Aerodynamic Boundary Conditions Research Articles

Numerical analysis of deformation and failure mechanism of metal truss structure of spacecraft under re-entry aerothermodynamic environment

A coupled aerodynamic-thermoelastic analysis framework is presented to simulate the deformation and failure mechanism of metal truss structure for spacecraft under re-entry aerothermodynamic environment. The three-dimensional dynamic thermo-mechanical coupling finite element algorithm is developed by computable modeling of thermoelasticity and heat conduction equations to solve the thermo-mechanical coupling response of spacecraft structure. The aerodynamic boundary conditions such as temperature, heat flux and pressure distribution are computed and provided by the gas-kinetic unified algorithm (GKUA) based on the Boltzmann model equation. The linear interpolation method is designed and employed to couple the boundary conditions between the external aerodynamic flow field and the solid structure interface. The present thermo-mechanical coupling algorithm is verified on one-dimensional and two-dimensional transient heat flow problems with analytical solutions, and good consistency is achieved. The reliability of the GKUA for solving the boundary conditions of the external flow field and the dynamic thermo-mechanical coupling algorithm for the thermal response behavior of the internal structure material is verified by the consistency calculation of the internal and external temperature distribution. Then, the aerodynamic-thermoelastic analysis framework is further applied to study the deformation behavior of structural thermo-mechanical responses and failure mechanism of a hollow sphere under re-entry aerothermodynamic environment with emphasis placed on the influence of flight speeds, rarefied effects, algorithm schemes and damping effects. The numerical results obtained by the presented algorithm including the internal temperature, thermal stress and deformation distribution law are compared and analyzed. Finally, the deformation and failure mechanism of spacecraft structure during re-entry process are solved and analyzed by the present algorithm framework. The displacement, temperature and stress distribution contours of the structure are presented to make a preliminary understanding and evaluation on the process of spacecraft deformation and failure disintegration. It is indicated that the present coupled aerodynamic-thermoelastic analysis framework can provide a reliable, applicable and efficient platform for the thermo-mechanical coupling responses and failure mechanism of re-entry spacecraft in the end of life.

Measurements of fuel thickness for prefilming atomisers at elevated pressure

This work describes an experimental study of the fuel flows on the prefilmer of an aerospace gas turbine airblast atomiser at elevated pressure. The work identifies the physics leading to contradictory findings within the literature. This concerns an important atomisation boundary condition, whether the thickness of the fuel film on the prefilming surface influences the downstream drop size distribution. Analysis of the experimental data shows that fuel film thickness becomes uncorrelated with the downstream drop size if surface tension forces dominate inertia at the prefilmer tip. Fuel film thickness however provides the initial length scale for primary atomization if fuel inertia exceeds surface tension forces. It is the high inertial conditions that are associated with gas turbine operation, but the low inertial conditions that are readily achievable at laboratory scale through momentum flux scaling. Additionally, a detailed statistical description of the fuel flow has been provided for the atomiser tested. This reveals the importance of upstream hydrodynamic and aerodynamic boundary conditions on the probability of a ligament forming. Surprisingly, operating pressure is shown to have limited effect on the probability of ligament formation, a significant advantage for future modelling of the primary atomization processes.

Advanced Endwall Contouring for Loss Reduction and Outflow Homogenization for an Optimized Compressor Cascade

The following paper deals with the development of an optimized non-axisymmetric endwall contour for reducing the total pressure loss and for homogenizing the outflow of a highly-loaded compressor cascade. In contrast to former studies using a NACA-65 K48 cascade airfoil this study starts with the design of a new high-performance airfoil which is based on the aerodynamic boundary conditions of the NACA-65 K48 cascade. This new airfoil is then used as a basis. Optimizations of the airfoil and of the endwall contour are performed using the German Aerospace Center (DLR) in-house tool AutoOpti and the RANS (Reynolds-averaged Navier-Stokes)-solver TRACE (Turbomachinery Research Aerodynamic Computational Environment). Three operating points at an inflow Mach number of 0.67 with different inflow angles are used to secure a wide operating range. The optimized endwall contour changes the secondary flow in such a way that the corner stall is reduced which, in turn, significantly reduces the total pressure loss. The endwall contour in the outflow region leads to a considerable homogenization of the outflow in the near wall region. Using non-axisymmetric endwall shaping demonstrates a valuable measure to further improve highly-efficient compressor blading on the vane level.

Using Meridional Streamline and Passage Shapes to Generate Radial Turbomachinery Geometry and Meshes

An important aspect for structural and aerodynamics design of radial flow turbomachinery is the definition of the geometry and the generation of meshes for computational analysis. Particularly in the area of computational design and optimization, the way the geometry is defined is important, as it can limit design space. Traditionally, radial compressors and radial turbine rotors are defined using a mechanical design approach. Effectively a hub and shroud profile, followed by a rotorblade geometry are defined and the shape is adjusted in order to meet certain aerodynamic boundary conditions. The current paper presents an alternative approach, in which the overall geometry is defined starting from an aerodynamic requirement. The corresponding rotor and blade geometry is generated automatically, based on certain constraints. The advantage of this approach is the ability to define directly the aerodynamic requirements, which may allow a simpler efficient optimization of the aerodynamics.

Gas turbine working grid profiles design via inverse problem

Application of existing methods of solution of direct and inverse problems for the 3D-flow at optimal design of turbomachinery grids is associated with big problems, and the method of solution of the inverse problem even for the Q3D-flow, which takes into account the characteristics of the organization of the process of optimization search, opens up new possibilities for the design. The causes for the development of a new method were also positive results of experimental studies of the grid of the nozzle profiles with wavy shapes of trough and back, designed via the inverse problem for the 2D-current, which were conducted by the author. The direct and inverse problems for the Q3D-flow were divided into separate subtasks that could have been solved in any order or simultaneously. This provided the possibility to control any part of the designed blade in terms of strength and technological effectiveness, which could significantly reduce design time. The individual subtasks were solved by the methods of nonlinear programming that did not require storage of significant amounts of information. The parameters of the flow in the interblade channels for the fixed and rotating working grid of a gas turbine were determined, using the direct problem. Applying the inverse problem the interblade channels of grids were designed and the effect of rotation of the form of the surface of current and the aerodynamic boundary conditions on the coefficient of the profile losses of the grid and the geometric characteristics of profiles was shown.

A non-linear integrated aeroelasticity method for the prediction of turbine forced response with friction dampers

The aim of this paper is to describe an integrated aeroelasticity model for turbine blade forced response predictions. Such an approach requires a successful integration of the unsteady aerodynamics with non-linear structural dynamics, the latter arising from the use of root friction dampers to dissipate energy so that the response levels can be kept as low as possible. The inclusion of friction dampers is known to raise the resonant frequencies by up to 20% from the standard assembly frequencies, a shift that is not known prior to the aeroelasticity calculations because of its possible dependence on the unsteady excitation. An iterative procedure was therefore developed in order to determine the resonance shift under the effects of both unsteady dynamic loading and non-linear friction dampers. The iterative procedure uses a viscous, non-linear time-accurate flow representation for evaluating the aerodynamic forcing, a look-up table for determining the aerodynamic boundary conditions at any speed, and a time-domain friction damping module for resonance tracking. The methodology was applied to a high-pressure turbine rotor test case where the resonances of interest were due to first torsion and second flap blade modes under 40 engine-order excitation. The forced response computations were conducted using a multi-bladerow approach in order to avoid errors associated with “linking” single bladerow computations since the spacing between the bladerows was relatively small. Three friction damper elements, representing one actual friction damper, were used for each rotor blade. The number of rotor blades was decreased by 2–90 to obtain a cyclic sector of 4 stator and 9 stator blades. Such a route allowed the analysis to be conducted on a much smaller domain, hence reducing the computational effort significantly. However, the stator blade geometry was skewed in order to adjust the mass flow rate. Frequency shifts of 3.2% and 20.0% were predicted for the 40 engine-order resonances in torsion and bending modes, respectively. The predicted frequency shifts and the dynamic behaviour of the friction dampers were found to be within the measured range. Furthermore, the measured and predicted blade vibration amplitudes showed a good agreement with available experimental data, indicating that the methodology can be applied to typical industrial problems.

An Integrated Time-Domain Aeroelasticity Model for the Prediction of Fan Forced Response due to Inlet Distortion

The forced response of a low aspect-ratio transonic fan due to different inlet distortions was predicted using an integrated time-domain aeroelasticity model. A time-accurate, nonlinear viscous, unsteady flow representation was coupled to a linear modal model obtained from a standard finite element formulation. The predictions were checked against the results obtained from a previous experimental program known as “Augmented Damping of Low-aspect-ratio Fans” (ADLARF). Unsteady blade surface pressures, due to inlet distortions created by screens mounted in the intake inlet duct, were measured along a streamline at 85 percent blade span. Three resonant conditions, namely 1F/3EO, 1T & 2F/8EO and 2S/8EO, were considered. Both the amplitude and the phase of the unsteady pressure fluctuations were predicted with and without the blade flexibility. The actual blade displacements and the amount of aerodynamic damping were also computed for the former case. A whole-assembly mesh with about 2,000,000 points was used in some of the computations. Although there were some uncertainties about the aerodynamic boundary conditions, the overall agreement between the experimental and predicted results was found to be reasonably good. The inclusion of the blade motion was shown to have an effect on the unsteady pressure distribution, especially for the 2F/1T case. It was concluded that a full representation of the blade forced response phenomenon should include this feature.

Holocene coastal dune formation in southern Australia: A case study

The 14,500 km long southern Australian coast is exposed to high-energy wave and wind regimes which have mobilized shelf-derived sediments to produce extensive Holocene shoreface deposits. The 8000 km of sandy coast are backed by over 2000 km 2 of Holocene dunes of varying degrees of stability. In places dune deposition above present sea level accompanied the post-glacial marine transgression commencing as early as 10 ka B.P., and has continued episodically to the present day. The dunes are widespread in extent but vary in type, size and evolution. The oldest and often most extensive (primary) dune fields accompanied the post-glacial marine transgression. They occurred, and in some places continue to form, in areas of high wave and wind energy and abundant sand supply. Many of these systems exhausted their sediment supply by mid-Holocene and remain as stranded clifftop dunes where sand ramps were eroded, or as eroding barriers. Secondary phases of dune activity occurred on moderate-to high-energy coasts following an initial period of post-stillstand barrier progradation. Diminished sediment supply resulted in shoreline erosion and (secondary) dune formation. The location, size and evolution both of dunes types is largely controlled by the breaker wave regime which in turn influences the rate and volume of sediment supply, beach morphodynamics and resulting aerodynamic boundary conditions, and foredune erosion. A study of dune formation along the 3700 km long South Australian coast provides a basis for an overview of dune formation throughout southern Australia.