Abstract: A fixed wing Vertical Take-Off and Landing (VTOL) aircraft for use as a Personal Air Vehicle (PAV) or unmanned vehicle. A first double-ended drive shaft engine is mounted sideways in the front of the fuselage to serve a first pair of ducted fans mounted at the ends of the front wing. A second double-ended drive shaft engine is mounted sideways in the rear of the fuselage to serve a second pair of ducted fans mounted on the rear fuselage. The ducted fans are rotatable from a horizontal orientation to a vertical orientation to permit the aircraft to take off and land as a VTOL or conventional aircraft, and to be flown as a conventional aircraft. A parachute is provided with inflation assistance to permit rapid low altitude deployment for a controlled descent of the aircraft in an emergency.

Abstract: A vertical takeoff and landing aircraft having a fuselage with three wings and six synchronously tilt-able propulsion units, each one mounted above, below, or on each half of the aforementioned three wings. The propulsion units are vertical for vertical flight, and horizontal for forward flight. The aircraft wings are placed such that the rear wing is above the middle wing which is placed above the front wing. The placement of each of the propulsion units relative to the center of gravity of the aircraft about the vertical axis inherently assures continued stability in vertical flight mode, following the loss of thrust from any one propulsion unit. The placement of the propulsion units, viewing the aircraft from the front, is such that each propulsion units' thrust wake does not materially disturb the propulsion unit to its rear. When engine driven propellers or rotors are utilized, flapped wing panels are attached outboard of the forward and/or rearward propulsion units to provide yaw control during vertical flight.

Abstract: A wingbeat forcing function and control method are presented that allow six-degree-of-freedom control of a flapping-wing micro air vehicle using only two actuators, each of which independently actuate a wing. Split-cycle constant-period frequency modulation with wing bias is used to produce nonzero cycle-averaged drag. The wing bias provides pitching-moment control and, when coupled with split-cycle constant-period frequency modulation, requires only independently actuated wings to enable six-degree-of-freedom flight. Wing bias shifts the cycle-averaged center-of-pressure locations of the wings, thus providing the ability to pitch the vehicle. Implementation of the wing bias is discussed, and modifications to the wingbeat forcing function are made to maintain wing position continuity. Instantaneous and cycle-averaged forces and moments are computed, cycle-averaged control derivatives are calculated, and a controller is developed. The controller is designed using a simplified aerodynamic model derived with blade-element theory and cycle averaging. The controller is tested using a simulation that includes blade-element-based estimates of the instantaneous aerodynamic forces and moments that are generated by the combined motion of the rigid-body fuselage and the flapping wings. Simulations using this higher-fidelity model indicate that the cycle-averaged blade-element-based controller is capable of achieving controlled flight.

Abstract: A Turboelectric Distributed Propulsion (TeDP) system differs from other propulsion systems by the use of electrical power to transmit power from the turbine to the fan. Electrical power can be efficiently transmitted over longer distances and with complex topologies. Also the use of power inverters allows the generator and motors speeds to be independent of one another. This decoupling allows the aircraft designer to place the core engines and the fans in locations most advantageous for each. The result can be very different installation environments for the different devices. Thus the installation effects on this system can be quite different than conventional turbofans where the fan and core both see the same installed environments. This paper examines a propulsion system consisting of two superconducting generators, each driven by a turboshaft engine located so that their inlets ingest freestream air, superconducting electrical transmission lines, and an array of superconducting motor driven fan positioned across the upper/rear fuselage area of a hybrid wing body aircraft in a continuous nacelle that ingests all of the upper fuselage boundary layer. The effect of ingesting the boundary layer on the design of the system with a range of design pressure ratios is examined. Also the impact of ingesting the boundary layer on off-design performance is examined. The results show that when examining different design fan pressure ratios it is important to recalculate of the boundary layer mass-average Pt and MN up the height for each inlet height during convergence of the design point for each fan design pressure ratio examined. Correct estimation of off-design performance is dependent on the height of the column of air measured from the aircraft surface immediately prior to any external diffusion that will flow through the fan propulsors. The mass-averaged Pt and MN calculated for this column of air determine the Pt and MN seen by the propulsor inlet. Since the height of this column will change as the amount of air passing through the fans change as the propulsion system is throttled, and since the mass-average Pt and MN varies by height, this capture height must be recalculated as the airflow through the propulsor is varied as the off-design performance point is converged.

Abstract: A multi-modal vehicle (“MMV”) 20 a - 20 d . The MMV 20 a - 20 d includes a fuselage 22 and a chassis 26 supporting at least three wheels 44 having deployed and stowed states. Extending away from the fuselage 22 is a canard wing system 28 and a main wing system 30 . The main wing system 30 includes an inboard portion 134 and an outboard portion 132 . The inboard portion 134 is pivotally connected to the fuselage 22 ; the outboard portion 132 is pivotally connected to the inboard portion 134 . The MMV 20 a - 20 d further includes a vertical thrust system 32 comprising a pair of ducted fans 100 that are incorporated into the fuselage 22 , and a dual-use thrust system 34 that is configured to transition between a first position for supplying vertical thrust and a second position for supplying a horizontal thrust. A controller 42 is configured to control the MMV operations, reconfigurations, or transitions.

Abstract: An innovative crashworthiness subfloor design concept was developed for middle or larger transport aircrafts. Closed-cell Rohacell-31 polymer foam blocks were introduced into the aircraft subfloor to enhance the energy absorption capability and mitigate peak acceleration during crash. Eight different configurations of foam blocks’ longitudinal distribution were considered. To provide a cost-effective method for assessing fuselage crashworthiness, compelling finite element models of an aircraft fuselage for these configurations were developed and vertical drops of the fuselage models were then simulated using the non-linear finite element code LS-DYNA. Impact responses, including failure behaviour, energy absorption and acceleration characteristics, were obtained and detailed effects of different foam block configurations on the crashworthiness was discussed. The analysis of the results showed that the optimal foam block length is 100–126 mm in our studied cases. The fuselage failure mode appeared to be sy...

Abstract: The GOAHEAD (Generation of an Advanced Helicopter Experimental Aerodynamic Database for CFD code validation) consortium was created in the frame of an EU-project in order to create an experimental database for the validation of 3D-CFD and comprehensive aeromechanics methods for the prediction of unsteady viscous flows This included the rotor dynamics for complete helicopter configurations, ie main rotor - fuselage - tail rotor configurations with emphasis on viscous phenomena like flow separation and transition from laminar to turbulent flow The wind tunnel experiments have been performed during two weeks in the DNW-LLF on a Mach-scaled model of a modem transport helicopter consisting of the main rotor, the fuselage, control surfaces and the tail rotor For the sake of controlled boundary conditions for later CFD validation, a closed test section has been used The measurement comprised global forces of the main rotor and the fuselage, steady and unsteady pressures, transition positions, stream lines, position of flow separation, velocity profiles at the test section inlet, velocity fields in the model wake, vortex trajectories and elastic deformations of the main and tail rotor blades

Abstract: Quadrangular tubes were used as a cabin floor's struts considering their excellent impact response performance to adequately utilise the energy absorbing ability of struts. The finite element model of civil aircraft under sound simplification was built to simulate its drop test. Several fuselage sections with different thicknesses and strut triggers were considered here. The acceleration history, failure behaviour and energy absorption ability of the civil aircraft were investigated. The result shows that the impact response of civil aircraft with different struts could be divided into two groups in these cases. The cabin floor collides with the hold floor for the first group, and there is a collision between the strut and the ground for the second one. The initial maximal peak acceleration increases with an increase in strut's thickness, and the triangular and quadrangular triggers are not recommended in these cases. When strut's thickness is 0.9 mm, the civil aircraft has the best impact performance, and it would get worse with increase or decrease in strut's thicknesses.

Abstract: A crash absorber element integrated in composite vertical (z-) struts of commercial aircraft fuselage structures was developed, which absorbs energy under crash loads by cutting the composite strut into stripes and crushing the material under bending. The design concept of this absorber element is described and the performance is evaluated experimentally in static, crash and fatigue test series on component and structural level under normal and oblique impact conditions. These tests highlight the robustness of the absorber design as this system worked under various conditions and angles with an impressively high reproducibility. The physics of the energy absorption by high rate material fragmentation are explained and numerical modelling methods in explicit finite element codes for the simulation of the crash absorber are assessed. The real physical fragmentation phenomena can just be approximated in simulations, emphasising that the numerical prediction of composite energy absorption for industrial use cases is still a big challenge.

Abstract: Collisions between birds and aircraft (bird strikes) pose safety risks to the public, cost airports and airlines money, and result in liability issues. Recent research suggests that aircraft visibility could be enhanced to increase detection and avoidance by birds. We questioned whether aircraft color scheme might play a role in bird-strike frequency. We used public records of bird strikes along with information on fl ights that were gathered by federal agencies in the United States. We estimated the bird-strike rates and compared them among airline companies using different fuselage color schemes, while controlling for aircraft type. Using an avian vision modeling approach, we fi rst corroborated the hypothesis that brighter colors would contrast more against the sky than darker colors. We found differences in bird-strike rates among airline companies with different color schemes in 3 out of the 7 aircraft types investigated: Boeing 737, DC-9, and Embraer RJ145. With each of these aircraft, we found that brighter aircraft were associated with lower bird-strike rates. Brighter fuselages might increase the contrast between the aircraft and the sky and enhance detection and avoidance behavior by birds. Our fi ndings are not conclusive but suggest a specifi c hypothesis and prediction about bird responses to aircraft with different color schemes that deserves empirical testing in the future.

Abstract: This paper documents a mixed-fidelity approach for the design of low-boom supersonic aircraft as a viable approach for designing a pract ical low-boom supersonic configuration. A low-boom configuration that is based on low-fidelit y analysis is used as the baseline. Tail lift is included to help tailor the aft portion of the g round signature. A comparison of low- and high-fidelity analysis results demonstrates the nec essity of using computational fluid dynamics (CFD) analysis in a low-boom supersonic configuration design process. The fuselage shape is modified iteratively to obtain a configuration with a CFD equivalent-area distribution that matches a predetermined low-boom target distribution. The mixed-fidelity approach can easily refine the low-fidelity low-boo m baseline into a low-boom configuration with the use of CFD equivalent-area analysis. The g round signature of the final configuration is calculated by using a state-of-the -art CFD-based boom analysis method that generates accurate midfield pressure distributions for propagation to the ground with ray tracing. The ground signature that is propagated from a midfield pressure distribution has a shaped ramp front, which is similar to the ground s ignature that is propagated from the CFD equivalent-area distribution. This result confi rms the validity of the low-boom supersonic configuration design by matching a low-boom equivalent-area target, which is easier to accomplish than matching a low-boom midfield pressure target.

Abstract: An arrangement for an aircraft includes a fuselage, a wing, and a propulsor. The wing may have a wing upper surface, a rear spar and a wing trailing edge. The propulsor may include at least one rotor having a rotor diameter and a rotor axis. The propulsor may be mounted such that the rotor is located longitudinally between the rear spar and the wing trailing edge. The propulsor may also be mounted such that a lowest point of the rotor diameter is located vertically above the wing upper surface.

Abstract: The present invention relates to an aircraft (110) for vertical take-off and landing. The aircraft comprises a first wing (101), a second wing (102) and a fuselage (103). The first wing (101) comprises a first longitudinal wing axis (104) and the second wing (102) comprises a second longitudinal wing axis (104). The first wing (101) extends along the first longitudinal wing axis (104) and the second wing (102) extends along the second longitudinal wing axis (104) from the fuselage (103). The first wing (101) is tiltable with a first rotational direction around the first longitudinal wing axis (104) and the second wing (102) is tiltable with a second rotational direction around the second longitudinal wing axis (104). The wings (101, 102) are adapted in such a way that, in a fixed- wing flight mode, the wings (101, 102) do not rotate around a second axis (105). The wings (101, 102) are further adapted in such a way that, in a hover flight mode, the wings (101, 102) are tilted around the longitudinal wing axis (104) with respect to its orientation in the fixed- wing flight mode and that the wing (100) rotates around the second axis (105).

Abstract: Composite materials are increasingly being used for aeronautic primary structures such as wing components or fuselage panels. However, their major drawback is their vulnerability against transversal impact loads, which may lead to internal delaminations or intralaminar fiber/matrix failure. Such loads may arise from numerous impact scenarios, with bird strikes being one of the most relevant load cases. The focus of the current study is on the numerical modeling and simulation of high velocity impact loads from soft body projectiles on composite structures with ABAQUS/explicit. At first, the impact on flat composite plates is studied in experiment and simulation, which allows for the validation of the modeling methods. Some of these plates have been preloaded in tension or compression in order to investigate the influence on the mechanical behavior. It could be shown that the preloading of the plate may have a significant influence on the structural response. As a second example, the bird impact on a composite wing leading edge is treated. Adequate modeling methods for the composite material (stacked shell model), delamination failure (cohesive elements), preloading (implicit-explicit coupling) and soft body impactor modeling (coupled Eulerian-Lagrangian approach) are assessed in this paper. The final simulation results correlate well with experimental test data.

Abstract: Random wall-pressure fluctuations due to the turbulent boundary layer (TBL) are a feature of the air flow over an aircraft fuselage under cruise conditions, creating undesirable effects such as cabin noise annoyance. In order to test potential solutions to reduce the TBL-induced noise, a cost-efficient alternative to in-flight or wind-tunnel measurements involves the laboratory simulation of the response of aircraft sidewalls to high-speed subsonic TBL excitation. Previously published work has shown that TBL simulation using a near-field array of loudspeakers is only feasible in the low frequency range due to the rapid decay of the spanwise correlation length with frequency. This paper demonstrates through theoretical criteria how the wavenumber filtering capabilities of the radiating panel reduces the number of sources required, thus dramatically enlarging the frequency range over which the response of the TBL-excited panel is accurately reproduced. Experimental synthesis of the panel response to high-speed TBL excitation is found to be feasible over the hydrodynamic coincidence frequency range using a reduced set of near-field loudspeakers driven by optimal signals. Effective methodologies are proposed for an accurate reproduction of the TBL-induced sound power radiated by the panel into a free-field and when coupled to a cavity.

Abstract: The development of a solar electric powered UAV (Unmanned Aerial Vehicle) is systematically introduced in this paper. The purpose of the project is to prove the feasibility of some crucial engineering techniques for future full-scale high altitude solar powered UAVs. The solar electric powered UAV adopted a normal configuration of a sailplane with wing of high aspect ratio, slender fuselage, and a tail boom. Aerodynamic geometry and longitudinal trim and stability performances were designed via lifting line theory, and verified using AVL code based on vortex lattice method. Composite and balsa combined structure was designed for the purpose of light weight and rigidness. Carbon-balsa-carbon laminated structure was used for the D-box of the wing and fuselage, which gave adequate strength and rigidness to the wing to carry solar array. Fragile photovoltaic cells were bended and mounted onto the surface of the wing, and would be working cooperatively with lithium-polymer secondary battery to drive the propulsion system. A miniature onboard power management device was developed to adjust and monitor the output power of PV cells and secondary battery. A set of ground tests and flight tests were carried out, the aerodynamic performance, power supply ability of solar array, function and efficiency of the power management device were all demonstrated and validated. The test results showed that the efficiency of the power management device was greater than 85%, the input power requirement for straight and level flight was less than 80 W, and the minimum solar radiation sufficient to support straight and level flight solely was 650 W/m 2 . All of the design objectives had been achieved.

Abstract: Aircraft fuselage section (32) subjected to impacts of external bodies, the aircraft fuselage having a curved shape with at least a vertical symmetry plane (A-A) and a central longitudinal axis and comprising a skin (35) and a plurality of frames (37) arranged perpendicularly to said longitudinal axis (33), the aircraft fuselage section also comprising at least an inner reticular structure (51, 53) mounted on a supporting structure (41, 43) comprising longitudinal beams (39) attached to the skin (35) and interconnected with said frames (37), said inner reticular structure (51, 53) being arranged for creating at least one closed cell (75) with the skin (35) for improving its resistance and its damage tolerance to said impacts. Said inner reticular structure (51, 53) can be formed by panels (61), rods (65, 65′), cables (67, 67′) or belts (69, 69′).

Abstract: The utility model relates to a twin-engine vertical take-off and landing fixed-wing unmanned aerial vehicle, which is characterized in that: the whole body adopts a canard pneumatic configuration, and canard winglets are positioned at the front side of a main fuselage; main wings are positioned at the back side of the fuselage, and two engines with propellers are arranged at the tips of the main wings; backward swept H-shaped vertical tails are simultaneously arranged on nacelles of the engines at the wing tips, and take-off and landing pillars are arranged at the tops of the H-shaped vertical tails; and movable rudder surfaces are arranged on the back edges of the H-shaped vertical tails, and elevons are arranged at the outer sides of the back edges of the main wings. The unmanned aerial vehicle can vertically take off and land in a very small space like a helicopter, also can hover over the air to complete special observation and reconnaissance missions, and also can fly as fast as a fixed-wing aircraft. The unmanned aerial vehicle does not need an airport runway normally used by the fixed-wing aircraft, without catapults and parachutes, so the use is convenient, and the cost is lower; and the whole structure is simple and the manufacturing cost is low, and so the unmanned aerial vehicle has obvious cost price advantages compared with the helicopter, a tiltrotor and the like.

Abstract: A sonar buoy includes a fuselage having a tube-like shape, one or more wings coupled to the fuselage, an engine coupled to the fuselage and operable to propel the sonar buoy through flight, and a guidance computer operable to direct the sonar buoy to a predetermined location. The sonar buoy further includes a sonar detachably coupled to the fuselage and forming at least a part of the fuselage, and a rocket motor detachably coupled to the fuselage. The one or more wings are operable to be folded into a position to allow the sonar buoy to be disposed within a launch tube coupled to a vehicle and to automatically deploy to an appropriate position for flight after the sonar buoy is launched from the launch tube. The rocket motor propels the sonar buoy from the launch tube and detaches from the fuselage after launch.

Abstract: The quantitative assessment of the effects of damage and material degradation on the dynamic aeroelastic reliability of composite aircraft has become more challenging in recent years due to the increased use of composites in main structural components of passenger aircraft. This paper presents a method for quantifying the flutter reliability of aircraft composite structures in the presence of multiple uncertainties. Automated rapid simulation tools for predicting flutter speeds of composite airframes subject to multiple uncertainties serve a key role and are used in Monte Carlo simulations. A thorough discussion is presented of the flutter uncertainty of composite aircraft covering material degradation, damage, repair, and design and certification practices, as well as maintenance procedures. The effectiveness of the method and its potential are illustrated using a composite tail/rudder structure representing a typical passenger aircraft structure. Conclusions are drawn and recommendations for future work are made.

Abstract: A systematic and general investigation about box wing aircraft is conducted, including aerodynamic and performance characteristics. The design of a promising medium range box wing aircraft based on the Airbus A320 taken as reference aircraft is performed. The design is taken through the general steps in aircraft preliminary design. The fuel consumption of the final aircraft is 9 % lower than that of the reference aircraft. The aircraft layout is well balanced regarding the position of the center of gravity and the travel of the center of gravity is minimized. This is necessary due to the aircraft’s particular characteristics concerning static longitudinal stability and controllability. The low wing tank capacity requires an additional fuselage tank. Because of its high span efficiency the aircraft has a glide ratio of 20,4. Its wing is about twice as heavy as the reference wing. This is partly compensated by a lighter fuselage.

Abstract: This paper describes the development and ground validation of a variable-span morphing wing intended be fitted to a small UAV prototype. The vehicle flies in the speed range 11m/s to 40m/s. The wing model is designed with the help of graphical CAD/CAM tools and then a full scale prototype is built for preliminary bench testing the wing/actuator system. The wing is built in composite materials and is made of two parts. The inboard part is fixed to the fuselage and uses a monocoque skin construction. The outboard part slides inside the inboard part to change the span of the wing and uses a typical structure made of spar, ribs and thin skin. An electro-mechanical actuation mechanism is developed using an aluminum rack and pinion system driven by two servomotors placed at center of the wings. Bench tests, performed to evaluate wing under load, showed that the system is capable of performing the required extension/retraction cycles and is suitable to be installed on a small UAV prototype.

Abstract: An over-the-wing-nacelle-mount airplane configuration is known to prevent the noise propagation from jet engines toward ground. However, the configuration is assumed to have low aerodynamic efficiency due to the aerodynamic interference effect between a wing and a nacelle. In this paper, aerodynamic design optimization is conducted to improve aerodynamic efficiency to be equivalent to conventional under-the-wing-nacelle-mount configuration. The nacelle and wing geometry are modified to achieve high lift-to-drag ratio, and the optimal geometry is compared with a conventional configuration. Pylon shape is also modified to reduce aerodynamic interference effect. The final wing-fuselage-nacelle model is compared with the DLR F6 model to discuss the potential of Over-the-Wing-Nacelle-Mount geometry for an environmental-friendly future aircraft.

Abstract: Applications of friction stir welding in a fuselage structure were studied. Samples with two different friction-stirweld orientations in the fuselage panelwere tested: one is along the fuselage longitudinal direction and the other one is along the fuselage circumferential direction. Then fatigue cracks were investigated that were set in three different types: parallel and perpendicular to friction stir welds and between doublewelds. Sample geometries weremachined from identical welds in order to remove the effect of theweld process on fatigue behavior. Tests were conducted onM (T) specimens with either longitudinal or transverse welds. Cracks growing into or growing away from the weld center, as well as in the nugget zone, were investigated. It is shown that fatigue crack growth for cracks growing away from the center seems similar to that of the parentmaterial; for crack growing in the nugget, crack grows slower than in the parent material; and for cracks starting between double welds, the crack grows slower than in the other two types. The virtual crack-closure technique method was used to calculate stress intensity factor from residual stress (Kres) and effective R ratio in an attempt to explain the experimental findings.

Abstract: A wide aerodynamic test campaign has been carried out on the tiltrotor aircraft ERICA at the Large Wind Tunnel of Politecnico di Milano by means of a modular 1:8 scale model in order to produce a dataset necessary to better understand the aerodynamic behaviour of the aircraft and to state its definitive design. The target of the tests was the measurement of the aerodynamic forces and moments in several different configurations and different attitudes. The test program included some conditions at very high incidence and sideslip angles that typically belong to the helicopter-mode flight envelope and measurements of forces on the tail and on the tilting wings. A large amount of data has been collected that will be very useful to refine the aircraft design. In general the aircraft aerodynamics do not present any critical problems, but further optimisation is still possible. From the viewpoint of drag in the cruise configuration, the sponsons of the landing gear seem to be worth some further design refinement since they are responsible for a 20% drag increase with respect to the pure fuselage configuration. On the contrary, the wing fairing has proved to work well when the aircraft longitudinal axis is aligned with the wind, providing just a slight drag increase. Two other interesting aspects are the quite nonlinear behaviour of the side force for the intermediate sideslip angles as well as the noticeable hysteresis in the moment coefficient at very high incidence angles.

Abstract: An aircraft having a variable geometry for adapting the flight characteristics to different flight situations includes a fuselage with a pair of wings projecting on both sides of the fuselage in the transverse direction (y), each of which wings has an inner wing section arranged stationarily with respect to the fuselage and an outer wing section adjacent thereto and pivotable about a pivot axis. The pivot axis is oriented in a direction deviating from the longitudinal direction (x) of the aircraft by a maximum of 40°.

Abstract: The increasing use of composite materials in aircraft structures aims in reducing the structural weight significantly. In order to exploit the advantages of composite materials especially within a large-scale optimization calculation, a model for a computationally efficient structural analysis needs to be developed. In this regard, algorithms need to be developed to rapidly compute stress distribution and critical loads for both strength and stability of the composite aircraft fuselage. Therefore, sizing methods for metallic and composite orthotropically stiffened fuselage structures have been reviewed. For critical load computation, fibre fracture and inter-fibre fracture need to be taken into consideration with respect to strength. Regarding stability, the critical buckling loads of skins and stringers as well as the critical crippling load need to be taken into account. The buckling of stringers often occurs after the skin buckling load is exceeded. Hence, the postbuckling behaviour needs to be analyzed and load redistributions in the postbuckling range have to be taken into consideration. These load redistributions can generally be calculated numerically using either the finite element method or the finite strip method (Mocker and Reimerdes in Compos Struct 73:237–243, 2006) as well as analytically. In order to minimize the computational time, the postbuckling behaviour of the skins regarded as composite plates is computed analytically within this work by means of a computation of effective stiffnesses for global analysis and local failure load computation. Even though the postbuckling behaviour of metallic and composite plates has been widely studied in literature, only few work has been spent on the analytical or semi-analytical derivation of methods for the common load case of combined compression and shear loading. As the preliminary design of an aircraft fuselage requires a rapid and sufficiently accurate description of the postbuckling behaviour, the postbuckling behaviour of an orthotropic composite plate under combined compression and shear loading is analytically analyzed within the present work. The derived rapid sizing method for postbuckling significantly reduces the computational time when compared to the computational time needed for a nonlinear finite element computation. In this regard, it even allows for the consideration of postbuckling behaviour within a large-scale optimization computation of complete fuselage structures.