Exit Flow Field Research Articles

Influence of Upstream and Downstream Compressor Stators on Rotor Exit Flow Field

Measurements acquired at the rotor exit plane illuminate the interaction of the rotor with the upstream vane row and the downstream vane row. The relative phase of the upstream and downstream vane rows is adjusted using vane clocking so that the effect of the upstream propagating potential field from the downstream stator can be distinguished from the effects associated with the wakes shed from the upstream stator. Unsteady absolute flow angle information shows that the downstream potential field causes the absolute flow angle to increase in the vicinity of the downstream stator leading edge. The presence of Stator 1 wake is also detected at this measurement plane using unsteady total pressure data. The rotor wakes are measured at different circumferential locations across the vane passage, and the influence of Stator 1 wake on the suction side of the rotor wake is evident. Also, the influence of the downstream stator is detected on the pressure side of the rotor wake for a particular clocking configuration. Understanding the role of the surrounding vane rows on rotor wake development will lead to improved comparison between experimental data and results from computational models.

Experimental research and numerical simulation on impeller exit flow field of bulb tubular pump

Three dimensional Laser Doppler Velocimetry (3D-LDV) technique was used to measure the impeller exit flow field of bulb tubular pump, the average velocity distribution of the exit section was obtained, and the energy performance of bulb tubular pump was obtained by model test. The flow field of tubular pump also simulated by CFD based on Navier-Stokes equations and RNG k-ϵ model. The calculated performance curve was consistent with the model test results, and the calculated velocities of impeller exit compared with the LDV measurement results, it was shown that the distribution of circumferential velocity and axial velocity were close to each other, but the radial velocity deviation between the LDV results and the CFD ones was obviously, the reasons of deviation were analyzed. The experimental results can provide the references for the impeller and diffuser design of tubular pumps, it also can improve the accuracy of calculation and perfect simulation.

Impact of Severe Temperature Distortion on Turbine Efficiency

This paper presents an experimental and computational study of the effect of severe inlet temperature distortion (hot streaks) on the efficiency of the MT1 HP turbine, which is a highly-loaded unshrouded transonic design. The experiments were performed in the Oxford Turbine Research Facility (OTRF) (formerly the TTF at QinetiQ Farnborough): an engine scale, short duration, rotating transonic facility, in which M, Re, Tgas/Twall and N/T01 are matched to engine conditions. The research formed part of the EU Turbine Aero-Thermal External Flows (TATEF II) program. An advanced second generation temperature distortion simulator was developed for this investigation, which allows both radial and circumferential temperature profiles to be simulated. A pronounced profile was used for this study. The system was novel in that it was designed to be compatible with an efficiency measurement system which was also developed for this study. To achieve low uncertainty (bias and precision errors of approximately 1.5% and 0.2% respectively, to 95% confidence), the mass flow rate of the hot and cold streams used to simulate temperature distortion were independently metered upstream of the turbine nozzle using traceable measurement techniques. Turbine power was measured directly with an accurate torque transducer. The efficiency of the test turbine was evaluated experimentally for a uniform inlet temperature condition, and with pronounced temperature distortion. Mechanisms for observed changes in the turbine exit flow field and efficiency are discussed. The data are compared in terms of flow structure to full stage computational fluid dynamics (CFD) performed using the Rolls Royce Hydra code.

Fan Root Aerodynamics for Large Bypass Gas Turbine Engines: Influence on the Engine Performance and 3D Design

The exit flow field of the fan root of large turbofan engines defines the inlet conditions to the core compressor. This in turn could have significant impact to the performance of the core compressor. This study is aimed to resolve two related issues concerning the impact of the fan root flow on the core compressor performance: to establish the effect of an increased loss at the inlet on the engine specific fuel consumption (SFC) and to assess the effect of the radial distribution of the fan root flow on the engine performance. With understanding of these issues, the geometric parameters and design details which can produce a more uniform core flow at the exit of the fan stage module can be identified. The fan root flow is analyzed with methods of different complexity and fidelity. A simple cycle analysis is used to assess the impact on engine SFC of a stagnation pressure deficit at the fan root; a throughflow code is used for the preliminary study of the curvature effect of the root flow path, and 3D RANS CFD calculations are then used to simulate the flow path from the inlet of the fan to the first stage of the core compressor. The adequacy of the application of the numerical code in this case has been assessed and confirmed by the comparison with the experimental data for two rig configurations. The results of this study show that the flow at the fan hub region is very complex and dominated by 3D effects. The interaction of the secondary flow with real geometries, such as leakage flows, is found to have a strong detrimental effect on the core performance. The curvature of the hub end wall is a key parameter controlling the fan root flow topology; it influences the strength of the secondary flow, the spanwise distribution of the flow, and its sensitivity to leakage flow. With this understanding, it is possible to redesign of the fan hub flow path to reduce the loss generation by a significant amount.

Investigation of Trailing Edge Cooling Concepts in a High Pressure Turbine Cascade—Aerodynamic Experiments and Loss Analysis

As part of a European research project, the aerodynamic and thermodynamic performance of a high pressure turbine cascade with different trailing edge cooling configurations was investigated in the wind tunnel for linear cascades at DLR in Go¨ttingen. A transonic rotor profile with a relative thick trailing edge was chosen for the experiments. Three trailing edge cooling configurations were applied, first central trailing edge ejection, second a trailing edge shape with a pressure side cut-back and slot equipped with a diffuser rib array, and third pressure side film cooling through a row of cylindrical holes. For comparison, aerodynamic investigations on a reference cascade with solid blades (no cooling holes or slots) were performed. The experiments covered the subsonic, transonic and supersonic exit Mach number range of the cascade while varying cooling mass flow ratios up to 2 %. This paper analyzes the effect of coolant ejection on the airfoil losses. Emphasis was given on separating the different loss contributions due to shocks, pressure, and suction side boundary layer, trailing edge, and mixing of the coolant flow. Employed measurement techniques are schlieren visualization, blade surface pressure measurements, and traverses by pneumatic probes in the cascade exit flow field and around the trailing edge. The results show that central trailing edge ejection significantly reduces the mixing losses and therefore decreases the overall loss. Higher loss levels are obtained when applying the configurations with pressure side blowing. In particular, the cut-back geometry reveals strong mixing losses due to the low momentum coolant fluid, which is decelerated by the diffuser rib array inside the slot. The influence of coolant flow rate on the trailing edge loss is tremendous, too. Shock and boundary layer losses are major contributions to the overall loss but are less affected by the coolant. Finally a parameter variation changing the temperature ratio of coolant to main flow was performed, resulting in increasing losses with decreasing coolant temperature.

High-temperature high-frequency turbine exit flow field measurements in a military engine with a cooled unsteady total pressure probe

The measurement of unsteady pressures within the hot components of gas turbine engines still remains a true challenge for test engineers. Several high-temperature pressure sensors have been developed, but so far, their applications are restricted to unsteady wall static pressure measurements. Because of the severe flow conditions such as turbine inlet temperatures of 1700 °C and pressures of 50 bar or more in the most advanced aero-engine designs, few (if any) experimental techniques exist to measure the time-resolved flow total pressure inside the gas path. This article describes the measurements performed at the turbine exit of a military engine with a cooled fast response total pressure probe. The probe concept is based on the use of a conventional miniature piezo-resistive pressure sensor, located in the probe tip to achieve a bandwidth of at least 40 kHz. Due to the extremely harsh conditions, the probe and sensor are heavily water cooled. The probe was designed to be continuously immersed into the hot gas stream to obtain time series of pressure with a high bandwidth and therefore statistically representative average fluctuations at the blade passing frequency (BPV). The experimental results obtained with a second-generation prototype are presented. The probe was immersed into the engine through the bypass duct between turbine exit and flame-holders of the afterburner of a Volvo RM12 engine, at exhaust temperatures above 900 °C. The probe was able to resolve the BPV (∼17 kHz) and several harmonics up to 100 kHz.

Experimental and computational fluid dynamics investigation of the efficiency of an unshrouded transonic high pressure turbine

This article presents an experimental and computational study of the efficiency of an unshrouded transonic turbine. This research formed part of the EU Turbine Aero-Thermal External Flows II programme. The experiments were performed in the Oxford Turbine Research Facility (previously the Turbine Test Facility at QinetiQ, Farnborough). This facility is an engine scale, short duration, rotating transonic facility, in which M, Re, [Formula: see text], and [Formula: see text] are matched to engine conditions. For these experiments, the MT1 turbine stage was installed. Historically, turbine efficiency measurements are conducted in steady state adiabatic facilities. However, short-duration facilities allow simultaneous aerodynamic and heat transfer measurements with a significant reduction in cost. An efficiency measurement system was developed for this investigation, and this is briefly described. The system allows efficiency to be evaluated to bias and precision errors of approximately ±1.45 per cent and ±0.16 per cent, respectively, to 95 per cent confidence. The results of accurate area surveys of the turbine inlet and exit flows are presented and discussed. At the turbine exit, data were taken at two traverse planes, approximately 0.5 and 4.5 rotor axial chords downstream of the rotor. The turbine efficiency was experimentally evaluated based on the data at both planes, using a number of mixing models, which are discussed and compared. The experimental result of turbine efficiency is also compared to that estimated from a mean-line prediction. Full-stage steady and unsteady computational fluid dynamics of the experiment using the Rolls-Royce HYDRA code was conducted and is also presented. The predicted and measured rotor exit flow-fields are compared at both downstream traverse planes.

Computational Predictions of Heat Transfer and Film-Cooling for a Turbine Blade With Nonaxisymmetric Endwall Contouring

Three-dimensional contouring of the compressor and turbine endwalls in a gas turbine engine has been shown to be an effective method of reducing aerodynamic losses by mitigating the strength of the complex vortical structures generated at the endwall. Reductions in endwall heat transfer in the turbine have been also previously measured and reported in literature. In this study, computational fluid dynamics simulations of a turbine blade with and without nonaxisymmetric endwall contouring were compared to experimental measurements of the exit flowfield, endwall heat transfer, and endwall film-cooling. Secondary kinetic energy at the cascade exit was closely predicted with a simulation using the SST k-ω turbulence model. Endwall heat transfer was overpredicted in the passage for both the SST k-ω and realizable k-ε turbulence models, but heat transfer augmentation for a nonaxisymmetric contour relative to a flat endwall showed fair agreement to the experiment. Measured and predicted film-cooling results indicated that the nonaxisymmetric contouring limits the spread of film-cooling flow over the endwall depending on the interaction of the film with the contour geometry.

Pneumatic measurements downstream of a radial turbine nozzle cascade

This paper deals mainly with pneumatic measurements on a radial turbine nozzle cascade. The full radial cascade guarantees the exit flow field periodicity downstream of it. A special traversing mechanism with a five — hole conical probe moving along a circular path behind the cascade was used for flow field investigation in this type of cascade with very low aspect ratio. The analyses of results of 2D and 3D pneumatic measurements including loss coefficient values are presented.

단일 경사열선을 이용한 원형노즐의 유동해석

The program using C/C?? which could calculate 3-D velocities from data of the flowfield obtained by a slant hot-wire has been developed. The characteristic of the exit flowfield of a circular nozzle with a curved surface was investigated. From the result, the inward-flow was verified at adjacent layers near the boundary of the exit flowfield of a circular nozzle in 1d and outward-flow was sensed gradually from down-flow of 1d. Similarity of the nondimensional velocity profile properlly could be obtained except near 1d and the virtual origin reasonably could be obtained.

Unsteady Entropy Measurements in a High-Speed Radial Compressor

The time-dependent relative entropy field at the impeller exit of a centrifugal compressor is measured. This study is part of a broader effort to develop comprehensive measurement techniques that can be applied in the harsh environment of turbomachines. A miniature unsteady entropy probe (diameter of 1.8 mm) is designed and constructed in the present study. The unsteady entropy probe has two components: a one-sensor fast-response aerodynamic probe and a pair of thin-film gauges. The time-dependent total pressure and total temperature are measured from the fast-response aerodynamic probe and pair of thin-film gauges, respectively. The time-dependent relative entropy derived from these two measurements has a bandwidth of 40 kHz and an uncertainty of ±2 J/kg. The measurements show that for operating Condition A, φ=0.059 and ψ=0.478, the impeller exit flowfield is highly three dimensional. Adjacent to the shroud there are high levels of relative entropy and at the midspan there are low and moderate levels. Independent measurements made with a two-sensor aerodynamic probe show that the high velocity of the flow relative to the casing is responsible for the high relative entropy levels at the shroud. On the other hand, at the midspan, a loss free, jet flow region and a channel wake flow of moderate mixing characterize the flowfield. At both the shroud and midspan, there are strong circumferential variations in the relative entropy. These circumferential variations are much reduced when the centrifugal compressor is operated at operating Condition B, φ=0.0365 and ψ=0.54, near the onset of stall. In this condition, the impeller exit flowfield is less highly skewed; however, the time-averaged relative entropy is higher than at the operating Condition A. The relative entropy measurements with the unsteady entropy probe are thus complementary to other measurements, and more clearly document the losses in the centrifugal compressor.

Experimental research on aerodynamic performance and exit flow field of low pressure axial flow fan with circumferential skewed blades

In this article, the low pressure axial flow fan with circumferential skewed rotor blade, including the radial blade, the forward-skewed blade and the backward-skewed blade, was studied with experimental methods. The aerodynamic performance of the rotors was measured. At the design condition at outlet of the rotors, detailed flow measurements were performed with a five-hole probe and a Hot-Wire Anemometer (HWA). The results show that compared to the radial rotor, the forward-skewed rotor demonstrates a wider Stable Operating Range (SOR), is able to reduce the total pressure loss in the hub region and make main loading of blade accumulating in the mid-span region. There is a wider wake in the upper mid-span region of the forward-skewed rotor. Compared to the radial rotor, in the backward-skewed rotor there is higher total pressure loss near the hub and shroud regions and lower loss in the mid-span region. In addition, the velocity deficit in the wake is lower at mid-span of the backward-skewed rotor than the forward-skewed rotor.

Investigation of the Flow Field in a High-Pressure Turbine Stage for Two Stator-Rotor Axial Gaps—Part I: Three-Dimensional Time-Averaged Flow Field

An extensive experimental analysis on the subject of unsteady flow field in high-pressure turbine stages was carried out at the Laboratorio di Fluidodinamica delle Macchine (LFM) of Politecnico di Milano. The research stage represents a typical modern HP gas turbine stage designed by means of three-dimensional (3D) techniques, characterized by a leaned stator and a bowed rotor and operating in high subsonic regime. The first part of the program concerns the analysis of the steady flow field in the stator-rotor axial gap by means of a conventional five-hole probe and a temperature sensor. Measurements were carried out on eight planes located at different axial positions, allowing the complete definition of the three-dimensional flow field both in absolute and relative frame of reference. The evolution of the main flow structures, such as secondary flows and vane wakes, downstream of the stator are here presented and discussed in order to evidence the stator aerodynamic performance and, in particular, the different flow field approaching the rotor blade row for the two axial gaps. This results set will support the discussion of the unsteady stator-rotor effects presented in Part II (Gaetani, P., Persico, G., Dossena, V., and Osnaghi, C., 2007, ASME J. Turbomach., 129(3), pp. 580–590). Furthermore, 3D time-averaged measurements downstream of the rotor were carried out at one axial distance and for two stator-rotor axial gaps. The position of the probe with respect to the stator blades is changed by rotating the stator in circumferential direction, in order to describe possible effects of the nonuniformity of the stator exit flow field downstream of the stage. Both flow fields, measured for the nominal and for a very large stator-rotor axial gap, are discussed, and results show the persistence of some stator flow structures downstream of the rotor, in particular, for the minimum axial gap. Finally, the flow fields are compared to evidence the effect of the stator-rotor axial gap on the stage performance from a time-averaged point of view.

Exit Flow Field and Performance of Axial Flow Fans

Flow fields near the exit and the global performance parameters of the various types of axial flow fans are studied with Particle Image Velocimetry and a standard AMCA 210 flow bench. The fans used in this study included the shrouded, shroudless, and winglet-blade types. The velocity vectors, streamlines, vorticity contours, velocity distributions, and performances are presented and discussed. The flow patterns on the radial and axial planes show that a vortex always exists near the exit of the fans at various impeller angles. The experimental results demonstrate that the shrouded fan with winglets has the most stable flow field and the best fan performance.

Effect of vane-rotor interaction on the unsteady flowfield downstream of a transonic high pressure turbine

This paper analyses the effect of the stator-rotor interference on the stage exit flow field of a transonic turbine operated under engine representative conditions in the von Karman Institute (VKI) compression tube facility. The test programme comprises three pressure ratios and two Reynolds numbers. The time-averaged radial distributions of total pressure and temperature downstream of the rotor are affected by the hub passage vortices and the interaction between the tip passage vortex and the rotor tip leakage vortex. The azimuthal distributions exhibit significant non-uniformities with a periodicity of one vane pitch. The amplitude of this non-uniformity is very sensitive to the pressure ratio. A simple model shows that, contrary to the common belief, the transport of the vane wake and secondary flows across the rotor is not enough to explain the magnitude of the variations. The pitch-wise vane exit static pressure distribution, which is significantly distorted owing to the transonic regime of the vane, should be taken into account. The amplitude of the time-resolved fluctuations of pressure and temperature, which are due to rotor blade passing events, also varies significantly depending on the probe location in a vane pitch. The measured time-accurate rotor relative inlet total pressure suggests that the rotor relative exit undergoes periodic excursions in the transonic regime.

Periodic Velocity Measurements in a Wide and Large Radius Ratio Automotive Torque Converter at the Pump∕Turbine Interface

The unsteady flow field due to blade passing at the pump∕turbine interface of a torque converter was studied. The current geometry is wide and has a large outer to inner radius ratio. A laser velocimeter was used to measure the periodic velocity components at four operating conditions determined by the speed ratios between the turbine and pump of 0.065 (near stall), 0.600, 0.800, and 0.875 (coupling point). The flow fields at the pump exit and turbine inlet planes were visualized and are presented. Using instantaneous pump and turbine blade positions with the velocity data, animations (“slow-motion movies”) are generated to effectively visualize and understand the unsteady behavior. The turbine inlet flow was markedly periodic due to the exiting jet∕wake from the upstream pump passage; however, the pump exit flow field showed little dependence on the turbine blade positions. The highest unsteadiness was seen for the highest speed ratios. Four “shots” from the sequences of one cycle for all speed ratios and each plane are presented herein. The results are also compared to unsteady results for a previously examined torque converter with a small radius ratio to determine the effect of parametric geometric changes on the flow field. Generally, the unsteady velocity fields show no significant difference for the two geometries—the trends are the same.

Effect of the Hub Endwall Cavity Flow on the Flow-Field of a Transonic High-Pressure Turbine

In high-pressure turbines, a small amount of cold flow is ejected at the hub from the cavity that exists between the stator and the rotor disk. This prevents the ingestion of hot gases into the wheel-space cavity, thus avoiding possible damage. This paper analyzes the interaction between the hub-endwall cavity flow and the mainstream in a high-pressure transonic turbine stage. Several cooling flow ratios are investigated under engine representative conditions. Both time-averaged and time-resolved data are presented. The experimental data is successfully compared with the results of a three-dimensional steady Navier-Stokes computation. Despite the small amount of gas ejected, the hub-endwall cavity flow has a significant influence on the mainstream flow. The Navier-Stokes predictions show how the ejected cold flow is entrained by the rotor hub vortex. The time-resolved static pressure field around the rotor is greatly affected when traversing the non-uniform vane exit flow field. When the cavity flow rate is increased, the unsteady forces on the rotor airfoil are reduced. This is linked to the decrease of vane exit Mach number caused by the blockage of the ejected flow.

Secondary Flows and Loss Caused by Blade Row Interaction in a Turbine Stage

A study of the three-dimensional stator-rotor interaction in a turbine stage is presented. Experimental data reveal vortices downstream of the rotor which are stationary in the absolute frame—indicating that they are caused by the stator exit flowfield. Evidence of the rotor hub passage vortices is seen, but additional vortical structures away from the endwalls, which would not be present if the rotor were tested in isolation, are also identified. An unsteady computation of the rotor row is performed using the measured stator exit flowfield as the inlet boundary condition. The strength and location of the vortices at rotor exit are predicted. A formation mechanism is proposed whereby stator wake fluid with steep spanwise gradients of absolute total pressure is responsible for all but one of the rotor exit vortices. This mechanism is then verified computationally using a passive-scalar tracking technique. The predicted loss generation through the rotor row is then presented and a comparison made with a steady calculation where the inlet flow has been mixed out to pitchwise uniformity. The loss produced in the steady simulation, even allowing for the mixing loss at inlet, is 10% less than that produced in the unsteady simulation. This difference highlights the importance of the time-accurate calculation as a tool of the turbomachine designer.

Evaluating Experimental Data Requirements for Spatially Varying Turbine Flowfields Through Uncertainty Analysis

Experimental data from two cold-airflow turbine tests were evaluated. The two tests had different, relatively high spatial variations in the turbine exit flowfields. The objective of the research was to evaluate data requirements, including the averaging techniques, the number of measurements (spatial resolution), and the types of measurements needed, to obtain high-quality performance and code validation data for these types of flowfields. Uncertainty analysis methodology was used, and new uncertainty techniques were developed for the data studies. The results were then used to establish guidelines for planning future tests of other turbines. The guidelines set forth a quantified method for determining data requirements to meet specified test goals. Implementation of these guidelines will improve the quality of experimental results while decreasing time and cost for experimental studies of highly nonuniform turbine flowfields.

Time-Resolved Vane-Rotor Interaction in a High-Pressure Turbine Stage

This paper describes the time-varying aerodynamic interaction mechanisms that have been observed within a transonic high-pressure turbine stage; these are inferred from the time-resolved behavior of the rotor exit flow field. It contains results both from an experimental program in a turbine test facility and from numerical predictions. Experimental data was acquired using a fast-response aerodynamic probe capable of measuring Mach number, whirl angle, pitch angle, total pressure, and static pressure. A 3-D time-unsteady viscous Navier-Stokes solver was used for the numerical predictions. The unsteady rotor exit flowfield is formed from a combination of four flow phenomena: the rotor wake, the rotor trailing edge recompression shock, the tip-leakage flow, and the hub secondary flow. This paper describes the time-resolved behavior of each phenomenon and discusses the interaction mechanisms from which each originates. Two significant vane periodic changes (equivalent to a time-varying flow in the frame of reference of the rotor) in the rotor exit flowfield are identified. The first is a severe vane periodic fluctuation in flow conditions close to the hub wall and the second is a smaller vane periodic fluctuation occurring at equal strength over the entire blade span. These two regions of periodically varying flow are shown to be caused by two groups of interaction mechanisms. The first is thought to be caused by the interaction between the wake and secondary flow of the vane with the downstream rotor; and the second is thought to be caused by a combination of the interaction of the vane trailing edge recompression shock with the rotor, and the interaction between the vane and rotor potential fields.