Airfoil Pressure Distribution Research Articles

Wind Tunnel Tests of a Thick Wind Turbine Airfoil

ABSTRACTThis article reports about a wind‐tunnel experiment carried out in the ONERA F2 low‐speed wind tunnel on a model of the DU 97‐W‐300Mod airfoil designed for wind turbine application. The wind tunnel, the airfoil model, and experimental techniques used are presented, with special emphasis on the data processing and corrections required to derive airfoil forces and pressure distribution. To better document the flow physics at play, the results are illustrated by infrared thermography and surface oil flow visualization. The test allowed investigating Reynolds number effects between 1 and 3.8 millions. To ameliorate the understanding of the benefits and limitations of such airfoil testing, one section is devoted to the comparison of present results with previous experiments in other wind tunnels. Some of the difficulties arising in airfoil testing are evidenced and discussed to contribute to the improvement of test methods.

Fast Prediction of Airfoil Aerodynamic Characteristics Based on a Combined Autoencoder

Aircraft airfoils are classified into two main categories: symmetrical and asymmetrical. Both types of airfoils have a significant impact on the flight performance and safety of the aircraft. The fast prediction of the aerodynamic coefficients and pressure distributions of airfoils is crucial for the design of aircraft. The traditional wind tunnel test and CFD methods have the disadvantages of high test cost and high time consumption. To solve these problems, a combined autoencoder (CAE) network is proposed in this paper, which can achieve the fast prediction of airfoil aerodynamic coefficients and pressure distributions. The network consists of an airfoil shape autoencoder (AE) network and a multilayer perceptron (MLP) network. Firstly, an autoencoder network reflecting the characteristics of the airfoil shape is established, and the effects of different latent variables on the performance of the autoencoder network are investigated. Then, the latent variables obtained from the autoencoder are concatenated with the inflow conditions such as the Reynolds number and the angle of attack to be used as inputs to the MLP network, and the aerodynamic coefficients of different airfoils in different inflow conditions are predicted. The effects of various latent variable inputs, as well as the direct input of the airfoil shape into the MLP network, on the prediction performance of aerodynamic coefficients are compared and analyzed. The optimal aerodynamic coefficient prediction network is then obtained. Finally, the CAE network is also applied to predict the pressure distributions of different airfoils in different inflow conditions and the effects of different latent variables and input conditions on the prediction performance of the pressure distributions are analyzed and compared with the advantages and disadvantages of the CAE network and the conditional variational autoencoder (CVAE) network. The results demonstrate that the proposed method is capable of accurately predicting aerodynamic characteristics in a shorter time, offering a valuable reference for the fast and efficient design of aircraft airfoils.

Influence of air flow field computation on an ice accretion simulation of airfoils

Ice accretion simulation is an important research technology for aircraft icing problems. Air flow field prediction of an iced aircraft is an essential step of ice accretion simulation. However, commonly used turbulence models have difficulty predicting the large flow separation of the iced aircraft, which makes air flow field calculation insufficiently accurate and further affects the ice accretion simulation. In this paper, the influence of air flow field computation on ice accretion simulation of airfoils is numerically studied. Several different turbulence models are applied to predict air flow fields of iced airfoils. Then, the Eulerian method of water droplets, integral boundary layer method and Myers model are used to simulate ice accretion. Different airfoils with several icing conditions, including glaze ice and rime ice cases, are numerically tested and compared with experimental data. The results show that the separating shear layer fixed k−ω model is more accurate than other turbulence models in predicting the pressure distribution of the iced airfoil, especially the leading-edge suction peak. Thus, a more reasonable droplet collection efficiency distribution and convective heat transfer distribution can be obtained, yielding a more consistent ice height compared with the experimental data.

Reynolds number effect correction of multi-fidelity aerodynamic distributions from wind tunnel and simulation data

Conventional wind tunnels cannot simulate actual flight Reynolds numbers because of testing condition constraints. The changes in aerodynamic characteristics caused by differences in the Reynolds number are especially sensitive on transonic supercritical airfoils and wings. Accordingly, this study proposes a correction method for the Reynolds number effect based on feature extraction of multi-fidelity aerodynamic distributions. Correcting wind tunnel test results at low Reynolds numbers with large amounts of numerical simulation data across various Reynolds numbers allows this method to achieve greater precision on airfoil pressure distribution predictions at high Reynolds numbers. The first step is to extract the primary features of the surface pressure distribution of the supercritical airfoil using the proper orthogonal decomposition technique. Next, a multi-fidelity neural network model is employed to relate the (low-fidelity) simulation data to the (high-fidelity) wind tunnel data acquired at low Reynolds numbers. Finally, the model is used to predict the airfoil pressure distributions at high Reynolds numbers, and the predictions are validated using wind tunnel test data collected from the RAE2822 airfoil. The results indicate that this approach can provide more accurate estimations than numerical simulations and single-fidelity models. Additionally, it can effectively reduce the impact of the Reynolds number difference on the transonic aerodynamic properties of supercritical airfoils.

Analysis of the unsteady surface pressure distribution of a pitching airfoil using modal decomposition

This work presents an exploratory attempt to understand how modal decomposition can be utilized for characterizing the behavior of the unsteady surface pressure distribution on a pitching airfoil. The unsteady aerodynamics of a thick airfoil geometry across a wide range of conditions were studied. Analyzing the pressure distributions across cases with different levels of flow separation yielded results that exhibited transitional behavior that encouraged modal analysis. The utility of the Proper Orthogonal Decomposition in capturing the distinct physical features of the unsteady pressure distribution was then evaluated. The first four global spatial modes extracted from surface pressure data resemble important and understandable attributes of the pressure distribution for an oscillating airfoil. The four-mode description sufficiently reconstructs the surface pressure data even for conditions with complex unsteady flow interactions. The resulting reconstructed pressure distributions can be used to approximate lift and moment coefficients with good accuracy.

Study on design characteristics of airfoil pressure distribution for low Mach number general aircraft

This paper studies several problems that may exist in the aerodynamic design of low Mach number general aircraft, and discusses the design characteristics of the pressure distribution of some related airfoils while part of the characteristics is verified by a multi-point laminar airfoil design case. Firstly, several issues which exist in the aerodynamic design of low Mach number general aircraft, such as relatively large zero-lift drag coefficient, high stall performance requirements, and large variation range of Reynolds number, are discussed and analyzed. Then, to address these issues, the paper extracts some useful design features that are beneficial to the airfoil aerodynamic performance from the study on the design characteristics of the pressure distribution of several turbulent flow airfoils and laminar flow airfoils that are designed for low Mach number general aircraft. Finally, by considering the typical design conditions of a light multi-purpose single-turboprop-engine general aircraft, including the cruise, climb and stall conditions, a multi-point laminar airfoil optimization is offered by using the GAW-1 airfoil as the baseline. The optimized foil has a 9.6 improvement in the lift-to-drag ratio of the cruise condition, a 16.3 improvement in the lift-to-drag ratio of the climb condition but a 0.115 decrease in the maximum lift coefficient of the stall condition. The results show that the requirements from the cruise, climb and stall conditions are contradictory to some degree in the considering multi-point optimization case, and the designers should deal with this trade-off carefully based on the features of the low Mach number general aircraft.

Knowledge discovery with computational fluid dynamics: Supercritical airfoil database and drag divergence prediction

Aerodynamic rules and knowledge are often obtained through theoretical research and experiments, which have contributed greatly to aircraft design. For example, Korn's equation predicts the airfoil drag divergence Mach number using the airfoil maximum thickness and the lift coefficient. It is very helpful in the aircraft initial design. However, it neither reveals the key factors of fluid features on the drag divergence nor contributes to the detailed design. This paper designs a supercritical airfoil database that covers the typical free stream Mach number, angle of attack, lift coefficient, and geometry of modern transonic commercial aircraft. Correlation screening and multivariate regression are carried out to discover knowledge about the airfoil drag divergence Mach number and pressure distribution features. A new linear correlation is discovered and validated by existed airfoil databases. Compared with Korn's equation, the discovered correlation reduces the maximum prediction error by approximately 40%. It indicates that the drag divergence Mach number can be increased by obtaining a shock wave that is further upstream in the detailed design. Furthermore, it enables the cruise performance and drag divergence Mach number to be predicted with only one simulation of the cruise point, which will greatly save the computational cost of optimizations.

Optimization and Design of a Flexible Droop Nose Leading Edge Morphing Wing Based on a Novel Black Widow Optimization (B.W.O.) Algorithm—Part II

This work presents an aerodynamic and structural optimization for a Droop Nose Leading Edge Morphing airfoil as a high lift device for the UAS-S45. The results were obtained using three optimization algorithms: coupled Particle Swarm Optimization-Pattern Search, Genetic Algorithm, and Black Widow Optimization algorithm. The lift-to-drag ratio was used as the fitness function, and the impact of the choice of optimization algorithm selection on the fitness function was evaluated. The optimization was carried out at various Mach numbers of 0.08, 0.1, and 0.15, respectively, and at the cruise and take-off flight conditions. All these optimization algorithms obtained effectively comparable lift-to-drag ratio results with differences of less than 0.03% and similar airfoil geometries and pressure distributions. In addition, an unsteady analysis of a Variable Morphing Leading Edge airfoil with a dynamic meshing scheme was carried out to study its flow behaviour at different angles of attack and the feasibility of leading-edge downward deflection as a stall control mechanism. The numerical results showed that the variable morphing leading edge reduces the flow separation areas over an airfoil and increases the stall angle of attack. Furthermore, a preliminary investigation was conducted into the design and sensitivity analysis of a morphing leading-edge structure of the UAS-S45 wing integrated with an internal actuation mechanism. The correlation and determination matrices were computed for the composite wing geometry for sensitivity analysis to obtain the parameters with the highest correlation coefficients. The parameters include the composite material qualities, thickness, ply angles, and the ply stacking sequence. These findings can be utilized to design the flexible skin optimization framework, obtain the target droop nose deflections for the morphing leading edge, and design an improved model.

Machine Learning to Predict Aerodynamic Stall

A convolutional autoencoder is trained using a database of airfoil aerodynamic simulations and assessed in terms of overall accuracy and interpretability. The goal is to predict the stall and to investigate the ability of the autoencoder to distinguish between the linear and non-linear response of the airfoil pressure distribution to changes in the angle of attack. After a sensitivity analysis of the learning infrastructure, we investigate the latent space identified by the autoencoder targeting extreme compression rates, i.e. very low-dimensional reconstructions. We also propose a strategy to use the decoder to generate new synthetic airfoil geometries and aerodynamic solutions by interpolation and extrapolation in the latent representation learned by the autoencoder.

익형 주변 박리 유동 예측 향상을 위한 데이터 기반 난류 모델 개선

A data-driven approach is investigated to improve the Spalart-Allmaras (SA) turbulence model. The field inversion and machine learning framework is chosen for the model improvement. Flow around the S809 airfoil is chosen for the test case. In this approach, a spatial-varying correction term is obtained using the discrete adjoint method in the field inversion process. Then, an artificial neural network is constructed to generalize the correction term with relevant flow features. This study shows that the corrected SA model reduces the turbulence production near the separation point, which leads to the improved prediction of the stalled airfoil at high angles of attack. Detailed investigation on modified flow fields and airfoil pressure distribution is conducted in this study to explain how the machine-learned model improves the turbulence model for separated flow.

Investigation of the Normal Blowing Approach to Controlling Wingtip Vortex Using LES

The characteristics and control of a wingtip vortex are of great significance when considering drag reduction and flight safety of transportation aircrafts. The associated aerodynamic phenomenon resulting from rolling up of a wingtip vortex includes boundary layer flow, shear layer separation, and vortex breakdown, while the interaction of a wingtip vortex with the airframe causes induced drag, wingtip noise, etc. This paper studies a normal blowing method utilized to control the wingtip vortex. Large eddy simulation (LES) technique applied to a straight NACA0012 wing having a chord length ( c ) of 0.4 m is adopted for this study. The Reynolds number based on the chord length is 1.6 × 10 6 and the angle of attack is 12°. The computational approach utilized the dynamic Smagorinsky-Lilly subgrid model for 3D simulations. Normal blowing from a high aspect ratio jet from the wingtip lower surface was used to control the wingtip vortex. From 0.05c to 0.30c, the blowing slit width was 1 mm, with the slit exit treated as a velocity inlet boundary condition which supplied the blowing jet with a momentum coefficient of 0.28%. Results of axial velocity and span-wise pressure distribution of the clean airfoil presented good agreement with known experimental data. LES results indicate that normal blowing suppresses the primary vortex strength, while the vortex core radius, maximum induced velocity, axial vorticity flux, and pressure peak of the primary vortex are reduced by 25%, 28%, 46%, and 52%, respectively. Flow field structures before and after blowing show that blowing suppresses the shedding, coiling, and convergence of the free vortex layers near the primary vortex. This study also shows that normal blowing generates a jet-induced vortex at the location of the secondary vortex, while backflow, volume expansion, and spiral burst can be observed in the jet-induced vortex. The bursting jet-induced vortex destroys the jet-like flow structure of the primary vortex at the trailing edge.

Effects of Geometrical Parameters of Internal Blown Flap and Its Optimal Design

The internal blown flap was numerically simulated. Firstly, a parameterization method was developed, which can properly describe the shape of the internal blown flap according to such geometrical parameters as flap chord length, flap deflection, height of blowing slot and its position. Then the reliability of the numerical simulation was validated through comparing the pressure distribution of the CC020-010EJ fundamental generic circulation control airfoil with the computational results and available experiment results. The effects of the geometrical parameters on the aerodynamic performance of the internal blown flap was investigated. The investigation results show that the lift coefficient increases with the increase of flap chord length and flap deflection angle and with the decrease of height of blowing slot and its front position. Lastly, a method of optimal design of the geometrical parameters of the internal blown flap was developed. The design variables include flap chord length, flap deflection, height of blowing slot and its position. The optimal design is based on maximum lift coefficient, the angle of attack of 5 degrees and the design constraint of stall angle of attack of less than 9 degrees. The optimization results show that the optimal design method can apparently raise the lift coefficient of an internal blown flap up to 1.7.

Aero-Propulsive and Propulsor Cross-Coupling Effects on a Distributed Propulsion System

This study aims to characterize the complex propulsive-airframe and cross-propulsor interactions that occur on an overwing distributed propulsion system with boundary-layer ingestion. Wind tunnel experiments were performed on a NACA airfoil model with an array of five ducted fans integrated across the upper-surface trailing edge. Net forces lift, drag, and pitching moment characteristics, as well as pressure distributions and particle image velocimetry (PIV) velocity fields, were experimentally characterized as a function of fan throttle setting. Additionally, various mixed-throttle cases were investigated with certain fans inoperative windmilling. The experimental results revealed nonlinearities in the changes of the streamwise and stream-normal forces lift, drag, and pitching moment polars as a function of and throttle setting. These nonlinear variations in the airfoil model forces and moments were attributed to variation in the magnitude and direction of the fan thrust vector, which is tied to changes in the airfoil pressure distribution and an induced circulation effect. The ingested boundary-layer height was seen to decrease linearly with a uniform increase in throttle setting. The fan-out cases showed that a windmilling inoperative fan on the edge of the array was more detrimental to performance than a windmilling inoperative fan located within the fan array, and a highly three-dimensional flow was observed from the PIV velocity fields and surface pressure distributions.

Experimental Investigation of Spoiler Deployment on Wing Stall

Upper surface wing flaps, known as spoiler, are typically used to reduce lift and increase drag at touchdown; however spoilers have been shown to increase lift and reduce drag at near-stall conditions. The purpose of this experiment was to determine the spoilers’ impact on lift, drag, moment, and aerodynamic efficiency of a NACA 2412 airfoil at angles of attack (α) from −8 ° to 32 °. The experiment was conducted in the Ryerson Low-Speed Wind Tunnel (closed-circuit, 1 m × 1 m test section) at Re=783761, Ma=0.136. The lift coefficient (Cl), drag coefficient (Cd), moment coefficient about the quarter-chord () were captured with a changing spoiler deflection angle (δ) and spoiler length (b in percent chord). It was found that deflecting the spoiler resulted in an increase maximum lift of up to 2.497%. It was found that deflecting the spoiler by 8° was optimal for the b=10 cases. Any larger deflection reduced the lift gain, and a deflection of 25° caused the maximum lift to be 2.786% less than the clean configuration. In the b=15 case, δ=15° was optimal (1.760% maximum lift coefficient increase). The b=10 cases increased maximum lift coefficient between 0.35% and 2.10% higher than the b=15 cases. The source of the lift gain at high angles of attack is apparent in an analysis of the airfoil pressure distribution. The spoiler increased the suction peak on the airfoil surface upstream of the spoiler, and increased the pressure downstream. However the suction increase upstream is larger than the pressure increase downstream, resulting in a net increase in lift. The spoiler increased the stall angle 37.658% to 87.658% higher than the clean configuration. Stall angle increased with both δ and with an increased spoiler length. The spoiler airfoil produced less drag than the clean configuration at high angles of attack. The combination of the increased lift, and reduced drag resulted in an increase in aerodynamic efficiency at high angle of attack.

Adaptive response prediction for aerodynamic shape optimization

PurposeThe purpose of this paper is to reduce the overall computational time of aerodynamic shape optimization that involves accurate high-fidelity simulation models.Design/methodology/approachThe proposed approach is based on the surrogate-based optimization paradigm. In particular, multi-fidelity surrogate models are used in the optimization process in place of the computationally expensive high-fidelity model. The multi-fidelity surrogate is constructed using physics-based low-fidelity models and a proper correction. This work introduces a novel correction methodology – referred to as the adaptive response prediction (ARP). The ARP technique corrects the low-fidelity model response, represented by the airfoil pressure distribution, through suitable horizontal and vertical adjustments.FindingsNumerical investigations show the feasibility of solving real-world problems involving optimization of transonic airfoil shapes and accurate computational fluid dynamics simulation models of such surfaces. The results show that the proposed approach outperforms traditional surrogate-based approaches.Originality/valueThe proposed aerodynamic design optimization algorithm is novel and holistic. In particular, the ARP correction technique is original. The algorithm is useful for fast design of aerodynamic surfaces using high-fidelity simulation data in moderately sized search spaces, which is challenging using conventional methods because of excessive computational costs.

Performance improvement of a wind turbine blade using a developed inverse design method

The purpose of this study is to improve the aerodynamic performance of wind turbine blades, using the Ball-Spine inverse design method. The inverse design goal is to calculate a geometry corresponds to a given pressure distribution on its boundaries. By calculating the difference between the current and target pressure distributions, geometric boundaries are modified so that the pressure difference becomes negligible and the target geometry can be obtained. In this paper, The Ball-Spine inverse design algorithm as a shape modification algorithm is incorporated into CFX flow solver to optimize a wind turbine airfoil. First, the presented inverse design method is validated for a symmetric airfoil in viscous incompressible external flows. Then, the pressure distribution of the asymmetric airfoil of a horizontal wind turbine is modified in such a way that its loading coefficient increases. The lift coefficient and lift to drag ratio for the new modified airfoil get 5% and 3.8% larger than that of the original airfoil. The improved airfoil is substituted by the original airfoil, respectively. in the wind turbine. Finally, the aerodynamic performance of the new wind turbine is calculated by 3-D numerical simulation. The results show that the power factor of the new optimized wind turbine is about 3.2% larger than that of the original one.

Aerodynamic optimization method of supercritical airfoil geared to the performance of swept and tapered wing

This paper presents an aerodynamic optimization method of supercritical airfoil geared to the performance of a swept and tapered wing based on a revised pressure coefficient transformation equation, which considers the tapered/swept effects and the influence of local swept angle, as well as the local curvature effects of supercritical wing. The airfoil optimized by the present method (2.75D method) is assembled to the outboard part of a supercritical wing of a wide body civil airplane. An outboard part of supercritical wing with good performance is achieved through this optimization method. The result of optimization design using 2.75D method is compared with traditional two-to three-dimensional transformation design method (2.5D method) and full three-dimensional optimization design method (3D method). The comparison results show that, when compared with 2.5D method, because of considering local swept angle and local curvature revision, the 2.75D method presented in this paper can reflect the tapered effects of three-dimensional wing in airfoil design, and has better agreement between the pressure distribution of airfoil and the corresponding section on the three-dimensional wing, so that better three-dimensional wing will be found in the optimization; on the other hand, when compared with full three-dimensional design method, the presented 2.75D method will greatly improve the optimization efficiency.

High-Lift Airfoil Interacting with a Vortical Disturbance: Wind-Tunnel Measurements

This paper describes experimental investigations on the effect of large-scale disturbances interacting with the flow around a two-element airfoil. Quasi two-dimensional transversal vortices are created in a closed wind-tunnel test section shortly after the nozzle. Convecting along the test section, these vortices interact with a two-element airfoil in high-lift configuration. Various measurement techniques are used to identify and characterize these complex interactions. The techniques are phase-locked particle image velocimetry, time-resolving five-hole probe, time-resolving static pressure measurements, and time-averaged measurements of the static pressure distribution of the high-lift airfoil. The vortex generation and evolution along the empty wind-tunnel test section is discussed before considering the vortex–airfoil interactions. As a result, the influence of the induced vortices on the time-resolving static pressure along the airfoil surface and the interaction with the airfoil flap flow is described. The findings are linked to flowfield images and the induced flow effects (the circulatory effect and the local velocity effect) are discussed.

Inverse airfoil design using variable-resolution models and shape-preserving response prediction

The paper presents a computationally efficient surrogate-based optimization algorithm for the inverse design of transonic airfoils. Our approach replaces the direct optimization of an accurate, but computationally expensive, high-fidelity airfoil model by an iterative re-optimization of two different surrogate models. Initially, for a few design iterations, a corrected physics-based low-fidelity model is employed, which is subsequently replaced by a response surface approximation model. The low-fidelity model is based on the same governing fluid flow equations as the high-fidelity one, but uses coarser mesh resolution and relaxed convergence criteria. The shape-preserving response prediction (SPRP) technique is utilized to predict the high-fidelity model response, here, the airfoil pressure distribution. In this prediction process, SPRP employs the actual changes of the low-fidelity model response (or the surface approximation model) due to the design variable adjustments. The SPRP algorithm is embedded into the trust region framework to ensure good convergence properties of the optimization procedure. Our algorithm is applied to constrained inverse airfoil design in inviscid transonic flow. A comparison with the basic version of the optimization algorithm, exploiting only a physics-based low-fidelity model, is carried out. While the performance of both versions is similar with respect to their ability to match the target pressure distribution, the improved algorithm offers substantial design cost savings, from 25 to 72 percent, depending on the test case.

Characterization of Low-Frequency Oscillations in the Flowfield About an Iced Airfoil

Wind-tunnel measurements were used to identify and characterize the low-frequency oscillation present in the flowfield about an NACA 0012 airfoil with a horn-ice shape. This low-frequency mode was identified in the unsteady content of the airfoil lift coefficient and the corresponding Strouhal numbers compared well with those reported in the literature. To study the relationship between the ice-induced separation bubble dynamics and the airfoil circulation, the unsteady shear-layer reattachment location was determined using measurements from a surface-mounted hot-film array. At the low-frequency mode, the lift coefficient led the reattachment location by a phase of . Additional characterization of the relationship between the unsteady shear-layer reattachment location and the unsteady airfoil pressure distribution was also completed, along with an analysis of the convective qualities of the low-frequency oscillation throughout the airfoil pressure distribution.