Abstract: Current turbulence models, such as those employed in Reynolds-averaged Navier-Stokes CFD, are unable to reliably predict the onset and extent of the three-dimensional separated flow that typically occurs in wing-fuselage junctions. To critically assess, as well as to improve upon, existing turbulence models, experimental validation-quality flow-field data in the junction region is needed. In this report, we present an overview of experimental measurements on a wing-fuselage junction model that addresses this need. The experimental measurements were performed in the NASA Langley 14- by 22-Foot Subsonic Tunnel. The model was a full-span wing-fuselage body that was configured with truncated DLR-F6 wings, both with and without leading-edge extensions at the wing root. The model was tested at a fixed chord Reynolds number of 2.4 million, and angles-of-attack ranging from -10 degrees to +10 degrees were considered. Flow-field measurements were performed with a pair of miniature laser Doppler velocimetry (LDV) probes that were housed inside the model and attached to three-axis traverse systems. One LDV probe was used to measure the separated flow field in the trailing-edge junction region. The other LDV probe was alternately used to measure the flow field in the leading-edge region of the wing and to measure the incoming fuselage boundary layer well upstream of the leading edge. Both LDV probes provided measurements from which all three mean velocity components, all six independent components of the Reynolds-stress tensor, and all ten independent components of the velocity triple products were calculated. In addition to the flow-field measurements, static and dynamic pressures were measured at selected locations on the wings and fuselage of the model, infrared imaging was used to characterize boundary-layer transition, oil-flow visualization was used to visualize the separated flow in the leading- and trailing-edge regions of the wing, and unsteady shear stress was measured at limited locations using capacitive shear-stress sensors. Sample results from the measurement techniques employed during the test are presented and discussed.

Abstract: This paper introduces the software TiGL: TiGL is an open source geometry modeler that is used in the conceptual and preliminary aircraft and helicopter design phase. It creates full three-dimensional models of aircraft from their parametric CPACS description. Due to its parametric nature, it is typically used for aircraft design analysis and optimization. First, we present the use-case and architecture of TiGL. Then, we discuss its geometry module, which is used to generate the B-spline based surfaces of the aircraft. The backbone of TiGL is its surface generator for curve network interpolation, based on Gordon surfaces. One major part of this paper explains the mathematical foundation of Gordon surfaces on B-splines and how we achieve the required curve network compatibility. Finally, TiGL’s aircraft component module is introduced, which is used to create the external and internal parts of aircraft, such as wings, flaps, fuselages, engines or structural elements.

Abstract: Shape control is a critical task in the composite fuselage assembly process due to the dimensional variabilities of incoming fuselages. To realize fuselage shape adjustment, actuators are used to pull or push several points on a fuselage. Given a fixed number of actuators, the locations of actuators on a fuselage will impact on the effectiveness of shape control. Thus, it is important to determine the optimal placement of actuators in the fuselage shape control problem. In current practice, the actuators are placed with equal distance along the edge of a fuselage without considering its incoming dimensional shape. Such practice has two limitations: (1) it is non-optimal and (2) larger actuator forces may be applied for some locations than needed. This paper proposes an optimal actuator placement methodology for efficient composite fuselage shape control by developing a sparse learning model and corresponding parameter estimation algorithm. The case study shows that our proposed method achieves the optimal actuator placement for shape adjustments of the composite fuselage.

Abstract: In the present paper, advanced numerical methodologies have been adopted to investigate the influence of impact angle on the crashworthiness behavior of a composite fuselage section. The analyzed fuselage section, made of unidirectional fiber-reinforced material, woven fabric material, and aluminum material, is representative of a regional aircraft fuselage. Two different angles of impact with rigid ground have been investigated and reported: Perpendicularly to the ground and with a pitch angle of 3 degrees with respect to the ground. The adopted numerical models have been preliminarily validated with experimental data from a drop test on a full-scale fuselage section, in terms of deformations and failure location and progression. The correlation between the numerical model and the experimental test has enabled evaluation of the effect of the impact angle on the deformation and damage in the sub-cargo floor area.

Abstract: A numerical study has been conducted to investigate the unsteady nature of afterbody vortices shed from an upswept region of a simple fuselage configuration. The upsweep angle considered is ϕ=28 d...

Abstract: Recent studies have focused on unsteady aerodynamics of an aircraft using computational fluid dynamics. It was found that a change in the density of air within the flight domain influences the aerodynamic performance of aircrafts in flight. This is as a result of the occurrence of turbulence and unsteadiness in the air velocity vectors. Limited work has focused on the impact of transient change in the density of air to lift, drag and moment coefficient acting on the aircraft fuselage during flight. This paper is aimed at determining how the unsteadiness in the density of air influences the aerodynamic performance of an aircraft with time. To determine the aerodynamic performance of an aircraft within this unsteady airflow condition, lift, drag coefficient and the wing design efficiency was used as the deterministic parameter for the investigation. A compressible simulation was performed for transonic flows with Mach number 0.84. The wing geometry was designed using SOLIDWORKS and the CFD simulation was performed using ANSYS Fluent version 18.1 software. A User Defined Function (UDF) was developed using a MATLAB code. This was aimed at, to introduce a time-dependent change in the density of air, in the Fluent environment. Results obtained showed that a 0.24 average change in air density caused a drop in lift coefficient by an amount 0.01014. The wing design efficiency achieved in this study was 56% and this value is low. Therefore, a change in the density of air will cause the aircraft to observe a poor aerodynamic behavior and eventually stall.

Abstract: The vibration loads of a helicopter rotor are transmitted to the fuselage through gearbox support struts and cause severe vibration of the fuselage. A piezoelectric stack actuator has the advantage...

Abstract: The flow taking place in the rear part of the fuselage during the emergency landing on water is investigated experimentally in realistic conditions. To this aim, tests on a double curvature specimen have been performed at horizontal velocities ranging from 21 m/s to 45 m/s. Test data highlight different cavitation and/or ventilation modalities which are strongly dependent on the horizontal velocity, with substantial changes in the flow features occurring with velocity variations of few meters per second. For the specimen considered here, the inception of the cavitation is found at about 30 m/s, confirming that scaled model tests performed at small horizontal velocities are unable to capture the hydrodynamics correctly. By analyzing pressure data, underwater movies, and force measurements, it is shown that the transition from the cavitation to ventilation condition has a significant effect on the longitudinal distribution of the loading which, together with inertia, aerodynamic loads, and engine thrust, governs the aircraft dynamics.

Abstract: Camber morphing of the wing outer mold line may improve aerodynamic performance and reduce noise emissions for some aircraft. However, not all outer mold lines are obtainable from a given reference...

Abstract: The use of fiber-metal laminates (FML) allows for substantial advantages over a fuselage skin made of monolithic aluminum materials. The combination of glass-fiberreinforced plastic and aluminum is characterized by low fatigue, high load tolerance and the resistance to residual stress. For this reason, FML, and GLARE in particular, have been identified as superior materials for aerospace applications. It has already been used extensively in the wide body aircraft of the Airbus Group A380, specifically on the upper fuselage shells. FML possess the potential to become the baseline material for next-generation single-aisle aircrafts. The development of a new production chain that will allow automated fuselage production for future short-haul aircrafts is the focus of the studies that make up the joint project AUTOGLARE. As part of the fifth call-up for the German Aeronautical Research Programme (LuFo), DLR is working with its project partners Airbus Operations, Premium Aerotech (PAG) and the Fraunhofer Gesellschaft (FhG). The development of a production chain for stiffened fuselage panels based on fiber-metal-laminates as a material is aimed at allowing a scaling-up to 60 aircrafts per month. This study contains the research work of the DLR and FhG regarding the automated and quality assured process for chain stiffened FML fuselages. In Addition to a detailed explanation of the systems that were set up, this paper covers the planned tests, the completed demonstration models and the findings derived from them.

Abstract: In the case of catastrophic events, such as an emergency landing, the fuselage structure is demanded to absorb most of the impact energy preserving, at the same time, a survivable space for the passengers. Moreover, the increasing trend of using composites in the aerospace field is pushing the investigation on the passive safety capabilities of such structures in order to get compliance with regulations and crashworthiness requirements. This paper deals with the development of a numerical model, based on the explicit finite element (FE) method, aimed to investigate the energy absorption capability of a full-scale 95% composite made fuselage section of a civil aircraft. A vertical drop test, performed at the Italian Aerospace Research Centre (CIRA), carried out from a height of 14 feet so to achieve a ground contact velocity of 30 feet/s in according to the FAR/CS 25, has been used to assess the prediction capabilities of the developed FE method, allowing verifying the response under dynamic load condition and the energy absorption capabilities of the designed structure. An established finite element model could be used to define the reliable crashworthiness design strategy to improve the survival chance of the passengers in events such as the investigated one.

Abstract: This paper focuses on the multiscale fatigue life analysis of an aircraft fuselage panel in order to investigate the damage tolerance performance. The two-dimensional multiscale analyses were carried out under different levels of external loads and based on linear elastic fracture mechanics. For this, a micro model consisting of a square plate with a central circular hole and two cracks, from pre-established initial damages and subjected to uniaxial tension, was modelled using the computer program BemCracker2D–an academic computational package for crack growth analysis based on the dual boundary element method (DBEM). The objective is to compute the number of fatigue cycles for each load increment and relate them to their respective compliance in local micro elements. Then, the results were treated statistically with Monte Carlo simulation to ensure the integrity of the fuselage and, therefore, avoiding reaching a Limit State during the design lifespan. The main advantages of using the DBEM formulation implemented in BemCracker2D are easy pre- and post-processing, simplicity of meshing and satisfactory accuracy.

Abstract: Composite laminate stiffened panels are often used in aircraft fuselage design because of their favorable properties. To assess the failure load of these thin-walled structures and to exploit their...

Abstract: An aircraft has a fuselage, a wing assembly coupleable to the fuselage, and an empennage including a pair of tail booms configured to be removably coupled to the wing assembly. The wing assembly includes a pair of boom interfaces located on laterally opposite sides of the fuselage. Each tail boom has a boom forward end configured to be mechanically attached to one of the boom interfaces using an externally-accessible mechanical fastener.

Abstract: During the impact process, fuselage frame structures often experience severe crushing-induced kinematic deterioration. To improve cabin safety, this paper developed an analysis and design algorithm...

Abstract: This paper proposes an optimisation process for the design of aircraft fuselage stiffeners using evolutionary optimisation. A new design problem is developed to find a layout for fuselage stiffeners (rings and stringers) such that the structural mass, compliance, and the first-mode natural frequency can be optimised, subject to structural constraints. The stiffeners are modelled as beam elements. Three multiobjective meta-heuristics are employed to solve the problem, and a comparative study of the results of these optimisers is carried out. It is found that the proposed layout synthesis problem for aircraft fuselage stiffeners leads to a set of efficient structural solutions, which can be used at the decision-making stage. It is an automated design strategy with high potential for further investigation.

Abstract: Fatigue plays a significant role in the crack growth of the fuselage skin structures. In addition, the fuselage may suffer also from the corrosion damage, and the wear defects. The proper maintenance and scheduled test intervals can avoid the sudden skin failure. Therefore, the inspection interval has to be shortened. Nevertheless, the young machines may be also suffering from the unexpected skin rupture. The cracks are emanating from the rivets and the holes under cyclic loading. The stress concentration around the notch has an effective role under the effect of cyclic loading. The cracks propagate toward the high stressed area such as the notches or other crack locations. The propagation into a critical crack size is rather fast and causes a sudden aircraft fuselage cracking. Hence, the number of cycles to failure will be decreased dramatically. During the last decades, the fracture toughness, design, and the new alloying element have been enhanced. The previous fuselage failures show that the inspections against the cracking are recommended even after a few thousand of cycles. To prevent the crack extending, the crack arresting is recommended to use around the fuselage.

Abstract: In this paper, a turbo-electric propulsion system was analyzed, and its performance was assessed. The aircraft considered here was a single-aisle, medium-range configuration targeting a capacity of 150 Pax. The propulsion concept comprised two boosted geared turbofan engines mounted under-wing. Those main engines were supported by an electrically driven aft-propulsor contributing to the thrust generation and by taking advantage of ingesting the boundary layer of the fuselage for potentially higher levels of propulsive efficiency and allowing for the improved operation of the main engines. The performance assessment as carried out in the context of this paper involved different levels: Firstly, based on the reference aircraft and the detailed description of its major components, the engine performance model for both main engines, as well as for the electrically driven aft-propulsor was set up. The methodology, as introduced, has already been applied in the context of hybrid-electric propulsion and allowed for the aforementioned aircraft sizing, as well as the subsequent gas turbine multi-point synthesis (simulation). A geared turbofan architecture with 2035 technology assumptions was considered for the main engine configuration. The present trade study focused on the design and performance analysis of the aft-propulsor and how it affected the performance of the main engines, due to the electric power generation. In order to allow for a more accurate description of the performance of this particular module, the enhanced streamline curvature method with an underlying and pre-optimized profile database was used to design a propulsor tailored to meet the requirements of the aft propulsor as derived from the cycle synthesis and overall aircraft specification; existing design expertise for novel and highly integrated propulsors could be taken advantage of herein. The resulting performance characteristics from the streamline curvature method were then fed back to the engine performance model in a closely coupled approach in order to have a more accurate description of the module behavior. This direct coupling allowed for enhanced sensitivity studies, monitoring different top-level parameters, such as the thrust/power split between the main engines and the aft propulsor. As a result, different propulsor specifications and fan designs with optimal performance characteristics were achieved, which in return affected the performance of all subsystems considered.

Abstract: Electric propulsion (EP) systems offer considerably more degrees of freedom (DOFs) within the design process of aircraft compared to conventional aircraft engines. This requires large, computationally expensive design space explorations (DSE) with coupled models of the single components to incorporate interdependencies during optimization. The purpose of this paper is to exemplarily study these interdependencies of system key performance parameters (KPIs), e.g., system mass and efficiency, for a varying DC link voltage level of the power transmission system considering the example of the propulsion system of the CENTRELINE project, including an electric motor, a DC/AC inverter, and the DC power transmission cables. Each component is described by a physically derived, analytical model linking specific subdomains, e.g., electromagnetics, structural mechanics and thermal analysis, which are used for a coupled system model. This approach strongly enhances model accuracy and simultaneously keeps the computational effort at a low level. The results of the DSE reveal that the system KPIs improve for higher DC link voltage despite slightly inferior performance of motor and inverter as the mass of the DC power transmission cable has a major share for a an aircraft of the size as in the CENTRELINE project. Modeling of further components and implementation of optimization strategies will be part of future work.

Abstract: Aim of this paper is to investigate the effects of the turbulent flow developing over a fuselage on fan noise for Boundary Layer Ingestion (BLI) embedded propulsion systems. Such engine configurations can suffer from inlet flow distortions and ingestion of turbulence at the fan plane with consequent impact on both broadband and tonal fan noise. The analysis is performed by considering a modified version of the Low-Noise configuration of the NASA Source Diagnostic Test (SDT) integrated into the Nextgen ONERA Versatile Aircraft (NOVA) fuselage in order to reproduce the NOVA BLI configuration. The numerical flow solution is obtained by solving the explicit, transient and compressible lattice-Boltzmann equation implemented in the high-fidelity CFD/CAA solver Simulia PowerFLOW R®. The acoustic far-field is computed by using the Ffwocs-Williams & Hawkings integral solution applied to a permeable surface encompassing the fan-stage. Simulations are performed for an operating condition representative of a take-off with power cutback. Installation effects due to the BLI configuration are quantified by comparison with an isolated configuration of the modified Low-Noise SDT fan-stage geometry at same operating conditions. Comparisons are carried out in terms of fan-stage intake/interstage velocity fields, fan blades section air-loads and far-field noise; correlations between the fan-stage velocity field and noise emission for the BLI configuration are outlined. It is found that the BLI fan-stage is characterized by strong azimuthal fan blade loading unsteadiness, less periodic and coherent rotor wake tangential velocity variations and higher levels of in-plane velocity fluctuations compared to the isolated engine, resulting in far-field noise spectra with no distinct tonal components and higher broadband levels. This study represents the first high-fidelity CFD/CAA simulation of a full-scale aircraft geometry comprehensive of a BLI fan/Outlet Guide Vane (OGV) stage.