a Boeing 737 airplane fuselage. The results of this simulation are compared with experimental results and those obtained by Potyondy [1].

Fatigue Testing of a Stiffened Lap Joint Curved Fuselage Structure

In April 1988, Aloha Airlines flight 243 experienced an explosive midair decompression that resulted in the separation of an 18-foot section of the fuselage crown of the Boeing 737 airplane. Investigations revealed that the linkup of small cracks emanating from multiple rivet holes in a debonded lap joint contributed to the catastrophic failure. This cracking scenario, known as multiple-site damage, is one of two sources of widespread fatigue damage; a type of structural degradation characterized by the simultaneous presence of fatigue cracks at multiple structural elements that are of sufficient size and density whereby the structure will no longer meet its damage tolerance requirement This study, sponsored by the National Aging Aircraft Research Program initiated by the Federal Aviation Administration in response to the Aloha accident, investigates multiple-site damage initiation and growth behavior in a pristine narrow-body fuselage panel. The test panel, a curved 6 x 10 ft stiffened structure containing six frames, seven stringers, and a longitudinal lap joint, was tested at the Federal Aviation Administration Full-Scale Aircraft Structural Test Evaluation and Research facility. The panel was subjected to a fatigue test with constant-amplitude cyclic loading, simulating the major modes of load associated with fuselage pressurization. Nondestructive inspections were conducted during the fatigue test to detect and monitor crack formation and growth. Multiple-site damage cracks were visually detected after about 80% of the fatigue life. Cracks developed and linked in the upper rivet row of the lap joint in the outer skin layer and formed a long fatigue crack before the termination of the fatigue test A residual strength test was then conducted by subjecting the panel to quasi-static loads until catastrophic failure. Fractographic examinations were conducted to reconstruct crack growth history. Preliminary results show multiple crack origins and significant subsurface crack growth.

Effects of crack closure on fatigue crack-growth predictions for 2024-T351 aluminum alloy panels under spectrum loading

Effects of plasticity induced crack closure on the calculated fatigue crack-growth behaviour of middle-crack tension specimens made of 2024-T351 alloy subjected to variable amplitude fatigue loading were examined. The specified loadings were four distinct aircraft spectra: (1) “modified” center-wing spectrum performed as part of this study; (2) fighter aircraft maneuver spectrum; (3) lateral gust spectrum; and (4) pressurized fuselage gust and maneuver spectrum. The experimentally observed crack-growth lives were compared with crack-growth predictions obtained using the crack closure concept in FASTRAN. Fatigue crack-growth material data extracted from the NASGRO materials database and the literature follows two distinct paths possibly due to different transitions from flat-to-slant (single or double shear) fracture. The material baselines for the FASTRAN code were selected to fit the two different paths. The two baseline curves were used to predict crack-growth life under four spectra to assess how the observed differences in the crack-growth rate trajectories affect the results. Using the lower bound baseline curve, which is suspected to be characterized by the double shear fracture mode, the FASTRAN model was able to predict crack- growth behaviour within ±25% of the test data for all four spectra.

Actuator Disk for Helicopter Rotors in an Unstructured Flow Solver

Motivated by the demand for a fast flow simulation tool that takes the interaction between a rotor and a helicopter fuselage into account, an actuator disk boundary condition suited for helicopter rotors in forward flight has been implemented in the unstructured grid DLR TAU code. The time-averaged effect of the rotor, which accelerates the flow and adds energy to the fluid, is imposed using source terms in the Navier-Stokes equations. The actuator disk is located in the grid. The transfer of this approach, previously implemented in the structured grid DLR FLOWer code, involved adapting the strategy to the unstructured framework. It is shown that propeller simulation results are in agreement with FLOWer results and simple one-dimensional theory predictions. Moreover, the rotor in forward flight cases prove the robustness of the implementation and resemble FLOWer results. Further development involved testing the implementation in parallel mode, and a more sophisticated rotor force distribution is applied instead of a constant pressure jump. Finally, a comparison of the viscous flow field around the EC145 helicopter computed by TAU and FLOWer is performed. It shows that there is good agreement between the two codes in predicting the effect of the actuator disk on the fuselage pressure distribution.

Effect of Interference on the Mechanics of Load Transfer in Aircraft Fuselage Lap Joints

Much of the fatigue damage in aircraft structures can be linked to the stress concentration arising at the rivet/skin interface in fuselage lap-joints. Fatigue damage can degrade the strength of the structure and reduce structural integrity. The stress distribution around the rivet holes, which depends on several loading conditions, is therefore of prime importance. Critical manufacturing process variations must be taken into account to observe the effect on local stresses at the hole. This paper presents three-dimensional (3D) nonlinear finite element analyses to investigate the stress state at rivet holes in fuselage lap joints. Initially, a 3D single rivet model of the riveting process was developed to characterize the unsymmetric residual stress distribution resulting from rivet installation. Then a global three-rivet model of the fuselage lap-joint, which takes into account the residual stresses from rivet installation and fuselage pressurization, was analyzed and compared to observations available from teardown inspection. The models were then implemented to observe the effects of rivet interference, sealant, and drill shavings on the stress state. A multiaxial fatigue criterion was implemented to predict cycles to crack nucleation for the modeled parameters. The effect of underdriven rivets and sealant were observed to be the most critical on the stress state of the fuselage splice. Excellent comparison with the damage characterization of the fuselage lap-joint provides validation to the finite element model.

Residual strength prediction of damaged composite fuselage panel with R-curve method

This study applies the crack growth resistance curve (R-curve) method to predict the residual strength of a composite fuselage panel with discrete source damage. The R-curve is constructed from the energy release rates computed from cracked tensile plate test specimens at their failure loads. To predict the residual strength of a curved fuselage panel with discrete source damage using the R-curve method, G-curves of the energy release rate of the damaged fuselage panel need to be established. These G-curves are generated for various fuselage pressures over a range of crack lengths. Since this study found that the membrane stiffening effect could significantly reduce the energy release rate at the crack tip of a pressurized fuselage, geometrically nonlinear behavior was considered in the calculations. The residual strength of the damaged fuselage panel was determined by comparing the energy release rates with the “allowable-like” R-curve. The correlation between the crack growth predictions and the damage progression results obtained from experiments is very good. Therefore, the R-curve method has the potential to be a practical engineering method for predicting the residual strength of damaged fuselage panels.