ASME Conference Presenter Attendance Policy and Archival Proceedings

2014;():V001T00A001. doi:10.1115/IMECE2014-NS1.

This online compilation of papers from the ASME 2014 International Mechanical Engineering Congress and Exposition (IMECE2014) represents the archival version of the Conference Proceedings. According to ASME’s conference presenter attendance policy, if a paper is not presented at the Conference, the paper will not be published in the official archival Proceedings, which are registered with the Library of Congress and are submitted for abstracting and indexing. The paper also will not be published in The ASME Digital Collection and may not be cited as a published paper.

Commentary by Dr. Valentin Fuster

Advances in Aerospace Technology: Advances in Aerodynamics

2014;():V001T01A001. doi:10.1115/IMECE2014-36051.

The centrifugal compressors are used in a wide variety of turbo machines ranging from low-pressure fans for cooling of electric motors to high-pressure ratio gas turbine compressors, from tiny cryogenic coolers to large industrial petrochemical compressor stations. An automated design scheme making use of Navier Stokes equation and Artificial Neural Network is envisaged as in figure 1. The present work describes a user friendly scheme to carry out blade geometry generation and to carry out preliminary designs that are two of the most important steps of the automated design scheme. Present paper describes a robust method to carry out these two steps. This paper also describes a method adopted to decide on viable design space. In the blade generation steps, present method was applied to that obtained by other methods reported in literature. It was found that the present scheme could replicate the centrifugal compressor geometry with just six parameters as compared to many more by others. In Mihai[1], use of adjoint method is used for optimization of NASA rotor. The design had 28 parameters. In Trigg[2], a systematic approach for optimization of a two dimensional blade design is reported. Two dimensional profile is defined by seventeen parameters. In Wang Hong liang[3] an optimization based on particle swarm principle is made use of to optimize the blade angle distribution at hub and shroud and had seven control points to do this. In order to minimize the number of control points, Asimara and Goto[4], used inverse method. Here blade loading at hub and shroud is the input to a inverse method to generate the blade geometry. Optimizer works on the blade loading. The generated geometry is analyzed by CFD calculations. From the above references, it is clear that the present scheme is a user friendly general method.

Commentary by Dr. Valentin Fuster
2014;():V001T01A002. doi:10.1115/IMECE2014-36936.

Synthetic jet actuators use oscillating motion near a fixed orifice to produce a net axial momentum flux with zero net mass flux. Through strategic application, these devices can provide flow control, propulsive thrust, and impingement cooling. To improve this performance, a new actuator has been designed with a variable orifice size, which can potentially increase exit flow speeds. The jet is generated using a pneumatic cylinder, which is oscillated linearly near an orifice. The opening consists of a camera aperture, whose diameter can decrease by a factor of 18 with the aid of a second pneumatic cylinder. The system is capable of operating at frequencies up to 5 Hz while maintaining full piston stroke, and the phase between the piston and orifice motion can be varied from 0 to 180 degrees. The flow structure is investigated using phase-locked particle image velocimetry (PIV), which shows that simultaneous constriction of the exit can substantially increase the exit speed. The initial design is used with air flow but will be extended to water applications in the future.

Topics: Actuators
Commentary by Dr. Valentin Fuster
2014;():V001T01A003. doi:10.1115/IMECE2014-37593.

This paper provides unambiguous velocity-field proof that a strong edge vortex occurs on a rotating blade in reverse flow. The rotor blades of a helicopter encounter reverse flow during flight at high advance ratio. Reverse flow is a limiter in rotorcraft design with excursions in pitch link loads and bending moments. Stereo particle-image velocimetry is used on the flow under the retreating blade of a two-bladed teetering rotor system in a low speed wind tunnel. The results are correlated with earlier aerodynamic loads and flow visualization data using the same rotor blade planform placed in a yawed, fixed-wing position. Results obtained with the blade held fixed at several rotor azimuths and angles of attack are used to ascertain the rotation effects on the flowfield by comparison with rotating blade results. Initial results suggest that radial velocity due to rotation hinders separation and delays the formation of an attached vortex, compared to the static case. The circulation of the reverse flow attached vortex is of the same order of magnitude as the bound circulation of the airfoil section, proving that the vortex contributes significantly to the lift force in addition to the pitching moment.

Commentary by Dr. Valentin Fuster
2014;():V001T01A004. doi:10.1115/IMECE2014-37699.

The frequency response of a trailing vortex downstream the free end of a NACA0015 wing section has been studied in a closed loop wind tunnel. The wing spanning approximately half height of the wind tunnel section was fixed to a dynamic load cell at the base. The flow cases studied have constant angle of attack of 10 degrees and constant Reynolds number of 160,000 based on the wing chord length while two levels of 0.5 and 4.6% free-stream turbulence were introduced to the flow. Lift and drag forces acting on the wing section were quantified using a load cell while an X-probe hot-wire was used to measure the turbulence parameters in the near-field wake region. The tip vortex was captured with sufficient accuracy. The effect of free-stream turbulence on the vortex structure has been discussed in detail by characterizing several turbulence parameters via time-averaging and frequency analysis of the turbulence signal.

Commentary by Dr. Valentin Fuster
2014;():V001T01A005. doi:10.1115/IMECE2014-38291.

A cyclorotor consists of a set of blades rotating about an horizontal axis that is parallel to the blade span. The designation of cycloidal rotor is related to the cycloidal path described by the rotating blades during forward flight. In the following paper we study, trough the use of numerical tools, the PECyT (Plasma Enhanced Cycloidal Thruster) system as a way of improving the performance of classical cycloidal rotors. PECyT consists in the introduction of Dielectric Barrier Discharge (DBD) plasma actuators in the CR blades. Such system act as an active flow control that is able to delay the stall onset at high angles of attack, thus increasing the aerodynamic efficiency of each blade.

Commentary by Dr. Valentin Fuster
2014;():V001T01A006. doi:10.1115/IMECE2014-38915.

The paper presents a study on a Coanda nozzle with applications in vectorized propulsion. The nozzle is able to change the exist flow angle as a function of a differential two-stream incoming flow rate. Herein we demonstrate that by using Dielectric Barrier Discharge actuators we are able to extend the range of attainable exit flow angles. First the analysis is performed using a numerical approach; afterwards an experimental facility is implemented to study this same effect. We include a comparison between the experimental testing on the Coanda thruster and CFD computations. Following an analysis of the results we demonstrate that it is possible to achieve a higher exit thrust angle, with the DBD plasma actuators active, and this is shown to be important in order to be able to keep the desired angles under several swirl velocities incoming from the feeding turbofans.

Commentary by Dr. Valentin Fuster
2014;():V001T01A007. doi:10.1115/IMECE2014-38966.

The Cycloidal Vertical Axis Wind Turbine is based on a self-adjustable pitch concept that automatically pitches the blade to improve the azimuthal load distribution and energy conversion. In nominal operating point, theoretical and CFD analysis have demonstrated that a cycloidal turbine can be more efficient than classical VAWTs, being able to produce more energy at low and intermediate Tip-Speed-Ratios. This innovative cycloidal wind turbine can also eliminate some problems of self-starting in VAWTs. In the following paper we propose the introduction of a new generation of wind energy converter system. We will show, trough the use of numerical tools, that the proposed system is more efficient than classical VAWTs.

Topics: Rotors , Wind energy
Commentary by Dr. Valentin Fuster
2014;():V001T01A008. doi:10.1115/IMECE2014-39146.

Potential analogs between dynamics induced by periodic passage through a bifurcation critical value and the nonlinear dynamics associated with the aerodynamic dynamic stall problem are presented for the first time. Koopman operator methods are used to study the spectral features of a streamwise oscillating cylinder which exhibits wake dynamics due to externally forced oscillations through a Hopf bifurcation critical value. Koopman decomposition results show that the system transitions to a more continuous spectrum compared to the discrete spectrum associated with a stationary cylinder in post-critical flow. Finally, Fourier analysis of flow variables associated with an oscillating airfoil under dynamic stall conditions were compared with the oscillating cylinder spectra. The spectral characteristics of the two systems exhibited similar frequency broadening behavior induced by the externally forced oscillations. Therefore, the results indicate that the nonlinear dynamics associated with dynamic stall appear to have strong linkages to a system oscillating through a bifurcation critical value.

Commentary by Dr. Valentin Fuster
2014;():V001T01A009. doi:10.1115/IMECE2014-39354.

Using computational methods, an investigation was performed on the physical mechanisms leading to vortex breakdown in high angle of attack flows over delta wing geometries. For this purpose, the Second International Vortex Flow Experiment (VFE-2) 65° sweep delta wing model was studied at a root chord Reynolds number (Recr) of 6 × 106 at various angles of attack. The open-source computational fluid dynamics (CFD) solver OpenFOAM was used in parallel with the commercial CFD solver ANSYS® FLUENT. For breadth, a variety of classic closure models were applied, including unsteady Reynolds-averaged Navier-Stokes (URANS) and detached eddy simulation (DES). Results for all cases are analyzed and flow features are identified and discussed. The results show the inception of a pair of leading edge vortices originating at the apex for all models used, and a region of steady vortical structures downstream in the URANS results. However, DES results show regions of massively separated helical flow which manifests after vortex breakdown. Analysis of turbulence quantities in the breakdown region gives further insight into the mechanisms leading to such phenomena.

Commentary by Dr. Valentin Fuster
2014;():V001T01A010. doi:10.1115/IMECE2014-39759.

In this study an academic Computational Fluid Dynamics (CFD) code, named Galatea-I, is described, which employs the Reynolds Averaged Navier–Stokes (RANS) equations along with the artificial compressibility method and the SST (Shear Stress Transport) turbulence model for the prediction of incompressible viscous flows. For the representation of the computational domain unstructured hybrid grids are utilized, composed of tetrahedral, prismatic and pyramidical elements, while for its discretization a node-centered finite-volume scheme is implemented. Galatea-I is enhanced with a parallelization method, which employs spatial domain decomposition, while the data exchange between processors/processes is performed with the use of the Message Passing Interface (MPI) protocol. In addition, a parallel agglomeration multigrid methodology has been incorporated to improve further its computational performance. The proposed code is validated against steady-state flow benchmark test cases, concerning laminar flow over a cubic cavity and a cylindrical surface, as well as turbulent flow over a rectangular wing with a NACA0012 airfoil. The obtained results, compared with these of corresponding reference solvers, reveal Galatea-I’s potential for simulation of inviscid, viscous laminar and turbulent incompressible flows.

Commentary by Dr. Valentin Fuster

Advances in Aerospace Technology: Aeromechanics and Aeroelasticity

2014;():V001T01A011. doi:10.1115/IMECE2014-36967.

Aeroelastic instabilities such as flutter, limit cycle oscillation (LCO), and divergence are traditionally considered undesirable. Designers try to avoid these instabilities by adding enough stiffness or damping to structures. A new approach to suppressing these instabilities is to use smart material to harvest energy from airflow. In this way not only are the aeroelastic instabilities avoided, but also some energy will be harvested. The harvested energy can be used for powering sensors, morphing parts of the structure, and ultimately increasing the performance of the aircraft. Energy harvesting from aeroelastic phenomena can also be used in designing small wind energy harvesters for home use. In this paper we will explore both capabilities. Piezoelectric materials are among the attractive smart materials for energy harvesting. Piezoelectric materials generate electric potential as they deform. We will explore the use of these materials in aeroelastic harvesting. Ref. 1 has a general overview of different forms of vibrational energy harvesting, including the use of piezoelectric materials. Harvesting energy from aeroelastic instabilities is a relatively new area; therefore, the body of literature on this subject is relatively young. Most of the analysis is limited to a 2-D cross-sectional analysis with steady or quasi-steady flow. We will use a 2-D model with an unsteady aerodynamic model as the preliminary result. More realistic cases with a beam model will be added to the final version of the paper. For the beam model, we will use fully intrinsic equations.

Commentary by Dr. Valentin Fuster
2014;():V001T01A012. doi:10.1115/IMECE2014-37473.

A multidisciplinary design optimization (MDO) method for turbine mortise assembly structure is presented. This method takes tenon and mortise as a whole to carry on the design optimization. With the adoption of Design of Experiment (DOE) method and Collaborative Optimization (CO) Strategy, this method has realized the coupling parallel optimization of the mortise assembly structure. To verify the effectiveness of the proposed method a typical fir-tree mortise structure design optimization is provided as an example. The results show that the mass and Von Mises Stress of the mortise assembly structure were reduced remarkably, which means that the proposed method has a significant role in enhancing structural performance.

Commentary by Dr. Valentin Fuster
2014;():V001T01A013. doi:10.1115/IMECE2014-37484.

For rotating critical parts of aero engines, such as turbine disks, it is essential to perform reliable life predictions. Probabilistic methods are ideal to investigate these life predictions. Beside other system parameters, the distribution of geometrical parameters has a strong effect on the system behavior. However, as there are always so many geometry parameters, it always takes lots of time to complete such a probabilistic analysis considering geometry.

Within this paper, a probabilistic method based on two surrogate models is proposed and applied to an analysis of a turbine disk. In order to save the computation time as well as to get accurate results, this process is divided into two cycles. The purpose of the first cycle is to filter the parameters which have little influence on the life of the disk. In this cycle, DOE method is used and a normal response surface is created as a surrogate model to calculate the sensitivities of all the input parameters. With the sensitivities some key parameters can be selected as the inputs for the second cycle. In the second cycle DACE method is used and a more accurate Kriging model is created as the surrogate model. By conducting MCS on the calculated Kriging model, the reliability of the turbine disk can be get. In this way a huge number of computations can be avoided, thus much time can be saved and the computational efficiency can be improved.

Topics: Turbines , Disks , Geometry
Commentary by Dr. Valentin Fuster

Advances in Aerospace Technology: Aerospace Structures and Materials

2014;():V001T01A014. doi:10.1115/IMECE2014-36316.

In this study, the behavior of repaired cracks, located in aluminum alloy sheets 2024T3, with bonded composite patch is analyzed experimentally and numerically. The experimental study has been conducted through fatigue tests on aluminum cracked plate repaired with Carbon/epoxy patch. In the numerical analysis, the stress intensity factor at the crack front is computed using three-dimensional finite element method. The obtained results show that the stress intensity factor at the crack front is highly reduced by the presence of the patch repair. Therefore, the fatigue life of the damaged structure can be significantly improved especially if the patch repair is applied at small crack lengths.

Commentary by Dr. Valentin Fuster
2014;():V001T01A015. doi:10.1115/IMECE2014-36397.

In this paper, a novel numerical-experimental methodology is outlined to predict delamination in pristine as well as isothermally aged (in air) polymer matrix composites. A rate-dependent viscoelastic cohesive layer model was implemented in an in-house test-bed finite element analysis (FEA) code to simulate the delamination initiation and propagation in unidirectional polymer composites before and after aging. This unified model is fully rate-dependent and does not require a pre-assigned traction-separation law. The actual shape of traction separation law depends on: (a) the strain rate via the viscoelastic constitutive relationship, (b) the degree of thermo-oxidative aging via the changes in the experimentally measured creep compliance due to oxidation, and (c) the evolution of the internal state variable defining the state of damage. To determine the model parameters, double cantilever beam (DCB) experiments were conducted on both pristine and isothermally aged IM-7/bismaleimide (BMI) composite specimens. The J-Integral approach was adapted to extract cohesive stresses near the crack tip. A principal-stretch dependent internal damage state variable defines the damage in the cohesive layer. Within the cohesive layer, pristine and cohesive stresses were compared to estimate the damage parameters. Once the damage parameters had been characterized, the test-bed FEA code employed a micromechanics based viscoelastic cohesive layer model to simulate interlaminar delamination. From a numerical stability standpoint, the viscous regularization effect of the viscoelastic constitutive equations in the cohesive layer helps mitigate numerical instabilities caused by elastic energy released due to crack growth, thereby enabling the FEA model to simulate the load-deflection response of the composite structure well beyond peak load. The present cohesive-layer based FEA model was able to accurately predict not only the macro level load-displacement curve, but also the micro level crack growth history in IM-7/BMI laminate before and after thermal aging, using only three parameters.

Topics: Delamination
Commentary by Dr. Valentin Fuster
2014;():V001T01A016. doi:10.1115/IMECE2014-36452.

Landing gear is the one of the key components for improving aircraft crashworthiness because its primary function is the energy absorption. But, in general, the shock absorbers are designed to have best efficiency for normal landing cases and can be ineffective when faced with very high sink speed. Thus special design and implementation are necessary for landing gear to have crashworthiness. For this purpose, various concepts have been studied and put to practical use such as structural pin, pressure relief valve and additional energy absorbing devices, etc. In this paper, the composite tube is investigated as an extra energy absorber and adopted to landing gear to increase shock absorbing performance in case of crash. To do this, first the quasi-static and impact test of composite tubes are conducted and the analysis model is tuned to explain the test results. During the correlation process, the failure modes and the specific energy absorption of the composite tubes are analyzed and the optimal configurations are searched.

The overall performance of landing gear including the composite tube is analyzed by developing a simplified dynamic model. Each force-stroke relation of oleo-pneumatic shock absorber, tire and composite tube are modeled as spring and damper, then the equation of motion is solved to obtain the crash responses. In this model, after the bottoming of shock absorber, the crushing of composite tube is activated for additional energy absorption. Numerical solutions show that the enhanced shock absorbing capability in case of crash when the composite tube adopted. For practical use, the landing gear performance should be verified by drop tests and this is author’s future research project.

Commentary by Dr. Valentin Fuster
2014;():V001T01A017. doi:10.1115/IMECE2014-36583.

In recent times, composite materials have gained mainstream acceptance as a structural material of choice due to their tailorability and improved thermal, specific strength/stiffness and durability performance [1–3]. For high temperature applications, which include exit nozzle for rockets, leading edge for missiles, nose cones, brake pads etc. Carbon-Carbon composites (C/C composite) are found suitable [4–6]. Mechanical property estimation of C/C composites is challenging due to their highly heterogeneous microstructure. The highly heterogeneous microstructure consists of woven C-fibers, C-matrix, irregularly shaped voids, cracks and other inclusions. Predicting the mechanical behavior of complex hierarchical materials like C/C composites is of interest which forms the motivation for the present work. A systematic study to predict the effective mechanical properties of C/C composite using numerical homogenization has been undertaken in this work. The Micro-Meso-Macro (MMM) principle of ensemble averages for estimating the effective properties of the composite has been adopted. The hierarchical length scales in C/C composites were identified as micro (single fiber with matrix), meso (fabric) and macro (laminate). Comparisons have been made with mechanical testing of C/C composites at different length scales.

Commentary by Dr. Valentin Fuster
2014;():V001T01A018. doi:10.1115/IMECE2014-36767.

The surface machining of Carbon Fiber Reinforced Plastics (CFRP) materials is a challenging process, given the heterogeneity and anisotropic nature of these composites, which, combined with the abrasiveness of the fibers involved, can produce some surface damage and extensive tool wear. The cutting temperature is one of the most important factors associated with the tool wear rate and machinability of these materials, which are also affected by the mechanical and thermal properties of the work material and the cutting conditions. In this work, the cutting temperature, forces and surface roughness were measured under different cutting conditions during the ball-end milling of unidirectional CFRP. Cutting speeds ranging from 200 to 350 m/min, a feed rate of 0.063 mm/rev, fiber orientation of (the angle between carbon fibers and feed direction) 0, 45, 90 and 135 degrees, and a 0.5 mm depth of cut were used. The results show that the cutting speed and fiber orientation have a significant influence on the cutting temperature and cutting force. The maximum and minimum cutting forces and temperature were achieved for fiber orientations of 90 and 0 degrees, respectively.

Commentary by Dr. Valentin Fuster
2014;():V001T01A019. doi:10.1115/IMECE2014-36827.

As part of an ongoing research development at Carleton University in ceramic matrix composites (CMCs) for high-temperature gas turbine applications, it was recognized that the performance of an oxide matrix could be improved by incorporating a metal reinforcement material. For this reason, a low cost CMC was created by reinforcing a yttria-stabilized zirconia (7YSZ) ceramic matrix with a Hastelloy X (HX) wire mesh. The CMC was manufactured by coating the HX mesh with a NiCrAlY bond coat, and then 7YSZ ceramic matrix, both using plasma spraying. The bond coat was employed to improve bonding and also to act as an oxygen diffusion barrier. In order to evaluate the performance of the HX/7YSZ composite at high temperatures, isothermal and cyclic oxidation tests were carried out for 1000 hours at 1050°C. The results showed that oxidation resistance was improved by vacuum heat treatment prior to testing due to the formation of stable thermally grown oxides (TGO) on the NiCrAlY bond coat. In the cyclic oxidation test, differences in thermal expansion coefficients caused cracking at interfaces between mesh/bond coat and bond coat/7YSZ. Minimizing the effect of thermal expansion by better material combination, as well as modifying manufacturing methods will allow for improved performance of metal mesh reinforced CMCs.

Commentary by Dr. Valentin Fuster
2014;():V001T01A020. doi:10.1115/IMECE2014-36882.

Today’s rocket engines regeneratively cooled using high energy cryogenic propellants (e.g. LOX and LH2, LOX and LCH4) play a major role due to the high combustion enthalpy (10–13.4 kJ/kg) and the high specific impulse of these propellants. In the frame of the HYPROB/Bread project, whose main goal is to design build and test a 30 kN regeneratively cooled thrust chamber, a breadboard has been conceived in order to:

• investigate the behavior of the injector that will be employed in the full scale final demonstrator,

• to obtain a first estimate of the heat flux on the combustion chamber for models validation,

• to implement a “battleship” chamber for a first verification of the stability of the combustion

The breadboard is called HS (Heat Sink) and it is made of CuCrZr (Copper Chromium Zirconium alloy), Inconel 718 and TZM (Titanium Zirconium Molybdenum alloy). The aim of the present paper is to illustrate the thermostructural design conducted on the breadboard by means of a Finite Element Method code taking into account the viscoplastic behavior of the adopted materials. An optimization process has been carried out in order to keep the structural integrity of the breadboard maximizing the life cycles of the component. Heat fluxes generated by combustion gases have been evaluated by means of CFD quick analyses, while convection and radiation with the external environment have not been considered in order to be as conservative as possible from a thermostructural point of view. Transient thermal analyses and static structural analyses have been performed by means of ANSYS code adopting an axisymmetric model of the chamber. These analyses have demonstrated that the Breadboard can withstand the design goal of 3 thermo-mechanical cycles with a safety factor equal to 4 considering a firing time equal to 3 seconds.

Topics: Design , Heat sinks
Commentary by Dr. Valentin Fuster
2014;():V001T01A021. doi:10.1115/IMECE2014-36966.

Modern helicopter blades are designed as thin-walled hollow structures in form of either C-spar or D-spar cross-sections. With the advent of new materials hollow designs have been implemented to reduce the overall weight of the structure. A D-spar is a rotor blade cross-section that is hollow in nature with a single vertical spar used to carry a large portion of the stresses otherwise carried by the skin [1]. The vertical spar is normally located between the leading edge and half of the chord length. The remaining volume aft of the vertical spar can either be hollow or filled with a honeycomb structure. The honeycomb structure increases the cross-sectional stiffness. Figure 1. shows an example of a common D-spar with a honeycomb structure aft of the vertical spar [2]. Due to new manufacturing methods the D-spar has now become common place in helicopter design [3]. A C-spar cross-section is very similar to the D-spar cross-section in design and construction. The C-spar cross-section does not have the honeycomb structure and the spar. The structural load is offset by more lamina layers towards the leading edge of the cross-section [4,5]. The thin-walled structure is comprised of many layers of composite materials such as fiberglass or carbon fibers. There has been extensive research into D-spar cross-section while there is a lack of studies for C-spar cross-sections [1,3,4].

Commentary by Dr. Valentin Fuster
2014;():V001T01A022. doi:10.1115/IMECE2014-37377.

This paper will present a numerical method for solving fully coupled dynamic problems of the mechanical behavior of electrically conductive composite plates in the presence of an electromagnetic field. The mechanical behavior of electrically conductive materials in the presence of an electromagnetic field is described by the system of nonlinear partial differential equations (PDEs), including equations of motion and Maxwell’s equations that are coupled through the Lorentz ponderomotive force. In the case of thin plates, the system of governing equations is reduced to the two-dimensional (2D) time-dependent nonlinear mixed system of hyperbolic and parabolic PDEs. This paper discusses a numerical solution method for this system, which consists of a sequential application of the Newmark finite difference time integration scheme, spatial (with respect to one coordinate) integration scheme, method of lines (MOL), quasilinearization, and a finite difference spatial integration of the obtained two-point boundary-value problem. The final solution is obtained by the application of the superposition method followed by orthonormalization.

Commentary by Dr. Valentin Fuster
2014;():V001T01A023. doi:10.1115/IMECE2014-37961.

One-dimensional models are widely used in mechanical design. Classical models, Euler-Bernoulli or Timoshenko, ensure a low computational cost but are limited by their assumptions, many refined models were proposed to overcome these limitations and extend one-dimensional models at the analysis of complex geometries or advanced materials. In this work a new approach is proposed to couple different kinematic models. A new finite element is introduced in order to connect one-dimensional elements with different displacement fields. The model is derived in the frameworks of the Carrera Unified Formulation (CUF), therefore the formulation can be written in terms of fundamental nuclei. The results show that the use variable kinematic models allows the computational costs to be reduced without reduce the accuracy, moreover, refined-one dimensional models can be used in the analysis of complex structures.

Topics: Kinematics
Commentary by Dr. Valentin Fuster
2014;():V001T01A024. doi:10.1115/IMECE2014-38479.

This study reports dynamic aeroelastic analyses of an aircraft wing with an attached mass subjected a lateral follower force in an incompressible flow. A swept thin-walled composite beam with a biconvex cross-section is used as the structural model that incorporates a number of non-classical effects such as material anisotropy, transverse shear deformation and warping restraint. A symmetric lay-up configuration i.e. circumferentially asymmetric stiffness (CAS) is further adapted to this model to generate the coupled motion of flapwise bending-torsion-transverse shear. For this beam model, the unsteady aerodynamic loads are expressed using Wagners function in the time-domain as well as using Theodorsen function in the frequency-domain. The flutter speeds are evaluated for several ply angles and the effects of follower force, transverse shear, fiber-orientation and sweep angle on the aeroelastic instabilities are further discussed.

Commentary by Dr. Valentin Fuster
2014;():V001T01A025. doi:10.1115/IMECE2014-38992.

Due to high working temperature and rotating speed, turbine disks are crucial parts in gas turbine engines. The weight of disks is always heavy in order to increase the reliability and structural integrity. So optimization design of disks could bring a significant reduction in engine weight. Focusing on a typical Low Pressure Turbine (LPT) disk, this paper improves its design in both static and dynamic characteristics with ANSYS Workbench platform. Based on a 2D parameterized model, the sensitivity of different structural parameters was investigated quickly. Then the optimization process to minimize the mass was conducted by NLPQL (Non-Linear Programming by Quadratic Lagrangian) method with 3D parameterized model. The equivalent stress of disk was limited in static optimization and resonance frequency was also restricted to a safe level through a Campbell diagram in dynamic optimization. A new design plan was acquired through optimization process, which reduces 13.6% of total weight under static and dynamic criteria.

Commentary by Dr. Valentin Fuster
2014;():V001T01A026. doi:10.1115/IMECE2014-39039.

A single platform D&A (Design & Analysis) tool is outlined in this paper that allows a user, through a single GUI (Graphical User Interface), to create a turbine rotor fixing as well as analyze the structural integrity of the fixing. This is done through the integration of CAD (Computer Aided Design) and FEA (Finite Element Analysis) software running in batch mode, driven by the GUI. This SPIE (Single Platform Integration Environment) captures the strength of CAD software to create a fully parameterized fixing that is able to model legacy, current designs and provides flexibility to design fixings not yet conceived. Using the automated use of FEA software through a secure and reliable gateway, stress analysis can be performed and the results displayed back to the user through the GUI. This tool provides a significant increase in quality and time savings to design a fixing when compared to the previous design methodologies. What used to take hours to design and analyze through the use of isolated specialist built and owned tools with little communication between them and non-ideal data management, now takes minutes; a reduction of up to 10 fold in the time taken.

Topics: Design , Turbines
Commentary by Dr. Valentin Fuster
2014;():V001T01A027. doi:10.1115/IMECE2014-39092.

Constant and variable stiffness strategies have been developed to design a composite laminate. With the former, each layer is designed with straight fibers that have the highest stiffness and strength in the fiber direction. With the latter, on the other hand, the stiffness can change within each layer by placing the fibers along a curvilinear fiber path. A variable stiffness design results in improved structural performance, as well as opens up opportunities to search for trade-off among structural properties. During the manufacture of a variable stiffness design with Automated Fiber Placement, certain defects in the form of gaps and overlaps could appear within the laminate and affect the laminate performance. In this study, we use the first-order shear deformation theory to assess the effect of transverse shear stresses on the critical buckling load, free and forced vibration of a variable stiffness laminate with embedded defects, an issue so far rarely examined in literature. The governing differential equations for the static analysis are first derived. A semi-analytic solution is then obtained using the hybrid Fourier-Galerkin method and the numeric time integration technique. The eigenvalue analysis is also conducted to determine the fundamental frequency and critical buckling load of the plate. It is found that the behavior of a variable stiffness plate is much more affected by the shear stresses than a constant stiffness plate. Ignoring the effect of transverse shear stresses results in 34% error in the predicted buckling load of a variable stiffness laminate with overlaps and a length-to-thickness ratio of 10.

Commentary by Dr. Valentin Fuster
2014;():V001T01A028. doi:10.1115/IMECE2014-39259.

This paper is devoted to the dynamic modeling of micropolar gyroelastic beams and explores some of the modeling and analysis issues related to them. The simplified micropolar beam torsion and bending theories are used to derive the governing dynamic equations of micropolar gyroelastic beams from Hamilton’s principle. Then these equations are solved numerically by utilizing the finite element method and are used to study the spectral and modal behaviour of micropolar gyroelastic beams.

Commentary by Dr. Valentin Fuster
2014;():V001T01A029. doi:10.1115/IMECE2014-39345.

Thermal Barrier Coating (TBC) system has been increasingly applied to gas turbine engine because of its outstanding ability to effectively protect substrate materials against impinging hot gas. However, the durability issue has still existed due to the growth of oxidation layer called thermal growth oxidation (TGO) and depletion of aluminum. Therefore, more detailed investigation is required to understand the effects of these two factors on the durability of TBC. In this study, a finite element (FE) model was developed to calculate the stress state near TGO interfaces by considering both the morphology of TGO and aluminum depletion effect. For acquiring the mechanical properties at the depletion area of bond coating layer, indentation tests were carried out, particularly, to the TBC specimens isothermally exposed at 1000 degree Celsius for 100h. Based on the proposed FE model with specific parameters obtained from experiments, parametric studies were performed in a variety of conditions of mechanical properties and TGO thickness. Simulation results clearly showed the influence of aluminum depletion and TGO growth on the durability of TBC system. Finally, optimal design criteria can be suggested to minimize the stress of TBC system.

Commentary by Dr. Valentin Fuster
2014;():V001T01A030. doi:10.1115/IMECE2014-39817.

Because of the high strength-to-weight and stiffness-to-weight ratios, the fiber-reinforced epoxy composites are the most common materials for many industrial applications. Nevertheless, these engineered materials can lose their properties when subjected to the high moisture, UV light and temperatures through oxidation and other decomposition processes. In order to minimize the weakness of the epoxy composites, graphene thin films were fabricated after the acid treatment process, and then applied onto the surface of fiber reinforced epoxy composites to act as heat shields. The 45 degree burn tests and surface paint adhesion tests were conducted on the prepared samples in accordance with the guidelines of the FAA Regulations, ASTM, SAE, and AMS specifications. The test results revealed significant improvements on the flame retardancy of the composites by incorporating graphene oxide thin films. Overall, this study may improve the fire resistance properties of the composites for different high temperature applications of composites.

Commentary by Dr. Valentin Fuster
2014;():V001T01A031. doi:10.1115/IMECE2014-39818.

This report presents the development of graphene-based nanocomposite coatings on the fiber reinforced composites to improve the coating resistance against the corrosion and other environmental weathering. Graphene nanoflakes were initially functionalized through a silanization process, and then dispersed well into the polyurethane primer and top coats at 0, 2, 4 and 8wt% using high speed agitation and sonication processes. The dispersed nanocomposite coatings were an air sprayed on the surfaces of the composite coupons at different thicknesses, and cured prior to the alternative UV and salt fog exposure tests for 20 days. The performance analyses of the nanocomposite coatings were carried out using atomic force microscopy (AFM), Fourier transform infrared spectrometer (FTIR), thickness measurements, water contact angle, and electro-chemical impedance spectroscopy tests. The test results indicated that the silanization process on the graphene nanoflakes significantly improved the corrosion resistances of the nanocomposite coatings when compared to the non-functionalized graphenes. This study may be useful for the performance improvements of many coatings on the composite aircraft, wind turbines and other applications.

Commentary by Dr. Valentin Fuster
2014;():V001T01A032. doi:10.1115/IMECE2014-39956.

It is widely understood that moisture can have a detrimental effect on the strength of composite structures. Traditional analysis often focuses on the effects to solid laminates or on the facesheets of composite sandwich structures. However, this focus is often not sufficient to ensure material strength and performance. It has been found that moisture effects on sandwich structures can also have a detrimental effect on secondary failure modes such as shear crimping and facesheet wrinkling, and that these effects can be significant, especially at temperature. A proper assessment of moisture effects on composite sandwich structures involves five key components: development of moisture diffusion constants, prediction of structural moisture levels, development of material allowables at predicted moisture levels, analysis of structure, and modification of the design, when warranted. This paper describes each component of this process, and introduces a simple algorithm to integrate the analysis.

Commentary by Dr. Valentin Fuster
2014;():V001T01A033. doi:10.1115/IMECE2014-39986.

Strength analysis of aircraft structures focuses on static strength, fatigue, and damage tolerance of the materials. Countless hours are invested in quantifying the static strength, which is allowed to fully yield under ultimate loads so that the structure can be designed as light and efficient as possible. This analysis is critical to a lightweight initial design, but is also of great importance in the evaluation of engineering modifications and repairs. Analysis of the critical sections often resorts to plastic bending analysis, and use of E.F. Bruhn’s Iterative Slice Method, or Cozzone’s Simplified Method for Symmetric Sections, is often employed. Yet these methods fall short when the critical section includes thin flanges that buckle or cripple prior to ultimate failure, as is generally the case for frames, floor beams, stringers, and other structural members used on aircraft. This paper presents a solution to this shortfall, and introduces a hybrid procedure for calculating the ultimate strength of a cross section that accounts for material non-linearity, flange stability, and other effects.

Commentary by Dr. Valentin Fuster
2014;():V001T01A034. doi:10.1115/IMECE2014-39999.

The purpose of this research was to determine the influence of material properties on the impact response of a laminate, whereby specimens were fabricated and cured under a vacuum and high temperature using three types of pre-impregnated (prepreg), carbon fibers, namely unidirectional fiber, plain weave woven fiber, and non-crimp fiber (NCF). Each carbon fiber panel, usually known for its low-impact properties, of 16 plies underwent impact testing using a low-velocity impactor and visual damage inspection by C-scan in order to measure the damage area and depth, before and after impact testing. These panels were treated with UV exposure and moisture conditioning for 20 days each. Water contact angles were taken into consideration to determine the hydrophobicity and hydrophillicity of the respective prepreg materials. Experimental results and damage analysis showed that UV exposure and moisture conditioning showcased the variation in impact response and behavior, such as load-carrying capacity, absorbed energy, and impact energy of the carbon fiber panels. This study illustrates that non-crimp carbon fiber laminates were far more superior relative to load capacity than woven and unidirectional laminates, with the NCF-AS laminate exhibiting the highest load capacity of 17,244 lb/in (pre-UV) with only 0.89% decrease after UV exposure. This same laminate also had a 1.54% decrease in sustaining impact and 31.4% increase in wettability of the panel. Moreover, the study shows how symmetric and asymmetric stacking sequences affect the impact behavior of non-crimp fiber laminates. These results may be useful for expanding the capacity of carbon fiber, lowering costs, and growing new markets, thus turning carbon fiber into a viable commercial product.

Commentary by Dr. Valentin Fuster

Advances in Aerospace Technology: Aircraft Modeling and Simulation

2014;():V001T01A035. doi:10.1115/IMECE2014-36676.

The Blended Wing Body (BWB) aircraft is based on the flying wing concept. For this aircraft the literature has reported performance improvements compared to conventional aircraft. However, most BWB studies have focused on large aircraft and it is not sure whether the gains are the same for smaller aircraft. The main objective of this work is to perform the conceptual design of a 200 passengers BWB and compare its performance against an equivalent conventional A320 aircraft in terms of payload and range. Moreover, an emphasis will be placed on obtaining a stable aircraft, with the analysis of static and dynamic stability. The design of BWB was carried out under the platform called Computerized Environment for Aircraft Synthesis and Integrated Optimization Methods (CEASIOM). This design platform, suitable for conventional aircraft design, has been modified and additional tools have been integrated in order to achieve the aerodynamic analysis, performance and stability of the BWB aircraft.

Commentary by Dr. Valentin Fuster
2014;():V001T01A036. doi:10.1115/IMECE2014-37236.

The solution to the problem on building mathematical models of technical objects through the approximation of various experimental dependences is offered in the paper. This approach is especially true for modeling aircraft because the aerodynamic coefficients of their models can be obtained either by full-scale study or by computer simulation only. Currently, the experimental simulation is performed either through the regression analysis (RGA) methods, or through spline approximation. However, the RGA has a significant disadvantage, namely a poor approximability of piecewise and multiextremal dependencies. The RGA gives a rough approximation of the experimental data for similar curves. Spline approximation is free from this disadvantage. However, a high degree of discretization, a strict binding to the number of spline points, and a large number of equations, make this approach inconvenient for application when a compact model building and an analytic transformation are required.

A problem solution combining the advantages of both approaches and clearing up the troubles is offered in the paper. The proposed approach is based on the regression construction of the mathematical models of the dependence fragments, the multiplicative excision of these fragments in the local functional form, and on the additive combining of these local functions into a single analytic expression. The effect is achieved by using special “selection” functions multiplicatively limiting a nonzero definition domain for each of the approximating functions. The method is named “cut-glue” by the physical analogy of the approximation techniques. The order and structure of the approximating function for each segment can be arbitrary. A significant advantage of the “cut-glue” approximation is in a single analytic expression of the whole piecewise function instead of a cumbersome system of equations. The analytical and numerical studies of the properties and operational experience of the proposed method are resulted.

Commentary by Dr. Valentin Fuster
2014;():V001T01A037. doi:10.1115/IMECE2014-37619.

The paper describes the application of a morphing wing technology on the wing of an Unmanned Aerial System (UAS). The morphing wing concept works by replacing a part of the rigid wing upper and lower surfaces with a flexible skin whose shape can be dynamically changed using an actuation system placed inside the wing structure. The aerodynamic coefficients are determined using the fast and robust XFOIL panel/boundary-layer codes, as the optimal displacements are calculated using an original, in-house optimisation tool, based on a coupling between the relatively new Artificial Bee Colony Algorithm, and the classical, gradient-based Broyden-Fletcher-Goldfarb-Shanno (BFGS) method. All the results obtained by the in-house optimisation tool have been validated using robust, commercially available optimization codes. Three different optimization scenarios were performed and promising results have been obtained for each. The numerical results have shown substantial aerodynamic performance increases obtained for different flight conditions, using the proposed morphing wing concept.

Topics: Wings
Commentary by Dr. Valentin Fuster
2014;():V001T01A038. doi:10.1115/IMECE2014-38174.

Airships have the intrinsic advantages of Lighter-Than-Air (LTA) vehicles: minimal energy consumption and Vertical Take-Off and Landing (VTOL) characteristics. Due to these advantages, significant efforts are being taken in order to investigate new applications and technical improvements. More specifically, there is a renewed interest in large airships for heavy payload transportation and for stratospheric airships. The design of large airships is a big challenge, especially when considering the structural point of view, since big volumes imply high loads, and since light weight is a major requirement for this type of vehicles. In this context, a light-weight structure is proposed by applying the structural Tensairity concept. A Tensairity beam consists of a rigid air beam designed on the basis of complete functional separation of the different structural elements, allowing for a maximum optimization. In this paper, the justification of the feasibility of applying Tensairity components in airships is discussed based on two criteria. The first criterion is the justification of the need of a lightweight structure by a state of the art analysis and a study of the principal characteristics of the existing types of LTA vehicles structures. The second criterion is a preliminary technical analysis, which aims to clarify if the load bearing behavior of airships is suited for the application of the Tensairity concept. Moreover, the bases for the development of the concept for the LTA vehicles structures are established.

The advantages and drawbacks of the traditional rigid airships structure in comparison with a non-rigid structure has been analyzed, which conclusion is that the use of a rigid structure is convenient for large airships, since it reduces significantly the stresses of the envelope, but at the same time decreases the payload efficiency due to the addition of the structure’s weight. Moreover, the analysis of the load bearing behavior suggests the technical feasibility of applying Tensairity components, since airships have to withstand high bending moments and Tensairity structures are appropriate for withstanding such loads. Finally, the principal guidelines for defining the various load cases and for modeling Tensairity beams have been defined. In order to confirm the hypothesis of the suitability of Tensairity structures on airships, extensive research on design, analysis and optimization of Tensairity beam grids in typical airship loading conditions is needed.

Commentary by Dr. Valentin Fuster
2014;():V001T01A039. doi:10.1115/IMECE2014-38420.

This paper addresses the aerodynamic response of damaged delta wings using steady-state Computational Fluid Dynamics simulations. Two types of delta wings are investigated: a High Speed Civil Transport (HSCT) wing and a F16 Block 40 Wing. These types of analyses are required to help predict wings’ remaining flight capability, after damage is inflicted (during battle). The damage is represented by a hole in the CFD model of both wings. Variations in the shape, size, location and orientation of holes are investigated. The lift and drag (at relatively low angles of attack) of the undamaged and damaged wings are predicted and compared. The obtained numerical results indicate that the location of the hole has a significant effect on the performance of the wing. Furthermore, straight-edged holes seem to have a larger impact on the wing’s aerodynamics as opposed to cylindrical-shaped holes. To make the shape of the hole as realistic as possible, petals emerging above the surface of the wing are introduced and their effect is also investigated. Results show a greater increase in drag compared to smooth cylindrical holes. Finally, and to better simulate the jet in cross flow mainly the strong-jet phenomenon, preliminary time-accurate high angles of attack simulation results will be presented.

Commentary by Dr. Valentin Fuster
2014;():V001T01A040. doi:10.1115/IMECE2014-38865.

The determination of aerodynamic coefficients such as pressure distributions and aerodynamic coefficients (lift, drag and moment) from the known parameters (angle of attack, Mach number …) in real time is still not achievable easily by methods of numerical analysis in aerodynamics and aeroelasticity domains. For this reason, we propose a flight parameters control system. This approach is based on new optimization methodologies with neural networks (NN) and extended great deluge (EGD). The validation of this new method is realized by experimental tests using a model installed in the wind tunnel to determine the pressure distribution. For lift, drag and moment coefficient, the results of our approach are compared to the XFoil results for different angles of attack and Mach number. The main purpose of this control system is to improve the aircraft aerodynamic performance.

Commentary by Dr. Valentin Fuster
2014;():V001T01A041. doi:10.1115/IMECE2014-39756.

Nowadays, the research in the aerospace scientific field relies strongly on CFD (Computational Fluid Dynamics) algorithms, avoiding (initially at least) a large fraction of the extremely time and money consuming experiments in wind tunnels. In this paper such a recently developed academic CFD code, named Galatea, is presented in brief and validated against a benchmark test case. The prediction of compressible fluid flows is succeeded by the relaxation of the Reynolds Averaged Navier-Stokes (RANS) equations, along with appropriate turbulence models (k-ε, k-ω and SST), employed on three-dimensional unstructured hybrid grids, composed of prismatic, pyramidical and tetrahedral elements. For the discretization of the computational field a node-centered finite-volume method is implemented, while for improved computational performance Galatea incorporates an agglomeration multigrid methodology and a suitable parallelization strategy. The proposed algorithm is evaluated against the Wing-Body (WB) and the Wing-Body-Nacelles-Pylons (WBNP) DLR-F6 aircraft configurations, demonstrating its capability for a good performance in terms of accuracy and geometric flexibility.

Commentary by Dr. Valentin Fuster

Advances in Aerospace Technology: Combustion and Engine Operation

2014;():V001T01A042. doi:10.1115/IMECE2014-36255.

The effect of secondary air inlet conditions on natural gas combustor is investigated numerically. Secondary air inlet conditions include its amount, position, total number of inlet ports and its arrangement along the combustor. The secondary air is introduced normally through inlet ports at different levels along the combustor. Each level includes a number of ports distributed around the combustor periphery. The number of ports levels varied from four up to sixteen and the number of ports in each level varied from four up to sixteen ports. Thus, the total number of ports varied from 16 up to 256. The combustor used has an air swirler at its upstream. Primary air, secondary air and fuel lines are also included. The sheer-stress transport (SST) k-omega model was used to simulate the turbulent isothermal flow and the non-premixed combustion model was used to simulate the turbulent reacting flow. For validating the model, a comparison between the measured and the calculated axial temperature distribution is made which show a reasonable agreement. Primary air swirl number of 0.87 and air to fuel ratio of 30 are used in this study. Secondary air leads to a decrease in flame size. For secondary to primary air ratio (SPAR) greater than 0.3, the flame became narrower in diameter and shorter in length. For certain secondary air configuration, NO, CO, CO2 are decreased with secondary air and are further decreased when increasing the value of SPAR.

Commentary by Dr. Valentin Fuster
2014;():V001T01A043. doi:10.1115/IMECE2014-36830.

This paper presents numerical study of high-speed combustion and its relationship with thermal stress distribution on a cavity combustion chamber. First, a physical model is established to describe high speed compressible turbulent reacting flow as well as thermal transport in combustor structure. It is then applied to a model combustor with two-staged fuel injections to examine the effects of fuel flow rate and inflow conditions on the heat flux intensity and thermal stress distributions across the thickness of the combustor wall. The result shows that the injection method of the first stage has a great influence on the flow field near the second one, and it affects combustion and heat release distribution inside the combustor. The intensity of heat flux passing through the combustor wall changes along the downstream of the flow, and large thermal stresses are generated in the vicinity of the injector, the leading edge and the trailing edge of the cavity.

Topics: Combustion
Commentary by Dr. Valentin Fuster
2014;():V001T01A044. doi:10.1115/IMECE2014-36998.

The HYPROB Program, developed by the Italian Aerospace Research Centre, has the aim to increase the system design and manufacturing capabilities on liquid oxygen-methane rocket engines. It foresees the designing, manufacturing and testing of a ground engine demonstrator of three tons thrust. The demonstrator baseline concept is featured by 18 injectors and it is regeneratively cooled by using liquid methane. In particular, the cooling system is made by a constant number of axial channels and the counter-flow architecture has been chosen; methane enters the channels in the nozzle region in supercritical liquid condition, is heated by the combustion gases along the cooling jacket and then injected into the combustion chamber as a supercritical gas by means of the injection head.

The goal of this paper is to describe the thermo-structural and the thermo-fluid dynamic analyses that have been performed in order to support the design activities aiming at identifying the optimal configuration of the cooling jacket in terms of number of channels, rib height and width. In fact, a fully 3-D model, regarding a single channel, heated by the design input heat flux has been considered in order to perform CFD simulations aiming at describing the thermo-fluid dynamic behavior of methane. The results in terms of convective heat transfer coefficients have been taken into account as inputs for the thermo-structural simulations on the most critical sections of the cooling jacket. The thermo-structural activity has been conducted on the demonstrator by means of a Finite Element Method code taking into account the visco-plastic behavior of the adopted materials. In particular, transient thermal analyses and static structural analyses have been performed using ANSYS code on a 2-D model. These analyses have demonstrated that the cooling jacket can withstand the design goal of 5 thermo-mechanical cycles with a safety factor equal to 4 considering a firing time equal to 30 seconds.

Commentary by Dr. Valentin Fuster
2014;():V001T01A045. doi:10.1115/IMECE2014-37385.

This work is focused on investigation of thermal efficiency of a Hypersonic scramjet engine and propose some improvement of thermal efficiency based on thermodynamic and fluid flow analysis. Thermal management system is one of the main research fields in scramjet design. As it has no moving parts, the total thermal efficiency depends on inlet conditions, conditions of combustor exit and conditions of the engine exit. A combustor exit condition dictates the velocity and temperature after combustion. we concentrate our focus on this section. The first part of the paper, we tried to describe the fundamental exergy relationship for scramjet and we developed the relation of exergy distribution and exergy delivery rate. From an extensive literature review, we have found the relations between fluid velocity, pressure and temperature, which is described in the later part of the paper. Our main focus is to develop a combined relation of thermal efficiency in terms of engine exit velocity, temperature and air-fuel ratio. Different characteristic parameters such as overall efficiency, thermal efficiency, specific impulse have been determined at different inlet temperature ratio or the cycle static temperature ratio (T3/T0) and an optimum inlet temperature ratio is proposed for maximum overall efficiency.

Topics: Scramjets
Commentary by Dr. Valentin Fuster
2014;():V001T01A046. doi:10.1115/IMECE2014-37554.

Aeroengine bearing chambers typically contain bearings, seals, shafts and static parts. Oil is introduced for lubrication and cooling and this creates a two phase flow environment that may contain droplets, mist, film, ligaments, froth or foam and liquid pools. Efficient and effective liquid removal from a bearing chamber is a functional requirement and in recent years the University of Nottingham Technology Centre in Gas Turbine Transmission Systems has been conducting an experimental and computational research program one strand of which is investigating bearing chamber off-take flows. Initial investigations focussed on a chamber where there was a relatively deep pocket for oil collection below the chamber [1, 2]. In more recent studies Chandra et al have investigated a shallower geometry [3]. In both sets of studies, chamber residence volume and wall film thickness data have been obtained for a range of shaft speeds, scavenge ratios and liquid supply rates. Two methods of introducing liquid to the chamber have been used: a film generator that puts liquid directly onto the chamber wall and a droplet inlet system that distributes droplets from the rotating shaft.

During some of the previous investigations, visual data relating to the two phase flow in the outlet pipe immediately below the chamber was gathered together with data from pressure transducers one located in this pipe and one on the chamber itself. It has been observed that for some parameter combinations the chamber flow is gravity dominated whereas for others (typically at higher shaft speeds) the flow is shear dominated. During transition between regimes a pressure spike on the pipe pressure transducer is observed and this may be linked to a change in two phase flow regime within the outlet pipe. A study by Baker et al [4] on transient effects in gas-liquid separation has shown pressure spikes during transitions to new equilibrium conditions for two-phase pipe flow where the gas flow rate is suddenly increased.

In this paper outlet visualisation, chamber visualisation and pressure data are combined and conclusions are drawn relating to the parameters controlling whether shear or gravity dominate. The effect of the chamber flow regime on the outlet flow regime is assessed and presented. An implication of the analysis is that during transitional conditions a bearing chamber may contain a different quantity of liquid than in steady state conditions.

Commentary by Dr. Valentin Fuster
2014;():V001T01A047. doi:10.1115/IMECE2014-39748.

Non-dimensional simulation models are often used to interpret experimental data in order to better understand the combustion processes. Recent investigation has presented a thermodynamic non-dimensional model which is used for evaluating the emitted greenhouse gases of a turbofan engine at flight altitude. Combustion chamber flow has been modeled as a non-dimensional flow in provided model. Required data for this model are combustion chamber inlet, turbine inlet and outlet together with exit nozzle pressure and temperature which are calculated by gas turbine non-dimensional modeling software. In this regard mass flow and rotational speed compatibility, conservation of mass and energy equations have been solved. (GSP and KineTechs software (chemical kinetic solving software) have been used for solving the mentioned equations as well as modeling the non-dimensional flow in the combustion chamber, respectively.) This paper investigates the JT9D engine. The results have been validated with experimental data published by International Civil Aviation Organization (ICAO). Noteworthy is that the obtained experimental data have been considered for low flight altitude phases and therefore the presented model takes into account the Greenhouse Gases estimation values for the mentioned phases. After result validation and conformity of experimental data, engine cruise altitude conditions have been modeled. Subsequently, New York – London airway has been studied and emitted Greenhouse gases quantity at cruise altitude has been calculated in one year while verified with published official Statistical data.

Topics: Gases , Engines , Flight , Turbofans
Commentary by Dr. Valentin Fuster

Advances in Aerospace Technology: Next Generation Aircraft Technologies

2014;():V001T01A048. doi:10.1115/IMECE2014-36510.

For long flights, the cruise is the longest phase and is where the largest proportion of fuel is consumed. A new flight trajectory calculation method utilizing genetic algorithms is proposed here. The lateral and vertical navigation profiles are analyzed to obtain the optimal cruise trajectory in terms of fuel consumption. With a complete analysis of the wind currents, a 3D grid is created for all along the cruise phase, including latitudes, longitudes and altitudes. Different flight trajectories are calculated using genetic algorithms. To improve calculation time and precision, the flight trajectories are calculated using the performance databases for commercial aircraft, databases which are used in actual flight management system platforms. This optimization process indicates that fuel cost savings of up to 5.6% can be achieved.

Commentary by Dr. Valentin Fuster
2014;():V001T01A049. doi:10.1115/IMECE2014-36905.

The effectiveness of a novel actuation architecture developed to control flutter and post-flutter is investigated in this paper. To this purpose, the performance of an active control strategy in various operational conditions is experimentally examined. A physical prototype, consisting of a wing section with multiple spoilers mounted on an aeroelastic apparatus, has been designed and assembled to carry out open- and closed-loop operations. Wind tunnel aeroelastic testing are performed with a plunging and pitching apparatus specifically designed to simulating wing sections with prescribed stiffness characteristics, including torsional structural nonlinearities responsible of a stable nonlinear post-flutter limit cycle behavior. Five surface mounted spoilers located at 15% of the chord from the leading edge are used to control aeroelastic vibrations in pre- and post-flutter. The spoilers design, including selection of best size and chord position and considering the geometrical constraints, has been carried out by CFD simulation, with the objective of maximizing the aerodynamic pitching moment used to stabilize the lifting surface at the various speeds. The spoiler actuations are commanded by an active control system as to extend the flight region in the natural post-flutter condition. A simple PID algorithm is implemented to test the efficiency of the control system design to suppress flutter oscillation. A trial and error tuning of the gain has been executed on-site during the experimental campaign. Only the pitch angle is used as state feedback in the control laws to stabilize the system above the open-loop flutter velocity. Results and pertinent conclusions are outlined.

Commentary by Dr. Valentin Fuster
2014;():V001T01A050. doi:10.1115/IMECE2014-37024.

Semi-active suspensions are able to adapt their behavior to the specific necessities of each operation phase of the vehicle. Their application in the aerospace industry is very attractive, specifically in landing gear actuators. However, according to Ref. [1], no examples of commercial use of this technology in this industry are reported in recent reviews of the state of the art.

In this paper, the real capabilities of a magneto-rheological semi-active suspension implemented on an UAV nose landing gear designed and manufactured by CESA are introduced by the analysis of the performances obtained by test when it is submitted to quasi-static and dynamic working conditions.

Topics: Gears
Commentary by Dr. Valentin Fuster
2014;():V001T01A051. doi:10.1115/IMECE2014-37029.

The Weight on Wheel system on an aircraft gives an indication of whether the aircraft has weight on its wheels, meaning the airborne is on the ground.

This system monitors the loads appearing on each Landing Gear structure. It sends a status signal to Landing gear computer that manage the redundant signals coming from each WOW systems. The management of these signals gives the information needed to activate “on ground” configuration of aircraft for braking control laws. “On ground” configuration specifies maximum load conditions and requirements for shock absorbers design for each type of aircraft (heavy freighters, military transport …)

Based on CESA experience on this field, a new approach oriented to fatigue health monitoring, is presented here. Current WOW system could be modified allowing a continuous monitoring of loads experienced by Landing Gear structure. This will optimize the life cycle of the aircraft by checking individually the status of different structural parts of each Landing Gear. Therefore, it will improve maintenance actions that will be performed only when needed. Load detection and monitoring of landing gears provides a valuable and useful information data to predict accurately the accumulated fatigue damage on landing gear components.

Topics: Fatigue , Stress , Design , Gears
Commentary by Dr. Valentin Fuster
2014;():V001T01A052. doi:10.1115/IMECE2014-37542.

The Pulse Detonation Engine (PDE) is now offering the Aviation Industry a new, efficient and cheap mode of propulsion. Outside of the valving of the fuel and the air, the basic design of the PDE contains no moving parts: it is merely a tube in which some fuel is detonated, and the resultant shock wave used for propulsion. It suffers, however, from the lack of an appropriate ignition system designed especially for this propulsion technique. This paper discusses the possibility of using microwave radiation to initiate detonation in the PDE. Background information regarding the PDE, the merits of detonation over deflagration, and extant techniques for initiating detonation is included. The merits of this technique over the more traditional methods are emphasized. A practical technique of producing and controlling microwave radiation is subsequently presented. To prove viability of the central idea, a list of public patents related to the previous work done regarding the use of microwave radiation to initiate ignition is presented, along with a short summary related to each entry. This area of research is still new and unorthodox, as far as both the PDE and microwave ignition are concerned, and no work has been done until now that involves both of these. Further experiments involving realistic fuels and conditions to demonstrate the viability and practical use of this technique are required. It is expected that this research will do for the PDE what invention of spark plugs did for the gasoline (or Spark Ignition) engine.

Commentary by Dr. Valentin Fuster
2014;():V001T01A053. doi:10.1115/IMECE2014-37568.

Vertical Navigation (VNAV) trajectory optimization has been identified as a means to reduce fuel consumption. Due to the computing power limitations of devices such as Flight Management Systems (FMSs), it is very desirable to implement a fast method for calculating trajectory cost using optimization algorithms. Conventional trajectory optimization methods solve a set of differential equations called the aircraft equations of motions to find the optimal flight profile. Many FMSs do not use these equations, but rather a set of lookup tables with experimental, or pre-calculated data, called a Performance Database (PDB). This paper proposes a method to calculate a full trajectory flight cost using a PDB. The trajectory to be calculated is composed of climb, acceleration, cruise, descent and deceleration flight phases. The influence of the crossover altitude during climb and step climbs in cruise were considered for these calculations. Since the PDB is a set of discrete data, Lagrange linear interpolations were performed within the PDB to calculate the required values. Given a takeoff weight, the initial and final coordinates and the desired flight plan, the trajectory model provides the Top of Climb coordinates, the Top of Descent coordinates, the fuel burned and the flight time needed to follow the given flight plan. The accuracy of the trajectory costs calculated with the proposed method was validated for two aircraft; one with an aerodynamic model in FlightSIM, software developed by Presagis, and the other using the trajectory generated by the reference FMS.

Commentary by Dr. Valentin Fuster
2014;():V001T01A054. doi:10.1115/IMECE2014-37570.

Optimizing the flight trajectory is a goal that will minimize fuel consumption and time related costs. Lateral Navigation (LNAV) has been investigated as part of identifying optimal trajectories. Winds and temperature have an important influence in the cost of a flight. Tail winds and low temperatures are desired, as both reduce flight costs. Implementing algorithms to locate where these favorable conditions exist close to the defined trajectory of a given flight will help to achieve optimal flight trajectories. These algorithms are to be implemented in an FMS using an aircraft model which is normally given in the form of a Performance Database (PDB). The approach given in this paper uses Dijsktra’s algorithm. This method is part of the graph-search techniques. The search area is defined by discretizing the cruise trajectory and defining adjacent waypoints, forming a grid where the possible trajectories are created. The algorithm requires the aircraft’s gross weight at the top of climb (TOC), the location of the top of descent (TOD), and the desired cruise speed and altitude. The related costs are calculated using the PDB’s model for two different commercial aircraft at a constant altitude and at a constant indicated mach. To minimize the costs, the algorithm considers the fuel burned, the flight time and the cost index (CI). The temperature and winds in the trajectory are obtained from the Canadian weather forecast (Environment Canada). Wind influence is taken into account by adding it to the ground speed, based on its direction regarding the aircraft’s trajectory heading. The effect of temperature is considered in the PDB. Generated trajectories are compared against the geodesic (or great circle) route.

Commentary by Dr. Valentin Fuster
2014;():V001T01A055. doi:10.1115/IMECE2014-38215.

We examine a new robust nonlinear flight control technology that employs an array of synthetic-jet micro-actuators embedded in UAV wing design in order to completely eliminate moving parts (such as ailerons) thus greatly enhancing maneuverability required for small fixed-wing air vehicles operating, e.g., in tight urban environments. Estimated fast response times are critical in mitigating gust effects while greatly improving flight stability and control. The new controller design is particularly advantageous for high levels of uncertainty and nonlinearity present both in the unsteady flow-path environment and in the embedded actuator’s response. The current work focuses on a benchmark case of flutter control of 2-DOF elastically-mounted airfoil entering limit-cycle oscillations (LCO) due to impinging upstream flow disturbance. Preliminary parametric studies conducted for various SJA excitation amplitudes and frequencies examine the thresholds of the actuator’s control authority to produce a desirable impact.

Topics: Airfoils
Commentary by Dr. Valentin Fuster
2014;():V001T01A056. doi:10.1115/IMECE2014-38581.

Micro-Air-Vehicles (MAV) flight regimes differs significantly from larger scales airplanes. They are operating at low Reynolds numbers of approximate 104, cruising at speed about 12m/s, and are capable of agile maneuvers in limited space environment. They are compact and easily stowable to facilitate transportation. However, due to the small size, they are usually more vulnerable to the wind gusts with significant complexities associated to their flight mechanics, stability and control, which also makes difficult to quantify flight qualities and performances. Furthermore, complex aerodynamics can produce loading scenarios leading to the destruction of the vehicle during flight operation. To minimize the size of the MAV when not in use, their wings are stowed within the body of the vehicle, and are deployed during operation. To supplement the bulk of knowledge in MAV aero-mechanics, the study of the aerodynamic characteristics of a deformable membrane MAV wing is carried out in this paper. The analysis of the membrane airfoil is performed using a fluid-structure interaction 2D model, to select a set of optimal airfoil parameters for the intended flight regime. Numerical simulations are supplied and validated with an M AV model tested in the wind tunnel.

Topics: Aircraft , Membranes , Wings
Commentary by Dr. Valentin Fuster
2014;():V001T01A057. doi:10.1115/IMECE2014-38851.

The lifting surfaces of next generation of flying vehicle exhibits enhanced flexibility, particularly for high aspect-ratio solutions needed for high-altitude long endurance aircrafts and for 24/7 operations. Often the wing of these vehicles is designed as slender body with an aeroelastic behavior distinctive of cantilever beam. Based on this typical assumption the governing equations of a thin-walled beam modified to account for surface mount piezoelectric elements and subjected to deterministic external gust loads have been derived and its dynamic behavior examined. This paper assesses the effectiveness of piezoelectric elements to carry out load alleviation function over the slender structure invested by atmospheric disturbances along with the evaluation of the amount of the power density harvested via a suitable electric circuit connected to the piezoelectric elements.

Commentary by Dr. Valentin Fuster

Advances in Aerospace Technology: Peridynamics Modeling

2014;():V001T01A058. doi:10.1115/IMECE2014-39620.

We present some studies of numerical schemes for nonlocal peridynamic and nonlocal diffusion models. We describe asymptotically compatible (AC) schemes recently developed for robust discretizations of nonlocal models. The AC schemes for peridynamic models provide convergent approximations to nonlocal models associated with fixed horizon parameter as well as their limiting local models. We illustrate what quadrature based discretizations can be AC schemes and what may fail to be AC.

Commentary by Dr. Valentin Fuster
2014;():V001T01A059. doi:10.1115/IMECE2014-39887.

While multiple peridynamic material models capture the behavior of solid materials, not all structures are conveniently analyzed as solids. Finite Element Analysis often uses 1D and 2D elements to model thin features that would otherwise require a great number of 3D elements, but peridynamic thin features remain underdeveloped despite great interest in the engineering community. This work presents non-ordinary state-based peridynamic models for the simulation of thin features. Beginning from an example non-ordinary state-based model proposed by Silling in 2007, lower dimensional peridynamic models of beams and plates are developed and shown to be energy equivalent to classical models for well-behaved deformations. The resulting plate model is initially restricted to a Poisson’s ratio of ν = 1/3, but is extended to arbitrary Poisson’s ratio via a bending state decomposition. Simple test cases demonstrate the model’s performance.

Commentary by Dr. Valentin Fuster

Advances in Aerospace Technology: Poster

2014;():V001T01A060. doi:10.1115/IMECE2014-38710.

In this work a study of tool life of uncoated WC-Co inserts — utilized in the dry turning of Ti6A14V (UNS R56400) alloy — has been achieved. This study has been developed on the basis of the tool wear time monitoring as a function of cutting speed (V) for fixed feed (f) and cutting depth (d). Tool life (T) has been evaluated through the flank wear (VB) measurement, making use of ISO 3685 standard.

VB (T,V) surfaces have allowed defining the cutting speed range where T is maximum. On the other hand T(V) curves have allowed establishing the influence of cutting speed on tool life. From these curves a potential model, in the way of Taylor’s model, has been proposed.

Topics: Alloys , Turning
Commentary by Dr. Valentin Fuster

Advances in Aerospace Technology: Propulsion

2014;():V001T01A061. doi:10.1115/IMECE2014-36898.

The growing need to drastically reduce aircrafts CO2 emissions has led engineers and scientists in the last years to develop a clean, renewable and sustainable energy system. Hydrogen as a fuel in aviation has shown to be a good choice, since it has an energy release much higher than common and long chain hydrocarbons (119.96 MJ/kg vs ∽42.8 MJ/kg, respectively), a wide flammability limits, a high diffusivity and a very short ignition time. Its thermal conductivity, the highest among all gases, its high heat capacity and its very low dynamic viscosity provide superior cooling properties for operation at high flight speeds and at combustor high temperatures. Furthermore, its low molecular weight makes it the fuel with the higher specific impulse (ISP), ∼450 s: this means that burning 1 kg/s of hydrogen with oxygen produces a thrust of 450 kg-force.

However, for its high combustion temperature, it has been demonstrated to be disadvantageous in terms of NOx production. Although NOx pollution is a relatively small part of global human pollution (less than 4%), a particular feature of air transport is that pollutants from air traffic are emitted at high altitudes, in the upper troposphere/lower stratosphere (8 to 12 km), where they are of greater influence than those emitted at ground level. Moreover, further emission reductions need to be achieved by the air transport community, since air traffic has a growth (3% to 5% per year) which exceeds the technology improvement rate.

Emissions may be controlled by operating at lean or very lean equivalence ratios (thanks to the wider flammability limits of the hydrogen-air flames compared to kerosene-air flames), or reducing the combustor length (thanks to the higher flame speed of hydrogen compared to other fuels), or via innovative strategies. In this paper, the RQL (Rich-Quench-Lean) strategy for the NOx abatement will be proposed for a high speed hydrogen fuelled vehicle.

Commentary by Dr. Valentin Fuster
2014;():V001T01A062. doi:10.1115/IMECE2014-36989.

This paper presents a conceptual study of two alternative inlet concepts for the United States Air Force B-1B bomber to provide for improved supersonic performance with expansion of capabilities to high-altitude, high-speed flight at Mach 2.0. The two inlet concepts are two-dimensional, variable-ramp inlet systems designed to replace the current fixed-geometry, pitot inlets of the B-1B. One inlet incorporates a two-ramp system, while a second inlet incorporates a two-ramp system containing an isentropic contour. The entire inlet system including the supersonic diffuser, throat, cowl lip, and subsonic diffuser sections was designed to maximize the total pressure recovery at the engine fan face to achieve maximum thrust by the engine at Mach 2.0 conditions. Analytic methods implemented into the MATLAB and the NASA SUPIN codes are used to design and analyze the two-dimensional inlet concepts. In addition, high-fidelity WIND-US computational fluid dynamics (CFD) simulations were used to verify the results of the analytic design methods. The results suggest that at Mach 2.0, the total pressure recovery of the inlets could increase from 0.70 to 0.94. The inlet capture area and cowl drag increased; however, the overall improvements resulted in a 98% thrust increase over the existing inlet at the design point.

Commentary by Dr. Valentin Fuster
2014;():V001T01A063. doi:10.1115/IMECE2014-37612.

A parametric, two-dimensional, computational study examining steady-state plug micronozzle performance has been conducted. As part of the study, a new method for plug contour construction is proposed. The performance of several different nozzle geometries is compared to that of a traditional plug nozzle geometry designed using the Method of Characteristics (MOC). New nozzle designs are derived from the MOC based design and geometric transformations are used to produce plug nozzles of reduced length. Spike lengths corresponding to 60, 50, 40, and 27% of the MOC nozzle’s length are examined. The throat Reynolds number is varied from 80–820. Thrust is used a metric to assess nozzle performance. The geometry which maximizes performance is found to vary with Reynolds number. It is observed that reducing the plug length improves thrust production for the range of Reynolds number examined.

Topics: Micronozzles
Commentary by Dr. Valentin Fuster
2014;():V001T01A064. doi:10.1115/IMECE2014-38097.

This article presents a new fan testing concept which was developed by the University of Braunschweig. The aim of the concept is to provide an experimental setup for integrated jet engine fan investigations. This integrated setup includes a full bypass model with nacelle geometry in order to obtain all aerodynamic interaction effects. Since these interactions have a strong dependency from fan inlet conditions the facility provides wind tunnel capacities. Generating both uniform and angle of attack inlet conditions it is possible to investigate fan performance during the most critical operating points of a civil flight mission. The main innovation of the wind tunnel is the crosswind concept for generating angle of attack conditions in case of a non-pitchable fan unit. Numerical investigations in this paper are intended to validate this new concept by means of typical angle of attack simulations. The results show that the crosswind concept is able to generate equivalent cp-characteristics as obtained for the reference simulations at the nacelle 6 o‘clock and 12 o‘clock positions.

Commentary by Dr. Valentin Fuster
2014;():V001T01A065. doi:10.1115/IMECE2014-38438.

Two mathematical models (a one-dimensional and a two-dimensional) were adopted to study, numerically, the thermal hydrodynamic characteristics of flow inside the cooling channels of a Nuclear Thermal Rocket (NTR) engine. In the present study, only a single one of the cooling channels of the reactor core is simulated. The one-dimensional model adopted here assumes the flow in this cooling channel to be steady, compressible, turbulent, and subsonic. The physics based mathematical model of the flow in the channel consists of conservation of mass, momentum, and energy equations subject to appropriate boundary conditions as defined by the physical problem stated above. The working fluid (gaseous hydrogen) is assumed to be compressible through a simple ideal gas relation. The physical and transport properties of the hydrogen is assumed be temperature dependent. The governing equations of the compressible flow in cooling channels are discretized using the second order accurate MacCormack finite difference scheme. Convergence and grid independence studies were done to determine the optimum computational cell mesh size and computational time increment needed for the present simulations. The steady state results of the proposed model were compared to the predictions by a commercial CFD package (Fluent.) The two-dimensional CFD solution was obtained in two domains: the coolant (gaseous hydrogen) and the ZrC fuel cladding. The wall heat flux which varied along the channel length (as described by the nuclear variation in the nuclear power generation) was given as an input.

Numerical experiments were carried out to simulate the thermal and hydrodynamic characteristics of the flow inside a single cooling channel of the reactor for a typical NERVA type NTR engine where the inlet mass flow rate was given as an input. The time dependent heat generation and its distribution due to the nuclear reaction taking place in the fuel matrix surrounding the cooling channel. Numerical simulations of flow and heat transfer through the cooling channels were generated for steady state gaseous hydrogen flow. The temperature, pressure, density, and velocity distributions of the hydrogen gas inside the coolant channel are then predicted by both one-dimensional and two-dimensional model codes. The steady state predictions of both models were compared to the existing results and it is concluded that both models successfully predict the steady state fluid temperature and pressure distributions experienced in the NTR cooling channels. The two dimensional model also predicts, successfully, the temperature distribution inside the nuclear fuel cladding.

Commentary by Dr. Valentin Fuster

Advances in Aerospace Technology: Turbine and Blade Aerodynamics and Performance

2014;():V001T01A066. doi:10.1115/IMECE2014-36263.

The film cooling performance using novel sister shaped single-hole (SSSH) schemes are numerically investigated in the present study. The downstream, upstream and up/downstream SSSH configurations are formed by merging the discrete sister holes to the primary injection hole through a series of specific orientations. The obtained results are compared with a conventional cylindrical hole and a forward diffused shaped hole. The RANS simulations are performed using the realizable k-ε model with the standard wall function. Results are presented for low and high blowing ratios of 0.25 and 1.5, respectively. The film cooling effectiveness is notably increased for the novel shaped holes, particularly at the high blowing ratio of 1.5. Furthermore, a considerable decrease in the jet lift-off has been achieved for the proposed film hole geometries, wherein fully attached flow to the wall surface is observed for the upstream and up/downstream SSSH schemes.

Commentary by Dr. Valentin Fuster
2014;():V001T01A067. doi:10.1115/IMECE2014-36866.

A correlation based approach for estimation of the turbulence length scale lT at the inflow boundary is proposed and presented. This estimation yields reasonable turbulence decay, supporting the transition model in accurately predicting the laminar-turbulent transition location and development. As an additional element of the approach, the sensitivity of the turbulence model to free stream values is suppressed by limiting the eddy viscosity in non-viscous regions. Therefore a criterion to detect those regions, based only on local variables, is derived. The method is implemented in DLR’s turbomachinery flow solver TRACE in the framework of the k–ω turbulence model by Wilcox 1988 [1] and the γ–Reθ transition model by Langtry and Menter [2] in combination with a cross flow (CF) induced transition criterion after Müller [3]. The improved model is tested to the T161 turbine test case [4], [5] and validated at the Durham turbine Cascade [6] and an outlet guide vane for low pressure turbine configurations [7].

Commentary by Dr. Valentin Fuster
2014;():V001T01A068. doi:10.1115/IMECE2014-36903.

Analytical Target Cascading (ATC), a multilayer multidisciplinary design optimization (MDO) formulation employed on a transonic fan design problem. This paper demonstrates the ATC solution process including the specific way of initializing the problem and handling system level and discipline level targets. High-fidelity analysis tools for aerodynamics, structure and dynamics disciplines have been used. A multi-level parameterization of the fan blade is considered for reducing the number of design variables. The overall objective is the transonic fan efficiency improvement under structure and dynamics constraints. This design approach is applied to the redesign of the NASA Rotor 67. The overall study explores the key points of implementation of ATC on transonic fan design practical problem.

Topics: Design , Optimization , Blades
Commentary by Dr. Valentin Fuster
2014;():V001T01A069. doi:10.1115/IMECE2014-37790.

Axial flow compressors have a limited operation range due to the difficulty controlling the secondary flow. Vortex generators are considered to control the secondary flow losses and consequently enhance the compressor’s performance. In the present work, a numerical simulation of three-dimensional unsteady compressible flow has been developed in order to gain insight into the nature of this flow. Based on the numerical simulation, the effects of vortex generators with variable geometrical parameters and their application inside the cascade are investigated. The predicted flow fields with and without the vortex generators are presented and discussed. For each configuration of vortex generator, the total pressure and loss coefficient are calculated. The predicted velocity and pressure distributions at different locations are compared with the predicted and measured values available in the literatures.

Commentary by Dr. Valentin Fuster
2014;():V001T01A070. doi:10.1115/IMECE2014-39079.

A multi-objective and multi-point optimization methodology is developed for aerodynamic design of transonic fan blades. The optimization method aims to increase design efficiency, near stall efficiency and stall margin while maintaining the required design pressure ratio and high speed choke margin. Numerical analyses are performed by solving three-dimensional Reynolds-Averaged Navier-Stokes equations combined with shear stress turbulence model. A multi-level blade parameterization is employed to modify the blade geometry. The proposed method is applied to redesign NASA rotor 67. First, an optimization case with considering two operating conditions at peak efficiency and near stall is performed to demonstrate the relation between near stall efficiency and stall margin. An investigation on Pareto optimal solutions of this optimization shows that the stall margin is increased with improving near stall efficiency. Then, in order to maintain the required choke margin, an operating point at high speed choked flow is added to the optimization process. A final optimized design is selected by considering the interaction of design requirements at all three operating points. The new design presents higher efficiency and stall margin without any reduction in the chocking mass flow rate.

Topics: Optimization , Blades
Commentary by Dr. Valentin Fuster
2014;():V001T01A071. doi:10.1115/IMECE2014-39750.

In this study, a 3D design procedure for axial flow tandem stage is developed, and then a highly loaded stage is designed with this method. Designed stage is numerically investigated with a CFD code and the stage characteristic map is reported. In order to investigate the effect of highly loaded design, a conventionally loaded compressor stage is designed with single blade for rotor and stator rows. For a better comparison, Chord lengths and hub/shroud geometries are selected same between highly and conventionally loaded compressors. Characteristic map of the conventional compressor is reported too and performance of highly loaded and conventionally loaded compressors are compared.

To validate the CFD method, another compressor stage is presented that its geometry and experimental results are available. Performance of the recent compressor stage is numerically investigated and compared with experimental results.

Topics: Compressors , Design
Commentary by Dr. Valentin Fuster
2014;():V001T01A072. doi:10.1115/IMECE2014-39768.

Thermal stresses and failure of components in the gas turbine passages are common due to exposure of the passage walls to extremely hot combustion gasses. In addition, the gas turbine engine efficiency suffers due to secondary flows and passage vortex formations in turbine passages. Investigations are being conducted to reduce the thermal stresses and secondary flows in the turbine passages with endwall modifications and film cooling. The present investigation employs two types of fillets at the junction of blade leading edge and endwall in a low speed linear blade cascade to control the effects of passage vortex. Blade and endwall static pressure, axial vorticity and air temperature near endwall, and Nusselt numbers on endwall and blade wall are measured and presented in the cascade passage with and without the leading edge fillets. A constant Reynolds number of 233,000 based on the blade chord and the inlet velocity is employed for the measurements. The blade profile is obtained from the hub-side of first stage blade in the GE-E3 gas turbine and scaled 10 times in the present cascade. One of the two fillet shapes has a linear profile from the blade to endwall (Filet 1), and the other has a parabolic profile (Fillet 2) from the blade to endwall. Fillets are employed only at the blade leading edge at the bottom endwall. Results on axial vorticity and air temperature indicate that the fillets weaken passage vortex and reduce heat transfer from the endwall. This occurs as the pitchwise endwall pressure difference from the pressure side to the suction side is reduced near the filleted region. However, the distributions of wall static pressure coefficients and Nusselt numbers along the blade surface are about the same with and without the fillets and indicate no effects from the fillets on the blade surface. The blade-loadings are thus unaffected and require no design modifications with the fillets. The endwall Nusselt number distributions show lower values for the fillets than for the baseline (without fillets) which also indicate reduced heat transfer to or from the endwall. The results of the present investigations thus can be applied in designing the blade passages where the secondary flow effects are passively controlled and endwall thermal stresses are reduced.

Commentary by Dr. Valentin Fuster
2014;():V001T01A073. doi:10.1115/IMECE2014-39796.

Multi-Objective Optimization Problems (MOP) are very usual and complicated subjects in Turbomachinery and there are several methodologies for optimizing these problems. Genetic Algorithm (GA) and Artificial Neural Network (ANN) are the most popular ones to solve MOP. In this study, optimization was done for leaned rotor blades to achieve maximum performance parameters including specifically stage pressure ratio, efficiency and operating range.

By bending an existing transonic rotor which is well-known as NASA rotor-67 in tangential direction, effect of leaning on performance and aerodynamic parameters of transonic axial-flow compressor rotors was studied. To understand all effects of lean angle, an organized investigation including numerical simulation of basic rotor, implementation of curvatures on basic rotor, numerical simulation of leaned blades and optimization were applied. Various levels of lean angles were implemented to basic rotor and by employing a three dimensional compressible turbulent model, the operating parameters were achieved. Afterwards, the results were used as input data of optimization computer code.

Finally, the ANN optimization method was used to achieve maximum stage pressure ratio, efficiency and safe operating range. it was found that the Optimized leaned blades according to their target function had positive or negative optimized angles and the optimized lean angles effectively increased the safe operating range about 12% and simultaneously increase the pressure ratio and efficiency by 4% and 5%, respectively.

Topics: Design , Rotors , Axial flow
Commentary by Dr. Valentin Fuster
2014;():V001T01A074. doi:10.1115/IMECE2014-39881.

There is a severe tendency to reduce weight and increase power of gas turbine. Such a requirement is fulfilled by higher pressure ratio of compressor stages. Employing tandem blades in multi-stage axial flow compressors is a promising methodology to control separation on suction sides of blades and simultaneously implement higher turning angle to achieve higher pressure ratio. The present study takes into account the high flow deflection capabilities of the tandem blades consisting of NACA-65 airfoil with fixed percent pitch and axial overlap at various flow incidence angles. In this regard, a two-dimensional cascade model of tandem blades is constructed in a numerical environment. The inlet flow angle is varied in a wide range and overall loss coefficient and deviation angles are computed. Moreover, the flow phenomena between the blades and performance of both forward and afterward blades are investigated. At the end, the aerodynamic flow coefficient of tandem blades are also computed with equivalent single blades to evaluate the performance of such blades in both design and off-design domain of operations. The results show that tandem blades are quite capable of providing higher deflection with lower loss in a wide range of operation and the base profile can be successfully used in design of axial flow compressor. In comparison to equivalent single blades, tandem blades have less dissipation because the momentum exerted on suction side of tandem blades confines the size of separation zone near trailing edges of blades.

Commentary by Dr. Valentin Fuster
2014;():V001T01A075. doi:10.1115/IMECE2014-39883.

This paper presents the design procedure of a ducted contra-rotating axial flow fan and investigates the flow behavior inside it using ANSYS CFX-15 flow solver. This study investigates parameters such as pressure ratio, inlet mass flow rate and efficiency in different operating points. This system consists of two rotors with an outer diameter of 434 mm and an inner diameter of 260 mm which rotate contrary to each other with independent nominal rotational speeds of 1300 rpm. Blades’ maximum thickness and rotational speeds of each rotor will be altered as well as the axial distance between the two rotors to investigate their effect on the overall performance of the system. Designed to deliver a total pressure ratio of 1.005 and a mass flow rate of 1.8 kg/s at nominal rotational speeds, this system proves to be much more efficient compared to the conventional rotor-stator fans. NACA-65 airfoils are used in this analysis with the necessary adjustments at each section. Inverse design method is used for the first rotor and geometrical constraints are employed for the second one to have an axial inlet and outlet flow without using any inlet or outlet guide vanes. Using free vortex swirl distribution method, characteristic parameters and the necessary data for 3D generation of this model are obtained. The appropriate grid is generated using ATM method in ANSYS TurboGrid and the model is simulated in CFX-15 flow solver by employing k-ε turbulence model in the steady state condition. Both design algorithm and simulation analysis confirm the high anticipated efficiency for this system. The accuracy of the design algorithm will be explored and the most optimum operating points in different rotational speed ratios and axial distances will be identified. By altering the outlet static pressure of the system, the characteristic map is obtained.

Commentary by Dr. Valentin Fuster
2014;():V001T01A076. doi:10.1115/IMECE2014-40190.

In this work, the 3D design of the stator, rotor of a turbine is performed. A one way coupling between a detailed physicochemical box model and multidimensional Navier-Stokes solver (FLUENT software) is used. Various series of three-dimensional calculations including approximately 500,000 elements are carried out to calculate aero-thermodynamics fields for a first stage of high-pressure turbine of the CFM56 aero-engine. The results show that blades of early turbine stages, directly downstream of combustor are subjected to relatively high levels of unsteadiness generated from complex significant three dimensional shear layers. The latter causes the formation of large-scale turbulent. By consequence, the complex interactions between the geometrical parameters, thermodynamical and chemical processes involving aerosol precursor formation in the turbine are analyzed and investigated.

Commentary by Dr. Valentin Fuster

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