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The numerical simulation of ice melting process on an iced helicopter rotor blade is presented. The ice melting model uses an enthalpy-porosity formulation, and treats the liquid-solid mushy zone as a porous zone with porosity equal to the liquid fraction. The ice shape on the blade section is obtained by the icing code with a dynamic mesh module. Both of the temperature change and the ice-melting process on the rotor blade section surface are analyzed. The phenomenon of ice melting is analyzed through the change of temperature and liquid fraction on the abrasion/ice interface. The liquid fraction change as with time on the abrasion/ice surface is observed, which describes the ice-melting process well. The numerical results show that the ice melting process can be simulated effectively by the melting model. The de-icing process can be monitored by observing the change of the liquid fraction of the area around the abrasion/ice interface.

NASA Glenn Research Center’s Propulsion Systems Laboratory (PSL) allows ice-crystal ice accretion tests on jet engines. This pressurized wind tunnel facility allows engines to be operated at flight altitudes and temperatures. Steady state and unsteady computational fluid dynamics simulations were performed for the PSL geometry, including the spray bars with their supports, and the converging duct section. These simulation results help to characterize the performance of the tunnel and are important for understanding the flow and particle behavior leading up to the engine test section. The results indicate complex flow structures, with vortex shedding and non-uniform flow features. Flow separation is observed in several regions. Several flow features and vortices are seen to persist to the duct exit plane where the fan section of a jet engine would be mounted for testing.

Numerical investigation of airflow at a transonic speed over the wing of the NASA high-lift Common Research Model (CRM) with and without a conformal vortex generator (CVG), placed on the airfoil suction side has been performed. The objective of the investigation was to assess the impact of CVG on the wing’s lift to drag (L/D) ratio and tip vortices. The wing has aspect and taper ratios of respectively 9, and 0.275, and a leading-edge sweep angle of 37.24 degrees. The root and tip chords were respectively 11.81m and 2.73 m with an approximate mean chord of 6.62 m. The angle of attack was 2.5 degrees. The CVG was distorted V-shaped with a base distance of approximately 4.8 cm, a depth of 8.8 cm, and a tip-to root angle of approximately 30.20. The CVG is on both sides of the tape pointing in opposite directions. The tape is 2 mm thick, 83 cm wide, spanning the entire length of the wing surface.

Hybrid (bolted/bonded) joining is becoming one of the innovative joining processes for light weight structures in the transport industry, especially in the aerospace industry where weight reduction and high joining requirements are permanent challenges. Combining the adhesive bonding with the mechanical joining -riveting for instance- can lead to an enhancement of the properties of the joint compared to the wide established riveting, as a result of a synergistic load bearing interaction between the fastener and the adhesive bondline. The influence of the rivet installation process on a hybrid joint regarding the joint stress state, the change of the bondline thickness as well as its effects on the joint performance and load transfer are some of the factors that drive the users to a better understanding of the hybrid joining process.

The automotive underbody diffuser is an expansion device which works by speeding up the air flowing underneath a vehicle. This reduces the pressure below the vehicle thereby increasing downforce. When designed properly, it can lead to a massive gain in downforce and even a reduction in drag. However, a majority of the research and development is restricted to motorsport teams and supercar manufacturers and is highly secretive. Most of the publicly available research has been done for very simple shapes (bluff bodies) to study the effects of ground clearance and rake angle. Very little research has been done for complex geometries with vanes, flaps and vortex generators. This paper aims to investigate the effects of the addition of vanes/strakes and flaps, their location as well as angle, on diffuser performance. Computational Fluid Dynamics simulations have been carried out using three dimensional, steady state RANS equations with the k-ε turbulence model on STAR CCM+ V9.06.

This paper describes the application of a finite volume procedure for a fluid network to predict thermofluid transients during chilldown of cryogenic transfer lines. The conservation equations of mass, momentum, and energy and the equation of state for real fluids are solved in a fluid network consisting of nodes and branches. The numerical procedure is capable of modeling phase change and heat transfer between solid and fluid. This paper also presents the numerical solution of pressure surges during rapid valve opening without heat transfer. The numerical predictions of the chilldown process have been compared with experimental data.

The use of a magnetic field applied to a gaseous conducting boundary layer flow as been proposed as a mean to control boundary layer losses. The solution of the Navier-Stokes equations together with the induction equation for the magnetic field was numerically obtained using a Finite Volume code. The effects of the magnetic field are introduced into the momentum equations by means of a Lorentz force source term. Further, several turbulence models were modified in order to include the magnetic field perturbation. The computation of the transitional flow on a flat plate was made for several values of the magnetic field, and a relaminarization of the flow was obtained for sufficient high values of the magnetic field. The code was also applied for the computation of the flow around an airfoil as an example application of an high speed high-altitude UAV.

A CFD simulation methodology for the inclusion of the post-impact trajectories of splashing/bouncing Supercooled Large Droplets (SLDs) and film detachment is introduced and validated. Several scenarios are tested to demonstrate how different parameters affect the simulations. Including re-injecting droplet flows due to splashing/bouncing and film detachment has a significant effect on the accuracy of the validations shown in the article. Validation results demonstrate very good agreement with the experimental data. This approach is then applied to a full-scale twin-engine turboprop to compute water impingement on the wings and the empennage.

This paper details the process involved in the numerical optimisation of a helicopter engine inlet electrothermal ice protection system. Although the process was developed using a production aircraft, it is demonstrated here using a generic intake and flight conditions, due to confidentiality of the actual design. The process includes adherence to the overall system design objectives (maximum power demand), including tolerances required to account for an industrial system (aircraft voltage variation, manufacturing tolerances). The numerical optimisation was performed using a combination of 2D and 3D methods to define the required heated area, power density, locations and settings for temperature control sensors. The use of 2D design tools allows a rapid iteration process to be performed, leading to the possibility of a higher level of optimisation within the allowable time-frame compared to the use of full 3D methods.

This paper presents the basic test data obtained from tests of a cabin air distribution system in a simulated Space Station Man-Tended Capability (MTC) configuration and correlations of some of this data with the results from analytical modeling of the test setup flow conditions. The MTC configuration simulated in the test setup includes: Lab-A, the Node, the Cupola, and the Pressurized Module Adaptor (PMA). The test data and analytical data presented are confined to those for the Lab module. The cabin air distribution system controls the flow of air in the open space of a Space Station module. In order to meet crew comfort criteria the local velocities for this cabin air are required to be distributed within a specified range with upper and lower limits.

This paper presents the results of a study in which the free rolling behavior of a F-16 tire was numerically modeled. The tire contact patch normal and shear stresses as well as the displacement distributions were obtained from a three dimensional finite element computer program used at the Wright-Patterson Air Force Base, Ohio. It is shown how the predicted deflections are in reasonable agreement with the rated load vs. deformation characteristics, while predicting the effective rolling radius using a theoretical solution. A significant development of this work is the formulation and execution of a finite difference algorithm to evaluate the contact patch slip velocity distribution by methodically manipulating the above computer program results. Slip velocities are then utilized in assessing the rate of slip energy generation at the contact patch, which directly contributes to tire wear. Finally, it is shown how even a low brake slip ratio can increase the contact patch slip energy.

Sealant is applied between joined aircraft parts in the final stage of the assembly, before installation of permanent fasteners. In this paper a novel approach for aircraft assembly simulation is suggested, which allows to resolve the transient interaction between parts and sealant in the course of airframe assembly process. The simulation incorporates such phenomena as compliance of parts, contact interaction between them and fluidity of sealant with presence of free surface. The approach based on fluid-structure interaction techniques consists of two basic steps: at the first one the pressure of sealant is found after corresponding fluid dynamics problem is solved and at the second the displacements of parts and sealant are calculated through the solving of contact problem. Iterations between structural and fluid dynamics solvers are performed to achieve convergence. The developed approach is demonstrated on example of joining of two test aircraft panels.

A 3D CFD methodology is presented to simulate ice build-up on propeller blades exposed to known icing conditions in flight, with automatic blade pitch variation at constant RPM to maintain the desired thrust. One blade of a six-blade propeller and a 70-passenger twin-engine turboprop are analyzed as stand-alone components in a multi-shot quasi-steady icing simulation. The thrust that must be generated by the propellers is obtained from the drag computed on the aircraft. The flight conditions are typical for a 70-passenger twin-engine turboprop in a holding pattern in Appendix C icing conditions: 190 kts at an altitude of 6,000 ft. The rotation rate remains constant at 850 rpm, a typical operating condition for this flight envelope.

A numerical approach based on a finite element method has been developed to simulate the two-dimensional blade-vortex interaction in viscous flows. In the scheme, the incompressible Navier-Stokes equations in the vorticity-stream form are used. When the passing vortex is approaching the airfoil from upstream, two different models for the vortex are used. Firstly, the vortex is modeled as an ivicsid vortex convecting at the speed of the local fluid, and the total flow is the superposition of the vortex-induced flow and the viscous flow. Secondly, when the vortex is close to the airfoil, it is distributed in a weighted manner at the nodes of the element where it is located, and then the total flow is governed by the Navier-Stokes equations. Some numerical results are presented, and the potential of the present work is discussed.

Numerical simulations of compressible inviscid flows are carried out for the complete configuration of experimental aircraft ‘ASKA’ which adopts the USB technology to increase the amount of lift force. Three different grid systems corresponding to different configurations are generated by a newly developed interactive grid generation method. Euler equations are solved by the second order upwind biased finite volume method. A planar Gauss-Seidel relaxation method is adopted to realize a rapid convergence to steady solutions. Computations are made to see the influences of different arrangements of engine nacelles over the interfered flow fields.

Electro-thermal deicing process was an unsteady heat transfer process including phase change. Based on the investigation of such a process, a code was developed to numerically simulate electro-thermal deicing process. Phase change was performed by an enthalpy method. A staircase approach was used to describe the variable ice thicknesses along the icing surface. The control volume method was adopted to discretize the governing equations. Tri-diagonal matrix method, alternating direction implicit method and block-correction technique were used to solve the discrete equations. Results of temperature distribution in this investigation were compared with experimental results of previous study. Their good agreements indicate the validity of our simulation. The effects of icing conditions, such as ambient temperature, liquid water content (LWC) and flight velocity, etc., were analyzed through a case. Some useful conclusions were achieved.

In helicopter, the icing rotor blades will decrease the effectiveness of the helicopter and endanger the lives of the pilots. The asymmetrical ice break-up and shedding could also lead to severe vibrations of the rotor blade. Ice break-up from the main rotor may strike the fuselage and tail rotor, even worse, find its way into the engine, which may cause serious aircraft accidents. An understanding of the mechanisms responsible for ice shedding process is necessary in order to optimize the helicopter rotor blade design and de-icing system to avoid hazardous ice shedding. In this paper, the ice shedding model is improved by introducing a bilinear cohesive zone model (CZM) to simulate the initiation and propagation of ice/blade interface crack. A maximum stress criterion is used to describe the failure occurred in the ice.

The objective of this paper is to present a numerical simulation method to calculate the gasification of porous particles of different sizes and composition which takes into account morphological changes of the inner pore structure. Different relationships for the evolution of the inner surface dependent on the degree conversion are assessed. In a first approach the gasification mechanism is represented by a one-step reaction. The method operates on the one-dimensional and transient conservation equation for mass and energy for spherical particles with external heat and mass transfer. The solution yields the distribution of temperature, inner surface, porosity, concentration of solid material and gaseous components versus radius and time. One of the main novelties of this approach is that it covers the entire reaction regime between reacting and shrinking core model. Satisfactory agreement is obtained between measured and calculated results.

Ice accretion is a phenomenon in which supercooled water droplets impinge and accrete on a body. In the present study, we focus on a jet engine because it is well known that ice accretion on blades and vanes leads to performance degradation and has caused severe accidents. Although various anti-icing and deicing systems have been developed, such accidents still occur. Therefore, it is important to clarify the phenomenon of ice accretion in a jet engine. However, flight tests for ice accretion are very expensive, and in the wind tunnel it is difficult to reproduce every climate condition where ice accretion occurs. Therefore, it is expected that computational fluid dynamics (CFD), which can estimate ice accretion in various climate conditions, will be a useful way to predict the ice accretion phenomenon. The characteristic phenomena of supercooled large droplets (SLD) are splash and bounce.

A CFD simulation methodology is presented to calculate blockage due to ice crystal icing of the IGV passages of a gas turbine engine. The computational domain consists of six components and includes the nacelle, the full bypass and the air induction section up to the second stage of the low-pressure compressor. The model is of a geared turbofan with a fan that rotates at 4,100 rpm and a low-pressure stage that rotates at 8,000 rpm. The flight conditions are based on a cruising speed of Mach 0.67 in Appendix-D icing conditions with an ice crystal content is 4.24 g/m3. Crystal bouncing, and re-entrainment is considered in the calculations, along with variable relative humidity and crystal melting due to warmer temperatures within the engine core. Total time of icing is set to 20 seconds. The CFD airflow and ice crystal simulations are performed on the full 6-stage domain. The initial icing calculation determines which stage will be chosen for a more comprehensive analysis.