Supersonic Wing Research Articles

ВИЗНАЧЕННЯ АЕРОДИНАМІЧНИХ ХАРАКТЕРИСТИК ЛІТАЛЬНОГО АПАРАТА З НАДЗВУКОВИМ ШВИДКОСТЯМИ ПОЛЬОТУ

The aircraft aerodynamic characteristics at supersonic flight modes can be determined with a high level of accuracy through testing the aircraft in wind tunnels or through time-consuming three-dimensional and analytical computational methods. However, in certain cases when the aircraft aerodynamic characteristics are already known within a given Mach number range for supersonic flight, it is possible to develop a relatively simple mathematical model for the aircraft aerodynamic characteristics at supersonic speeds using the aerodynamic properties of the supersonic wing leading edge for flight speeds Mf ≥ 1.2M. The object of the study is the aircraft aerodynamic characteristics at supersonic flight speeds. The subject of the study is the impact of the aircraft geometric features on the mathematical model, which is used to determine its aerodynamic characteristics in the supersonic flight speed range. The study hypothesizes that to determine the aerodynamic characteristics of an aircraft at Mf ≥ 1.2, it is sufficient to identify the values of the aircraft aerodynamic parameters at a flight speed of Mf =1.2 and then use aerodynamic analytical dependencies for the supersonic wing leading edge to determine these parameters at Mf ≥ 1.2. A mathematical model of the aircraft aerodynamic characteristics in the supersonic flight speed range was developed. Coefficients of drag polar and zero-lift drag for the reference mode Mf = 1.2 have been identified by the developed mathematical model. Then, the Mach number range for the aerodynamic characteristics of the NASA 1044 configuration has been extended to supersonic speeds of Mf = 1.8…4.0. The resultsof the aerodynamic calculations, obtained in the form of dependencies of the drag polar and zero-lift drag coefficients on Mach number, and the aircraft polar plot, have been compared with known NASA data for the 1044 configuration aircraft. The average modeling error of the aircraft aerodynamic characteristics in the supersonic speed range is less than 3%.

Nonlinear metamaterial enabled aeroelastic vibration reduction of a supersonic cantilever wing plate

The violent vibration of supersonic wings threatens aircraft safety. This paper proposes the strongly nonlinear acoustic metamaterial (NAM) method to mitigate aeroelastic vibration in supersonic wing plates. We employ the cantilever plate to simulate the practical behavior of a wing. An aeroelastic vibration model of the NAM cantilever plate is established based on the mode superposition method and a modified third-order piston theory. The aerodynamic properties are systematically studied using both the timedomain integration and frequency-domain harmonic balance methods. While presenting the flutter and post-flutter behaviors of the NAM wing, we emphasize more on the pre-flutter broadband vibration that is prevalent in aircraft. The results show that the NAM method can reduce the low-frequency and broadband pre-flutter steady vibration by 50%–90%, while the post-flutter vibration is reduced by over 95%, and the critical flutter velocity is also slightly delayed. As clarified, the significant reduction arises from the bandgap, chaotic band, and nonlinear resonances of the NAM plate. The reduction effect is robust across a broad range of parameters, with optimal performance achieved with only 10% attached mass. This work offers a novel approach for reducing aeroelastic vibration in aircraft, and it expands the study of nonlinear acoustic/elastic metamaterials.

Mitigating aeroelastic vibration of strongly nonlinear metamaterial supersonic wings under high temperature

Mitigating aeroelastic vibration of strongly nonlinear metamaterial supersonic wings under high temperature

Wind Tunnel Experiment and Numerical Simulation of Secondary Flow Systems on a Supersonic Wing

Wind Tunnel Experiment and Numerical Simulation of Secondary Flow Systems on a Supersonic Wing

Attenuation of boundary-layer instabilities for natural laminar flow design on supersonic highly swept wings

To meet the challenge of drag reduction for next-generation supersonic transport aircraft, increasing attention has been focused on Natural Laminar Flow (NLF) technology. However, the highly swept wings and high-Reynolds-number conditions of such aircraft dramatically amplify Crossflow (CF) instabilities inside boundary layers, making it difficult to maintain a large laminar flow region. To explore novel NLF designs on supersonic wings, this article investigates the mechanisms underlying the attenuation of Tollmien–Schlichting (TS) and CF instabilities by modifying pressure distributions. The evolution of TS and CF instabilities are evaluated under typical pressure distributions with different leading-edge flow acceleration region lengths, pressure coefficient slopes and pressure coefficient deviations. The results show that shortening the leading-edge flow acceleration region and using a flat pressure distribution are favorable for suppressing CF instabilities, and keeping a balance of disturbance growth between positive and negative wave angles is favorable for attenuating TS instabilities. Based on the uncovered mechanisms, a strategy of supersonic NLF design is proposed. Examination of the proposed strategy at a 60° sweep angle and Ma = 2 presents potential to exceed the conventional NLF limit and achieve a transition Reynolds number of 17.6 million, which can provide guidance for NLF design on supersonic highly swept wings.

Data-driven hierarchical collaborative optimization method with multi-fidelity modeling for aerodynamic optimization

The aerodynamic optimization of aircraft based on evolutionary algorithms generally requires thousands or even tens of thousands of iterations. Moreover, it faces challenges such as high-dimensional design variables and high computational costs, leading to low efficiency in algorithm search and poor optimization results. In order to enhance the efficiency of aerodynamic optimization, a data-driven collaborative optimization and multi-fidelity hierarchical modeling method is proposed. Firstly, a Clustering-Enhanced Teaching-Learning-Based Optimization (CETLBO) algorithm is introduced to improve its exploration and exploitation capabilities. By employing a collaborative optimization strategy, a divide-and-conquer paradigm is utilized to efficiently solve high-dimensional, nonlinear, and complex engineering optimization problems. Secondly, a two-stage hierarchical aerodynamic optimization framework is designed to reduce computational costs by integrating data from different fidelity levels. Stage 1 adopts a data-driven approximate optimization method to provide prior knowledge for the high-fidelity optimization of stage 2. Finally, the proposed method is applied to the optimization design of supersonic wing shapes. Compared to traditional optimization design systems, it demonstrates high efficiency in high-dimensional aerodynamic optimization problems, yielding better optimization results.

Flutter investigation and deep learning prediction of FG composite wing reinforced with carbon nanotube

Abstract The flutter of a composite wing reinforced with functionally graded carbon nanotubes (CNTs) has been investigated. A rectangular plate models a supersonic wing with cantilever boundary conditions. To determine displacement fields of a moderately thick plate, shear deformation theory is used. Using the Hamilton principle, a first-order piston theory was used to simulate supersonic airflow. This study examines four types of CNT thickness. Also, four different CNT distribution patterns are investigated. In a two-layer asymmetric composite, the effects of patch mass, mass distribution, fiber orientation angle, and distribution of CNTs were examined. Moreover, the results are compared and verified with other studies. A greater mass ratio led to a smaller flutter boundary, while a longer added mass increased the flutter boundary. A variation in the distribution pattern in CNT fiber orientation results in a distinct behavior of the flutter boundary for asymmetric composites with increasing orientation angles. The artificial neural network is utilized to predict the damping ratio, and the results showed great accuracy compared to the study results. Hyperparameter tuning is employed for better optimizing the predictive models.

Static Aeroelastic Research of Supersonic Wing

The thin airfoil design and the supersonic flight condition make the aeroelastic deformation have an important effect on the aerodynamic characteristics of the supersonic aircraft. In this paper, a supersonic Unmanned Aerial Vehicle (UAV) wing is taken as the research object, and the static aeroelastic weak coupling calculation method (CFD/CSD) is used to carry out the modification of the aerodynamic coefficient of the wing. The analysis results show that the aeroelastic deformation reduces the lift coefficient, and the pitch moment produces a head-up increment. At the same time, the aerodynamic focus of the wing moves forward, which reduces the static stability margin of the aircraft. The influence of aeroelastic deformation on aerodynamic characteristics increases with increasing Mach number.

Aerodynamic Optimization Design of Supersonic Wing Based on Discrete Adjoint

Reducing fuel consumption and improving the economy by effectively reducing cruising drag is the main objective of the aerodynamic design of supersonic civil aircraft. In this paper, the aerodynamic optimization design system based on the Reynolds-Averaged Navier–Stokes (RANS) equation and discrete adjoint theory is applied to supersonic wing design. Based on this system, a single-point optimization design study of aerodynamic drag reduction in cruise conditions was carried out for two typical supersonic wing layouts, subsonic leading edge and supersonic leading edge, and the drag reduction reached 3.78% and 4.53%, respectively. The aerodynamic design characteristics of different types of supersonic wings were explored from the perspectives of wing load, twist angle distribution, pressure distribution, airfoil shape characteristics, and flow field characteristics. The optimization results show that the drag reduction of the subsonic leading edge configuration is dominated by the induced drag, while the optimizer mainly focuses on reducing the shock wave drag for the supersonic leading edge configuration. By comparing the sensitivity analysis of lift and drag coefficients to airfoil deformation with the optimization results, the optimized dominant directions of two types of supersonic wings are qualitatively analyzed. The derivatives obtained from discrete adjoint equations are useful to elaborate the design tendency and the reason for the trade-off generation of supersonic wings under specific layouts and engineering constraints, which provides a reference for the design of supersonic wings in the future.

Measurement and Identification of Supersonic Stationary Crossflow Waves Based on Sublimation Method

Aiming at the crossflow transition problem of supersonic swept wings, an experimental study on the measurement and identification of boundary-layer stationary crossflow waves was carried out at Mach number 2.0. A 65 deg straight-swept-wing model with simple airfoil geometry was designed to capture the crossflow transition phenomenon. The sublimation method for a large-scale cold supersonic wind tunnel was developed to acquire a transition pattern. The stationary crossflow wave identification technology is based on image processing, a vertical and horizontal decomposition method is established, and the quantitative characteristic information (such as the vortex axis spacing and direction) can be identified from high-resolution sublimation images. The experimental results show that the transition position caused by crossflow instability is closed to the leading edge for supersonic wings with large sweep angles at high Reynolds numbers, and the transition position moves upstream with the increase of the unit Reynolds number. The identified vortex axis spacing and direction are basically consistent with the linear stability theory. The stationary disturbance wave information at different Reynolds numbers is compared and analyzed. It is found that the unit Reynolds number plays an important role in the selection and amplification of the unstable stationary modes with different wavelengths in the transition process.

Analytical Study of Double Wedge Airfoil for Supersonic Applications and Shape Modification for Subsonic Applications

A Double Wedge airfoil is an airfoil for supersonic blades and wings, with a wedge like tapered sharp leading and trailing edges. The shock waves and the expansion waves govern the supersonic characteristics of the double wedge airfoil. The present work is based on the design and aerodynamic analysis of double wedge airfoil at supersonic regime and shape modification of the airfoil for subsonic regime. The flow analysis is carried out using ANSYS Fluent, which is CFD based software. The supersonic analysis is performed for Mach number of 2.0. The airfoil was modified for subsonic applications and the analysis was performed for Mach number 0.7. Blowing technique was incorporated to further improve the aerodynamic performance of the modified double wedge airfoil. From the simulated results, it was found that the co-efficient of drag of double wedge airfoil at Mach number of 2.0 agreed with the manual calculations. The modified airfoil gave optimum performance with the addition of a blowing stream at subsonic speeds.

Supersonic Forward-Swept Wing Design Using Multifidelity Efficient Global Optimization

In this paper, the ability of the forward-swept supersonic wing to simultaneously reduce aerodynamic drag and sonic boom under cruise conditions is investigated via the optimization of airfoil distribution. The forward-swept wing is superior to the conventional backward-swept delta wing for reducing sonic booms. To realize optimum aerodynamic characteristics, airfoil distributions for the forward-swept and backward-swept wings were acquired and compared for low-drag, low-boom wing abilities. For sonic boom evaluations, the augmented Burgers equation and multipole analysis were applied to near-field pressure distributions calculated by Euler simulations of both configuration samples. This process was time-consuming, so a multifidelity approach was introduced with a multi-additional sampling. Low-drag and low-boom solutions were obtained for both planforms; the forward-swept wing was found to reduce sonic boom and aerodynamic drag more efficiently than the backward-swept wing. Based on functional analysis of variance, different design variables were noted to contribute toward reductions of the various objective functions. In case of sonic boom reduction via comparison of the low-boom solutions, for the given airfoil geometries at the wing tips for the forward-swept and backward-swept wings, the former had a larger twisted-down angle and the latter had a stronger camber shape.

Analysis of Onset of Transition Region on a Supersonic Wing

Abstract: Climate change is due to the sole effect of global warming, which is an outcome of Green House Gas (GHG) emissions. The aviation industry is considered to be one of the fastest-growing sources of these emissions. Mitigation of these emissions is a challenging task since it requires the entire world to work hand-in-hand. The International Civil Aviation Organisation (ICAO), an international governing body, has come up with a challenging target of net 50% reduction in global aviation emission by 2050, relative to the emission levels in 2005. To achieve this target, there must at least be 2% improvement in fuel efficiency every year. This 2% efficiency improvement is further split into Aircraft technology improvement (1.1%) and Airport infrastructure improvement (0.9%). This research aims to achieve the 2% improvement by aircraft technology improvement. Over the years, it was evidenced that mitigation of skin-friction drag on the wing would contribute to a majority of the fuel efficiency improvement. Hence, the primary aim is narrowed down to wing skin friction drag. The theory of boundary-layer has been studied thoroughly in this research work. Also, the interaction between the shock-waves and boundary-layer will be briefed in this paper as the baseline aircraft is supersonic. As it was understood as understood over the research that extension of the laminar boundary-layer favours the reduction of skin-friction drag, it is the primary objective of this research work. A Computational Fluid Dynamics (CFD) study on the of the wing aerodynamics has been conducted to enhance the performance of the aircraft and ANSYS Fluent has been used to serve this purpose. The results obtained from ANSYS fluent will be validated in one of the NASA’s open-source tool VSPAero which uses Vortex Lattice Method to model the flow. Turbulent Kinetic Energy technique is used to predict the onset of the transition of the flow. Initial study is performed on the 2D aerofoil and one the most optimum aerofoil is obtained, the study is continued in 3D to understand the effect of cross-flow instabilities. Free-stream velocity profile variation is another reliable technique in predicting the transition and will be discussed in the next volume of the paper. Finally, aerodynamic performance analysis will be performed to obtain the best wing configuration.

Bayesian Optimization of a Low-Boom Supersonic Wing Planform

A surrogate-assisted methodology to accelerate the design space exploration process and multi-objective optimization of a notional low-drag, low-boom supersonic transport planform using Euler computational fluid dynamics with an empirical parasite drag addition and an augmented Burgers equation solver is demonstrated. The use of Kriging-based surrogate models, coupled with expected hypervolume improvement in a Kriging believer framework, allows efficient global optimization of the selected wing design parameters. The use of effective nondominated sampling is used to further improve the solutions found. The effects of parameterizing the planform of a baseline model (NASA’s 69 deg wing–body) with 6- and 11-variables are explored. Genetic and local optimizers are used to search the Kriging surrogates. The results show the tradeoff between planform shapes that efficiently use compression lift to increase the lift-to-drag ratio and a smooth undertrack pressure signature to reduce ground-level noise. Furthermore, correctly positioned wingtips may also be used to smooth the undertrack signature, reducing the sonic boom.

Aero-structural optimization of supersonic wing under thermal environment using adjoint-based optimization algorithm

For solving the problems of jig-shape design and aerodynamic shape design at the same time, based on the aero-thermo-structural (ATS) analysis model, the paper presents a coupling optimization method, in which the adjoint equations of coupling system are developed. In order to demonstrate this methodology, a supersonic wing is optimized, and its objective function is the lift-to-drag ratio considering the elastic deformation of the wing structure induced by aerodynamic load and aerodynamic heating. And independent variables are the geometric twist angles of wing sections along the span, the thickness of skin, and the cross-sectional geometrical parameters of spar and stringer. The constraint is the strength of the wing structure. When the maximum allowable mass is 700 kg, the results show that the lift-to-drag ratio is increased from 6.0505 to 6.2875, and all constraints are satisfied. Therefore, the coupling optimization method is effective. And this method can control the increment of structural mass through the maximum mass constraint.

Influence of Different Thermal Boundary on Numerical Analysis of Aerodynamic Heat of Supersonic Wing

Influence of Different Thermal Boundary on Numerical Analysis of Aerodynamic Heat of Supersonic Wing

Planform dependency of optimum cross-sectional geometric distributions for supersonic wing

The difference in the supersonic aerodynamic characteristic of different wing planforms is evaluated to determine the optimum parameters for the cross-sectional geometry of a supersonic wing. The supersonic performance of wing with integrated engine intakes was evaluated for a quadruple-tapered wing with a large sweep angle and for a single-tapered wing with a small sweep angle. To reduce the required computation time, the design problems were solved using a multi-fidelity approach consisting of a hybrid surrogate model assisted by evolutionary computation. To evaluate aerodynamic performance, the compressible Euler equation and the linearized compressible potential equation were employed as high- and low-level fidelity solvers, respectively. Through design optimizations, the contributions of different cross-sectional parameters to drag reduction were determined. It was found that the shape of the forward camber and the twist angle around the middle of the wing had the most noteworthy influence on the drag reduction for both wing planforms, because most of the aerodynamic force was generated near the wing mid-span. For a wing with a large sweep angle, a cross-sectional geometry involving a small positive camber at the leading edge, and a small twisted angle was optimum. For a wing with a small sweep angle, a cross-sectional geometry involving a negative camber at the leading edge, and a thinner leading edge, and higher twisted angle than those for a large swept-back wing was optimum because of the ready generation of a shock wave at the leading edge.

Comparison of the reliability-based and safety factor methods for structural design

Abstract Reliability-based design optimization theory has been widely acknowledged as an advanced and advantageous methodology for complex structural system design. Comparatively, the traditional safety factor design method has fallen out of favor with designers since it is simply just a comprehensive expression of all the uncertainties existing in a practical engineering structure, which has been verified to be unreasonable. Moreover, there is no description of the method for searching for the optimal design. In this paper, a comparison of the two approaches is performed for the optimization results and process. It is demonstrated that the weight of the designed structure by the reliability-based optimization is not heavier than that by the safety factor design method under the same reliability requirements. The time efficiency, error estimates and sensitivity analysis are also compared to discuss the advantages and disadvantages of the two approaches. Eventually, the advancement of the reliability-based design optimization is illustrated with optimization designs of a two-bar truss and one practical supersonic wing.

Computational aerodynamic optimization of wing-design concept at supersonic conditions by means of the response surface method

In this paper, by means of computational fluid dynamics, a significant study has been made on the effects of geometric parameters of wing with capability of flying efficiently and cost-effectively at supersonic condition. Multi-objective optimization has been performed for the aerodynamic shape optimization of the wing configuration. The three-dimensional wing shape defined by four design variables is optimized. For achieving the most desirable aerodynamic efficiency (lift-to-drag ratio), Response Surface Method and Genetic Algorithm is utilized. To ensure the reliability of the solution and validating the numerical evaluation, flow around a delta wing is simulated and results are compared with credible numerical works. Furthermore, the particular design variables, which have serious effects on the objective functions, are found. Wing sweep angle along with aspect ratio has strong effects on the main outputs. Given that the current flow regime is supersonic, counteracting the negative effects of shock waves is one the most important design points. Among the studied parameters, leading edge sweep angle has the greatest impact on the main objectives of this research and it is the primary factor for delaying the formation of shock waves over the wing surface. Lift and drag coefficients, as primary objective functions, have higher sensitivity to changes in aspect ratio. By comparison to best geometry among initial cases, aerodynamic efficiency is increased by approximately 15% as a result of optimization study of this supersonic wing geometry.

An efficient single-loop strategy for reliability-based multidisciplinary design optimization under non-probabilistic set theory

Non-probabilistic reliability based multidisciplinary design optimization (NRBMDO) offers a powerful tool for making reliable decisions with the consideration of uncertain-but-bounded uncertainties for complex engineering systems. However, the prohibitive computation and convergence difficulties caused by the directly coupling of uncertainty based multidisciplinary analysis (UMDA), non-probabilistic reliability analysis (NRA) and MDO would seriously hamper the application of NRBMDO. In this paper, an efficient single loop strategy for NRBMDO (SLS_NRBMDO) is developed to decouple the nested issue and thus improve the computational efficiency. The key idea of the proposed strategy is decoupling NRBMDO with several cycles of sequential MDO, UMDA, NRA and translating distance calculation (TDC). For UMDA, three methods, i.e., the first order interval Taylor expansion method, the interval vertex theorem, the direct optimization approach are formulated. Besides, NRA is conducted on the basis of the expanded non-probabilistic stress–strength interference model and the volume ratio thought, which provides a clear and definite assessment criterion for the structural safety with uncertain-but-bounded parameters. Furthermore, the translating strategy based on the performance measure approach is proposed to shift and update the constraints, and the expression of the translating distance is mathematically derived to accelerate the design procedure. Eventually, the effectiveness and efficiency of the proposed method are illustrated with one numerical case and one practical supersonic wing optimization design problem.