Aircraft Geometry Research Articles

Effects of Detailed Aircraft Geometry on Wake Vortex Dynamics During Landing

The final approach and touchdown of a landing aircraft is studied numerically with different hybrid Reynolds-averaged Navier–Stokes (RANS)/large-eddy simulations (LES) approaches. Particular emphasis is put on the ground effect and the wake vortex evolution before and after touchdown. Additionally, the effect of the landing gear on the aircraft’s aerodynamics during landing is analyzed. The near field around the landing aircraft is generated using two different RANS methods. In a number of steady-state simulations the altitude of the aircraft above ground was varied in fixed steps. Furthermore, a simplified continuous landing maneuver was simulated by applying unsteady RANS simulations. The resulting RANS fields are then coupled with an LES simulation using three different approaches. The computed vortex circulation is compared with aircraft measurements. As a result it can be seen that the new transient hybrid approach provides physically reasonable results. The landing gears have a notable effect on the aircraft’s drag, as well as on the wake footprint. However, the effect on the wake vortex circulation evolution is moderate. Remarkably, the simulations with landing gear show a strong coherent vortex, merging late with the main vortex. A particular emphasis is put on the comparison with light detection and ranging (LIDAR) measurement data collected in a recent campaign at Vienna airport. The vortex structures predicted by the simulations are clearly visible in those measurements. Simulated tangential velocity profiles are compared with vortex model profiles. It is observed that standard vortex models fail to capture the complex vortex structures emerging during the landing process but can be improved including mirror vortices.

Aircraft Geometry and Meshing with Common Language Schema CPACS for Variable-Fidelity MDO Applications

This paper discusses multi-fidelity aircraft geometry modeling and meshing with the common language schema CPACS. The CPACS interfaces are described, and examples of variable fidelity aerodynamic analysis results applied to the reference aircraft are presented. Finally, we discuss three control surface deflection models for Euler computation.

Knowledge-based automatic Airframe Design using CPACS

Abstract The CPACS data format [Nagel et al. (2012), CPACS Homepage (2017)] has long been established as the primary means of data exchange in preliminary aircraft design projects within DLR. As described by Scherer et al. (2016), it contains a wide range of options for describing the structural layout of a design including frames and stringers, floors, bulkheads, etc. Based on these descriptions, several finite element model generators comparable to the one described by Walther et al. (2017) have been implemented, which can provide detailed computational structural models of a given design. However, all model generators require the information on the structural layout to be available in CPACS upfront. Within a larger aircraft design context, this necessitates the augmentation of the description of the structure to a given plain aircraft geometry. So far, this has been accomplished through a manual process, which not only results in an increased risk of errors, but also prohibits the exposure of parameters to a larger multidisciplinary optimization. In the presented paper, a newly developed knowledge-based airframe augmentation module will be introduced. Implemented in Python, it provides methods to automatically initialize a full structural layout on a given CPACS geometry, based on a manageable number of control parameters. In addition to an outline of the governing design rules, several application cases will also be given.

Wind-Tunnel and CFD Investigations Focused on Transition and Performance Predictions of Laminar Wings

This paper is a continuation of the paper recently published by Hue et al. (“Experimental and Numerical Methods for Transition and Drag Predictions of Laminar Airfoils,” AIAA Journal, Vol. 53, No. 9, Sept. 2015, pp. 2694–2712) that focused on transition and drag predictions of laminar airfoils. The extension of such studies to three-dimensional configurations representative of modern civil aircraft is a further step toward the implementation of natural laminar-flow technologies. The present work, therefore, focuses on validating the laminar design of a low-swept wing for business jet applications. In 2015, an experimental campaign was carried out in the European Transonic Windtunnel, and included laminar/turbulent transition measurements with temperature-sensitive paint at Mach and Reynolds numbers typical of cruise flight conditions. Subsequently, fluid dynamics computations were performed on this aircraft geometry either with a Reynolds-averaged Navier–Stokes solver using both Tollmien–Schlichting and crossflow transition criteria, or with a boundary-layer code using either database methods for transition location or exact stability analyses. In this paper, experimental and numerical transition predictions are compared for three representative cases corresponding to different angles of attack. The agreement that is achieved is satisfactory, and extended regions of laminar flow are observed on the wing at cruise lift levels. In these conditions, the drag reduction can account for 10–15% of the aircraft drag.

Commercial space operations noise and sonic boom issues

Commercial space transportation is a rapidly developing industry worldwide. The expansion of the space transportation infrastructure creates additional challenges to the National Airspace System. New launch facilities have and are being developed some of which co-locate with commercial airports. Rocket launch community noise impact requires adequate assessment and mitigation. New space vehicles encompass a wide range of design, aircraft geometry, and flight parameters. A database of operational launch profiles for rockets does not exist at this moment. The acoustical characteristic of these new vehicles may also differ from that of existing rockets or conventional aircraft. Spacecraft ascent sonic boom signatures have higher overpressure, longer duration and a bow shock. There are no standard methodologies for the environmental assessment of launch vehicles and sites. Developing and validating these models is an emerging field. In addition, the metrics used for community noise assessment for commercial aircraft are being used for commercial space operations and may not be appropriate for such use. The existing gaps in rocket launch community noise assessment, technical and regulatory requirements, and current steps the FAA is undertaking to establish an environmental regulatory framework for the commercial space operations will be discussed.

Computational-Fluid-Dynamics Best Practices for Aircraft Wing-Tip Vortex Roll-Up

Best practices for simulating aircraft wing-tip vortex roll-up with Reynolds-averaged Navier–Stokes computational fluid dynamics were developed using a build-up approach with the NASA OVERFLOW flow solver. Methods were tested for accuracy and computational expense beginning with a simple case, then building up step by step in complexity toward a realistic aircraft geometry. The cases tested started two-dimensionally with an inviscid vortex, then a viscous vortex, then three-dimensionally with a basic rectangular finite wing, a semispan wing that included validating wind-tunnel test data, and finally to a realistic wing/body aircraft configuration that also included wind-tunnel data. Comparisons were also made with algebraic vortex models used to initialize large-eddy simulations of wake decay.

Efficient Parameterization of Waverider Geometries

This paper summarizes the results of investigations into the development of parametric waverider geometry models, with emphasis on their efficiency, in terms of their ability to cover a large feasible design space with a sufficiently small number of design variables to avoid the “curse of dimensionality.” The work presented here is focused on the parameterization of idealized waverider forebody geometries that provide the baseline shapes upon which more sophisticated and realistic hypersonic aircraft geometries can be built. Three different aspects of rationalizing the decisions behind the parametric geometry models developed using the osculating cones method are considered. Initially, three different approaches to the design method itself are discussed. Each approach provides direct control over different aspects of the geometry for which very specific shapes would be more complex to obtain indirectly, thus enabling the geometry to more efficiently meet any related design constraints. Then, a number of requirements and limitations are investigated that affect the available options for the parametric design-driving curves of the inverse design method. Finally, the performance advantages that open up with increasing flexibility of the design-driving curves in the context of a design optimization study are estimated. This allows one to reduce the risk of overparameterizing the geometry model, while still enabling a variety of meaningful shapes. Although the osculating cones method has mainly been used here, most of the findings also apply to other similar inverse design algorithms.

Responses and Injuries to PMHS in Side-Facing and Oblique Seats in Horizontal Longitudinal Sled Tests per FAA Emergency Landing Conditions.

The objective of the present exploratory study is to understand occupant responses in oblique and side-facing seats in the aviation environment, which are increasingly installed in modern aircrafts. Sled tests were conducted using intact Post Mortem Human Surrogates (PMHS) seated in custom seats approximating standard aircraft geometry. End conditions were selected to represent candidate aviation seat and restraint configurations. Three-dimensional head center-of-gravity linear accelerations, head angular velocities, and linear accelerations of the T1, T6, and T12 spinous processes, and sacrum were obtained. Three-dimensional kinematics relative to the seat were obtained from retroreflective targets attached to the head, T1, T6, T12, and sacrum. All specimens sustained spinal injuries, although variations existed by vertebral level. While the tension mechanism was associated with cervical spine injuries, complex distraction-coupled with bending and tension was attributed to thoracolumbar spine injuries. Skeletal fractures to the ribcage were attributed to compression induced by the restraint from the seatbelt, the presence of the armrest, and/or severe motions of the unconstrained torso. Pelvic injuries were also attributed to restraint offered by the lap belt on the accelerating torso-pelvis complex in the absence of the armrest. Lower extremity injuries occurred due to the unconstrained motion (flailing mechanism). These results serve as an initial dataset to understand the kinematics of different body regions, injuries and patterns, and potential injury mechanisms describing PMHS responses in the aviation environment.

An unstructured quadrilateral mesh generation algorithm for aircraft structures

Because commercial aircraft are built with thin-walled structures, their structural performance is well-modeled using shell-element meshes. However, creating these meshes for the full aircraft configuration can be challenging and presents a bottleneck in the design process, especially in a configuration-level design space. This paper presents an algorithm that automatically creates unstructured quadrilateral meshes for the full airframe based on just the description of the desired structural members. The approach consists in representing each node in the mesh as a linear combination of points on the geometry so that the structural mesh morphs as the geometry changes, as it would, for example, in aerostructural optimization. The algorithm divides the aircraft skin into 4-sided domains based on the underlying B-spline representation of the geometry. It meshes each domain independently using an algorithm based on constrained Delaunay triangulation, triangle merging and splitting to obtain a quadrilateral mesh, and elliptical smoothing. Examples of full-configuration structural meshes are provided, and a mesh convergence study is performed to show that element quality is maintained as the structural mesh is refined. The algorithm is available as part of the open-source aircraft geometry tool suite, GeoMACH.

Aerodynamic Loads over Arbitrary Bodies by Method of Integrated Circulation

A novel numerical method for the conversion of unstructured surface vorticity into directional integrated circulation has been presented. This method allows for the reduction of vorticity distributions on arbitrary three-dimensional bodies into their mathematically equivalent formulations in the Prandtl lifting-line theory. A discussion of its implementation into a practical numerical solver has been presented. This new solver was applied to simple wing shapes as well as complex aircraft geometries to demonstrate the applicability of such an approach to flow solvers. It was shown that the reduction to lifting-line theory was exact for simple shapes such as rectangular and elliptical wings. For complex aircraft geometries, such as leading-edge root extension and canard-equipped aircraft, it was shown that the method of integrated circulation allowed for accurate prediction of aircraft loads and the analysis of flow physics, in keeping with the Prandtl original concepts.

Buckling Alleviation for Joined-Wing Aircraft

A novel mechanism is described that increases the critical buckling load of the rear wing of joined-wing aircraft in an effort to reduce structural weight. The benefits of the joined-wing configuration are briefly discussed, along with its current limitations. The behavior of the modified structure under lift is investigated by conducting a number of linear structural analyses for varying geometric arrangements. It is shown that the mechanism, called the buckling alleviation component, is effective in reducing structural weight for a significant proportion of aircraft geometries, even though there is a tradeoff between the rear wing buckling load and front wing root bending moment. A more detailed structural model is used to investigate the effect of geometric structural nonlinearities. The nonlinear analysis shows that the buckling alleviation component delays the onset of nonlinear behavior and that the critical buckling loads of the conventional configuration are drastically overestimated by the linear analysis, which means that the buckling alleviation component performs much more effectively than the initial linear analysis suggests.

Crash Testing of a CFRP Commercial Aircraft Sub-Cargo Fuselage Section

The paper addresses the testing at the ONERA (Office National d’Etudes et de Recherches Aérospatiales) crash tower of a sub-cargo demonstrator representative of new generation CFRP (carbon fiber reinforced polymer) commercial aircraft. The project was conducted within a technological platform coordinated by AIRBUS Germany, which aimed at designing, simulating and testing an innovative crashworthy sub-cargo structure. Technical activities were shared between AIRBUS Germany for the demonstrator design and manufacturing, DLR (German Aerospace Center) for the numerical analysis and ONERA for the crash test. The demonstrator was based on single aisle aircraft geometry and comprised 2 Integrated Cargo Units (ICU) equipped with Triggered Tube Segments (TTS) dedicated to energy absorption and CFRP stringer-stiffened skin. The crash concept was based on an integrated structural design which applied the “bend-frame-concept” where the cargo cross-beam acts as a bend frame and withstands the dynamic loads introduced by the TTS components. The testing configuration - loading system and instrumentation - was defined on the basis of numerical analysis performed by DLR at the fuselage section level. In that frame, a kinematic model with a 2-frames typical fuselage section and ICUs involving the “bend-frame” concept was numerically simulated, with the main objective to identify the loading conditions that apply at specific sections, notably those surrounding the ICU-frame coupling areas where the test fixtures were to be implemented. As the outcomes of these numerical works showed that bending/compression loading, at a specific ratio, shall be targeted in priority, the accordingly designed loading system thus consisted of articulated rigs maintaining both ends of the demonstrator. The testing was performed at the ONERA-Lille crash tower at a 6,7m/s impact velocity, with a 1050 kg trolley mass. The acquisition system cumulated a total of 48 channels, including force sensors (6), strain gauges (36), displacement laser sensors (5) and an accelerometer (1). Besides, 4 high-speed cameras were implemented to visualize the rupture phenomenon likely to develop during the crash test. Results confirmed the expected crash scenario, with the bending of the cargo cross-beams and the resulting progressive crushing of the TTS components.

Spatial Simulation of Contrail Formation in Near-Field of Commercial Aircraft

During this study, three-dimensional numerical simulations of contrails generated by commercial aircraft in cruise conditions were performed. The objective was to study the early development of contrails in the near field of an aircraft, including engine core and bypass flows. Computational fluid dynamics simulations, based on the three-dimensional Reynolds-averaged Navier–Stokes approach, were carried out on a realistic aircraft geometry. A microphysical model was implemented in the computational fluid dynamics code to simulate particle growth using an Eulerian approach. Results showed that the mixing processes in the bulk plume were in good agreement with the literature. Then, the early growth of a contrail in the near wake of the aircraft was investigated for two ambient relative-humidity conditions. This parameter played a significant role in contrail local properties. Increased relative humidity led to an increase in the fraction of particles under the supersaturated condition and condensation rate. This resulted in a higher mean particle radius and, as a consequence, in a higher optical depth, making contrails more distinct. Particle-size distribution shifted toward larger particles due to an increase in available vapor and got narrower as the relative humidity increased because more particles were in favorable conditions for ice growing.

A parametric study of fixed-wing aircraft perching maneuvers

This paper presents a trajectory optimization-based study for identifying the influences of different fixed-wing aircraft parameters on perching maneuvers. Perching trajectories, which involve climbing for landing with a near-zero speed, are optimized to minimize the unwanted altitude gain, herewith referred to as undershoot. Undershoot has a complex relationship not only with the control inputs, but also with the aircraft geometry and features, such as stability characteristics, controllability, and thrust available. To facilitate this parametric study, an aerodynamic modeling tool is developed, which combines a nonlinear vortex correction method with the Leishman's state-space form of Wagner's unsteady model. This tool is used for in-flight aerodynamics evaluation during the optimization of the perching maneuvers. The perching solutions – minimum undershoot and trajectory length – are then computed, and the effects of key parameters of the maneuver are studied. It is found that a high aspect ratio, high thrust available, optimal placement of the center of gravity, relaxed terminal velocity constraints, low zero-lift drag, and headwind reduce the spatial requirements for the maneuver. Furthermore, it is confirmed that a high angle-of-attack maneuver is not critical for very low speed perching.

Comparison of Digital DATCOM and Wind Tunnel Data of a Winged Hybrid Airship’s Generic Model

Use of low fidelity tools in designing subscale generic wind tunnel models is usually required to get first-hand knowledge about general trends of aerodynamic and stability parameters. Most of such tools are limited to well-defined conventional aircraft configurations. In the present work, aerodynamic and stability characteristics of a winged hybrid airship is explored at low subsonic speed by using Aircraft Digital DATCOM, which is based on semi-empirical methods for preliminary aircraft geometries. Spheroidal ellipsoidal shaped hull of the airship is modeled in DATCOM along with the geometrical details of wing and empennages. The prediction of zero lift drag coefficient, coefficient of lift and pitching moment is the focus of this paper. Except the drag coefficients, trends of analytical results compare well with experimental data.

A methodology for vehicle and mission level comparison of More Electric Aircraft subsystem solutions: Application to the flight control actuation system

As part of the More Electric Initiative, there is a significant interest in designing energy-optimized More Electric Aircraft, where electric power meets all non-propulsive power requirements. To achieve this goal, the aircraft subsystems must be analyzed much earlier than in the traditional design process. This means that the designer must be able to compare competing subsystem solutions with only limited knowledge regarding aircraft geometry and other design characteristics. The methodology presented in this work allows such tradeoffs to be performed and is driven by subsystem requirements definition, component modeling and simulation, identification of critical or constraining flight conditions, and evaluation of competing architectures at the vehicle and mission level. The methodology is applied to the flight control actuation system, where electric control surface actuators are likely to replace conventional centralized hydraulics in future More Electric Aircraft. While the potential benefits of electric actuation are generally accepted, there is considerable debate regarding the most suitable electric actuator – electrohydrostatic or electromechanical. These two actuator types form the basis of the competing solutions analyzed in this work, which focuses on a small narrowbody aircraft such as the Boeing 737-800. The competing architectures are compared at both the vehicle and mission levels, using as metrics subsystem weight and mission fuel burn, respectively. As shown in this work, the use of this methodology aids the decision-making process by allowing the designer to rapidly evaluate the significance of any performance advantage between the competing solutions.

Aerodynamic Interference Issues in Aircraft Directional Control

This work investigates the aerodynamic interference among airplane components caused by rudder deflection for a typical turboprop aircraft geometry through the computational fluid dynamics technique. At no sideslip, an airplane is in symmetric flight conditions. The rudder deflection creates a local sideslip angle close to the vertical tailplane, and this effect is increased by fuselage and horizontal tail. Typical semiempirical methods, such as United States Air Force Stability and Control Data Compendium (USAF DATCOM), do not take into account for these effects, proposing the same corrective parameters both for pure sideslip and rudder deflection conditions. Numerical analyses executed on several aircraft configurations with different wing and horizontal tailplane positions show that the interference factors are smaller than those predicted by the USAF DATCOM procedure, providing guidelines for a more accurate aircraft directional control analysis and hence rudder preliminary design.

From Spreadsheets to Simulation-Based Aircraft Conceptual Design

An automated, simulation-based aircraft design process allows for the prediction of unanticipated problems early inthe design stage, leading to reduced turn-around time and development cost. Having reliable, and affordable (fast)design tools is crucial to achieving this level of automation in design process. An example of this is illustrated fora jet trainer aircraft using two aircraft design codes: Jet Designer and CEASIOM. A set of aerodynamic methodswith different degrees of fidelity and computational expense is considered, with the limitations of each methodprovided. In particular, this paper examines the challenges that CFD-based aircraft design poses to a designer,including: a) the cost of generation of large data tables, b) the automated handling of geometry, c) treating controlsurface deflections, d) and calculation of dynamic derivatives using CFD. A Kriging-based sampling approachwas used for generating aerodynamic tables with a reasonable computational cost compared with a brute-forceapproach. For Euler calculations, an automated CAD and mesh generation approach from a geometry descriptionwas used. It is demonstrated that application of Euler solutions to low fidelity aircraft geometry shows the expecteddesign trends. Also, results show that the wave drag at transonic speeds can be predicted with Euler equations, butnot with vortex lattice or Digital DATCOM. The treatment of control surface deflections was also investigated forthe vortex lattice solver and the Euler code. Transpiration boundary condition approach was used in the Euler codeto model the flap surface movements, although this approach is limited to small control surface deflections. Thecalculated aero tables form each aero source were used next to study the vehicle flying qualities. Results presenteddemonstrate the validity and feasibility of the simulation-based approach for aircraft conceptual design.

Sonic-Boom Mitigation Through Aircraft Design and Adjoint Methodology

This paper presents a novel approach to design of the supersonic aircraft outer mold line by optimizing a A-weighted loudness-based objective of the sonic-boom signature predicted on the ground. The optimization process uses the sensitivity information obtained by coupling the discrete adjoint formulations for the augmented Burgers equation and computational-fluid-dynamics equations. This coupled formulation links the loudness of the ground boom signature to the aircraft geometry, thus allowing efficient shape optimization for the purpose of minimizing the loudness. The accuracy of the adjoint-based sensitivities is verified against sensitivities obtained using an independent complex-variable approach. The adjoint-based optimization methodology is applied to a configuration previously optimized using alternative state-of-the-art optimization methods and produces additional loudness reduction. The results of the optimizations are reported and discussed.

Study on Aircraft Lift Estimation

This chapter describes how lift can be derived from a known aircraft geometry and conversely, how the selection of aircraft layout influences the generation of lift. As the fuselage shape is predetermined by the payload requirements, the nacelle geometry by the engine size, and the tail surfaces by stability considerations, the wing parameters are the principal variables to be considered in respect to lift aspects. To this end, guidance is given on the selection of wing section profile and platform geometry to provide specified performance requirements. A lift estimation method covering all the aircraft components is also provided. Such methods are suitable for use in the conceptual design phase and provide an introduction to more detailed estimation procedures.