Film-cooling jets behavior in a combustor chamber is deeply affected by swirling flow interactions and unsteadiness; on the other hand, the jets behavior has a direct impact on different phenomena such as cooling capabilities and ignition. For these reasons, an in-depth characterization of the film-cooling flows in the presence of a swirling main flow and demands dedicated time-resolved analyses. The experimental setup consists of a nonreactive single-sector linear combustor simulator installed in an open-loop wind tunnel. It is equipped with a swirler and a multiperforated plate to simulate the effusion cooling system of the liner. The rig is scaled with respect to the engine configuration to increase spatial resolution and to reduce the characteristic frequencies of the unsteady phenomena. Time-resolved particle image velocimetry (TRPIV) was exploited for the investigation testing different values of liner pressure drop. In addition, numerical investigations were carried out to gain a deeper insight of the behavior highlighted by the experiments and to assess the capability of computational fluid dynamics (CFD) in predicting the flow physics. In this work, the stress-blended eddy simulation (SBES) approach implemented in ansys fluent was adopted. Oscillations of the jets and intermittent interactions of the mainstream with the wall of the liner and hence with the film development have been investigated in detail. The results demonstrate how an unsteady analysis of the flow structures that characterize the jets, the turbulent mixing of coolant flows, and the interaction between mainstream and cooling jets is strictly necessary to have a complete knowledge of the behavior of the coolant, which in turn affects combustor operability and life time.

Internal crossflow, or internal flow that is perpendicular to the overflowing mainstream, reduces film cooling effectiveness by disrupting the diffusion of coolant at the exit of axial shaped holes. Previous experimental investigations have shown that internal crossflow causes the coolant to bias toward one side of the diffuser and that the severity of the biasing scales with the inlet velocity ratio, VR i , or the ratio of crossflow velocity to the jet velocity in the metering section of the hole. It has been hypothesized and computationally predicted that internal crossflow produces an asymmetric swirling flow within the hole that causes the coolant to bias in the diffuser and that biasing contributes to ingestion of hot mainstream gas into the hole, which is undesirable. However, there are no experimental measurements as of yet to confirm these predictions. In the present study, in- and near-hole flow field and thermal field measurements were performed to investigate the flow structures and mainstream ingestion for a standard axial shaped hole fed by internal crossflow. Three different inlet velocity ratios of VR i = 0.24, 0.36, and 0.71 were tested at varying injection rates. Measurements were made in planes normal to the nominal direction of coolant flow at the outlet plane of the hole and at two downstream locations—x/d = 0 and 5. The predicted swirling structure was observed for the highest inlet velocity ratio and flow within the hole was shown to scale with VR i . Ingestion within the diffuser was significant and also scaled with VR i . Downstream flow and thermal fields showed that increased biasing contributed to more severe jet detachment and coolant dispersion away from the surface.

In modern lean-burn aero-engine combustors, highly swirling flow structures are adopted to control the fuel-air mixing and to provide the correct flame stabilization mechanisms. Aggressive swirl fields and high turbulence intensities are hence expected in the combustor-turbine interface. Moreover, to maximize the engine cycle efficiency, an accurate design of the high-pressure nozzle cooling system must be pursued: in a film-cooled nozzle, the air taken from last compressor stages is ejected through discrete holes drilled on vane surfaces to provide a cold layer between hot gases and turbine components. In this context, the interactions between the swirling combustor outflow and the vane film cooling flows play a major role in the definition of a well-performing cooling scheme, demanding for experimental campaigns at representative flow conditions. An annular three-sector combustor simulator with fully cooled high-pressure vanes has been designed and installed at THT Lab of University of Florence. The test rig is equipped with three axial swirlers, effusion-cooled liners, and six film-cooled high-pressure vanes passages, for a vortex-to-vane count ratio of 1:2. The relative clocking position between swirlers and vanes has been chosen in order to have the leading edge of the central airfoil aligned with the central swirler. In this experimental work, adiabatic film effectiveness measurements have been carried out in the central sector vanes, in order to characterize the film-cooling performance under swirling inflow conditions. The pressure-sensitive paint (PSP) technique, based on heat and mass transfer analogy, has been exploited to catch highly detailed 2D distributions. Carbon dioxide has been used as coolant in order to reach a coolant-to-mainstream density ratio of 1.5. Turbulence and five-hole probe measurements at inlet/outlet of the cascade have been carried out as well, in order to highlight the characteristics of the flow field passing through the cascade and to provide precise boundary conditions. Results have shown a relevant effect of the swirling mainflow on the film cooling behavior. Differences have been found between the central airfoil and the adjacent ones, both in terms of leading edge stagnation point position and of pressure and suction side film coverage characteristics.

In turbines, secondary vortices and tip leakage vortices form in the blade passage and interact with each other. In order to understand the flow physics of this vortices interaction, the effects of incoming vortex on the downstream tip leakage flow are investigated by experimental, numerical, and analytical methods. In the experiment, a swirl generator was used upstream of a linear turbine cascade to generate the incoming vortex, which could interact with the downstream tip leakage vortex (TLV). The swirl generator was located at ten different pitchwise locations to simulate the quasi-steady effects. In the numerical study, a Rankine-like vortex was defined at the inlet of the computational domain to simulate the incoming swirling vortex (SV). The effects of the directions of the incoming vortices were investigated. In the case of a positive incoming SV, which has a large vorticity vector in the same direction as that of the TLV, the vortex mixes with the TLV to form one major vortex near the casing as it transports downstream. This vortices interaction reduces the loss by increasing the streamwise momentum within the TLV core. However, the negative incoming SV has little effects on the TLV and the loss. As the negative incoming SV transports downstream, it travels away from the TLV and two vortices can be identified near the casing. A triple-vortices-interaction kinetic model is used to explain the flow physics of vortex interaction, and a one-dimensional mixing analytical model are proposed to explain the loss mechanism.

At the large scale turbine rig (LSTR) at Technische Universität Darmstadt, Darmstadt, Germany, the aerothermal interaction of combustor exit flow conditions on the subsequent turbine stage is examined. The rig resembles a high pressure turbine and is scaled to low Mach numbers. A baseline configuration with an axial inflow and a swirling inflow representative for a lean combustor is modeled by swirl generators, whose clocking position toward the nozzle guide vane (NGV) leading edge can be varied. A staggered double-row of cylindrical film cooling holes on the endwall is examined. The effect of swirling inflow on heat transfer and film cooling effectiveness is studied, while the coolant mass flux rate is varied. Nusselt numbers are calculated using infrared thermography and the auxiliary wall method. Boundary layer, turbulence, and five-hole probe measurements as well as numerical simulations complement the examination. The results for swirling inflow show a decrease of film cooling effectiveness of up to 35% and an increase of Nusselt numbers of 10–20% in comparison to the baseline case for low coolant mass flux rates. For higher coolant injection, the heat transfer is on a similar level as the baseline. The differences vary depending on the clocking position. The turbulence intensity is increased to 30% for swirling inflow.

This paper presents an experimental investigation on the performances of a new film cooling structure design, in which a ramp is placed upstream of a cylindrical film hole and a cylindrical cavity with two diagonal impingement holes is set at the inlet of the film hole to generate a swirling coolant flow entering the film hole. The experiments are carried out by two undisturbed measurement techniques, planar laser induced fluorescence (PLIF) and time-resolved particle image velocimetry (TR-PIV) in a water tunnel. The effects of the upstream ramp angle, blowing ratio (BR), and coolant impingement angle on the film cooling performances of a flat plate are studied at three ramp angles (0 deg, 15 deg, and 25 deg), two coolant swirling directions (clockwise and counterclockwise), two impingement angles (15 deg and 30 deg), and three BRs (0.6, 1.0, and 1.4). The experimental results show that at high BRs, the combination structures of the upstream ramp with the swirling coolant flow generated by the impingement angles can significantly improve film cooling performances; the best combination is at a 30 deg impingement angle and a 25 deg ramp angle. This can be explained by the fact that the swirling flow is significantly pressed on to the wall by means of the upstream ramp. Using the analogous analysis of heat and mass transfer, the adiabatic film effectiveness averaged over a cross section is obtained; the analysis indicates that at high BRs, the combined effect of a ramp with a large angle of 25 deg with 30 deg impingement angle can increase the film effectiveness up to 30% when compared to the test case without a ramp at the exit of the film hole. The images captured by PLIF exhibit an interesting phenomenon, i.e., the swirling of the coolant in different directions can influence the counter vortex pair (CVP) in rotating layers, and the coolant swirling in a clockwise direction enhances the right mixing of the CVP with coolant ejection, whereas the coolant swirling in a counterclockwise direction enhances the left-mixing of the CVP with coolant ejection.

A study examining the internal cooling of turbine blades by swirling flow is presented. The sensitivity of swirling flow is investigated with regard to Reynolds number, swirl intensity, and the common geometric features of blade-cooling ducts. The flow system consists of a straight and round channel that is attached to a swirl generator with tangential inlets. Different orifices and 180-deg bends are employed as channel outlets. The experiments were carried out with magnetic resonance velocimetry (MRV) for which water was used as flow medium. As the main outcome, it was found that the investigated flows are highly sensitive to the conditions at the channel outlet. However, it was also discovered that for some outlet geometries the flow field remains the same. The associated flow features a favorable topology for heat transfer; the majority of mass is transported in the annular region close to the channel walls. Together with its high robustness, it is regarded as an applicable flow type for the internal cooling of turbine blades. A large eddy simulation (LES) was conducted to analyze the heat transfer characteristic of the associated flow for S 0 = 3 and Re = 20 , 000 . The simulation showed an averaged Nusselt number increase of factor 4.7 compared to fully developed flow. However, a pressure loss increase of factor 43 must be considered as well.

This paper describes the experimental results of a new film cooling method that utilizes swirling coolant flow through circular and shaped film cooling holes. The experiments were conducted by using a scale-up model of a film-cooling hole installed on the bottom surface of a low-speed wind tunnel. Swirling motion of the film coolant was induced inside a hexagonal plenum using two diagonal impingement jets, which were inclined at an angle of α toward the vertical direction and installed in staggered positions. These two impingement jets generated a swirling flow inside the plenum, which entered the film-cooling hole and maintained its angular momentum until exiting the film-cooling hole. The slant angle of the impingement jets was changed to α = 0 deg, 10 deg, 20 deg, and 30 deg in the wind tunnel tests. The film cooling effectiveness on a flat wall was measured by a pressure sensitive paint (PSP) technique. In addition, the spatial distributions of the nondimensional concentration (or temperature) and flow field were measured by laser-induced fluorescence (LIF) and particle image velocimetry (PIV), respectively. In the case of a circular film-cooling hole, the penetration of the coolant jet into the mainstream was suppressed by the swirling motion of the coolant. As a result, although the coolant jet was deflected in the pitch direction, the film cooling effectiveness on the wall maintained a higher value behind the cooling hole over a long range. Additionally, the kidney vortex structure disappeared. For the shaped cooling hole, the coolant jet spread wider in the spanwise direction downstream. Thus, the pitch-averaged film cooling effectiveness downstream was 50% higher than that in the nonswirling case.

The feasibility of a drag management device that reduces engine thrust on approach by generating a swirling outflow from the fan (bypass) nozzle is assessed. Deployment of such “engine air-brakes” (EABs) can assist in achieving slower and/or steeper and/or aeroacoustically cleaner approach profiles. The current study extends previous work from a ram air-driven nacelle (a so-called “swirl tube”) to a “pumped” or “fan-driven” configuration and also includes an assessment of a pylon modification to assist a row of vanes in generating a swirling outflow in a more realistic engine environment. Computational fluid dynamics (CFD) simulations and aeroacoustic measurements in an anechoic nozzle test facility are performed to assess the swirl-flow-drag-noise relationship for EAB designs integrated into two NASA high-bypass ratio (HBPR), dual-stream nozzles. Aerodynamic designs have been generated at two levels of complexity: (1) a periodically spaced row of swirl vanes in the fan flowpath (the “simple” case), and (2) an asymmetric row of swirl vanes in conjunction with a deflected trailing edge pylon in a more realistic engine geometry (the “installed” case). CFD predictions and experimental measurements reveal that swirl angle, drag, and jet noise increase monotonically but approach noise simulations suggest that an optimal EAB deployment may be found by carefully trading any jet noise penalty with a trajectory or aerodynamic configuration change to reduce perceived noise on the ground. Constant speed, steep approach flyover noise predictions for a single-aisle, twin-engine tube-and-wing aircraft suggest a maximum reduction of 3 dB of peak tone-corrected perceived noise level (PNLT) and up to 1.8 dB effective perceived noise level (EPNL). Approximately 1 dB less maximum benefit on each metric is predicted for a next-generation hybrid wing/body aircraft in a similar scenario.

The high pressure (HP) turbine is subject to inlet flow nonuniformities resulting from the combustor. A lean-burn combustor tends to combine temperature variations with strong swirl and, although considerable research efforts have been made to study the effects of a circumferential temperature nonuniformity (hot-streak), there is relatively little known about the interaction between the two. This paper presents a numerical investigation of the transonic test HP stage MT1 behavior under the combined influence of the swirl and hot-streak. The in house Rolls-Royce HYDRA numerical computational fluid dynamics (CFD) suite is used for all the simulations of the present study. Baseline configurations with either hot-streak or swirl at the stage inlet are analyzed to assess the methodology and to identify reference performance parameters through comparisons with the experimental data. Extensive computational analyses are then carried out for the cases with hot-streak and swirl combined, including both the effects of the combustor-nozzle guide vane (NGV) clocking and the direction of the swirl. The present results for the combined hot-streak and swirl cases reveal distinctive radial migrations of hot fluid in the NGV and rotor passages with considerable impact on the aerothermal performance. It is illustrated that the blade heat transfer characteristics and their dependence on the clocking position can be strongly affected by the swirl direction. A further computational examination is carried out on the validity of a superposition of the influences of swirl and hot-streak. It shows that the blade heat transfer in a combined swirl and hot-streak case cannot be predicted by the superposition of each in isolation.

Previous experimental investigations revealed the existence of acoustic modes in the side cavities of a high-pressure centrifugal compressor. These modes were excited by pressure patterns which resulted from rotor/stator-interactions (often referred to as Tyler/Sofrin-modes). The acoustic modes were significantly influenced by the prevailing flow in the side cavities. The flow field in such rotor/stator-cavities is characterized by a high circumferential velocity component. The circumferential velocity of the flow and the phase velocity of the acoustic eigenmode superimpose each other, so that the frequencies of the acoustic eigenmodes with respect to the stator frame of reference follow from the sum of both velocities. In the previous study the circumferential velocity was estimated based on existing literature and the phase velocities of the acoustic modes were calculated via an acoustic modal analysis. Based on these results the rotational speeds of the compressor, where acoustic modes were excited in resonance, were determined. The present paper is based on these results and focuses on the influence of the swirling flow and the coupling of the excited acoustic modes between the two side cavities. Such a coupling has been predicted in previous numerical studies but no experimental evidence was available at that time. In this study the circumferential velocities of the flow are determined by measuring the actual radial pressure distribution in the side cavities and assuming radial equilibrium. The determined values are directly used for the prediction of the rotational speeds at resonance. The values for the rotational speeds at resonance predicted that way are compared to the resonance speeds found in the experiments. Further on, simultaneously measured pressure fluctuations in the shroud and hub side cavities with respect to the rotor frame of reference give evidence about the coupling of the acoustic modes between the two side cavities in case of resonance. If the experimentally determined swirling flow velocity is accounted for in the prediction of acoustic resonances, the calculated rotational speeds of resonance are in good agreement with the experimental findings in most cases. Neglecting the flow in the cavities, however, leads to large deviations between calculated and experimentally determined rotational speeds. Varying the operating point of the compressor results in changes of the circumferential velocities in the side cavities and, therefore, in changes of the rotational speeds of resonance. Contrary to the acoustic modes calculated via a finite element analysis by the authors of this paper in previous studies the excited acoustic modes in the experiments are mostly not coupled between the two side cavities, but are localized to one of both cavities. This finding is assumed to be caused by the flow field in the compressor.

The mainstream flow past the stationary nozzle guide vanes and rotating turbine blades in a gas turbine creates an unsteady nonaxisymmetric variation in pressure in the annulus, radially outward of the rim seal. The ingress and egress occur through those parts of the seal clearance where the external pressure is higher and lower, respectively, than that in the wheel-space; this nonaxisymmetric type of ingestion is referred to here as externally induced (EI) ingress. Another cause of ingress is that the rotating air inside the wheel-space creates a radial gradient of pressure so that the pressure inside the wheel-space can be less than that outside; this creates rotationally induced (RI) ingress, which—unlike EI ingress—can occur, even if the flow in the annulus is axisymmetric. Although the EI ingress is usually dominant in a turbine, there are conditions under which both EI and RI ingress are significant, these cases are referred to as combined ingress. In Part I of this two-part paper, the so-called orifice equations are derived for compressible and incompressible swirling flows, and the incompressible equations are solved analytically for the RI ingress. The resulting algebraic expressions show how the nondimensional ingress and egress vary with Θ 0 , which is the ratio of the flow rate of sealing air to the flow rate necessary to prevent ingress. It is shown that ε , the sealing effectiveness, depends principally on Θ 0 , and the predicted values of ε are in mainly in good agreement with the available experimental data.

Ingress of hot gas through the rim seals of gas turbines can be modeled theoretically using the so-called orifice equations. In Part I of this two-part paper, the orifice equations were derived for compressible and incompressible swirling flows, and the incompressible equations were solved for axisymmetric rotationally induced (RI) ingress. In Part II, the incompressible equations are solved for nonaxisymmetric externally induced (EI) ingress and for combined EI and RI ingress. The solutions show how the nondimensional ingress and egress flow rates vary with Θ 0 , the ratio of the flow rate of sealing air to the flow rate necessary to prevent ingress. For EI ingress, a “saw-tooth model” is used for the circumferential variation of pressure in the external annulus, and it is shown that ε , the sealing effectiveness, depends principally on Θ 0 ; the theoretical variation of ε with Θ 0 is similar to that found in Part I for RI ingress. For combined ingress, the solution of the orifice equations shows the transition from RI to EI ingress as the amplitude of the circumferential variation of pressure increases. The predicted values of ε for EI ingress are in good agreement with the available experimental data, but there are insufficient published data to validate the theory for combined ingress.

Experiments and numerical computations are performed to investigate the convective heat transfer characteristics of a gas turbine can combustor under cold flow conditions in a Reynolds number range between 50,000 and 500,000 with a characteristic swirl number of 0.7. It is observed that the flow field in the combustor is characterized by an expanding swirling flow, which impinges on the liner wall close to the inlet of the combustor. The impinging shear layer is responsible for the peak location of heat transfer augmentation. It is observed that as Reynolds number increases from 50,000 to 500,000, the peak heat transfer augmentation ratio (compared with fully developed pipe flow) reduces from 10.5 to 2.75. This is attributed to the reduction in normalized turbulent kinetic energy in the impinging shear layer, which is strongly dependent on the swirl number that remains constant at 0.7 with Reynolds number. Additionally, the peak location does not change with Reynolds number since the flow structure in the combustor is also a function of the swirl number. The size of the corner recirculation zone near the combustor liner remains the same for all Reynolds numbers and hence the location of shear layer impingement and peak augmentation does not change.

A novel air-brake concept for next-generation, low-noise civil aircraft is introduced. Deployment of such devices in clean airframe configuration can potentially reduce aircraft source noise and noise propagation to the ground. The generation of swirling outflow from a duct, such as an aircraft engine, is demonstrated to have high drag and low noise. The simplest configuration is a ram pressure-driven duct with stationary swirl vanes, a so-called swirl tube. A detailed aerodynamic design is performed using first principles based modeling and high-fidelity numerical simulations. The swirl-drag-noise relationship is quantified through scale-model aerodynamic and aeroacoustic wind tunnel tests. The maximum measured stable flow drag coefficient is 0.83 at exit swirl angles close to 50 deg. The acoustic signature, extrapolated to full-scale, is found to be well below the background noise of a well-populated area. Vortex breakdown is found to be the aerodynamically and acoustically limiting phenomenon, generating a white-noise signature that is about 15 dB louder than a stable swirling flow.

An advanced technique for establishing pressure boundary conditions in annular sector cascade experiments has been developed. This novel technique represents an improvement over previous methods and provides the first means by which annular sector boundary conditions that are representative of those which develop in an annular cascade can be established with a high degree of satisfaction. The technique will enable cascade designers to exploit the obvious advantages of annular sector cascade testing: the reduced cost of both facility manufacture and facility operation and the use of engine parts in place of two-dimensional counterparts. By employing an annular sector of deswirl vanes downstream of the annular sector of test vanes, the radial pressure gradient established in the swirling flow downstream of the test vanes is not disturbed. The deswirl vane exit flow—which has zero swirl velocity—can be exhausted without unsteadiness, and without the risk of separation, into a plenum at constant pressure. The pressure ratio across the annular sector of test vanes can be tuned by adjusting the throat area at the deswirl vane exit plane. Flow conditioning systems which utilize the Oxford deswirl vane technology have previously been used to set pressure boundary conditions downstream of fully annular cascades in both model and engine scale (the Isentropic Light Piston Facility at Farnborough) experimental research facilities ( Povey, T., Chana, K. S., Oldfield, M. L. G., Jones, T. V., and Owen, A. K., 2001, Proceedings of the ImechE Advances in Fluid Machinery Design Seminar, London, June 13 ; Povey, T., Chana, K. S., Jones, T. V., and Oldfield, M. L. G., 2003, Advances of CFD in Fluid Machinery Design, ImechE Professional Engineering, London, pp. 65–94 ). The deswirl vane is particularly suited to the control of highly whirling transonic flows. It has been demonstrated by direct comparison of aerodynamic measurements from fully annular and annular sector experiments that the use of a deswirl vane sector for flow conditioning at the exit of an annular sector cascade represents an attractive novel solution to the boundary condition problem. The annular sector technique is now described.

A systematic and rational methodology for the calculation of an equilibrium state from an initial nonuniform flow field, is presented with particular emphasis on the underlying assumptions and their attendant justifications. The imposed conservation criteria that are used to define a final state from the initial one depend on the coordinate system and flow configuration being analyzed. The imposition of these criteria for flow in parallel-walled annular ducts defines a state of complete (mechanical and thermal) equilibrium, for which radial profiles of the velocity components, static pressure and temperature assume a specific form for a perfect gas. A robust method for solving the system of equations that define the state of complete equilibrium (or mixed-out state) is presented using an efficient algorithm. The procedure is applied to the swirling flow exiting an isolated transonic compressor, and comparisons are made with other available methods of averaging flow fields.

Experimental measurements from a new single stage turbine are presented. The turbine has 26 vanes and 59 rotating blades with a design point stage expansion ratio of 2.5 and vane exit Mach number of 0.96. A variable sealing flow is supplied to the disk cavity upstream of the rotor and then enters the annulus through a simple axial clearance seal situated on the hub between the stator and rotor. Measurements at the annulus hub wall just downstream of the vanes show the degree of circumferential pressure variation. Further pressure measurements in the disk cavity indicate the strength of the swirling flow in the cavity, and show the effects of mainstream gas ingestion at low sealing flows. Ingestion is further quantified through seeding of the sealing air with nitrous oxide or carbon dioxide and measurement of gas concentrations in the cavity. Interpretation of the measurements is aided by steady and unsteady computational fluid dynamics solutions, and comparison with an elementary model of ingestion.

In most gas turbines, blade-cooling air is supplied from stationary preswirl nozzles that swirl the air in the direction of rotation of the turbine disk. In the “cover-plate” system, the preswirl nozzles are located radially inward of the blade-cooling holes in the disk, and the swirling airflows radially outward in the cavity between the disk and a cover-plate attached to it. In this combined computational and experimental paper, an axisymmetric elliptic solver, incorporating the Launder–Sharma and the Morse low-Reynolds-number k–ε turbulence models, is used to compute the flow and heat transfer. The computed Nusselt numbers for the heated “turbine disk” are compared with measured values obtained from a rotating-disk rig. Comparisons are presented, for a wide range of coolant flow rates, for rotational Reynolds numbers in the range 0.5 X 10 6 to 1.5 X 10 6 , and for 0.9 < β p < 3.1, where β p is the preswirl ratio (or ratio of the tangential component of velocity of the cooling air at inlet to the system to that of the disk). Agreement between the computed and measured Nusselt numbers is reasonably good, particularly at the larger Reynolds numbers. A simplified numerical simulation is also conducted to show the effect of the swirl ratio and the other flow parameters on the flow and heat transfer in the cover-plate system.

The swirling motion of the shroud-to-housing leakage flow in pumps is known to have an adverse impact on the impeller rotordynamic stability. Swirl brakes, under such circumstances, would enhance the stability margin by reducing or, ideally, eliminating, the prerotation at the leakage passage inlet station. The numerical analysis outlined in this paper provides a quantitative means of predicting the effectiveness of such devices. The computed results also illustrate the mechanism with which the fluid/rotor interaction, with the aid of a typical brake, is altered towards relative overall rotordynamic stability. This is done through a comparative examination of the pressure perturbation distribution over the shroud surface for a wide range of backward and forward impeller-whirl frequencies. The conclusions in this study are consistent with recent experimental findings and have important design implications.