Axial compressor response to inlet flow distortions by a CFD analysis

The usual approach to compressor design considers uniform inlet flow characteristics. Especially in aircraft applications, the inlet flow is quite often non uniform, and this can result in severe performance degradation. The magnitude of this phenomenon is amplified in military engines due to the complexity of inlet duct configurations and the extreme flight conditions. CFD simulation is an innovative and powerful tool for studying inlet distortions and can bring this inside the very early phases of the design process. This project attempts to study the effects of inlet flow distortions in an axial flow compressor trying to minimize the use computer resources and computational time. The first stage of a low bypass ratio compressor has been analyzed and its clean and distorted performance compared outlining the principal changes due to uneven flow distribution: drop in mass flow, increase in pressure and temperature ratios, decrease in surge margin. Three different studies have then been conducted to better understand the effects of the level, the type and the frequency of the distortion.

Movement of deposited water on turbomachinery rotor blade surfaces

A theoretical approach for calculating the movement of liquid water following deposition onto a turbomachine rotor blade is described. Such a situation can occur during operation of an aero-engine in rain. The equation of motion of the deposited water is developed on an arbitrarily oriented plane triangular surface facet. By dividing the blade surface into a large number of facets and calculating the water trajectory over each one crossed in turn, the overall trajectory can be constructed. Apart from the centrifugal and Coriolis inertia effects, the forces acting on the water arise from the blade surface friction, and the aerodynamic shear and pressure gradient. Non- dimensionalisation of the equations of motion provides considerable insight and a detailed study of water flow on a flat rotating plate set at different stagger angles demonstrates the paramount importance of blade surface friction. The extreme cases of low and high blade friction are examined and it is concluded that the latter (which allows considerable mathematical generalisation) is the most likely in practice. It is also shown that the aerodynamic shear force, but not the pressure force, may influence the water motion. Calculations of water movement on a low-speed compressor blade and the fan blade of a high bypass ratio aero-engine suggest that in low rotational speed situations most of the deposited water is centrifuged rapidly to the blade tip region. Copyright © 2006 by ASME.

Large-eddy simulation of complex geometry jets

Computations are made for chevron and coflowing jet nozzles. The latter has a bypass ratio of 6:1. Also, unlike the chevron nozzle, the core flow is heated, making the inlet conditions reminiscent of those for a real engine. A large-eddy resolving approach is used with circa 12 × 10 6 cell meshes. Because the codes being used tend toward being dissipative the subgrid scale model is abandoned, giving what can be termed numerical large-eddy simulation. To overcome near-wall modeling problems a hybrid numerical large-eddy simulation-Reynolds-averaged Navier-Stokes related method is used. For y + ≤ 60 a Reynolds-averaged Navier-Stokes model is used. Blending between the two regions makes use of the differential Hamilton-Jabobi equation, an extension of the eikonal equation. For both nozzles, results show encouraging agreement with measurements of other workers. The eikonal equation is also used for ray tracing to explore the effect of the mean flow on acoustic ray trajectories, thus yielding a coherent solution strategy. © 2011 by Cambridge University.

Large eddy simulation of a hot, high speed coflowing jet

Computations are made of a short cowl coflowing jet nozzle with a bypass ratio 8 : 1. The core flow is heated, making the inlet conditions reminiscent of those for a real engine. A large eddy resolving approach is used with a 12 × 106 cell mesh. Since the code being used tends towards being dissipative the sub-grid scale (SGS) model is abandoned giving what can be termed Numerical Large Eddy Simulation (NLES). To overcome near wall modelling problems a hybrid NLES-RANS (Reynolds Averaged Navier-Stokes) related method is used. For y+ ≤ 60 a κ-l model is used. Blending between the two regions makes use of the differential Hamilton-Jabobi (HJ) equation, an extension of the eikonal equation. Results show encouraging agreement with existing measurements of other workers. The eikonal equation is also used for acoustic ray tracing to explore the effect of the mean flow on acoustic ray trajectories, thus yielding a coherent solution strategy. Copyright © 2011 by ASME.

Large-eddy simulations and measurements of a small-scale high-speed coflowing jet

Measurements and predictions are made of a short-cowl coflowing jet with a bypass ratio of 8:1. The Reynolds number is 300,000, and the inlet Mach numbers are representative of aeroengine conditions. The low Reynolds number of the measurements makes the case well suited to the assessment of large-eddy-simulation-related strategies. The nozzle concentricity is carefully controlled to deal with the emerging metastability issues of jets with coflow. Measurements of mean quantities and turbulence statistics are made using both laser Doppler anemometry and particle image velocimetry. The simulations are completed on 6× 106, 12× 106, and 50 × 106 cell meshes. To overcome near-wall modeling problems, a hybrid large-eddy-simulation-Reynolds-averaged-Navier-Stokesrelated method is used. The near-wall Reynolds-averaged-Navier-Stokes layer is helpful in preventing nonphysical separation from the nozzle wall.Copyright © 2010 by the American Institute of Aeronautics and Astronautics, Inc.

Comparison of les to LDA and PIV measurements of a small scale high speed coflowing jet

Measurements and predictions are made of a short cowl co-flowing jet with a bypass ratio of 8:1. The Reynolds number for computations and measurements are matched at 300,000 and the Mach numbers representative of realistic jet conditions with core and co flow velocities of 240m/s and 216m/s respectively. The low Reynolds number of the measurements makes the case well suited to the assessment of large eddy resolving computational strategies. Also, the nozzle concentricity was carefully controlled to deal with the emerging metastability issues of jets with coflow. Measurements of mean quantities and turbulence statistics are made using both two dimensional coincident LDA and PIV systems. The computational simulations are completed on a modest 12×106 mesh. The simulation is now being run on a 50×106 mesh using hybrid RANSNLES (Numerical Large Eddy Simulation). Close to the nozzle wall a k-l RANS model is used. For an axisymmetric jet, comparison is made between simulations which use NLES, RANSNLES and also a simple imposed velocity profile where the nozzle is not modeled. The use of a near wall RANS model is shown to be beneficial. When compared with the measurements the NLES results are encouraging. Copyright © 2008 by the American Institute of Aeronautics and Astronautics, Inc. All rights reserved.

A novel MPT noise prediction methodology for highly-integrated propulsion systems with inlet flow distortion

Embedded propulsion systems, such as for example used in advanced hybrid-wing body aircraft, can potentially offer major fuel burn and noise reduction benefits but introduce challenges in the aerodynamic and acoustic integration of the high-bypass ratio fan system. A novel approach is proposed to quantify the effects of non-uniform flow on the generation and propagation of multiple pure tone noise (MPTs). The new method is validated on a conventional inlet geometry first. The ultimate goal is to conduct a parametric study of S-duct inlets in order to quantify the effects of inlet design parameters on the acoustic signature. The key challenge is that the mechanism underlying the distortion transfer, noise source generation and propagation through the non-uniform flow field are inherently coupled such that a simultaneous computation of the aerodynamics and acoustics is required. The technical approach is based on a body force description of the fan blade row that is able to capture the distortion transfer and the MPT noise generation mechanisms while greatly reducing computational cost. A single, 3-D full-wheel unsteady CFD simulation, in which the Euler equations are solved to second-order spatial and temporal accuracy, simultaneously computes the MPT noise generation and its propagation in distorted mean flow. Several numerical tools were developed to enable the implementation of this new approach. Parametric studies were conducted to determine appropriate grid and time step sizes for the propagation of acoustic waves. The Ffowcs-Williams and Hawkings integral method is used to propagate the noise to far field receivers. Non-reflecting boundary conditions are implemented through the use of acoustic buffer zones. The body force modeling approach is validated and proof-of-concept studies demonstrate the generation of disturbances at both blade-passing and shaft-order frequencies using the perturbed body force method. The full methodology is currently being validated using NASA's Source Diagnostic Test (SDT) fan and inlet geometry. Copyright © 2009 by Jeff Defoe, Alex Narkaj & Zoltan Spakovszky.

Flexible conversion ratio fast reactors: Overview

Conceptual designs of lead-cooled and liquid salt-cooled fast flexible conversion ratio reactors were developed. The performance achievable by the unity conversion ratio cores of these reactors was compared to an existing supercritical carbon dioxide-cooled (S-CO2) fast reactor design and an uprated version of an existing sodium-cooled fast reactor. All concepts have cores rated at 2400 MWt. The cores of the liquid-cooled reactors are placed in a large-pool-type vessel with dual-free level, which also contains four intermediate heat exchangers (IHXs) coupling a primary coolant to a compact and efficient supercritical CO2 Brayton cycle power conversion system. The S-CO2 reactor is directly coupled to the S-CO2 Brayton cycle power conversion system. Decay heat is removed passively using an enhanced reactor vessel auxiliary cooling system (RVACS) and a passive secondary auxiliary cooling system (PSACS). The selection of the water-cooled versus air-cooled heat sink for the PSACS as well as the analysis of the probability that the PSACS may fail to complete its mission was performed using risk-informed methodology. In addition to these features, all reactors were designed to be self-controllable. Further, the liquid-cooled reactors utilized common passive decay heat removal systems whereas the S-CO2 uses reliable battery powered blowers for post-LOCA decay heat removal to provide flow in well defined regimes and to accommodate inadvertent bypass flows. The multiple design limits and challenges which constrained the execution of the four fast reactor concepts are elaborated. These include principally neutronics and materials challenges. The neutronic challenges are the large positive coolant reactivity feedback, small fuel temperature coefficient, small effective delayed neutron fraction, large reactivity swing and the transition between different conversion ratio cores. The burnup, temperature and fluence constraints on fuels, cladding and vessel materials are elaborated for three categories of material - materials currently available, available on a relatively short time scale and available only with significant development effort. The selected fuels are the metallic U-TRU-Zr (10% Zr) for unity conversion ratio and TRU-Zr (75% Zr) for zero conversion ratio. The principal selected cladding and vessel materials are HT-9 and A533 or A508, respectively, for current availability, T-91 and 9Cr-1Mo steel for relatively short-term availability and oxide dispersion strengthened ferritic steel (ODS) available only with significant development. © 2009 Elsevier B.V. All rights reserved.

Plane-strain discrete dislocation plasticity with climb-assisted glide motion of dislocations

A small-strain two-dimensional discrete dislocation plasticity (DDP) framework is developed wherein dislocation motion is caused by climb-assisted glide. The climb motion of the dislocations is assumed to be governed by a drag-type relation similar to the glide-only motion of dislocations: such a relation is valid when vacancy kinetics is either diffusion limited or sink limited. The DDP framework is employed to predict the effect of dislocation climb on the uniaxial tensile and pure bending response of single crystals. Under uniaxial tensile loading conditions, the ability of dislocations to bypass obstacles by climb results in a reduced dislocation density over a wide range of specimen sizes in the climb-assisted glide case compared to when dislocation motion is only by glide. A consequence is that, at least in a single slip situation, size effects due to dislocation starvation are reduced. By contrast, under pure bending loading conditions, the dislocation density is unaffected by dislocation climb as geometrically necessary dislocations (GNDs) dominate. However, climb enables the dislocations to arrange themselves into lower energy configurations which significantly reduces the predicted bending size effect as well as the amount of reverse plasticity observed during unloading. The results indicate that the intrinsic plasticity material length scale associated with GNDs is strongly affected by thermally activated processes and will be a function of temperature. © 2013 IOP Publishing Ltd.

Angle-of-attack effects on counter-rotating propellers at take-off

Due to their potential for significant fuel consumption savings, Counter-Rotating Open Rotors (CRORs) are currently being considered as an alternative to high-bypass turbofans. When CRORs are mounted on an aircraft, several 'installation effects' arise which are not present when the engine is operated in isolation. This paper investigates how flow features arising from one such effect - The angle-of-attack of the engine centre-line relative to the oncoming flow - can influence the design of CROR engines. Three-dimensional full-annulus unsteady CFD simulations are used to predict the time-varying flow field experienced by each rotor and emphasis is put on the interaction of the frontrotor wake and tip vortex with the rear-rotor. A parametric study is presented that quantifies the rotorrotor interaction as a function of the angle-of-attack. It is shown that angle-of-attack operation significantly changes the flow field and the unsteady lift on both rotors. In particular, a frequency analysis shows that the unsteady lift exhibits sidebands around the rotor-rotor interaction frequencies. Further, a non-linear increase in the total rear-rotor tip unsteadiness is observed for moderate and high angles-of-attack. The results presented in this paper demonstrate that common techniques used to mitigate CROR noise, such as modifying the rotor-rotor axial spacing and rear-rotor crop, can not be applied correctly unless angle-of-attack effects are taken into account. Copyright © 2012 by ASME.

Models for acoustic propagation through turbofan exhaust flows

Turbomachinery noise radiating into the rearward arc is an important problem. This noise is scattered by the trailing edges of the nacelle and the jet exhaust, and interacts with the shear layers between the external flow, bypass stream and jet, en route to the far field. In the past a range of relevant model problems involving semi-infinite cylinders have been solved. However, one limitation of previous solutions is that they do not allow for the jet nozzle to protrude a finite distance beyond the end of the nacelle (or in certain configurations being buried a finite distance upstream). In this paper we use the matrix Wiener-Hopf technique, which will allow precisely the finite nacelle-jet nozzle separation to be included. The crucial step in our work is to factorise a certain matrix as a product of terms analytic and invertible in the upper/lower halves of the complex plane. The way we do this matrix factorisation is quite different in the buried and protruding nozzle cases. In the buried case our solution method is the so-called pole-removal technique. In the technically more demanding protruding case, however, we must first use Pade approximants to generate a uniformly-valid, meromorphic representation of a certain function, before the same pole-removal method can be applied. Sample results are presented, investigating in particular the effects of exit plane stagger. © 2007 by B Veitch and N Peake.

Aerodynamic design of high end wall angle turbine stages-part I: Methodology development

A design methodology is presented for turbines in an annulus with high end wall angles. Such stages occur where large radial offsets between the stage inlet and stage outlet are required, for example in the first stage of modern low pressure turbines, and are becoming more prevalent as bypass ratios increase. The turbine vanes operate within s-shaped ducts which result in meridional curvature being of a similar magnitude to the bladeto-blade curvature. Through a systematic series of idealized computational cases, the importance of two aspects of vane design are shown. First, the region of peak end wall meridional curvature is best located within the vane row. Second, the vane should be leant so as to minimize spanwise variations in surface pressure-this condition is termed "ideal lean." This design philosophy is applied to the first stage of a low pressure turbine with high end wall angles. © 2014 by ASME.

Improving intermediate pressure turbine performance by using a nonorthogonal stator

Intermediate pressure (IP) turbines in high bypass ratio civil aeroengines are characterized by a significant increase in radius and a low aspect ratio stator. Conventional aerodynamic designs for the IP turbine stator have had leading and trailing edges orthogonal to the hub and casing end walls. The IP turbine rotor, however, is stacked radially due to stress limits. These choices inevitably lead to a substantial gap between the IP stator and rotor at the outer diameter in a duct that is generally diffusing the flow due to the increasing radius. In this low Mach number study, the IP stator is redesigned, incorporating compound sweep so that the leading and trailing edges are no longer orthogonal to the end walls. Computational investigations showed that the nonorthogonal stator reduces the flow diffusion between the stator and rotor, which yields two benefits: The stator trailing edge velocity was reduced, as was the boundary layer growth on the casing end wall within the gap. Experimental measurements confirm that the turbine with the nonorthogonal stator has an increased efficiency by 0.49%, while also increasing the work output by 4.6%, at the design point. © 2014 by ASME.

Measurements of intake separation hysteresis in a model fan and nacelle RIG

A 1/20-scale, low speed model rig representing the fan and nacelle of a high bypass ratio jet engine has been tested under crosswind conditions. The flow conditions under which the intake flow separates and reattaches have been found to exhibit considerable hysteresis. This phenomenon has been examined by a careful test procedure in which the crosswind angle has been slowly increased and then decreased. Measurements of the hysteresis associated with separation and reattachment are presented for independent variations in stream-tube contraction ratio, ground clearance, fan operating point and Reynolds number. The results reveal that particular care must be taken to allow for any hysteresis when testing intakes under crosswind conditions. They also indicate that separation hysteresis is particularly sensitive to fan operating point and the position of the ground plane. These findings suggest that it is important for high Reynolds number intake tests and calculations to include a ground plane and a model of the downstream turbomachinery. © 2002 by the author(s).

Development of a turning mid turbine frame with embedded design-part II: Unsteady measurements

© 2014 by ASME. This paper, the second of two parts, presents a new setup for the two-stage two-spool facility located at the Institute for Thermal Turbomachinery and Machine Dynamics (ITTM) of Graz University of Technology. The rig was designed to reproduce the flow behavior of a transonic turbine followed by a counter-rotating low pressure stage such as those in high bypass aero-engines. The meridional flow path of the machine is characterized by a diffusing S-shaped duct between the two rotors. The role of wide chord vanes placed into the mid turbine frame is to lead the flow towards the low pressure (LP) rotor with appropriate swirl. Experimental and numerical investigations performed on this setup showed that the wide chord struts induce large wakes and extended secondary flows at the LP inlet flow. Moreover, large deterministic fluctuations of pressure, which may cause noise and blade vibrations, were observed downstream of the LP rotor. In order to minimize secondary vortices and to damp the unsteady interactions, the mid turbine frame was redesigned to locate two zero-lift splitters into each vane passage. While in the first part of the paper the design process of the splitters and the time-averaged flow field were presented, in this second part the measurements performed by means of a fast response probe will support the explanation of the time-resolved field. The discussion will focus on the comparison between the baseline case (without splitters) and the embedded design.