The effect of wake passing on a flow separation in a low pressure turbine cascade

This paper describes an investigation into the effect that passing wakes have on a separation bubble that exists on the pressure surface and near the leading edge of a low pressure turbine blade. Previous experimental studies have shown that the behaviour of this separation is strongly incidence dependent and that it responds to its disturbance environment. The results presented in this paper examine the effect of wake passing in greater detail. Two dimensional, Reynolds averaged, numerical predictions are first used to examine qualitatively the unsteady interaction between the wakes and the separation bubble. The separation is predicted to consist of spanwise vortices whose development is in phase with the wake passing. However, comparison with experiments shows that the numerical predictions exaggerate the coherence of these vortices and also overpredict the time-averaged length of the separation. Nonetheless, experiments strongly suggest that the predicted phase locking of the vortices in the separation onto the wake passing is physical.

Numerical simulations of unsteady wet steam flows with non-equilibrium condensation in the nozzle and the steam turbine

Numerical techniques for non-equilibrium condensing flows are presented. Conservation equations for homogeneous gas-liquid two-phase compressible flows are solved by using a finite volume method based on an approximate Riemann solver. The phase change consists of the homogeneous nucleation and growth of existing droplets. Nucleation is computed with the classical Volmer-Frenkel model, corrected for the influence of the droplet temperature being higher than the steam temperature due to latent heat release. For droplet growth, two types of heat transfer model between droplets and the surrounding steam are used: a free molecular flow model and a semi-empirical two-layer model which is deemed to be valid over a wide range of Knudsen number. The computed pressure distribution and Sauter mean droplet diameters in a convergent-divergent (Laval) nozzle are compared with experimental data. Both droplet growth models capture qualitatively the pressure increases due to sudden heat release by the non-equilibrium condensation. However the agreement between computed and experimental pressure distributions is better for the two-layer model. The droplet diameter calculated by this model also agrees well with the experimental value, whereas that predicted by the free molecular model is too small. Condensing flows in a steam turbine cascade are calculated at different Mach numbers and inlet superheat conditions and are compared with experiments. Static pressure traverses downstream from the blade and pressure distributions on the blade surface agree well with experimental results in all cases. Once again, droplet diameters computed with the two-layer model give best agreement with the experiments. Droplet sizes are found to vary across the blade pitch due to the significant variation in expansion rate. Flow patterns including oblique shock waves and condensation-induced pressure increases are also presented and are similar to those shown in the experimental Schlieren photographs. Finally, calculations are presented for periodically unsteady condensing flows in a low expansion rate, convergent-divergent (Laval) nozzle. Depending on the inlet stagnation subcooling, two types of self-excited oscillations appear: a symmetric mode at lower inlet subcooling and an asymmetric mode at higher subcooling. Plots of oscillation frequency versus inlet sub-cooling exhibit a hysteresis loop, in accord with observations made by other researchers for moist air flow. Copyright © 2006 by ASME.

An experimental study of unsteady separated shock wave boundary layer interactions

An experimental investigation of the unsteady interaction between a turbulent boundary layer and a normal shock wave of strength M∞ = 1.4 subject to periodic forcing in a parallel walled duct has been conducted. Emphasis has been placed on the mechanism by which changes in the global flow field influence the local interaction structure. Static pressure measurements and high speed Schlieren images of the unsteady interaction have been obtained. The pressure rise across the interaction and the appearance of the local SBLI structure have been observed to vary during the cycle of periodic shock wave motion. The magnitude of the pressure rise across the interaction is found to be related to the relative Mach number of the unsteady shock wave as it undergoes periodic motion. Variations in the upstream Influence of the interaction are sensitive to the magnitude and direction of shock wave velocity and acceleration and it is proposed that a viscous lag exists between the point of boundary layer separation and the shock wave position. Further work exploring the implications of these findings is proposed, including studies of the variation in position of the points of boundary layer separation and reattachment.

Improving the performance of a valveless pulse combustor using unsteady fuel injection

This paper describes an experimental investigation into the effect of unsteady fuel injection on the performance of a valveless pulse combustor. Two fuel systems were used. The first delivered a steady flow of ethylene through choked nozzles, and the second delivered ethylene in discrete pulses using high-frequency fuel injectors. Both fuel systems injected directly into the combustion chamber. The high-frequency fuel injectors were phase locked to the unsteady pressure measured on the inlet pipe. The phase and opening pulse width of the injectors and the time-averaged fuel mass flow rate through the injectors were independently varied. For a given fuel mass flow rate, it is shown that the maximum pressure amplitude occurs when fuel is injected during flow reversal in the inlet pipe, i.e. flow direction is out of the combustor. The optimal fuel injection pulse width is shown to be approximately 2/9th of the cycle. It should, however, be noted that this is the shortest time in which the injectors can reliably be fully opened and closed. It is shown that by using unsteady fuel injection the mass flow rate of fuel needed to achieve a given amplitude of unsteady pressure can be reduced by up to 65% when compared with the steady fuel injection case. At low fuel mass flow rates unsteady fuel injection is shown to raise the efficiency of the combustor by a factor of 7 decreasing to a factor of 2 at high fuel mass flow rates. Copyright © 2008 by the American Institute of Aeronautics and Astronautics, Inc.

Behaviour of unsteady transonic shock / boundary layer interactions with three-dimensional effects

An experimental investigation into the response of transonic SBLIs to periodic down-stream pressure perturbations in a parallel walled duct has been conducted. Tests have been carried out with a shock strength of M ∞ = 1.5 for pressure perturbation frequencies in the range 16-90 Hz. Analysis of the steady interaction at M∞ = 1.5 has also been made. The principle measurement techniques were high speed schlieren photography and laser Doppler anemometry. The structure of the steady SBLI was found to be highly three-dimensional, with large corner flows and sidewall SBLIs. These aspects are thought to influence the upstream transmission of pressure information through the interaction by affecting the post-shock flow field, including the extent of regions of secondary supersonic flow. At low frequency, the dynamics of shock motion can be predicted using an inviscid analytical model. At increased frequencies, viscous effects become significant and the shock exhibits unexpected dynamic behaviour, due to a phase lag between the upstream transmission of pressure information in the core flow and in the viscous boundary layers. Flow control in the form of micro-vane vortex generators was found to have a small impact on shock dynamics, due to the effect it had on the post-shock flow field outside the viscous boundary layer region. The relationship between inviscid and viscous effects is developed and potential destabilising mechanisms for SBLIs in practical applications are suggested. Copyright © 2009 by Paul Bruce and Holger Babinsky.

Unsteady ejectors: The effect of driver jet mark-space ratio

mark Unsteady ejectors can be driven by a wide range of driver jets. These vary from pulse detonation engines, which typically have a long gap between each slug of fluid exiting the detonation tube (mark-space ratios in the range 0.1-0.2) to the exit of a pulsejet where the mean mass flow rate leads to a much shorter gap between slugs (mark-space ratios in the range 2-3). The aim of this paper is to investigate the effect of mark-space ratio on the thrust augmentation of an unsteady ejector. Experimental testing was undertaken using a driver jet with a sinusoidal exit velocity profile. The mean value, amplitude and frequency of the velocity profile could be changed allowing the length to diameter ratio of the fluid slugs L/D and the mark-space ratio (the ratio of slug length to the spacing between slugs) L/S to be varied. The setup allowed L/S of the jet to vary from 0.8 to 2.3, while the L/D ratio of the slugs could take any values between 3.5 and 7.5. This paper shows that as the mark-space ratio of the driver jet is increased the thrust augmentation drops. Across the range of mark-space ratios tested, there is shown to be a drop in thrust augmentation of 0.1. The physical cause of this reduction in thrust augmentation is shown to be a decrease in the percentage time over which the ejector entrains ambient fluid. This is the direct result ofthe space between consecutive slugs in the driver jet decreasing. The one dimensional model reported in Heffer et al. [1] is extended to include the effect of varying L/S and is shown to accurately capture the experimentally measured behavior ofthe ejector. Copyright © 2010 by the American Institute of Aeronautics and Astronautics, Inc.

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.

The propagation of acoustic waves in a slowly varying duct with radially sheared axial mean flow

We consider the propagation of acoustic waves along a cylindrical duct carrying radially sheared axial mean flow, in which the duct radius is allowed to vary slowly along the axis. In previous work [A.J. Cooper & N. Peake, Journal of Fluid Mechanics 445 (2001) 207-234.] radially sheared axial mean flow with nonzero swirl in a slowly varying duct was considered, but in this paper we set the swirl to zero, thereby allowing simplification of the calculations of both the mean and unsteady flows. In this approach the acoustic wavenumber and corresponding eigenfunction are determined locally, while the wave amplitude is found by solving an evolution equation along the duct. Sample results are presented, including a case in which, perhaps surprisingly, the number of cut-on modes increases as the duct radius decreases. © 2013 Elsevier Ltd. All rights reserved.

Unsteady penetration of a target by a liquid jet.

It is widely acknowledged that ceramic armor experiences an unsteady penetration response: an impacting projectile may erode on the surface of a ceramic target without substantial penetration for a significant amount of time and then suddenly start to penetrate the target. Although known for more than four decades, this phenomenon, commonly referred to as dwell, remains largely unexplained. Here, we use scaled analog experiments with a low-speed water jet and a soft, translucent target material to investigate dwell. The transient target response, in terms of depth of penetration and impact force, is captured using a high-speed camera in combination with a piezoelectric force sensor. We observe the phenomenon of dwell using a soft (noncracking) target material. The results show that the penetration rate increases when the flow of the impacting water jet is reversed due to the deformation of the jet-target interface--this reversal is also associated with an increase in the force exerted by the jet on the target. Creep penetration experiments with a constant indentation force did not show an increase in the penetration rate, confirming that flow reversal is the cause of the unsteady penetration rate. Our results suggest that dwell can occur in a ductile noncracking target due to flow reversal. This phenomenon of flow reversal is rather widespread and present in a wide range of impact situations, including water-jet cutting, needleless injection, and deposit removal via a fluid jet.

Modelling and validation of wet steam flow in a low pressure steam turbine

Results of numerical investigations of the wet steam flow in a three stage low pressure steam turbine test rig are presented. The test rig is a scale model of a modern steam turbine design and provides flow measurements over a range of operating conditions which are used for detailed comparisons with the numerical results. For the numerical analysis a modern CFD code with user defined models for specific wet steam modelling is used. The effect of different theoretical models for nucleation and droplet growth are examined. It is shown that heterogeneous condensation is highly dependent on steam quality and, in this model turbine with high quality steam, a homogeneous theory appears to be the best choice. The homogeneous theory gives good agreement between the test rig traverse measurements and the numerical results. The differences in the droplet size distribution of the three stage turbine are shown for different loads and modelling assumptions. The different droplet growth models can influence the droplet size by a factor of two. An estimate of the influence of unsteady effects is made by means of an unsteady two-dimensional simulation. The unsteady modelling leads to a shift of nucleation into the next blade row. For the investigated three stage turbine the influence due to wake chopping on the condensation process is weak but to confirm this conclusion further investigations are needed in complete three dimensions and on turbines with more stages. Copyright © 2011 by ASME.

Wave propagation in a slowly-varying duct with mean vortical swirling flow

The propagation of unsteady disturbances in a slowlyvarying cylindrical duct carrying mean swirling flow is investigated using a multiple-scales technique. This is applicable to turbomachinery flow behind a rotor stage when the swirl and axial velocities are of the same order. The presence of mean vorticity couples acoustic and vorticity equations which produces an eigenvalue problem that is not self-adjoint unlike that for irrotational mean flow. In order to determine the amplitude variation along the duct, an adjoint solution for the coupled system of equations is derived. The solution breaks down where a mode changes from cut on to cut off. In this region the amplitude is governed by a form of Airy's equation, and the effect of swirl is to introduce a small shift in the origin of the Airy function away from the turning-point location. The variation of axial wavenumber and amplitude along the duct is calculated. In hard-walled ducts mean swirl is shown to produce much larger amplitude variation along the duct compared with a nonswirling flow. Mean swirl also has a large effect in ducts with finite-impedance walls which differs depending on whether modes are co-rotating with the swirl or counter rotating. © 2001 by A.J. Cooper, Published by the American Institute of Aeronautics and Astronautics, Inc.

On the laminar-turbulent transition of the rotating-disk flow: The role of absolute instability

© 2014 Cambridge University Press. This paper describes a detailed experimental study using hot-wire anemometry of the laminar-turbulent transition region of a rotating-disk boundary-layer flow without any imposed excitation of the boundary layer. The measured data are separated into stationary and unsteady disturbance fields in order to elaborate on the roles that the stationary and the travelling modes have in the transition process. We show the onset of nonlinearity consistently at Reynolds numbers, R, of ∼ 510, i.e. at the onset of Lingwood's (J. Fluid Mech., vol. 299, 1995, pp. 17-33) local absolute instability, and the growth of stationary vortices saturates at a Reynolds number of ∼ 550. The nonlinear saturation and subsequent turbulent breakdown of individual stationary vortices independently of their amplitudes, which vary azimuthally, seem to be determined by well-defined Reynolds numbers. We identify unstable travelling disturbances in our power spectra, which continue to grow, saturating at around R=585, whereupon turbulent breakdown of the boundary layer ensues. The nonlinear saturation amplitude of the total disturbance field is approximately constant for all considered cases, i.e. different rotation rates and edge Reynolds numbers. We also identify a travelling secondary instability. Our results suggest that it is the travelling disturbances that are fundamentally important to the transition to turbulence for a clean disk, rather than the stationary vortices. Here, the results appear to show a primary nonlinear steep-fronted (travelling) global mode at the boundary between the local convectively and absolutely unstable regions, which develops nonlinearly interacting with the stationary vortices and which saturates and is unstable to a secondary instability. This leads to a rapid transition to turbulence outward of the primary front from approximately R=565 to 590 and to a fully turbulent boundary layer above 650.

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.

Effect of Turbine Inlet Temperature on Rotor Blade Tip Leakage Flow and Heat Transfer

One of the most critical gas turbine engine components, the rotor blade tip and casing, is exposed to high thermal load. It becomes a significant design challenge to protect the turbine materials from this severe situation. The purpose of this paper is to study numerically the effect of turbine inlet temperature on the tip leakage flow structure and heat transfer. In this paper, the effect of turbine inlet temperature on the tip leakage flow structure and heat transfer has been studied numerically. Uniform low (LTIT: 444 K) and high (HTIT: 800 K) turbine inlet temperature, as well as non-uniform inlet temperature have been considered. The results showed the higher turbine inlet temperature yields the higher velocity and temperature variations in the leakage flow aerodynamics and heat transfer. For a given turbine geometry and on-design operating conditions, the turbine power output can be increased by 1.33 times, when the turbine inlet temperature increases 1.80 times. Whereas the averaged heat fluxes on the casing and the blade tip become 2.71 and 2.82 times larger, respectively. Therefore, about 2.8 times larger cooling capacity is required to keep the same turbine material temperature. Furthermore, the maximum heat flux on the blade tip of high turbine inlet temperature case reaches up to 3.348 times larger than that of LTIT case. The effect of the interaction of stator and rotor on heat transfer features is also explored using unsteady simulations. The non-uniform turbine inlet temperature enhances the heat flux fluctuation on the blade tip and casing.