Natural Laminar-Flow Airfoil Optimization Design Using a Discrete Adjoint Approach

In the present work, the aerodynamic shape design of an advanced high-lift system for a natural laminar flow (NLF) wing, based on the combination of a morphing droop nose and a single slot trailing edge flap, is presented. The paper presents both the aerodynamic design and optimization of the NLF wing and the high-lift configuration considering the mutual effects of both flap devices. Concerning the morphing droop nose (DN), after defining the parameterization techniques adopted to describe the geometry in terms of morphing shape and flap settings, the external configuration is obtained by an aerodynamic shape optimization procedure able to meet geometrical constraints and the skin structural requirements due to the morphing. The final performance assessment of the three-dimensional high-lift configurations is performed by high-fidelity aerodynamic analyses. The design procedure is applied to a twin-prop regional aircraft equipped with a natural laminar flow wing. The morphing droop nose is compatible with an NLF wing that requires the continuity of the skin and, at the same time, extends the possibilities to improve the performances of the class of regional aircraft which usually are not equipped with conventional leading edge devices. Additionally, the morphing technology applied to the flap allows the design of a tracking system fully integrated inside the airfoil geometry, leading to a solution without external fairings and so with no extra friction drag penalty for the aircraft.

Aerodynamic shape optimisation technology is presented, comprising an efficient variable fidelity shape parameterisation method, an efficient andhigh quality mesh deformation scheme, and a parallel optimisation algorithm. The objective of the research presented here is the comparison of truly three-dimensional optimisations of aircraft wings in both aerodynamic and aeroelastic environments. The novel shape parameterisation technique allows various fidelities of design parameters, ranging from detailed surface changes to novel truly three-dimensional planform adjustments. An efficient interpolation scheme, using radial basis functions, transfers domain element movements into direct deformations of the design surface and corresponding CFD mesh, thus allowing total independence from the grid generation package and type (structured or unstructured). Optimisation is independent from the CFD flow solver by obtaining sensitivity information for an advanced parallel gradient-based optimiser by finite-differences. This ‘wrap-around’ optimisation technique is applied to a modern large transport aircraft wing in the cruise flight condition for minimum drag with stringent constraints in lift, volume, and two root moments. The objective of all optimisations is aerodynamic, however the static aeroelastic deflection provided by an aeroelastic solver will give that particular optimisation increased accuracy and real world relevance. The result of a constrained inviscid aerodynamic optimisation is presented and has a significant reduction in drag when compared to the initial wing with no violation of any constraints. The shape parameterisation method demonstrates that only a low number of design variables are necessary to achieve innovative planform and surface geometries with dramatically improved performance. Computational fluid dynamics (CFD) methods are now commonplace in aerospace industries, and at the forefront of analysis capabilities, providing a fast and effective method of predicting a design’s aerodynamic performance. However, with ever increasing complexity of designs, engineers can often struggle to interpret the intricacies of the CFD results sufficiently to be able to manually alter the geometry to improve performance. Hence, there has been an increase in demand for intelligent and automatic shape optimisation schemes. This requires combining geometry control methods with numerical optimisation algorithms, to provide a mechanism to mathematically seek improved and optimum designs, using CFD as the analysis tool. Optimisation requires consideration of three issues, each of which have numerous solutions: shape parameterisation including CFD surface and volume mesh deformation, computation of design variable derivatives, and effective use of these derivatives to improve design. Geometry parameterisation is critical for effective shape optimisation. This is the method of representing the design surface, and defines the degrees of freedom in which the geometry can be altered and, ideally, this should be linked with an effective method of deforming the CFD surface and volume mesh in a corresponding fashion. Parameterising complex shapes is a problem that remains a serious obstacle to both manual and automatic CFD-based optimisation. A wide variety of shape control and morphing methods have been developed, but many do not allow sufficiently

Adjoint-based aerodynamic shape optimization including transition to turbulence effects

Effect of biofouling roughness on a marine propeller's performance including cavitation and underwater radiated noise (URN)

Numerical noise is an inevitable by-product of Computational Fluid Dynamics (CFD) simulations which can lead to challenges in finding optimum designs. This article draws attention to the issue, illustrating the difficulties it can cause for road vehicle aerodynamics simulations. Firstly a benchmark problem is used to assess a range of turbulence models and grid types. Large noise amplitudes up to 22% are evident for solutions computed on unstructured tetrahedral grids whereas computations on hexahedral and polyhedral grid structures exhibit substantially less noise. The Spalart Allmaras turbulence model is shown to be far less susceptible to noise levels than two other commonly-used models for this application. Secondly, multi-objective aerodynamic shape optimization is applied to a fairing for a practical road vehicle which is parameterised in terms of three design variables. Moving Least Squares (MLS) metamodels are constructed from 50 high-fidelity CFD solutions for two objective functions. Subsequent optimization is successful for the first objective, however numerical noise levels in excess of 7% give rise to difficulties for the second one. A revision to the problem leads to success and the construction of a small Pareto front. Further analysis underlines the inherent capability of MLS metamodels in dealing with noisy CFD responses.

Adaptive morphing trailing-edge wings have the potential to reduce the fuel burn of transport aircraft. However, to take full advantage of this technology and to quantify its benefits, design studies are required. To address this need, the aerodynamic performance benefits of a morphing trailing-edge wing are quantified using aerodynamic design optimization. The aerodynamic model solves the Reynolds-averaged Navier–Stokes equations with a Spalart–Allmaras turbulence model. A gradient-based optimization algorithm is used in conjunction with an adjoint method that computes the required derivatives. The baseline geometry is optimized using a multipoint formulation and 192 shape design variables. The average drag coefficient is minimized subject to lift, pitching moment, geometric constraints, and a maneuver bending moment constraint. The trailing edge of the wing is optimized based on the multipoint optimized wing. The trailing-edge morphing is parameterized using 90 design variables that are optimized independently for each flight condition. A total of 407 trailing-edge optimizations are performed at different flight conditions to span the entire cruise flight envelope. A 1% drag reduction at on-design conditions and a 5% drag reduction near off-design conditions are observed. The effectiveness of the trailing-edge morphing is demonstrated by comparing it with the optimized results of a hypothetical fully morphing wing. In addition, the fuel-burn reductions for a number of flights are computed using the optimization results. A 1% cruise fuel-burn reduction is achieved using an adaptive morphing trailing edge for a typical long-haul twin-aisle mission.

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.

This paper aims to develop an efficient global optimization method for design of transonic natural-laminar-flow (NLF) airfoils and wings, based on high-fidelity computational fluid dynamics (CFD) solver. The CFD solver features functionality of automatic transition prediction, by coupling Reynolds-averaged Navier-Stokes (RANS) equations with the linear-stability-theory-based dual e N method for Tollmien-Schlichting and crossflow instabilities. An A320-sized transonic NLF wing with a laminar supercritical airfoil is designed for cruise condition at Mach=0.74, Re=20 million, CL=0.515. In order to further improve the cruise efficiency, this NLF wing is optimized at higher Mach number of 0.75 via an in-house surrogate-based optimizer. The optimization is formulated as a drag minimization problem with constraints on lift, pitching moment and geometric thickness. Through only 130 CFD evaluations, 12.1 counts drag reduction is obtained, while all constraints are strictly satisfied. Further study shows that the drag reduction is contributed by both of shock-wave weakening and laminar-flow extension. On suction side, the favorable pressure gradient is maintained while shock wave is weakened; on pressure side, the crossflow (CF) instability is effectively suppressed and thereby the laminar flow region is dramatically extended. The improvement of aerodynamic performance is observed not only at design point but also over a certain range of off-design lift coefficients.

Aerodynamic characteristics of a guided rocket projectile are numerically simulated. Simulations are performed in a range of flight conditions using a commercial CFD software FLUENT, and variations of lift coefficients and drag coefficients with Mach numbers and angles of attack were demonstrated. The numerical results also indicate that the angle of canard deflection has significant influence on the flow field near tailfins.

Propeller boss cap fin (PBCF) is an energy-saving device that enhances the efficiency of marine vessels by mitigating propeller hub vortex and improving propulsion performance. In this work, a design procedure of the PBCF is introduced. It involves airfoil and fin shape optimization as well as PBCF performance evaluation. The airfoil optimization is performed under given ship advance ratio and propeller rotating speed, using an engineering tool and surrogate-based optimization to achieve maximum net thrust. The optimized airfoil demonstrates improved performance compared to the baseline design. The pitch angle describing the orientation of the PBCF is then optimized through iterative adjustments. The adjustment is guided by performance data obtained from computational fluid dynamics (CFD) simulations. The thrust and torque on the fin, cap and propeller are analyzed for better understanding of the force contribution from each component. Both straight fin for constant pitch angle and twisted fin with linearly decreased pitch angle are investigated. Moreover, simulations are conducted to evaluate the performance with a convergent cap as a replacement of the straight cap. CFD simulations show that the efficiency of the propeller equipped with the optimized PBCF and a convergent cap can increase by nearly 1% compared to the propeller without PBCF. The increase in propeller efficiency with optimized straight fins is higher than that achieved with twisted fins. With this design optimization framework, propeller performance can be improved, resulting in better fuel efficiency and cost savings.

Multidisciplinary design optimization techniques become more and more applied in the field of aerodynamics due to the rapid development of computers high-performance, numerical methods and optimization algorithms. These techniques coupled with Computational Fluid Dynamics (CFD), which aims to incorporate mathematical relations and algorithms to analyze and solve fluid flow problems, involve the use of those numerical methods and algorithms to improve the fluid flow solutions. CFD analysis of an airfoil determines its ability by producing results such as lift and drag forces, and the application of an optimization algorithm involves improving the shape of this airfoil in order to manipulate the lift and drag coefficients according to the requirements. In this work, a numerical investigation, using ANSYS/Fluent, of two-dimensional transonic flow over a NACA 0012 airfoil was conducted at various Mach numbers and compared with the provided experimental data. The flow to be considered is compressible and turbulent and the solver used is the density based implicit solver, which gives good results for high speed compressible flows. Then a shape optimization algorithm, based on a Multi-Objective Genetic Algorithm, was used in order to obtain an improved performance control of the aerodynamic coefficients of the optimized airfoil.

PurposeThis paper aims to present a topology optimization (TopO) method for designing heat exchangers (HEx) with two working fluids to be kept apart. The introduction of cut–cells gives rise to the cut-cell TopO method, which computes the optimal distribution of an artificial impermeability field and successfully overcomes the weaknesses of the standard density-based TopO (denTopO) by computing the fluid–solid interface (FSI) at each cycle. This allows to accurately solve the flow and conjugate heat transfer (CHT) problem by imposing exact boundary conditions on the computed FSI and results to correct performances computed without the need to re-evaluate the optimized solutions on a body-fitted grid.Design/methodology/approachThe elements of an artificial impermeability distribution field defined on a background grid act as the design variables and allow topological changes to take place. Post-processing them yields two fields indicating the location of the two flow streams inside the HEx. At each TopO cycle, the FSIs computed based on these two fields are used as the cutting surfaces of the cut-cell grid. On the so-computed grid, the incompressible Navier–Stokes equations, coupled with the Spalart–Allmaras turbulence model, and the temperature equation are solved. The derivatives of the objective and constraint functions with respect to the design variables of TopO are computed by the continuous adjoint method, using consistent discretization schemes devised thanks to the “Think Discrete – Do Continuous” (TDDC) adjoint methodology.FindingsThe effectiveness of the cut-cell–based TopO method for designing HEx is demonstrated in 2D parallel/counter flow and 3D counter flow HEx operating under both laminar and turbulent flow conditions. Compared to the standard denTopO, its ability to compute FSIs along which accurate boundary conditions are imposed, increases the accuracy of the flow solver, which usually leads to optimal, rather than sub-optimal, solutions that truly satisfy the imposed constraints.Originality/valueThis work proposes a new/complete methodology for the TopO of two-fluid systems including CHT that relies on the cut-cell method. This successfully combines aspects from both TopO and Shape Optimization (ShpO) in a single framework thus overcoming the well-known downsides of standard denTopO regarding its accuracy or the need for a follow-up ShpO after TopO. Instead of adding the well-known Brinkman penalization terms into the flow equations, it computes the FSIs at each optimization cycle allowing the solution of the CHT problem on a cut-cell grid.

Flow and heat transfer in an aero-engine compressor disk cavity with radial inflow has been studied using computational fluid dynamics (CFD), large eddy simulation (LES), and coupled fluid/solid modeling. Standalone CFD investigations were conducted using a set of popular turbulence models along with 0.2 deg axisymmetric and a 22.5 deg discrete sector CFD models. The overall agreement between the CFD predictions is good, and solutions are comparable to an established integral method solution in the major part of the cavity. The LES simulation demonstrates that flow unsteadiness in the cavity due to the unstable thermal stratification is largely suppressed by the radial inflow. Steady flow CFD modeling using the axisymmetric sector model and the Spalart–Allmaras turbulence model was coupled with a finite element (FE) thermal model of the rotating cavity. Good agreement was obtained between the coupled solution and rig test data in terms of metal temperature. Analysis confirms that using a small radial bleed flow in compressor cavities is effective in reducing thermal response times for the compressor disks and that this could be applied in management of compressor blade clearance.

As part of the Hypersonic International Flight Research Experimentation (HIFiRE) Direct-Connect Rig (HDCR) test and analysis activity, three-dimensional computational fluid dynamics (CFD) simulations were performed using two Reynolds-Averaged Navier Stokes solvers. Measurements obtained from ground testing in the NASA Langley Arc-Heated Scramjet Test Facility (AHSTF) were used to specify inflow conditions for the simulations and combustor data from four representative tests were used as benchmarks. Test cases at simulated flight enthalpies of Mach 5.84, 6.5, 7.5, and 8.0 were analyzed. Modeling parameters (e.g., turbulent Schmidt number and compressibility treatment) were tuned such that the CFD results closely matched the experimental results. The tuned modeling parameters were used to establish a standard practice in HIFiRE combustor analysis. Combustor performance and operating mode were examined and were found to meet or exceed the objectives of the HIFiRE Flight 2 experiment. In addition, the calibrated CFD tools were then applied to make predictions of combustor operation and performance for the flight configuration and to aid in understanding the impacts of ground and flight uncertainties on combustor operation.

A computational fluid dynamic (CFD) investigation is presented that provides predictions of the aerodynamic impact of uniform and nonuniform coatings applied to the leading edge of a compressor airfoil in a cascade. Using a NACA 65(12)10 airfoil, coating profiles of varying leading edge nonuniformity were added. A nonuniform coating is obtained when a liquid coating is applied to a surface with high curvature, such as an airfoil leading edge. The CFD code used, RVCQ3D, is a Reynolds averaged Navier–Stokes solver, with a k-omega turbulence model. The code predicted that these changes in leading edge shape can lead to alternating pressure gradients in the first few percent of chord that create small separation bubbles and possibly early transition to turbulence. The change in total pressure loss and trailing edge deviation are presented as a function of a coating nonuniformity parameter. Results are presented over a range of negative and positive incidences and inlet Mach numbers from 0.6 to 0.8. A map is provided that shows the allowable degree of coating nonuniformity as a function of incidence and inlet Mach number.