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This paper examines the applicability of a triple layer of thermochromic liquid crystals on the evaluation of local heat transfer coefficients using the transient liquid crystal measurement technique. The experiments were carried out in a narrow impingement cooling channel typically found in turbine airfoils. Narrow-band liquid crystals of 35°C, 38°C and 41°C were individually painted on the target plate of the cooling cavity and the overall paint thickness was accurately measured. Therefore, the 1D transient heat conduction equation was implicitly solved for each individual liquid crystal layer on its realistic depth on the painted surface. As a result, local heat transfer coefficients were calculated three times for the same location improving the measurement accuracy, especially at regions where liquid crystal detection times are too short or wide. The results indicate that if multiple liquid crystals layers are used and the paint thickness is not considered, heat transfer coefficients can be underestimated.

A detailed study of the interaction between high-speed gas flows and surface dielectric barrier discharges (DBD) is presented. In the present paper, it is demonstrated that a DBD can be sustained in transonic airflows, up to isentropic Mach numbers of 1.1. The plasma is characterized electrically, as well as optically with a CCD camera and a photomultiplier tube. Different airflow velocities, plasma excitation frequencies and voltages are investigated. The airflow has a significant influence on the plasma characteristics: the glow component is reduced, the discharge becomes more filamentary and most importantly, the light emission duration from individual microdischarges is reduced by more than a factor of ten at high flow velocities. Large edge effects play a key role in the interaction between the flow and the plasma. These results offer new perspectives for the use of dielectric barrier discharges in transonic and supersonic gas flows and their applications to airflow control and to plasma-assisted combustion.

This paper examines the applicability of a triple layer of thermochromic liquid crystals (TLCs) for the determination of local heat transfer coefficients using the transient liquid crystal (LC) technique. The experiments were carried out in a narrow impingement channel, typically used for turbine blade cooling applications. Three types of narrow bandwidth LCs (1 °C range) of 35 °C, 38 °C, and 41 °C were individually painted on the target plate of the cooling cavity and the overall paint thickness was accurately determined with an integral coating thickness gauge. The 1D transient heat conduction equation is then implicitly solved for each individual TLC layer on its realistic depth on the painted surface. Local heat transfer coefficients are therefore calculated three times for the same location in the flow improving the measurement accuracy, especially at regions where the LC detection times are too short (stagnation points) or too long (wall-jet regions). The results indicate that if multiple LC layers are used and the paint thickness is not considered, the heat transfer coefficients can be significantly underestimated.

The transient liquid crystal technique is currently widely used for measuring the heat transfer characteristics in gas turbine applications. Usually, the assumption is made that the wall of the test model can be treated as a flat and semi-infinite solid. This assumption is correct as long as the penetration depth of the heat compared to the thickness of the wall and to the radius of curvature is small. However, those two assumptions are not always respected for measurements near the leading edge of a blade. This paper presents a rigorous treatment of the curvature and finite wall thickness effects. The unsteady heat transfer for a hollow cylinder has been investigated analytically and a data-reduction method, taking into account curvature and finite wall-thickness effects has been developed. Experimental tests made on hollow cylinder models have been evaluated using the new reduction method as well as the traditional semi-infinite flat-plate approach and a third method that approximately accounts for curvature effects. It has been found that curvature and finite thickness of the wall have, in some cases, a significant influence on the obtained heat transfer coefficient. The parameters influencing the accuracy of the semi-infinite flat-plate model and the approximate curvature correction are determined and the domains of validity are represented.

The ability of nanosecond pulsed DBD actuators to control flow separation was investigated on a NACA 0015 profile for velocities reaching Re = 230,000 (24 m/s). The optimal location for the actuator was determined from oil flow experiments. Moderate voltage levels were applied (3 and 6 kV) and the actuator was operated at frequencies ranging from 0.058 to 10 kHz, in pulse and burst modes. The peak effectiveness of the actuator occurred at reduced frequency values of around 1. Plasma influence was observed at all tested angles of attack (up to 26°), and the stall angle was delayed by 8°.

The stabilizing effect of grouping rotor blades in pairs has been assessed both, numerically and experimentally. The bending and torsion modes of a low aspect ratio high speed turbine cascade tested in the non-rotating test facility at EPFL (Ecole Polytechnique Fédérale de Lausanne) have been chosen as the case study. The controlled vibration of 20 blades in travelling wave form was performed by means of an electromagnetic excitation system, enabling the adjustement of the vibration amplitude and inter blade phase at a given frequency. Unsteady pressure transducers located along the blade mid-section were used to obtain the modulus and phase of the unsteady pressure caused by the airfoil motion. The stabilizing effect of the torsion mode was clearly observed both in the experiments and the simulations, however the effect of grouping the blades in pairs in the minimum damping at the tested frequency was marginal in the bending mode. A numerical tool was validated using the available experimental data and then used to extend the results at lower and more relevant reduced frequencies. It is shown that the stabilizing effect exists for the bending and torsion modes in the frequency range typical of low-pressure turbines. It is concluded that the stabilizing effect of this configuration is due to the shielding effect of the pressure side of the airfoil that defines the passage of the pair on the suction side of the same passage, since the relative motion between both is null. This effect is observed both in the experiments and simulations.

This paper presents and discusses unsteady experimental results of controlled vibration measurements of blade-cluster and single-blade test cases, all subjected to the same subsonic ow conditions. The experiments were performed in a non-rotating annular test facility with the aim to investigate the influence of cluster vibrations on the aerodynamic damping of turbine rotor blades. A frequently-used blade package type was selected: a two-blade-cluster in mixed mode, combing in a alternating manner a bending and a torsion vibration mode. In addition, single-blade test cases were investigated serving as reference for the cluster case. The actuation of the 20 turbine blades was obtained by an electro-magnetic excitation system, enabling individually the adjustment of vibration amplitude and phase for each blade at one vibration frequency. The unsteady measurements obtained include time-synchronous unsteady blade surface pressures and blade displacements resulting in local and global aerodynamic blade damping. A comparison of the cluster- and the single-blade-test cases was performed and differences were identified and analyzed.

The application of nanosecond pulse driven surface dielectric barrier discharge (DBD) actuators as control devices on airfoils and turbomachinery blades in transonic flow is investigated experimentally. Images acquired with a short exposure ICCD camera document the spatial temporal discharge development in both absence and presence of flow as a means of examining the effect of the flow on the actuator. In order to visualize the interaction of the DBD actuator with the flow Schlieren images were acquired. Furthermore a power supply capable of generating the required voltage pulses was built and characterized. The impact of the pulse transition time on the discharge development was also investigated.

There were analyzed the conditions under which the basal glycemia and/or glucose tolerance curve is not contaminated by stress induced changes in glycide metabolism. When the basal glycemia was monitored,blood was sampled to heparinized capillaries from retrobulbar plexus under the light ether anaesthesia or by decapitation without narcosis. The animal represented the control for itself. No differences was found in basal glycemia under the two mentioned blood sampling. In the second series of experiments glycemia was monitored in the time schedule which is used in glucose tolerance test, i.e., blood sampling was performed 30,60,120 and 180 min after the first sampling from retrobulbar plexus either under light ether anaesthesia (in 14 h starvated rats or in rats with free access to diet) or under Nembutal anaesthesia (in 14 h starvated rats). No differences were found in glycemia when two types of narcosis is compared. No signs of augmentation were detected. In the last series when blood sampling was taken in two sec intervals, time dependent augmentation of stress glucose elevation was found. The augmentation was more expressed in the rats with free access to diet than in starvated animals.

A novel transient measurement technique has been developed for determining the heat transfer characteristics in the presence of film cooling (heat transfer coefficient and adiabatic film-cooling effectiveness). The method is based on a transient heater foil technique, where a non-homogeneous surface heat flux is applied to the test surface. A regression analysis of multiple transient liquid crystal experiments is used to obtain the heat transfer characteristics. The method introduced here has the advantage that the (often not known) heat flux distribution at the surface is not needed for the analysis of the measured data. The method is used to study the influence of several heater foil configurations on a flat plate with film cooling, elucidating the effect of different thermal boundary conditions on film-cooling performance. The obtained data is also compared to results presented in literature and good agreement is found.

This paper describes a portion of a research effort to eliminate occurrences of flutter of low pressure turbine blades in aircraft engines. Fundamental unsteady aerodynamic experiments in an annular cascade plus correlating analyses are conducted to improve the understanding of the flutter mechanism in these blades and to identify the key flutter parameters. Mode shape is identified as the most important contributor to determining the stability of a blade design, while reduced frequency has a secondary influence. Low pressure turbine flutter is shown to be a linear, inviscid phenomenon.

Calculated results for tip flow around two different blade configurations are presented and compared with experimental data. The first configuration (case number 1) is a flat-plate profile tested in a linear transonic tunnel — the profile is an idealized representation of the aft-section of some highly curved turbine blades. The second configuration (case number 2) originates from the outer profile on the last-stage-blade of a steam turbine, however it is also reminiscient of a section from a turbine blade with supersonic exit flow. This configuration was tested in an annular cascade at Mach numbers representative of engine operating conditions. The computed results were obtained using a parallel 3D unstructured Navier-Stokes code. The code runs on a work-station cluster, as well as being optimized for the 256 processor Cray T3D at EPFL: the code is capable of gigaflop performance using more than 3 million cells — adaptive mesh refinement thus allows enhanced resolution within the tip gap region. For each configuration we have calculated two Runs. In both cases, Run-1 is similar to the experimental conditions, so that direct comparison between measured and calculated results is possible. With case number 1/Run-2 we re-calculated the flow without imposing a prescribed inflow boundary-layer along the sidewall. Comparison between the two runs helped reveal how free-stream total pressure can establish itself within the tip gap region. For the second configuration — in the annular cascade — we were interested in observing the influence of relative movement between the blade tip and adjacent sidewall. Hence for case number 2/Run-2 we imposed a circumferential velocity on the adjacent sidewall. This modified the effective sidewall boundary-layer and had a noticeable influence on the development of the tip-leakage flow.

This paper presents heat transfer measurements on a turbine airfoil in a linear cascade at various exit Reynolds and Mach numbers ranging from 3.2e5 to 1.6e6 and 0.2 to 0.8, respectively, which have been conducted with the transient liquid crystal technique. Two series were performed at turbulence intensities of 5.5% and 10%, the latter being created by a squared-bar mesh placed 10 meshsizes upstream of the turbine airfoils. While normally polished liquid crystals were used additional experiments were done at the high turbulence intensity with naturally rough liquid crystals. All measurements indicate a gradual increase in heat transfer and an upstream shift of the laminar-to-turbulent transition with increasing Reynolds number and turbulence intensity. The leading edge heat transfer agrees well with correlations if the turbulence length scale is taken into account. The measurements conducted with rough liquid crystals show an earlier transition on the suction side. Calculations with a two-dimensional boundary layer code agree well with the measurements.

Keywords: Unsteady Flows ; Turbomachinery ; GTT ; LTT Reference LTT-CONF-1989-004 Record created on 2007-04-18, modified on 2016-08-08

A computational study has been conducted to determine the variation of device drag with profile shape and angle of attack for aerofoil boundary-layer manipulators (LEBUs) operating at high subsonic Mach numbers. Calculations have been made at a free-stream Mach number of 0.80 for both symmetrical and asymmetrical NACA-00xx and 44xx series devices including an inverted cambered NACA-4409 profile. The LEBUs considered were located in a turbulent boundary-layer at a mid-chord heighth equivalent to 0.67δ from the wall. The present investigations sought to confirm suggestions based upon experimental observations that there may be some advantage in replacing a symmetrical device by an inverted asymmetrical profile form.

A quasi three-dimensional method has been developed to calculate the transonic flow in a compressor rotor. The method allows changes in the stream tube thickness and mean radius evolution through the rotor. This blade-to-blade method uses original concepts in order to be rapid (≈ 3–4 sec CPU on a HP 715 workstation), to find the precise location of the shocks (and thus the flow values on them) and the flow distribution around the profile, and finally to cover a large field of “Off-Design” operating points. Thus, it can be used in the process of optimization of a transonic compressor. The supersonic flow at the inlet as well as the oblique shock configuration inside the cascade are calculated by methods based on the characteristics theory. Conditions with attached shock wave (and thus unique incidence angle) or with detached shock wave can be calculated. Quasi-3D equations were developed (Bölcs and Tsamourtzis, 1991). The subsonic flow is calculated by a streamline curvature method, with some new concepts, and finally the shock in the inter-blade canal is found by a combination of the supersonic and subsonic flow values. This method was combined, in a S1 - S2 calculation, with a throughflow method (Sayari and Bölcs, 1995) in order to be validated by comparing the results with the measurements provided by NASA - Lewis Research Center, on a transonic compressor rotor (Strazisar T.; 1994).

Background: Gestational trophoblastic neoplasia (GTN) is a curable disease that involves the development of malignant tumor in the woman after a normal or molar pregnancy. The position of surgery in GTN is not properly specified and is changing due to new chemotherapy protocols. However, the role of surgery is highlighted in chemotherapy-resistant GTN. Other indications of surgery in trophoblastic diseases are drug toxicity and uterine perforation. Based on the fact that most women in certain age tend to preserve fertility, this study reported 4 cases of successful treatment after fertility sparing surgery. Case: A hospital-based case-report study was carried out to investigate the role of surgery in 4 patients with GTN. In this study, acute complications, such as intra-abdominal bleeding and liver dysfunction due to chemotherapy occurred in some patients. Surgery was performed and all cases underwent localized tumor removal while preserving the uterus. No hysterectomy surgery was performed. Conclusion: Surgery is supposed in specific cases of GTN, who desire preserving fertility.

This paper presents a new base pressure correlation for transonic and supersonic flow conditions obtained from systematic measurements in linear and annular turbine cascades and by considering the physics of the flow in choked cascades. The correlation is valid in the range from near sonic outlet flow velocity to limit loading. Most of the presently used base pressure correlations show deviations from experimental results for several profile types. One of the reasons is the uncertainty in the evaluation of the downstream pressure in the cascade used as a reference pressure. In the model proposed here, another reference value is defined instead, namely the Mach number of the recompression point at the end of the triangular base pressure region. This recompression Mach number is obtained in an indirect way from the measured pressure distribution on the profile, i. e., by applying characteristic relations. The advantage of this base pressure correlation is that a unique relationship between the base pressure and an arbitrarily defined reference pressure (for example, the back pressure) can be established.

Detailed surface measurements of the heat transfer coefficient and the film cooling effectiveness by application of the transient liquid crystal method were carried out on a heavily film cooled nozzle guide vane (NGV) in a linear cascade wind tunnel at the EPFL as part of the European Research Project TATEF2 (Turbine Aero-Thermal External Flows 2). The external cooling setup included a showerhead cooling scheme and suction and pressure side of the airfoil several rows of fan-shaped cooling holes. By testing two different cooling flow rates at a NGV exit Reynolds number of 1.46E+06, detailed aerodynamic and heat transfer measurement data were obtained that can be used for validation of numerical codes and design tools for cooled airfoils. The data include the NGV surface static pressure distribution and wall heat transfer and film cooling effectiveness obtained by application of the transient liquid crystal technique. An engine representative density ratio between the coolant and the external hot gas flow was achieved by using CO2 as coolant gas. For the coupled simulation of internal cooling and external flow the numerical model was composed of the cooling air feeding the internal plenum, the cooling holes, and the outer external flow domain. An unstructured mesh was generated for the simulations by applying two different commercial CFD codes (Fluent and CFX). Identical boundary conditions were chosen in order to allow for a direct comparison of both codes. The computations were carried in two ways, first using a built-in transition model and second by imposing fully turbulent flow starting at the leading edge. For both codes the same built-in turbulence models were applied. The computations were set up to solve for the aerodynamic flow quantities both within and around the test model and for the thermal quantities on the vane surface, i.e. heat transfer coefficient and film cooling effectiveness. The computational results from the two codes are compared and validated against the results from the experiments. The numerical results were able to confirm a suspicion that the cross flow in the feeding plenum causes an observed non-symmetry of the measured film cooling effectiveness at the outlet of some cooling holes.

Integrally cast turbine airfoils with wall-integrated cooling cavities are greatly applicable in modern turbines providing enhanced heat exchange capabilities compared to conventional cooling passages. In such arrangements, narrow impingement channels can be formed where the generated crossflow is an important design parameter for the achievement of the desired cooling efficiency. In this study, a regulation of the generated crossflow for a narrow impingement channel consisting of a single row of five inline jets is obtained by varying the width of the channel in the streamwise direction. A divergent impingement channel is therefore investigated and compared to a uniform channel of the same open area ratio. Flow field and wall heat transfer experiments are carried out at engine representative Reynolds numbers using particle image velocimetry (PIV) and liquid crystal thermography (LCT). The PIV measurements are taken at planes normal to the target wall along the centerline for each individual jet, providing quantitative flow visualization of jet and crossflow interactions. The heat transfer distributions on the target plate of the channels are evaluated with transient techniques and a multilayer of liquid crystals (LCs). Effects of channel divergence are investigated combining both the heat transfer and flow field measurements. The applicability of existing heat transfer correlations for uniform jet arrays to divergent geometries is also discussed.

An experimental study on a gas turbine cascade operating under transonic flow conditions is presented. The flow is compared for airfoil shapes corresponding to the design geometry and the geometry taken from a rotor blade, in an industrial gas turbine burning heavy oil, after a few thousand hours of operation. Steady-state data have been obtained in a linear cascade over a range of isentropic exit Mach numbers from 0.6 to 1.6. The flow field was determined by static pressure measurements on the side walls up- and downstream of the cascade, on one side wall in the blade passage, and on the blade surface at midspan. Furthermore, the flow was visualized by the methods of Schlieren and laser holography. The results show that the choked flow conditions are reached at different steady-state isentropic outlet Mach numbers for the two blade shapes. The deposit, typical for a gas turbine, does not however significantly modify the boundary layer separation point. The flow visualization indicates that the shock wave fluctuations have not been significantly influenced by the important roughness and thickness of the deposit. The experimental results on the two cascades are also compared with two-dimensional time-marching calculations after Denton. In the subsonic regime, good agreement was found for the “clean” blade. For the profile with deposit, the flow cannot be correctly predicted by the time-marching calculation, even in subsonic flow condition. The sonic line calculated by the numerical model under transonic outlet conditions (0.9 < M2S < 1.20) does not agree with the laser holography measurements for either of the two cascades.

Detailed surface measurements of the thermal performance of a film cooling system have been performed on the endwall of a nozzle guide vane (NGV) mounted in a linear cascade facility at EPFL. An external cooling scheme including several rows of fan-shaped and cylindrical cooling holes has been designed. By testing different cooling flow rates at a NGV exit Reynolds number of 1.7E+06 and Mach number of 0.88, detailed aerodynamic and heat transfer values were obtained destined to assess the design tools for film cooled platforms. The surface static pressure distribution and the film cooling effectiveness on the endwall surface have been experimentally determined. The measurements were obtained applying the pressure sensitive paint technique measuring the coolant gas concentration. An engine representative density ratio between the coolant and the external hot gas flow was achieved by the injection of CO2. The working conditions of the test case similar to realistic engine conditions allow for the validation of in-house CFD codes and the investigation of the reliability of modern commercial tools in such a complex cooling system. The numerical campaign has been performed on the same numerical grid, using the commercial codes FLUENT and CFX, used by EPFL and MTU respectively. A detailed analysis of the grid effects on the obtained results has been previously realised as well as the study of the influence of the modelling approximations. Three cooling mass flows have been simulated and the performance parameters of the film cooling system have been compared to the experimentally obtained data. Special emphasis has been put on the jet penetration effects and on the interaction of secondary flows with the coolant flow. The experimental and numerical efforts were part of the EU funded research project TATEF2 (Turbine Aero-Thermal External Flows 2).

This paper presents experimental and theoretical results for turbine cascades performing harmonic oscillations in transonic flow at design and off-design conditions. The experimental investigations were performed in an annular test facility where unsteady blade pressures were measured in two different test cascades, one operating at the nominal inlet flow angle, the other at an incidence angle exceeding the normal value by more than 20 degrees. The corresponding theoretical results were computed with a 2D Euler code which makes use of flux vector splitting in combination with a time-dependent grid generation. The present data were all obtained for tuned bending modes where the blades performed heaving oscillations with the same frequency and amplitude, but with a constant interblade phase angle. For the cascade operating at design conditions, the steady flow was purely subsonic. The other test cascade was run in transonic flow, and a normal shock appeared on the rear part of the blade’s suction surface. It was found that measured unsteady pressure and damping coefficients are well reproduced by the computed results for the first test cascade. In the case of steady off-design flow (the second test cascade), significant differences between experimental and theoretical results are observed.

Non-synchronous excitation under low volume operation is a major risk to the mechanical integrity of last stage moving blades (LSMBs) in low-pressure (LP) steam turbines. These vibrations are often induced by a rotating aerodynamic instability similar to rotating stall in compressors. Currently extensive validation of new blade designs is required to clarify whether they are subjected to the risk of not admissible blade vibration. Such tests are usually performed at the end of a blade development project. If resonance occurs a costly redesign is required, which may also lead to a reduction of performance. It is therefore of great interest to be able to predict correctly the unsteady flow phenomena and their effects. Detailed unsteady pressure measurements have been performed in a single stage model steam turbine operated with air under ventilation conditions. 3D CFD has been applied to simulate the unsteady flow in the air model turbine. It has been shown that the simulation reproduces well the characteristics of the phenomena observed in the tests. This methodology has been transferred to more realistic steam turbine multi stage environment. The numerical results have been validated with measurement data from a multi stage model LP steam turbine operated with steam. Measurement and numerical simulation show agreement with respect to the global flow field, the number of stall cells and the intensity of the rotating excitation mechanism. Furthermore, the air model turbine and model steam turbine numerical and measurement results are compared. It is demonstrated that the air model turbine is a suitable vehicle to investigate the unsteady effects found in a steam turbine.

A new quasi-3D characteristics method has been developed to calculate the supersonic flow in a compressor rotor on an axisymmetric stream surface with variable radius and stream tube thickness. The method is valid for supersonic inlet flow with an attached shock wave at the blade leading edge where the inlet Mach number and flow angle are related by the unique incidence condition. On the basis of the quasi-3D characteristics method, a simple quasi-3D method has also been derived that calculates the unique incidence condition and the Mach number distribution on the suction side of the profile for a given stream surface and stream tube thickness. The reliability of the simple method is established by comparing results with those obtained by Denton’s time-marching method as well as those from the quasi-3D characteristics method. Systematic calculations performed using the simple quasi-3D method show the effects of varying the geometrical parameters of the cascade for different stream tube thickness and stream surface radius evolutions through the rotor.

The prediction of flutter and forced response at normal flow conditions has become a standard procedure during the design of compressor airfoils. But at severe off-design conditions, the flow field becomes very complex, especially during the surge blow-down phase where reversed flow conditions occur. The correct prediction of the unsteady pressures and the resulting aerodynamic excitation or damping at these conditions remains an extremely challenging task. In the first part of the paper, basic investigations for these flow conditions are presented. Aeroelastic calculations during compressor surge are shown in the second part. Experimental investigations were performed in the Annular Test Facility for non-rotating cascades at EPF Lausanne. The test cascade was exposed to flow conditions as expected during the surge blow-down phase which is characterized by large separation regions. Measurements of the steady-state flow conditions on the blade surface, at the outer wall, upstream and downstream of the cascade provided detailed information about the steady flow conditions. The cascade was then subjected to controlled vibration of the blades with constant amplitudes and inter-blade phase angles. Unsteady pressure measurements on the blade surface and at the casing wall provided information about the resulting unsteady flow conditions. Analytical CFD calculations were performed. The steady flow field was calculated using a RANS code. Based on the steady-state flow field, unsteady calculations applying a linearized code were carried out. The agreement between measurements and calculations shows that the steady flow as well as the unsteady flow phenomena can be predicted quantitatively. In addition, knowing the blade vibration mode shape, which in this case is a torsion mode, the aerodynamic damping can be determined for the corresponding flow conditions.

Non-synchronous excitation under low volume operation is a major risk to the mechanical integrity of last stage moving blades (LSMBs) in low-pressure (LP) steam turbines. These vibrations are often induced by a rotating aerodynamic instability similar to rotating stall in compressors. Unsteady computational fluid dynamics (CFD) has been applied to simulate the rotating stall phenomenon in two model turbines. It is shown that the investigated flow field presents a challenge to conventional Reynolds-averaged Navier–Stokes equations simulations. The modelling has been enhanced by applying scale-resolving turbulence modelling, which can simulate large-scale turbulent fluctuations. With this type of simulation a qualitative and quantitative agreement between CFD and measurement for the unsteady and time averaged flow field has been achieved. The results of the numerical investigation allow for a detailed insight into the dynamic flow field and reveal information on the nature of the excitation mechanism. It is concluded that the CFD approach developed can be used to assess LSMB blade designs prior to model turbine tests to check whether they are subjected to vibration under LVF caused by rotating stall.

A two-dimensional section of a gas turbine cascade has been investigated experimentally in an annular non-rotating cascade facility as regards to its steady-state and time-dependent aerodynamic characteristics at off-design flow conditions. The blades vibrated in the first traveling wave bending mode. Steady-state and unsteady data were obtained for an off-design incidence angle of about 22° and for an isentropic outlet Mach number of M2s=1.19. At this flow condition, a separation bubble was present on the suction surface close to the leading edge. A shock appeared at trans- and supersonic outlet flow conditions on the suction surface. The data showed high unsteady loads close to the leading edge and in the shock region. It was found that the steady and the unsteady pressures in the shock region on the blade surface seemed to be very sensitive to small changes in the flow conditions. The periodicity and repetitivity of the steady and the unsteady pressures (σ=180°) was checked at several circumferential channel positions. This was done to figure out to which extend test data obtained in an annular ring channel can serve as a basis for the comparison with numerically obtained data. The aim of this paper is to show where problems may arise when comparing calculated results with test data.

The diversification of power generation methods within existing power networks has increased the requirement for operational flexibility of plants employing steam turbines. This has led to the situation where steam turbines may operate at very low volume flow conditions for extended periods of time. Under operating conditions where the volume flow through the last stage moving blades (LSMBs) of a low-pressure (LP) steam turbine falls below a certain limit, energy is returned to the working fluid rather than being extracted. This so-called “ventilation” phenomenon produces non-synchronous aerodynamic excitation, which has the potential to lead to high dynamic blade loading. The aerodynamic excitation is often the result of a rotating phenomenon, with similarities to rotating stall, which is well known in compressors. Detailed unsteady pressure measurements have been performed in a single stage model steam turbine operated with air under ventilation conditions. Detailed analysis revealed that the rotating excitation mechanism observed in operating steam turbines, is reproduced in the model turbine. 3D CFD has been applied to simulate the unsteady flow in the air model turbine. The numerical model consists of the single stage modeled as a full annulus, as well as the axial-radial diffuser. An unsteady CFD analysis has been performed for sufficient rotor revolutions such that the flow is globally periodic. It has been shown that the simulation reproduces well the characteristics of the phenomenon observed in the tests. The detailed insight into the flow field allows the drawing of conclusions as to the nature of the excitation mechanism. One result is that the LSMB tip clearance flow is found to have very little or no effect on the characteristics of mechanism for the case studied.

Keywords: Heat Transfer ; Turbomachinery ; GTT ; LTT Reference LTT-CONF-2005-002View record in Web of Science URL: http://www.euroturbo.eu/ Record created on 2007-04-18, modified on 2017-05-10

A unique non-rotating aerodynamic test facility is presented. It is designed for steady and unsteady transonic flow investigations in annular cascades. The fundamental advantage of this vehicle design, compared to a linear cascade facility, is the absence of boundary disturbances in the cascade and the self-adjustment of the flow periodicity. The downstream flow conditions can therefore be adjusted by the backpressure. The nozzle consists of a radial-axial section, in which spiral axisymetric flow is generated. With the adjustment of the total pressure and swirl angles in two separate distributors, a continuous variation of Mach number between 0 Further regulation of the flow quantities in the two separate distributors and boundary layer suction permits a continuous spanwise variation of the velocity and flow angle in the test section.