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A research program was conducted to evaluate the effectiveness of icing tunnel hybrid model design. A hybrid design is where the full-scale leading edge of a wing section is maintained only to a certain percentage of the local chord, while the aft section of the model is redesigned into a shortened or truncated planform. An initial study was conducted in 2020 where the ice shape geometries on a full-chord length version of the swept CRM65 wing model were compared to those from the hybrid version of CRM65 that were obtained in the NASA Icing Research Tunnel in 2015. The results were reported in a 2021 paper. For most test conditions, the overall size and shape of the ice shapes compared well. However, the ice shapes from the full-chord model were generally slightly smaller than those from the hybrid model.

Modifications have been implemented in the GlennICE software to accommodate a non-inertial reference frame. GlennICE accepts a flow solution from an external flow solver. It then introduces particles and tracks them through the flow field in a Lagrangian manner. Centrifugal and Coriolis terms were added to the GlennICE software to account for relative frame simulations. The objective of the present paper is twofold. First, to check that the new terms are implemented correctly and that the code still behaves as expected with respect to convergence. And second, to provide some initial insight into an upcoming propeller experiment in the NASA Icing Research Tunnel. The paper presents a description of the code modifications. In addition, results are presented for two operating conditions, and three particle sizes. Each case was simulated with four different grid densities to assess grid dependence.

Cold soaked fuel frost (CSFF) is frost that forms on aircraft wing surfaces following a flight because of cold excess fuel remaining in integrated fuel tanks. Previous investigations by Zhang et al. (2021a) and Zhang et al. (2021b) have focused on experimental measurements and correlation development for frost observed using a small frost wind tunnel employing a thermo-electric cooler to impose a surface temperature for a range of environmental conditions. To model the CSFF approach in more detail, an experimental facility was developed and described by McClain et al. (2020) using a thermal model of an integrated wing fuel tank placed inside of a climatic chamber. In this paper, experimental measurements of CSFF are presented using two aluminum wing skins. One of the skins was created using an aluminum rib structure, and the other skin was created without the rib.

In-flight icing is an important consideration that affects aircraft design, performance, certification and safety. Newer regulations combined with increasing demand to reduce fuel burn, emissions and noise are driving a need for improvements in icing simulation capability. To that end, this paper presents the results of additional ice accretion testing conducted in the NASA Icing Research Tunnel in January 2022 with a large swept wing section typical of a modern commercial transport. The model was based upon a section of the Common Research Model wing at the 64% semispan station with a streamwise chord length of 136 in. The test conditions were developed with an icing scaling analysis to generate similar conditions for a small median volumetric diameter (MVD) = 25 μm cloud and a large MVD = 110 μm cloud. A series of tests were conducted over a range of total temperature from -23.8 °C to -1.4 °C with all other conditions held constant.

Understanding the fundamental details of drop/wall interactions is important to improving engine performance. Most of the drop-wall interactions studies are based on the impact of a single drop on the wall. To accurately mimic and model the real engine conditions, it is necessary to characterize spray/wall interactions with different impingement frequencies at a wide range of wall temperatures. In this study, a numerical method, based on Smoothed Particle Hydrodynamics (SPH), is used to simulate consecutive droplet impacts on a heated wall both below and above the Leidenfrost temperature. Impact regimes are identified for various impact conditions by analyzing the time evolution of the post-impingement process of n-heptane drops at different impingement frequencies and wall surface temperatures. For wall temperature below the Leidenfrost temperature, the recoiled film does not leave the surface.

Numerical solutions have been generated which simulate flow inside an aircraft engine flying at altitude through an ice crystal cloud. The geometry used for this study is the Honeywell Uncertified Research Engine (HURE) which was recently tested in the NASA Propulsion Systems Laboratory (PSL) in January 2018. The simulations were carried out at predicted operating points with a potential risk of ice accretion. The extent of the simulation is from upstream of the engine inlet to downstream past the strut in the core and bypass. The flow solution is produced using GlennHT, a NASA in-house code. A mixing plane approximation is used upstream and downstream of the fan. The use of the mixing plane allows for steady state solutions in the relative frame. The flow solution is then passed on to LEWICE3D for particle trajectory, impact and breakup prediction. The LEWICE3D code also uses a mixing plane approximation at the boundaries upstream and downstream of the fan.

Accurate quantification of 3D shapes of the complex ice structures accreted on wind turbine blades is highly desirable to develop ice prediction models for more accurate prediction of the aerodynamic performance degradation and power reduction due to the ice accretion on wind turbine blades. In the present study, an experimental investigation was conducted to quantitatively characterize the 3D shapes of the ice structures accreted over a DU91-W2-250 wind turbine airfoil model in the Icing Research Tunnel available at Iowa State University (ISU-IRT). A glaze icing condition and a rime icing condition that wind turbines usually experience in winter were duplicated by using ISU-IRT. A high-resolution non-intrusive 3D scanning system was used to make detailed 3D-shape measurements to quantify the complicated ice structures accreted on the wind turbine airfoil model as a function of the ice accretion time.

Elastic soft material/surface, such as Polydimethylsiloxane (PDMS), is a perspective, useful and low-cost hydrophobic and icephobic coating. While it has been reported to have good mechanical durability, its erosion durability under the high impacting of water droplets pertinent to aircraft inflight icing phenomena has not been explored. In this study, the droplet imping erosion characteristics of an icephobic PDMS surface/material is evaluated systematically upon the dynamic impinging of water droplets at different impact velocities (~ up to 75m/s), in comparison with other state-of-the-art icephobic materials/surfaces, such as superhydrophobic surface (SHS) and slippery liquid-infused porous surface (SLIPS). Surprisingly, the contact angle (CA) of the elastic PDMS is shown to have an over 20° increase (from 105° to 128°), which represents better hydrophobicity, after the erosion test which is mainly contributed to the higher roughness of the eroded PDMS surface.

An explorative study was performed to demonstrate the feasibility of using a novel hybrid anti-/de-icing strategy for aircraft icing mitigation. The hybrid method was developed by combining the electro-thermal heating mechanism and specialized surfaces/coatings with different wettabilities. While an electrical film heater was utilized to provide thermal energy around the leading edge of a NACA0012 airfoil model, two different coating strategies, (i.e., (a). Superhydrophobic coating covering the entire airfoil surface to increase droplets bounce-off and accelerate surface water runback vs. (b). super-hydrophilic coating at the leading edge to increase evaporation area + superhydrophobic coating in downstream to prevent runback refreezing) were proposed and evaluated aiming at maximizing the anti-/de-icing efficiency of the hybrid method.

The goal of the present study is to investigate the dynamics of a thin water film on a flat plate and an airfoil using direct numerical simulation (DNS). The first case for a wetted flat plate is used to model former experiments and investigate the dynamics of a wind-driven water film. The second case for a thin film on a NACA 0012 airfoil of chord length 0.5 m is used to investigate the dynamics of a wind-driven water film on a curved surface. Particular attention is paid to the interaction between the liquid film and the air shear-layer above the film. As the incoming airflow moves over the thin water film, instability is triggered at the gas-liquid interface. Interfacial waves develop and are advected downstream. The interaction between the air flow and the interfacial waves induces shedding vortices near the interface, which in turn perturb the liquid film farther downstream. Simulations are performed using the open source multiphase flow solvers, Gerris and Basilisk.

Computational ice shapes were generated on the boundary layer ingesting engine nacelle of the D8 Double Bubble aircraft. The computations were generated using LEWICE3D, a well-known CFD icing post processor. A 50-bin global drop diameter discretization was used to capture the collection efficiency due to the direct impingement of water onto the engine nacelle. These discrete results were superposed in a weighted fashion to generate six drop size distributions that span the Appendix C and O regimes. Due to the presence of upstream geometries, i.e. the fuselage nose, the trajectories of the water drops are highly complex. Since the ice shapes are significantly correlated with the collection efficiency, the upstream fuselage nose has a significant impact on the ice accretion on the engine nacelle. These complex trajectories are caused by the ballistic nature of the particles and are thus exacerbated as particle size increases.

Recently developed dielectric barrier discharge (DBD) plasma-based anti-icing systems have shown great potential for aircraft icing mitigation. In the present study, the ice accretion experiments were performed on to evaluate the effects of different layouts of DBD plasma actuators on their anti-/de-icing performances for aircraft icing mitigations. An array of DBD plasma actuators were designed and embedded on the surface of a NACA0012 airfoil/wing model in different layout configurations (i.e., different alignment directions of the plasm actuators (e.g., spanwise vs. streamwise), width of the exposed electrodes and the gap between the electrodes) for the experimental study. The experimental study was carried out in the Icing Research Tunnel available at Iowa State University (i.e., ISUIRT).

Cold-soaked fuel frost (CSFF) is a form of aircraft wing contamination that occurs when a vehicle caries sufficient fuel for multiple trips or take-offs and landings. Following the first trip, which may reach altitudes above 10,000 m (33,000 ft), the fuel for the subsequent trips is carried in the wing tanks and may reach temperatures below -25 °C. In certain times of the year at some airports, temperatures and humidity levels will form CSFF on the aircraft wing surfaces over the fuel tanks. Unless an exemption is granted for the specific aircraft model, aircraft are not allowed to takeoff if the wing surfaces are contaminated by frost. Because aircraft operators desire to minimize vehicle time spent at airports, aircraft manufacturers are expected to pursue designs that safely operate with CSFF at takeoff and to pursue certification exemptions for aircraft models enabling CSFF takeoffs.

Artificial ice shapes of various geometric fidelity were tested on a wing model based on the Common Research Model. Low Reynolds number tests were conducted at Wichita State University’s Walter H. Beech Memorial Wind Tunnel utilizing an 8.9% scale model, and high Reynolds number tests were conducted at ONERA’s F1 wind tunnel utilizing a 13.3% scale model. Several identical geometrically-scaled ice shapes were tested at both facilities, and the results were compared at overlapping Reynolds and Mach numbers. This was to ensure that the results and trends observed at low Reynolds number could be applied and continued to high, near-flight Reynolds number. The data from Wichita State University and ONERA F1 agreed well at matched Reynolds and Mach numbers. The lift and pitching moment curves agreed very well for most configurations.

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.

The influence of freestream static temperature on roughness temporal evolution and spatial variation was investigated in the Icing Research Tunnel (IRT) at NASA Glenn Research Center. A 53.34 cm (21-in.) NACA 0012 airfoil model and a 152.4 cm (60-in.) HAARP-II business jet airfoil model were exposed to Appendix C clouds for fixed exposure times and thus fixed ice accumulation parameter. For the base conditions, the static temperature was varied to produce different stagnation point freezing fractions. The resulting ice shapes were then scanned using a ROMER Absolute Arm system and analyzed using the self-organizing map approach of McClain and Kreeger. The ice accretion prediction program LEWICE was further used to aid in interrogations of the ice accretion point clouds by using the predicted surface variations of local collection efficiency.

Structures in cold weather environments are susceptible to atmospheric ice formation. A fracture mechanics based approach is proposed for in situ characterization of the interfacial fracture energy of ice on different substrates. This paper summarizes the development of the experimental and analytical framework to measure the ice adhesion energy, calibrated on static ice. The testing configuration utilizes a shaft-loaded blister test to produce stable crack propagation, from a well-defined pre-crack at the interface of the ice layer and the substrate. Measurements of the fracture energy are taken over a range of ice thicknesses and surface roughnesses. The developed analytical framework to estimate adhesion energy are verified and calibrated via finite element numerical simulation of the proposed geometric configuration and employing cohesive surfaces along the interface to simulate the crack nucleation and propagation process.

Many studies have been performed to quantify the formation and evolution of roughness on ice shapes created in Appendix C icing conditions, which exhibits supercooled liquid droplets ranging from 1-50 µm. For example Anderson and Shin (1997), Anderson et al. (1998), and Shin (1994) represent early studies of ice roughness during short-duration icing events measured in the Icing Research Tunnel at the NASA Glenn Research Center. In the historical literature, image analysis techniques were employed to characterize the roughness. Using multiple images of the roughness elements, these studies of roughness focused on extracting parametric representations of ice roughness elements. While the image analysis approach enabled many insights into icing physics, recent improvements in laser scanning approaches have revolutionized the process of ice accretion shape characterization.

Changes in convection coefficient caused by the changes in surface roughness characteristics along an iced NACA 0012 airfoil were investigated in the 61-cm by 61-cm (24 in. by 24 in.) Baylor Subsonic Wind Tunnel using a 91.4-cm (36-in.) long heated aerodynamic test plate and infrared thermometry. A foam insert was constructed and installed on the wind tunnel ceiling to create flow acceleration along the test plate replicating the scaled flow acceleration the along the leading 17.1% (3.6 in.) of a 53.3-cm (21-in.) NACA 0012 airfoil. Two sets of rough surface panels were constructed for the study, and each surface used the same basic random droplet pattern created using the Lagrangian droplet simulator of Tecson and McClain (2013). For the first surface, the roughness pattern was replicated with the same geometry over the plate following a smooth-to-rough transition location noted in historical literature for the case being replicated.

This work presents the results of an experimental study of ice particle impacts on a flat plate made of glass. The experiment was conducted at the Ballistics Impact Laboratory of NASA Glenn Research Center in 2014 and is part of the NASA fundamental research efforts to study physics of ice particles impact on a surface, in order to improve understanding of ice crystal ingestion and ice accretion inside jet engines. The ice particles, which were nominally spherical ranging in initial diameter between 1 and 3.5 millimeters, were accelerated to velocities from 20 to 130 m/s using a pressure gun. High speed cameras captured the pre-impact particle diameter and velocity data as well as the post-impact fragment data. The initial stages of ice particle breakup were captured and studied at 1,000,000 frames per second with a high speed camera imaging at a plane normal to the impact surface.