Full Aircraft Configurations Research Articles

Efficient Global Aerodynamic Shape Optimization of a Full Aircraft Configuration Considering Trimming

Most existing aerodynamic shape optimization (ASO) studies do not take the balanced pitching moment into account and thus the optimized configuration has to be trimmed to ensure zero pitching moment, which causes additional drag and reduces the benefit of ASO remarkably. This article proposes an efficient global ASO method that directly enforces a zero pitching moment constraint. A free-form deformation (FFD) parameterization combing Laplacian smoothing method is implemented to parameterize a full aircraft configuration and ensure sufficiently smooth aerodynamic shapes. Reynolds-averaged Navier–Stokes (RANS) equations are solved to simulate transonic viscous flows. A surrogate-based multi-round optimization strategy is used to drive ASO towards the global optimum. To verify the effectiveness of the proposed method, we adopt two design optimization strategies for the NASA Common Research Model (CRM) wing–body–tail configuration. The first strategy is to optimize the configuration without considering balance of pitching moment, and then manually trim the optimized configuration by deflecting the horizontal tail. The second one is to directly enforce the zero pitching moment constraint in the optimization model and take the deflection angle of the horizontal tail as an additional design variable. Results show that: (1) for the first strategy, about 4-count drag-reducing benefits would be lost when manually trimming the optimal configuration; (2) the second strategy can achieve 3.2-count more drag-reducing benefits than the first strategy; (3) compared with gradient-based optimization (GBO), surrogate-based optimization (SBO) is more efficient than GBO for ASO problems with around 80 design variables, and the benefit of ASO achieved by SBO is comparable to that obtained by GBO.

Drag Decomposition Using Partial-Pressure Fields: Ventus-3 and NASA Common Research Model

The development of partial-pressure field (PPF) theory (Schmitz, S “Drag Decomposition Using Partial-Pressure Fields in the Compressible Navier–Stokes Equations,” AIAA Journal, Vol. 57, No. 5, 2019, pp. 2030–2038) and the application of the concept to conditions relevant to commercial transport aircraft, operating in both the subsonic and transonic regimes (Hart, P. L., and Schmitz, S., “Drag Decomposition using Partial-Pressure Fields: ONERA M6 Wing,” AIAA Journal, Vol. 60, No. 5, 2022 pp. 2941–2951), have shown promise as complements to classical far-field methods. In the present work, a further comprehensive analysis of drag decomposition methods is applied to wing and aircraft configurations. First, the Ventus-3 sailplane wing case demonstrates the ability of PPFs to successfully deconstruct drag on industry-relevant cases including transition modeling and winglets. A quantitative comparison to classical aerodynamic theory and far-field methods supports PPF theory and provides additional insight to viscous–inviscid interaction by means of a viscous interactional aerodynamics term in PPF theory. Second, the NASA Common Research Model is simulated in transonic flow to showcase drag decomposition of full aircraft configurations where PPF applications are used to provide further insight into wave drag sources acting on the aircraft in comparison to classical far-field decomposition methods.

Low Speed Aerodynamic Analysis of the N2A Hybrid Wing–Body

Reduction of atmospheric emissions is currently a mandatory requirement for aircraft manufacturers. Several studies performed on Blended Wing–Body configurations showed a promising capability of reducing fuel consumption by increasing, at the same time, passengers’ transport capabilities. Although several aerodynamic studies are available at transonic speeds, low-speed evaluations of aerodynamic performances of Blended Wing Body aircrafts are less investigated. In this framework, the present paper deals with the aerodynamic performance of the N2A aircraft prototype at low-Mach number conditions. Aircraft longitudinal aerodynamics is addressed at M∞=0.2 with steady state three-dimensional RANS simulations carried out at two Reynolds numbers equal to 6.60×106 and 1.27×108, respectively. The former refers to an experimental test campaign performed at NASA Langley 14-by-22 foot subsonic tunnel, while the latter is related to free-flight conditions close to an approach and landing phase. Flowfield simulations are performed using the Computational Fluid Dynamic code FLUENT and the SU2 open-source code, currently adopted for research applications. Numerical solutions are validated by using available experimental data with reference to lift, drag, pitching moment and drag polar estimations. Pre-stall and post-stall aerodynamic behaviour through mean flow-field visualization along with the comparison of pressure distributions at several AoAs is addressed. Furthermore, the effect of convective discretization on a numerical solution for SU2 is discussed. Results indicate a good agreement with available experimental predictions. The present study aims to bridge existing computations at a Eulerian low-Mach number, with RANS computations and constitutes a further test-case for SU2 code with respect to a full aircraft configuration.

Solid wall and open boundary conditions in hybrid recursive regularized lattice Boltzmann method for compressible flows

Complex geometries and open boundaries have been intensively studied in the nearly incompressible lattice Boltzmann method (LBM) framework. Therefore, only few boundary conditions for the high speed fully compressible LBM have been proposed. This paper deals with the definition of efficient boundary conditions for the compressible LBM methods, with the emphasis put on the newly proposed hybrid recursive regularized D3Q19 LBM (HRR-LBM) with applications to compressible aerodynamics. The straightforward simple extrapolation-based far-field boundary conditions, the characteristic boundary conditions, and the absorbing sponge layer approach are extended and estimated in the HRR-LBM for the choice of open boundaries. Moreover, a cut-cell type approach to handle the immersed solid is proposed to model both slip and no-slip wall boundary conditions with either isothermal or adiabatic behavior. The proposed implementations are assessed considering the simulation of (i) isentropic vortex convection with subsonic to supersonic inflow and outflow conditions, (ii) two-dimensional (2D) compressible mixing layer, (iii) steady inviscid transonic flow over a National Advisory Committee for Aeronautics (NACA) 0012 airfoil, (iv) unsteady viscous transonic flow over a NACA 0012 airfoil, and (v) three-dimensional (3D) transonic flows over a German Aerospace Center (DLR) F6 full aircraft configuration.

Active control of supersonic transport aeroelastic oscillations using high-fidelity equations

A higher order computational procedure to alleviate aeroelastic oscillations by using active control surfaces is addressed. Flow is modeled using the high-fidelity Navier–Stokes (NS) equations, compared to the current use of the low-fidelity linear theory to design active controls. An active control module is embedded into a general-purpose NS flow solver developed for full aircraft configurations. The structure is modeled using the modal equations of motion, and existing active control equations based on the energy approach are utilized. Previously established transonic small perturbation theory and wind tunnel experiments are considered for validating. The use of active controls for next generation supersonic transport aircraft is demonstrated.

DLR TAU Simulations for the Second AIAA Sonic Boom Prediction Workshop

The accurate prediction of the sonic boom is essential for the development of a future low-boom supersonic aircraft. In this paper, geometries and grids provided by the Second AIAA Sonic Boom Prediction Workshop are used to improve a best practice for predicting near-field pressure signatures with the DLR TAU code. The workshop-provided low-boom geometries with different complexities, such as an axisymmetric body, a delta-wing body, and a full aircraft configuration with a cruciform tail and an engine. Results for the DLR TAU simulations on the workshop-provided grids are presented for different grid refinement levels using both Euler and Reynolds-averaged Navier–Stokes simulations. In addition to the provided grids, mixed-element grids with varying setups are generated. The positions of the shocks and expansions are essentially identical for all medium to fine grids but vary for very coarse grids due to the surface triangulation. The pressure magnitude at shocks and expansions is higher for simulations on fine grids compared to coarse grids. It is shown that the influence of the viscosity on the pressure signatures is more significant than the influence of the grid resolution. The agreement between TAU solutions and other participants’ solutions is good.

Component-Based Geometry Manipulation for Aerodynamic Shape Optimization with Overset Meshes

Mesh generation for high-fidelity computational fluid dynamics simulations and aerodynamic shape optimization is a time-consuming task. Complex geometries can be accurately modeled using overset meshes, whereby multiple high-quality structured meshes corresponding to different aircraft components overlap to model the full aircraft configuration. However, from the standpoint of geometry manipulation, most methods operate on the entire geometry rather than on separate components, which diminishes the advantages of overset meshes. To address this issue, a geometry module is introduced that operates on individual components and automatically computes their intersections to update overset meshes during optimization. Reverse-mode automatic differentiation is applied to compute partial derivatives across this geometry module so that it fits into an optimization framework that uses a hybrid adjoint method (known as ADjoint) to efficiently compute gradients for a large number of design variables. By using these automatically updated meshes and the corresponding derivatives, the aerodynamic shape of the DLR-F6 geometry is optimized while allowing changes in the wing–fuselage intersection. Sixteen design variables control the fuselage shape, and 128 design variables determine the wing surface. Under transonic flight conditions, the optimization reduces drag by 15 counts (5%) as compared with the baseline design.

Framework of Multiblock Structured Grid Smoothing Algorithms for Aircraft Configurations

This paper presents the development of a new framework of multiblock structured elliptic grid smoothing for full-aircraft configurations. The framework is fully automated and applies different solvers on blocks based on their properties. Blocks with a solid wall are initialized by algebraic orthogonal grids and solved with a line SOR implicit elliptic solver. The remaining blocks are initialized by transfinite interpolation grids and solved with an SOR explicit elliptic solver. Both solvers apply spacing and curvature source terms. This combination results in orthogonal grids close to walls while imposing elliptic spacing and curvature control using variable relaxation factors in the field, including across the block boundaries. The orthogonality close to the solid walls is improved by using the properties of the initial algebraic orthogonal grids. The code is parallelized using Message Passing Interface for acceleration. The performance of the framework is demonstrated in a first step on 2D problems: a canonical test case and a multi-element high-lift airfoil. In a second step, the performance is analyzed on a full three-dimensional aircraft configurations (wing/body/engine) case at Bombardier Aerospace. The paper examines the effect of the proposed approach through various assessments of grid metrics and flow solver behaviors. In particular, the study highlights the improvements in reducing spurious drag through a far-field drag analysis.

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.

Nonlinear Modal Aeroservoelastic Analysis Framework for Flexible Aircraft

A nonlinear formulation in modal coordinates of the equations of motion of a flexible aircraft is presented. It relies on the projection of the intrinsic equations of geometrically nonlinear composite beams on the linear normal modes at a reference condition, which are coupled with two-dimensional unsteady aerodynamics. The resulting description is suitable for nonlinear dynamic analysis and control design, whereas the description in modal coordinates links directly to linear aeroservoelastic analysis methods. Results are presented on and compared to cantilever wings and full aircraft configurations previously studied in the literature. Linear control synthesis and closed-loop nonlinear simulations are finally explored on a highly flexible flying wing under large-amplitude discrete gusts. Results show the ability of the proposed framework to capture the nonlinear dynamics of the aeroelastic system, while providing a seamless integration with linear methods, as well as its strength in identification of the dominant contributors to the nonlinear response.

Aerodynamic Shape Optimization of Common Research Model Wing–Body–Tail Configuration

Wing shape is one of the main drivers of aircraft aerodynamic performance, so most aerodynamic shape optimization efforts have focused solely on the wing. However, the performance of the full aircraft configuration must account for the fact that the aircraft needs to be trimmed. Thus, to realize the full benefit of aerodynamic shape optimization, one should optimize the wing shape while including the full configuration and a trim constraint. To evaluate the benefit of this approach, we perform the aerodynamic shape optimization of the Common Research Model wing–body–tail configuration using gradient-based optimization with a Reynolds-averaged Navier–Stokes model that includes a discrete adjoint implementation. We investigate the aerodynamic shape optimization of the wing with a trim constraint that is satisfied by rotating the horizontal tail. We then optimize the same wing–body configuration without the tail but with an added trim drag penalty based on a surrogate model we created before the optimization. The drag coefficient is minimized subject to lift and trim constraints. We found that considering the trim during optimization is a better approach than using a fixed-wing moment constraint. We also show that the trim drag surrogate model we created yields a minimum drag coefficient that is within 1.2 counts of the minimum drag coefficient obtained by rotating the tail to satisfy the trim constraint. However, we recommend rotating the tail within the optimization process to obtain the best possible performance.

Multimission Aircraft Fuel-Burn Minimization via Multipoint Aerostructural Optimization

Aerodynamic shape and aerostructural design optimizations that maximize the performance at a single flight condition may result in designs with unacceptable off-design performance. While considering multiple flight conditions in the optimization improves the robustness of the designs, there is a need to develop a way of choosing the flight conditions and their relative emphases such that multipoint optimizations reflect the true objective function. In addition, there is a need to consider uncertain missions and flight conditions. To address this, a new strategy to formulate multipoint design optimization problems is developed that can maximize the aircraft performance over a large number of different missions. This new strategy is applied to the high-fidelity aerostructural optimization of a long-range twin-aisle aircraft with the objective of minimizing the fuel burn over all the missions it performs in one year. This is accomplished by determining 25 flight conditions and their respective emphases on drag and structural weight that emulate the fuel-burn minimization for over 100,000 missions. The design optimization is based on the computational fluid dynamics of a full aircraft configuration coupled to a detailed finite element model of the wing structure, enabling the simultaneous optimization of wing aerodynamic shape and structural sizing leading to optimal static aeroelastic tailoring. A coupled adjoint method in conjunction with a gradient-based optimizer enable optimization with respect to 311 design variables subject to 152 constraints. Given the high computational cost of the aerostructural analysis, kriging models are used to evaluate the multiple missions. The results show that the multipoint optimized design reduced the total fuel burn by 6.6%, while the single-point optimization reduced it by only 1.7%. This capability to analyze large numbers of flight conditions and missions and to reduce the multimission problem to a multipoint problem could be used with a few modifications to minimize the expected value of any objective function given the probability density functions of the flight conditions.

Boundary-Layer Stability Analysis of the Mean Flows Obtained Using Unstructured Grids

Boundary-layer stability analyses of mean flows extracted from unstructured-grid Navier–Stokes solutions have been performed. A procedure has been developed to extract mean flow profiles from the FUN3D unstructured-grid solutions for the purpose of stability analysis. Extensive code-to-code validations were performed by comparing the extracted mean flows as well as the corresponding stability characteristics to the predictions based on structured-grid mean flow solutions. Comparisons were made for a set of progressively complex geometric configurations ranging from a simple flat plate to a full aircraft configuration: a modified Gulfstream III with a natural laminar-flow glove. The results for the swept wing flow over the wing–glove assembly point to the need for stability analysis based on Navier–Stokes solutions or possibly fully three-dimensional boundary-layer codes when the underlying flow develops strong three-dimensionality. The effect of grid resolution, mean flow convergence, and low-order interpolation to a stability grid on metrics relevant to linear stability of the boundary-layer flow are also examined to provide guidelines for the use of both structured and unstructured grids in practical applications related to transition prediction for swept wing boundary layers.

통합된 상용 툴을 이용한 전투기급 항공기 개념설계

Automated design program using commercial process integration and optimization program was developed for conceptual design of fighter-class aircraft. Wind tunnel test data and performance analysis results were compared for the verification of analysis tool of this program, and the usefulness of the tool was found. After integration with radar cross section analysis tool, the correlation with configuration design variables of wing, tail and performance parameters was identified by design of experiment, and the optimized configuration for weight and RCS was derived from optimization of empty weight and average frontal RCS value. After parameter definition of fuselage, the program can be implemented for full aircraft configuration.

Scalable Parallel Approach for High-Fidelity Steady-State Aeroelastic Analysis and Adjoint Derivative Computations

Aeroelastic systems achieve the best performance when the aerodynamic shape and structural sizing are optimized concurrently, but such an optimization is challenging when high-fidelity aerodynamic and structural models are required. This paper addresses this challenge through several significant improvements. Fully coupled Newton–Krylov methods are presented for the solution of aerostructural systems and for the corresponding adjoint systems. The coupled adjoint method presented can compute gradients with respect to thousands of multidisciplinary design variables accurately and efficiently. This is enabled by several improvements in the computation of the multidisciplinary terms in the coupled adjoint. The parallel scalability of the methods is demonstrated for a full aircraft configuration using an Euler computational fluid dynamics model with more than state variables and a detailed structural finite element model of the wing with more than degrees of freedom. The coupled Newton–Krylov methods are shown to improve the convergence rate of both the aerostructural solution and the coupled adjoint derivative computations. Gradient computations of aerodynamic and structural functions with respect to both aerodynamic shape and structural sizing variables are verified, and scaling is demonstrated to variables. The accuracy and scalability of the presented methods make it possible to perform aerostructural optimizations of full aircraft configurations with respect to hundreds of external shape and structural sizing design variables, leading to optimal aeroelastic tailoring.

Linear-Frequency-Domain Predictions of Dynamic-Response Data for Viscous Transonic Flows

Determining the flutter boundaries for full aircraft configurations by time-accurately solving the Reynolds-averaged Navier–Stokes equations is prohibitive with respect to computational expense, as the unsteady aerodynamic loading must be predicted for a wide range of flight conditions, frequencies, and structural mode shapes. Nonetheless, there is an increasing demand to accurately predict flutter boundaries in the viscous transonic regime—a demand, which, until recently, could only be satisfied by high-fidelity Reynolds-averaged Navier–Stokes methods. Brought to application readiness over the last years, time-linearized/small-disturbance methods, however, have been shown to satisfy this demand as well. They retain the Reynolds-averaged Navier–Stokes method’s fidelity to a high degree, at a substantially reduced computational expense. Such a method is presented here on the basis of the TAU–Reynolds-averaged Navier–Stokes method. Denoted as the TAU linear-frequency-domain method, it is validated for both a standard transonic airfoil and a high-aspect-ratio-wing dynamic test case using rigid pitch modes. The response data obtained from the linear frequency domain are in good agreement with the experiment for a two-dimensional case. For the three-dimensional case, there are larger differences. More important, the linear-frequency-domain method is in excellent agreement to time-accurate Reynolds-averaged Navier–Stokes simulations. Depending on the linear-frequency-domain-employed-solution scheme, reductions in computational costs well beyond an order of magnitude are obtained. In addition, the limits of the so-called frozen-eddy-viscosity approach are established.

Improved methodologies for the design of maneuver for stability and control simulations

With many modern fighter aircraft experiencing unpredicted flight dynamics during flight tests, recent research has focused on developing methodologies for incorporating computational fluid dynamics into the aircraft development process. The goal of this approach is to identify configurations susceptible to stability and control issues early in the design process. Previous research has primarily focused on full aircraft configurations, however, to increase the rate of development the current study focused on a two-dimensional NACA0012 airfoil. The two-dimensional NACA0012 airfoil has the advantage of reducing the computational cost by orders of magnitude compared to full scale aircraft simulations, while still providing complicated aerodynamics at high angles of attack. Computationally predicted lift coefficients from a number of newly developed training maneuvers were used to generate reduced order aerodynamic loads models. For evaluation, these models were compared to generated static and dynamic validation data. Methods of improving both the computational training maneuver and the reduced order modeling approach are suggested.

Towards Real-Time Computational-Fluid-Dynamics-Based Aeroelastic Computations Using a Database of Reduced-Order Information

This paper describes a computational-fluid-dynamics-based computational methodology for fast on-demand aeroelastic predictions of the behavior of a full aircraft configuration at variable flight conditions and demonstrates its feasibility. The methodology relies on the offline precomputation of a database of reduced-order bases and models associated with a discrete set of flight parameters, and its training for an interpolation method suitable for reduced-order information. The potential of this near-real-time computational methodology for assisting flutter flight testing is highlighted with the aeroelastic identification of an F-16 configuration in the subsonic, transonic, and supersonic regimes.

Automatic Transition Prediction in Hybrid Flow Solver, Part 1: Methodology and Sensitivities

A hybrid Reynolds-averaged NavierA¢Â�Â�Stokes solver, a laminar boundary-layer code, and a fully automated local, linear stability code were coupled to predict the laminarA¢Â�Â�turbulent transition due to TollmienA¢Â�Â�Schlichting and crossflow instabilities using the eN method based on the two-N-factor approach. The coupled system was designed to be applied to three-dimensional aircraft configurations which are of industrial relevance. The transition prediction methodology provides two different approaches which are available to be used in different flow situations. Both approaches are described and tested in detail. The application of the complete coupled system to a two-dimensional two-element airfoil configuration and a three-dimensional generic full aircraft configuration is described and documented in this paper. The prediction of the laminarA¢Â�Â�turbulent transition lines was done in a fully automatic manner. It will be shown that complex aircraft configurations can be handled without a priori knowledge of the transition characteristics of the specific flow problem. The computational results are partially compared to experimental data. This article is the first of two companion papers: the first dealing with the transition prediction methodology and the second dealing with the practical application of the coupled system.

Comparisons of computational fluid dynamics solutions of static and manoeuvring fighter aircraft with flight test data

As the capabilities of computational fluid dynamics (CFD) to model full aircraft configurations improve, and the speeds of massively parallel machines increase, it is expected that CFD simulations will be used more and more to steer or in some cases even replace traditional flight test analyses. The mission of the US Air Force SEEK EAGLE office is to clear any new weapon configurations and loadings for operational use. As more complex weapons are developed and highly asymmetric loadings are requested, the SEEK EAGLE office is tasked with providing operational clearances for literally thousands of different flight configurations. High-fidelity CFD simulations employing the turbulent Navier—Stokes equations are in a prime position to help reduce some of the required wind-tunnel and/or flight test workload. However, these types of CFD simulations are still too time consuming to populate a full stability and control parameter database in a brute-force manner. This article reviews results previously published by the authors, which validate the ability of high-fidelity CFD techniques to compute static force and moment characteristics of aircraft configurations. A methodology to generate efficient but non-linear reduced-order aerodynamic loads models from dynamic CFD solutions, which in-turn may be used to quickly analyse various stability and control characteristics at a particular flight condition, is introduced, and the results based on the US Air Force F-16C fighter aircraft that exemplify the process are discussed.