Simulated Ice Shapes Research Articles

Bayesian inference of airfoil icing condition from simulated ice shapes

Bayesian inference of airfoil icing condition from simulated ice shapes

The Influence of Angle of Attack on the Icing Distribution Characteristics of DU97 Blade Airfoil Surface for Wind Turbines

This study explores the influence of angle of attack (AOA) on the icing distribution characteristics of asymmetric blade airfoil (DU97) surfaces for wind turbines under icing conditions by numerical simulation. The findings demonstrate a consistence between the simulated ice shapes and experimental data. The ice thickness distribution on the lower surface of the leading edge exhibits a trend of first rising and then declining along the chord direction while showing a gradually decreasing trend on the upper surface. The ice distribution range on the upper surface of the trailing edge is broader than that on the lower surface. The peak ice thickness at the trailing edge rises significantly as AOA increases from 5° to 10°, and at the leading edge raises dramatically at droplet sizes of 30–40 μm and wind speeds of 5–10 m/s. The peak ice thickness is more significantly influenced by AOA than by ambient temperature due to the combined effect of airflow characteristics induced by AOA and latent heat (phase change) and sensible heat (thermal convection and thermal radiation) caused by ambient temperature. The findings offer valuable insights into the flow and heat transfer physics, and can operate as references for wind turbine anti/de-icing technology.

Low-Dimensional Models for Aerofoil Icing Predictions

Determining the aero-icing characteristics is key for safety assurance in aviation, but it may be a computationally expensive task. This work presents a framework for the development of low-dimensional models for application to aerofoil icing. The framework builds on: an adaptive sampling strategy to identify the local, nonlinear features across the icing envelope for continuous intermittent icing; a classic technique based on Proper Orthogonal Decomposition, and a modern Neural Network architecture. The extreme diversity in simulated ice shapes, from smooth and streamlined to rugged and irregular shapes, motivated the use of an unsupervised classification of the ice shapes. This allowed deploying the Proper Orthogonal Decomposition locally within each sub-region, sensibly improving the prediction accuracy over the global model. On the other hand, the Neural Network architecture and the convolutional auto-encoder were found insensitive to the complexity in ice shapes. A strong correlation was found to exist between the ice shape, resulting ice mass and aerodynamic performance of the iced aerofoil, both in terms of the average and variance. On average, rime ice causes a loss of maximum lift coefficient of 21.5% compared to a clean aerofoil, and the average ice thickness is 0.9% of the aerofoil chord. For glaze ice, the average loss of maximum lift coefficient is 46.5% and the average ice thickness is 2.1%. Glaze ice was also found to have three times more surface coverage than rime ice.

Automatic roughness characterization of simulated ice shapes

During in-flight ice accretion, roughness plays an important role since it heavily influences the convective heat transfer and skin friction coefficients. This paper aims to assess the ability of existing ice accretion simulation tools to compute the growing ice’s roughness. To this purpose, a technique based on Self Organizing Maps is applied to numerical simulations of in-flight ice accretion to characterize the roughness. The numerical ice predictions are performed using a standard approach comprising RANS computations, Lagrangian particle tracking, the solution of the unsteady Stefan problem, and a morphogenetic model. Numerical simulations are performed on selected benchmark cases from the 1st AIAA Ice Prediction Workshop. Validation of roughness computation is performed on synthetic test cases, while ice roughness is compared directly to that extracted from ice scans. The results of simulated ice shapes compare reasonably well with experimental data. Computations can replicate the trend of the experimental mean ice shape and roughness distribution for both rime and glaze cases.

Research on Value-Seeking Calculation Method of Icing Environmental Parameters Based on Four Rotating Cylinders Array

This paper calculates and analyzes the collision and freezing characteristics of water droplets on the cylindrical conductors with different diameters and deduces the variation law of icing weight change rate on different diameters rotating cylindrical conductors with wind speed, ambient temperature, liquid water content, and median volume diameter. Then, a prediction and value-seeking method of icing environment parameters based on rotating circular conductors is determined. The laboratory test results show that the measured values of ambient temperature T, wind speed U, liquid water content w and median volume diameter MVD are in good conformity with the reference values measured by other instruments, and the average errors of each parameter are 3.41%, 5.44%, 6.08% and 7.20%, respectively. The field natural experiment was carried out at Hunan Xuefengshan Energy Equipment Safety National Observation and Research Station. The obtained icing environmental parameters were substituted into the icing simulation calculation model of the LDQ-100 glass insulator. Compared with the field experiment results, the simulated ice shape and the field ice shape were in good agreement. The error between the numerical calculation results of the icing weight and the field icing weight of the experimental insulator was less than 6%.

Improving aircraft aerodynamic performance with bionic wing obtained by ice shape modulation

For Unmanned Aerial Vehicles (UAVs) with limited electrical power to achieve effectively anti-/de-icing at the leading edge of the wing, a strategy of ice shape modulation was proposed. Isolated simulated ice shape pieces printed by 3D printing technology are mounted on a NACA0012 finite wing model, and its lift/drag coefficients and suction-side velocity fields are measured by the six-component force balance and the Particle Imaging Velocimetry (PIV), respectively. The ratio of the spanwise length of a single ice shape piece to chord length and the spanwise length of the non-icing area between the two adjacent single ice shape pieces are defined as dimensionless ice shape length (w/c) and dimensionless modulation ratio (w/λ), respectively. The results indicate that for a fixed w/λ, the wing lift coefficient first increases and then drops with increasing w/c, and a peak value exists when w/c is between 0.1 and 0.2. The lower the w/λ is, the higher the wing lift coefficient will be. The periodical variation of the flow separation area along the spanwise direction is attributed on the one hand to the acceleration effect of the flow field in the non-icing area which reduces the separation area, and on the other hand to the cross-flow caused by the streamwise vortices from the non-icing area to the icing area which promotes the mixing of the flow field (similar to vortex generators). The obtained modulation law is verified through flight tests and provides guidance for the use of ice shape modulation scheme for UAVs that cannot be completely anti-/de-icing under severe weather conditions.

Research and Development of a Small-Scale Icing Wind Tunnel Test System for Blade Airfoil Icing Characteristics

In order to study the icing mechanism and anti-icing technology, a small low-speed reflux icing wind tunnel test system was designed and constructed. The refrigeration system and spray system were added to the small reflux low-speed wind tunnel to achieve icing meteorological conditions. In order to verify the feasibility of the test system, the flow field uniformity, temperature stability, and liquid water content distribution of the test section were tested and calibrated. On this basis, the icing tests of an aluminium cylinder, an NACA0018 airfoil, and an S809 airfoil were carried out, and the two-dimensional ice shape obtained by the test was compared with the two-dimensional ice shape obtained by the numerical simulation software. The results show that in the icing conditions and icing time studied, the parameters of the test system are stable, and the experimental ice shape is consistent with the simulated ice shape, which can meet the needs of icing research.

Experimental Aerodynamic Simulation of a Scallop Ice Accretion on a Swept Wing

<div class="section abstract"><div class="htmlview paragraph">Understanding the aerodynamic impact of swept-wing ice accretions is a crucial component of the design of modern aircraft. Computer-simulation tools are commonly used to approximate ice shapes, so the necessary level of detail or fidelity of those simulated ice shapes must be understood relative to high-fidelity representations of the ice. Previous tests were performed in the NASA Icing Research Tunnel to acquire high-fidelity ice shapes. From this database, full-span artificial ice shapes were designed and manufactured for both an 8.9%-scale and 13.3%-scale semispan wing model of the CRM65 which has been established as the full-scale baseline for this swept-wing project. These models were tested in the Walter H. Beech wind tunnel at Wichita State University and at the ONERA F1 facility, respectively. The data collected in the Wichita St. University wind tunnel provided a low-Reynolds number baseline study while the pressurized F1 facility produced data over a wide range of Reynolds and Mach numbers with the highest Reynolds number studied being approximately <i>Re</i> = 11.9×10<sup>6</sup>. Past work focused on only three different fidelity variations for ice shapes based on multiple icing conditions. This work presents a more detailed investigation into several fidelity representations of a single highly three-dimensional scallop ice accretion. Sensitivity to roughness size and application technique on a low-fidelity smooth ice shape is described. The data indicate that the aerodynamic performance is not especially sensitive to the grit variations. An ice accretion code was also used to generate ice shapes for aerodynamic testing and comparisons. These ice shapes have a general appearance like the low-fidelity smooth ice shapes, but in this case, the computer-generated ice shape is significantly smaller. As such, the impact of that ice shape on the aerodynamic performance of the wing is reduced compared to the smooth ice shape based on the icing experiment for those same conditions. Spanwise discontinuities were also introduced to a low-fidelity ice shape in an attempt to quantify the impact of those variation in the high-fidelity ice shape. While the lift data indicate good agreement between the high-fidelity ice shapes and the low-fidelity ice shapes with spanwise discontinuities, a closer investigation of the data suggests potential, significant differences in the flowfield. These results were similar at both facilities over the wide range of test conditions utilized.</div></div>

Numerical Study of Aerodynamic Characteristics of a Symmetric NACA Section with Simulated Ice Shapes

To develop a numerical model of icing on wind turbine blades, a CFD simulation was conducted to investigate the effect of critical ice accretions on the aerodynamic characteristics of a 0.610 m chord NACA 0011 airfoil section. Aerodynamic performance coefficients and pressure profile were calculated and compared with the available measurements for a chord Reynolds number of 1.83x106. Ice shapes were simulated with flat plates (spoiler-ice) extending along the span of the wing. Lift, drag, and pressure coefficients were calculated in zero angle of attack through the steady state and transient simulations. Different approaches of numerical studies have been applied to investigate the icing conditions on the blades. The simulated separated flow over the sharp spoilers is challenging and can be seen as a worst test case for validation. It allows determining a reliable strategy to simulate real ice shapes [1] for which the detailed validation cannot easily be provided.

New 3-D ice accretion method of hovering rotor including effects of centrifugal force

A new numerical method for predicting 3-D ice accretion on a helicopter rotor in hover is proposed. The new method creatively takes the centrifugal force into account, and the prediction of ice accretion is fulfilled by loosely coupling the novel 3-D icing model with a CFD method. Considering the complexity of ice shapes on a rotor, orthogonally body-fitted grids around rotor blades are generated and modified by solving Poisson equations, and the CFD method for calculating the rotor flowfield in hover is developed by solving Reynolds-averaged Navier–Stokes (RANS) equations and employing the Spalart–Allmaras (S–A) turbulence model. Based upon an embedded grid system around the rotor, an Eulerian method is presented to obtain the water droplet impingement property, and the multistage Runge–Kutta scheme is adopted for temporal discretization. Subsequently, the new 3-D icing model which considers the effect of centrifugal forces is proposed in the freezing process. In this model, the shear stress, pressure gradient and the centrifugal force drive the water film to move on the blade surface. The ice accretion on a UH-1H helicopter rotor in hover is calculated to verify the accuracy of the new method, and the simulated ice shapes on the rotor blade (especially near the blade tip) correlate better with the experimental data than the numerical results of 2-D icing models. Finally, the influence of the centrifugal force in the icing process is analyzed, and numerical results indicate that the total ice amount would decrease when the centrifugal force is taken into account, which have better agreements with the experimental data.

EFFECT OF ICE FORMATION ON AIRFOILS PERFORMANCE (PART-I)

A numerical investigation was conducted to study the effect of simulated ice accreted on airfoilaerodynamics performance. The simulated ice shapes were tested on NACA 0012 airfoil wing atdifferent Mach numbers. The study includes the one of the famous types of ice accreted on theairfoils called the rime ice.The calculation of ice droplet trajectories was performed by solving the trajectories governing equations of the droplet using FVM. A numerical model based on staggered FVM is built up to solve the governing equations of a body fitted grid, trajectories equation, continuity equation and momentum equation using FORTRAN 97. The turbulence model of (k-ε) has been adopted in the programming to evaluate the turbulence effect. The program is valid for any type of 4 or 5 digits airfoil. The program is available to evaluate the rime ice accumulation. The pressure, lift, drag and pitching moment coefficients are computed and compared with that of clean case results. The program was run over different Mach numbers, to compare the results obtained at these Mach Numbers. The investigation of the work was tested on NACA 0012 in a range of angle of attack 0º to 6º, where stall starts at this angle of attack as demonstrated by the results. The results show that the severity of ice formation could be more dangerous with increasing the angle of attack or the Mach number.

Sensitivity of Aircraft Performance to Icing Parameter Variations

A IRCRAFT icing is widely recognized as a significant hazard to aircraft operations. For this reason, aircraft and ice protection systems must be certified for flight into icing conditions. The aircraft certification and icing research communities rely on icing wind tunnels as an efficient way to produce ice accretions in a controlled environment. The aerodynamic performance of aircraft with these ice accretions contributes to the certification of aircraft. It is therefore critical to ensure that the ice accretions are simulated within a known aerodynamic uncertainty. The aerodynamic performance of an airfoil containing an ice accretion is highly dependent on the geometry of the ice accretion, which is in turn dependent on icing conditions such as temperature, liquid water content (LWC), and median volume diameter (MVD). Icing wind tunnels have the capability to vary LWC and MVD, however, the accuracy in LWC and MVD required to create an aerodynamically representative ice accretion is not known. This research addressed the problem of “how good is good enough” by determining the relationship and sensitivity of iced-airfoil performance to these icing cloud parameters. In addition, these data were placed in perspective by relating measurable or significant aircraft performance changes to the underlying changes in airfoil aerodynamic performance. Recent NASA studies [1,2] in the Icing Research Tunnel (IRT) measured the effect of icing parameter variations on ice-accretion geometry. These studies showed that small variations in LWC and MVD corresponded to distinct changes in ice-accretion geometry. In addition, the effect of ice-accretion geometry on aerodynamic performance has been recently investigated [3–6]. Papadakis et al. [3,4] used spoilers to simulate horn ice and showed that Clmax degradation was related to the horn height (k=c). Kim and Bragg [5], and Broeren et al. [6] showed that k=c and surface location (s=c) had the biggest impact on airfoil performance degradation. This study used ice tracings from theNASA studies [1,2] as a basis to examine the sensitivity of aerodynamic performance to icing parameter variations. Eleven ice-accretion tracings were selected from the 39measured byMiller et al. [2] to reasonably span the range of LWC andMVD tested. The selected ice tracings were modeled as two-dimensional smooth simulated ice shapes for wind-tunnel testing. The experiments for this research were performed in the Illinois subsonic, low-turbulence, open-return wind tunnel. The airfoil model was an aluminum NACA 0012 airfoil with an 18 in. chord, 33.6 in. span, and a removable leading edge to facilitate installation of the ice simulations. Testing was performed at a Reynolds number of 1.8 million, and a Mach number of 0.18. The results of the aerodynamic testing were related to the corresponding icing parameters in the form of two sensitivities: airfoil performance to icing parameter variations, and derived aircraft performance to icing parameter variations. More details can be found in Campbell et al. [7,8].

Detached Eddy Simulations for Iced Airfoils

To improve the predictive ability of the stall behavior for iced airfoils, computations using the detached eddy simulation (DES) approach were conducted. The WIND code with the DES methodology was first validated with a well-documented flow past a circular cylinder. The validation results were in good agreement with the experimental and previous numerical results for the time-averaged aerodynamic forces, as well as for the unsteady flowfield pattern. DES computations were then performed for a series of angles of attack (around the stall angle) fo rN ACA 23012 and natural laminar flow (NLF) 0414 airfoils with a simulated (two-dimensional) ice shape on the upper surface. The spanwise domain for these simulations was scaled based on the expected height of the separation region, consistent with previous blunt-body large eddy simulation and DES. The DES for both iced airfoils provided the maximum lift coefficients and stall angles that were qualitatively consistent with experiments [whereas Reynolds-averaged Navier‐Stokes (RANS) simulations of the same cases failed to provide either such behavior]. However, the DES approach was much more computationally intensive, and even higher grid resolution within the separated flow region is suggested to improve the predictive performance for this technique.

Effect of Airfoil Geometry on Performance with Simulated Intercycle Ice Accretions

This paper presents the results of an experimental study designed to evaluate the performance effects of intercycle ice accretions on airfoils with different geometries. The intercycle ice accretions were simulated using combinations of various size grit roughness. These simulations were tested on three airfoils: NACA 23012, NACA 3415, and NLF 0414 at a Reynolds number of 1.8 × × 10 6 and a Mach number of 0.18. Results from the NACA 23012 airfoil tests closely matched those from a previous study, validating the ice-shape simulation method. This also showed that a simple geometric (chord-based) scaling of the ice was appropriate. The simulated ice effect, in terms of maximum lift performance, was most severe for the NACA 23012 airfoil. The maximum lift coefficients were in the range of 0.65 to 0.80 for the iced configuration compared to a clean value of 1.47 for the NACA 23012 airfoil at this Reynolds number. In contrast, the maximum lift coefficients for the NLF 0414 airfoil with the same ice simulations were in the range of 0.90 to 1.05, compared to a clean value of 1.34. The results for the NACA 3415 with the simulated intercycle ice shapes were between the other two airfoils.

Experimental Investigation of Simulated Large-Droplet Ice Shapes on Airfoil Aerodynamics

An experimental investigation was conducted to study the aerodynamic effect of simulated supercooled largedroplet ice accretion on a modie ed NACA 23012 airfoil. Forward-facing quarter-round simulations with heightto-chord ratios of 0.0083 and 0.0139 were used at a Reynolds number of 1 :8 £ 10 6 . When the simulated ice was placed at critical chordwise locations, a long separation bubbleformed downstream of thesimulated ice shape and effectively eliminated the formation of a large leading-edge suction peak that was observed on the clean NACA 23012 airfoil. This resulted in a dramatic reduction in the maximum lift coefe cient, as low as 0.27, when the larger simulated ice shape was located at 12% chord. Because the airfoil loading distribution was severely altered, large changes in airfoil pitching moments and e ap-hinge moments were also observed.

Computational Investigation of Simulated Large-Droplet Ice Shapes on Airfoil Aerodynamics

The objective of this research was to study numerically the effects of simulated spanwise-step-ice accretions (resulting from supercooled large-droplet icing conditions ) on subsonic aircraft aerodynamics. The investigation was performed with a high-resolution full Navier ‐Stokes code with a solution-adaptive unstructured grid for both non-icedandicedcone gurations.Theairfoilinvestigatedwasamodie ed NACA23012 witha simplee ap.Simulated ice shapesweretested on theairfoilto determinethe sensitivity oficeshapesizeon airfoil performanceand control. Predictions of sectional aerodynamic characteristics for quarter-round ice shape heights of 0.0083 and 0.0139 chords are presented and compared with experimental data. Signie cant reductions in lift were noted for these relatively small protuberances, which were consistent with experimental e ndings. The results also indicate good predictive performance for drag, pitching-moment, and hinge-moment variations.

Predicting Galloping Amplitudes

The use of wind tunnel data, lift and drag, together with a method of analysis, leads to estimates of gallop amplitudes. No damping is considered. The nonlinear analysis is similar to the describing‐function method used in control system design. The method is convenient to use because it can account for effects of Reynold's number directly. Two bluff body shapes are chosen to illustrate the prediction method. The first shape is the square section. This is also compared with earlier results. The second shape is the simulated ice shape on a round conductor. The results compare favorably with previous work.

Experimental aerodynamic characteristics of an NACA 0012 airfoil with simulated ice

An experimental program was conducted with an NACA 0012 airfoil to investigate the effect of Reynolds number on the aerodynamic coefficients with and without a leading-edge simulated ice shape. Six Reynolds numbers from 0.36 to 3.36 x 106 were investigated at both positive and negative angles of attack through stall. Values of ACd, AQ, ACF , and ACm were measured for the no-ice/generic-ice airfoil configurations. The experimental values have shown that the addition of the generic ice shape causes premature stall with a considerable reduction in C, and stall angle of attack; an increase in the drag values of 120-200% compared to the clean airfoil values; and a significant increase in Cmcl4 resulting in positive values of dCm /dQ. However, the aerodynamic coefficients of the airfoil with the leading-edge simulated ice showed little dependence on Reynolds number throughout the range tested, unlike that of the clean airfoil configuration.