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• A regularized multi-fidelity framework for robust aerodynamic design optimization is established. • MHK and MFEI are used simultaneously to improve the optimization efficiency. • A progressive multi-round multi-fidelity optimization strategy is developed for extremely expensive robust optimization. • Diverse and mixed uncertainties are considered in the robust aerodynamic design optimization of airfoil and wing configurations. • A complicated robust design optimization of M6 wing with 36 design variables and 10 stochastic variables (including geometric and operational uncertainties) is implemented, which verifies the proposed method. Robust aerodynamic design optimization (RADO) is developed to obtain aircraft designs with insensitive aerodynamic performance. The uncertainties in manufacturing, operating conditions, or freestream turbulence intensity are quantified and involved in the objective and constraint functions of RADO. The iterative uncertainty quantification (UQ) based on high-fidelity simulations for each design generated in the optimization greatly increases the computation cost, which is the main challenge of RADO. To alleviate this problem, an efficient robust RADO method based on a multi-level hierarchical Kriging (MHK) model and multi-fidelity expected improvement (MFEI) is proposed. A regularized multi-fidelity framework is established for RADO and a progressive multi-round multi-fidelity strategy is further developed to reduce the cost. The MHK surrogate is utilized for global optimization and UQ, which is constructed with very few high-fidelity simulations by taking advantage of cheap simulations of multiple lower fidelities. The MFEI can adaptively infill samples of arbitrary fidelity during the optimization to accelerate the convergence. The proposed RADO method is verified by drag minimization of RAE 2822 airfoil under operational uncertainties and then demonstrated by robust designs of a natural-laminal-flow airfoil NLF0416 at high lift under environmental uncertainty and ONERA M6 wing in transonic viscous flow under operational and geometric uncertainties. The results confirm that the proposed method significantly improves optimization efficiency and obtains robust aerodynamic design within an affordable computational budget. In contrast to the deterministic optimum, the robust configurations can retain a good performance when uncertainties exist. [ABSTRACT FROM AUTHOR]

Since the 2.75D method takes both wing sweep and taper into consideration, it can build a transformation relation for pressure distributions between a 2D airfoil and a 3D wing under the assumption of conical flow. However, it is observed that the classical 2.75D method proposed by Lock has insufficient transformation accuracy in subsonic flow regions on transonic and subsonic wings. This article proposes an improved 2.75D method to address this problem. For subsonic flow region(s), a method based on Kaman-Tsien compressibility correction is proposed; for supersonic flow region(s), the original method of enforcing an equal normal Mach number between an airfoil and the corresponding wing section is used. Lock's and the improved methods are demonstrated by test cases of transonic and subsonic wings. Results show the transformation accuracy of pressure distributions between 2D airfoils and 3D wings is dramatically improved using the proposed method. The application of these methods to the pressure distribution design for a tapered swept wing section is also demonstrated. The design optimization using the improved 2.75D method leads to a design result being more approximate to a 3D tapered swept wing with a computation effort similar to a 2D airfoil design. [ABSTRACT FROM AUTHOR]

Abstract: Variable-fidelity surrogate modeling offers an efficient way to generate aerodynamic data for aero-loads prediction based on a set of CFD methods with varying degree of fidelity and computational expense. In this paper, direct Gradient-Enhanced Kriging (GEK) and a newly developed Generalized Hybrid Bridge Function (GHBF) have been combined in order to improve the efficiency and accuracy of the existing Variable-Fidelity Modeling (VFM) approach. The new algorithms and features are demonstrated and evaluated for analytical functions and are subsequently used to construct a global surrogate model for the aerodynamic coefficients and drag polar of an RAE 2822 airfoil. It is shown that the gradient-enhanced GHBF proposed in this paper is very promising and can be used to significantly improve the efficiency, accuracy and robustness of VFM in the context of aero-loads prediction. [Copyright &y& Elsevier]

Aerodynamic shape optimization of a hypersonic transport aircraft over a wide Mach-number range is challenging. The difficulty is not only associated with the large computational cost of high-fidelity CFD (computational fluid dynamics) simulations but also linked to the reasonable compromise between aerodynamic performances of aircraft at different speed ranges. This article proposes to use efficient global optimization based on surrogate models to address this type of problems. A RANS (Reynolds averaged Navier-Stokes) flow solver is adopted to evaluate objective and constraint functions; Kriging surrogate model combined with a parallel infill-sampling method and a multi-round strategy are employed to find the global optimum. First, a profile optimization with baseline airfoil of NACA64A-204 is conducted to achieve high lift-to-drag ratio (L / D) at both hypersonic (Ma = 6.0) and transonic (Ma = 0.8) regimes with up to 18 design variables, and significant improvement has been observed. Then, multi-objective wing optimizations of maximizing both hypersonic and transonic L / D s with up to 54 design variables are performed and multiple optimizations with various sets of weight coefficients are investigated. The optimized wings are further evaluated and compared with the baseline wing at the flow regimes from subsonic to hypersonic speeds. Results show that by using the optimized profile, hypersonic and transonic L / D s of a typical wing configuration are increased by 13.85% and 7.32%, respectively. After 3-D optimizations, the L / D s at hypersonic and transonic design points are further improved by 4.47% and 3.19%, respectively, and a better performance over a wide Mach-number range is obtained. It is shown that a multi-round surrogated-based optimization is feasible and effective for aerodynamic shape optimization of hypersonic transport aircrafts over a wide Mach-number range. [ABSTRACT FROM AUTHOR]

Aerostructural design optimizations of helicopter rotors based-on high-fidelity numerical simulations present a great challenge due to extremely large computational costs. In order to improve optimization efficiency by introducing assistance from cheap low-fidelity simulations, this article proposes to use a multilevel hierarchical kriging (MHK) model, which can incorporate three or more levels of fidelity to accelerate the convergence of a high-fidelity aerostructural optimization of helicopter rotors towards the global optimum. The optimization strategy is implemented by three steps. Firstly, a single-fidelity optimization is conducted based on kriging models and low-fidelity simulations. Secondly, a bi-fidelity optimization is performed based on hierarchical kriging (HK) models with pre-calculated low-fidelity samples and additional medium-fidelity samples. Finally, a multi-fidelity optimization is conducted towards the high-fidelity optimum based on MHK models with pre-generated low- and medium-fidelity samples and additional high-fidelity samples. The high-fidelity analysis is defined as coupled computational structural dynamics (CSD) and computational fluid dynamics (CFD) simulations, and the medium- and low-fidelity analyses are defined as merely CFD simulations on fine and coarse computational grids, respectively. To verify the effectiveness of MHK models in rotor optimizations, a blade shape optimization for aeroacoustic noise reduction is carried out by using the proposed optimization strategy. Results show that the total cost of an MHK-based optimization is around 49% less than that of a kriging-based optimization. Besides, by using kriging- and HK-based optimizations the aeroacoustic noises are reduced by 6.84 dB after 175 evaluations and 102 evaluations, respectively, while an MHK-based optimization could achieve the same noise reduction with only 54 evaluations. This example demonstrates that an MHK-based multi-fidelity optimization method can significantly improve the efficiency of aerostructural optimizations for helicopter rotors. [ABSTRACT FROM AUTHOR]

Aerodynamic shape design considering both subsonic and hypersonic performance is a challenge in the field of aerospace. Many studies have shown that morphing wing is a promising technology to address this challenge. In this article, a new morphing mechanism which changes both the plane shape (such as span and sweep angle) and the profile (such as chord length and relative thickness) of the wing is proposed and a multi-objective optimization design for this new morphing mechanism is conducted based on surrogate-based optimization (SBO) algorithm. In the optimization design, three different configurations of the morphing wing which are the 36-, 45- and 70-degree sweep angle state, respectively and four different flow conditions which are subsonic flow (0.25 Ma), transonic flow (0.85 Ma), supersonic flow (1.2 Ma) and hypersonic flow (6 Ma) are considered. 36-degree sweep angle state is used for subsonic flow, 45-degree sweep angle state is used for transonic flow and supersonic flow and 70-degree sweep angle state is used for hypersonic flow. Subsonic lift and transonic, supersonic, hypersonic lift-to-drag ratios are objectives of this optimization design which has up to 64 design variables. To reduce computational cost of high-fidelity CFD (computational fluid dynamics) simulations and find the global optimum, Kriging surrogate model combined with a parallel infill-sampling method and a multi-round strategy are employed in this study. At last, the optimized morphing wing is evaluated and compared with the baseline and a fixed wing which is a moderate sweep wing. Results show that the optimized morphing wing features a very good aerodynamic performance improvement over the flow regimes from subsonic to hypersonic flow conditions. [ABSTRACT FROM AUTHOR]

Integrating components, as a new afterburner concept, is currently an important trend in afterburner development research. Spurred by this trend, this study proposes an inner cone cavity flame holder for an integrated afterburner. The experimental and numerical investigation evaluates the ignition performance under various pneumatic parameters, illustrating that stable flame is successfully achieved in the cavity of the integrated afterburner for a wide range of flow conditions. This is important, as an appropriate cone expansion angle benefits the ignition and flame stability. By comparing the experimental results, the optimum case is obtained, which reduces the ignition FAR. Furthermore, double fuel line can effectively broaden the ignition boundary. The successful completion of our ignition experiment demonstrates that the cavity flame holder of the integrated afterburner has a good structural design and ignition performance. [ABSTRACT FROM AUTHOR]

Double-ended airfoils with round leading and trailing edges outperform conventional round-nose and sharp trailing-edge airfoils when used for the inboard part of a rigid coaxial rotor operating in forward flight, as reverse flow can occupy 80% surface of the retreating blade. However, aerodynamic design optimization for this kind of airfoils is challenging, since it is required to achieve drag reduction over a wider lift range in both forward and reverse flows but there is no method available currently. To address this problem, this article proposes an efficient hybrid inverse/optimization design method that combines inverse design and direct optimization within a surrogate-based optimization (SBO) framework. The weighted objective function consists of two components: prescribed pressure distributions and total drag. The resulting constrained optimization problem is solved by using kriging surrogate model and dedicated infill-sampling method. The proposed method is demonstrated by aerodynamic shape optimizations of a DBLN-526 airfoil and an elliptical airfoil and compared with direct optimization method. For DBLN-526 airfoil case, 7%–30% drag reduction is achieved in the range from zero lift to maximum lift and the optimum airfoil features better stall characteristics. For the case of elliptical airfoil of 40% thickness, significant drag reduction is achieved in both forward and reverse flows and the optimum airfoil features more robust aerodynamic performance at off-design Mach numbers. The results confirm that the proposed method outperforms the conventional direct optimization method and more practical for aerodynamic shape optimization of double-ended airfoils. [ABSTRACT FROM AUTHOR]

A local-correlation-based laminar-turbulent transition model has been developed for external aerodynamic problems and coupled with the Spalart-Allmaras (S − A) turbulence model. This laminar-turbulent transition modeling scheme is then termed as the γ − S A model. It has been validated against a series of two- and three-dimensional test cases with a relatively wide range of Reynolds numbers and free-stream turbulence intensity levels. Results show that the γ − S A model can give accurate transition-onset-location predictions, as well as more accurate predictions on the force coefficients than the underlying S − A turbulence model at various angles of attack. Comparison with the classic S S T − γ − R e θ transition model shows that the γ − S A model can achieve comparable results, with the expect of less time or memory consumption, as the γ − S A model has two equations fewer than the former one. With a newly proposed local-farfield mixed turbulence intensity strategy, the turbulence intensity on the inlet or farfield boundary can be conveniently set directly with the value tested for the wind tunnel or the value as expected, while achieving local turbulence intensity variation. Moreover, the γ − S A model employs only local variables, which makes it totally modern-CFD-compatible and very appropriate for large-scale parallel running using unstructured grids. In the future, it can be modified into a robust and efficient DES-type model with the consideration of transition process. [ABSTRACT FROM AUTHOR]