Full-Annulus URANS Study on the Transportation of Combustion Inhomogeneity in a Four-Stage Cooled Turbine

The inhomogeneity of temperature in a turbine is related to the nonuniform heat release and air injections in combustors. In addition, it is influenced by the interactions between turbine cascades and coolant injections. Temperature inhomogeneity results in nonuniform flow temperature at turbine outlets, which is commonly measured by multiple thermal couples arranged in the azimuthal direction to monitor the operation of a gas turbine engine. Therefore, the investigation of temperature inhomogeneity transportation in a multistage gas turbine should help in detecting and quantifying the over-temperature or flameout of combustors using turbine exhaust temperature. Here the transportation of temperature inhomogeneity inside the four-stage turbine of a 300-MW gas turbine engine was numerically investigated using 3D CFD. The computational domain included all four stages of the turbine, consisting of more than 500 blades and vanes. Realistic components (N2, O2, CO2, and H2O) with variable heat capacities were considered for hot gas and cooling air. Coolants were added to the computational domain through more than 19,000 mass and momentum source terms. his was simple compared to realistic cooling structures. A URANS CFD run with over-temperature/flameout at 6 selected combustors out of 24 was carried out. The temperature distributions at rotor–stator interfaces and the turbine outlet were quantified and characterized by Fourier transformations in the time domain and space domain. It is found that the transport process from the hot-streaks/cold-streaks at the inlet to the outlet is relatively stable. The cold and hot fluid is redistributed in time and space due to the stator and rotor blades, in the region with a large parameter gradient at the inlet, strong unsteady temperature field and composition field appear. The distribution of the exhaust gas composition has a stronger correlation with the inlet temperature distribution and is less susceptible to interference.

Design and Test of a Novel Highly-Loaded Compressor

The blades of rear stages in small size core compressors are reduced to shorter than 20 mm or even less due to overall high pressure ratio. The growing of tip clearance-to-blade height ratio of the rear stages enhance the leakage flow and increase the possibility of a strong clearance sensitivity, thus limiting the compressor efficiency and stability. A new concept of compressor, namely diffuser passage compressor (DP), for small size core compressors was introduced. The design aims at making the compressors robust to tip clearance leakage flow by reducing pressure difference between pressure and suction surfaces. To validate the concept, the second stage of a two-stage highly loaded axial compressor was designed with DP rotor according to a diffuser map. The diffuser passage stage has the same inlet condition and loading as the conventional compressor (CNV) stage, of which the work coefficient is around 0.37. The predicted performance and flow field of the DP were compared with the conventional axial compressor in detail. The rig testing was supplemented with the numerical predictions. Results reveal that the throttle characteristic of DP indicates higher pressure rise and the loss reduction in tip clearance is mainly responsible for the performance improvement. For the compressor with DP, the pressure and flow angle are more uniform on exit plane. What’s more, the rotor with diffused passage reveals more robust than the conventional rotor at double clearance gap. Furthermore, the experimental data indicate that DP presents higher pressure rise at design and part speeds. At design speed, the stall margin was extended by 7.25%. Moreover, peak adiabatic efficiency of DP is also higher than that of CNV by about 0.7%.

Introduction and Commissioning of the New Darmstadt Transonic Compressor Test Facility

With the well-known Transonic Compressor Darmstadt (TCD) in operation since 1994, profound knowledge in designing and operating a sophisticated test-rig is available at the Institute of Gas Turbines and Aerospace Propulsion of TU Darmstadt. During this period, TCD has been subject to a vast number of redesigns within different measurement campaigns (see [1], [2], [3], [4], [5], [6], [7], [8]). To expand the capabilities and ensure a sustainable process of compressor research, a new test facility was designed and built by the institute. The new test rig Transonic Compressor Darmstadt 2 (TCD2) features increased power for higher pressure ratios and higher mass-flow, a state of the art control system, increased flexibility towards different compressor geometries and modern data acquisition hardware and software. Following the successful commissioning of the test-rig in March 2018, a first measurement campaign has been conducted. Early test results regarding aerodynamic performance and aeroelastic effects of the test compressor are presented together with a detailed overview of test-rig infrastructure and control systems as well as the test compressor and the measurement hardware.

Prediction of Reynolds Number Effects on Low-Pressure Turbines Using a High-Order ILES Method

A comprehensive comparison between Implicit Large Eddy Simulations (ILES) and experimental results of a modern highlift low-pressure turbine airfoil has been carried out for an array of Reynolds numbers (Re). Experimental data were obtained in a low-speed linear cascade at the Polithecnic University of Madrid using hot-wire anemometry and LDV. The numerical code is fourth order accurate, both in time and space. The spatial discretization of the compressible Navier-Stokes equations is based on a high-order Flux Reconstruction approach while a fourth order Runge-Kutta method is used to march in time the simulations. The losses, pressure coefficient distributions, and boundary layer and wake velocity profiles have been compared for an array of realistic Reynolds numbers. Moreover, boundary layer and wake velocity fluctuations are compared for the first time with experimental results. It is concluded that the accuracy of the numerical results is comparable to that of the experiments, especially for integral quantities such as the losses or exit angle. Turbulent fluctuations in the suction side boundary layer and the wakes are well predicted also. The elapsed time of the is about 140 hours on 40 Graphics Processor Units. The numerical tool is integrated within an industrial design system and reuses pre- and post-processing tools previously developed for another kind of applications. The trend of the losses with the Reynolds number has a sub-critical regime, where the losses scale with Re−1, and a supercrital regime, where the losses scale with Re−1/2. This trend can be seen both, in the simulations and the experiments.

Active Control of Unsteady Cavitating Flows in Turbomachinery

A preliminary 2D numerical investigation of the active control of unsteady cavitation by means of one single synthetic jet actuator (SJA) is presented. The SJA has been applied to hinder the intrinsic instabilities of a cloud cavitating flow of water around a NACA 0015 hydrofoil with an angle of attack of 8° and ambient conditions. It has been placed inside the inception region at a distance of 16% of the chord from the leading edge. Concerning the numerical approach, a Eulerian homogeneous mixture/mass transfer model has been used, in combination with an extended Schnerr-Sauer cavitation model and a Volume of Fluid (VOF) interface tracking method. The synthetic jet has been modeled by means of a user-defined velocity boundary conditions based on a sinusoidal waveform. A sensitivity analysis has been first performed in order to evaluate the influence of the main control parameters, namely the momentum coefficient Cμ, the dimensionless frequency F+ and the jet angle αjet. By combining the cavitating vapor content and the impact on the hydrodynamic performance, the best performing SJA configuration has been retrieved. Then, a deeper analysis of the vapor cavity dynamics and the vorticity field has been conducted in order to understand the modification of the main flow produced by the synthetic jet. The best SJA configuration was observed at Cμ = 0.0002, F+ = 0.309 and αjet = 90°, which led to a reduction of both the average vapor content and the average torsional load in the measure of 34.6% and 17.8% respectively. A reduction of the average pulsation frequency of the pressure upstream confirmed the beneficial effect of the SJA. The analysis of the coupled dynamics between vapor cavity-vorticity and their POD-based modal structures highlighted that the benefit of the SJA lies on preventing the growth of a thick sheet cavity which tends to cause the development of the highly cavitating cloud dynamics after the cavity breakup. This is mainly due to an additional vorticity close to the hydrofoil surface just downstream the SJA, as well as a local pressure modification close the SJA during the blowing stroke.

Modelling the Transient Behaviour of Gas Turbines

As a consequence of the increasing share of renewable energy sources in present-day electrical grid systems, time variations of the power demand for fossil fuel plants can become more sudden. Therefore, an ability to respond to sudden load changes becomes an important issue for power generation gas turbines. This paper describes a real-time model for predicting the transient performance of gas turbines. The method includes basic transient phenomena, such as volume packing and the heat transfer between the working fluid and the structural elements. The dynamics of components are quantified by solving ordinary differential equations with appropriate initial and boundary conditions. Compressor and turbine operating points are determined from corresponding performance maps previously calculated using sophisticated aerodynamic, through-flow codes. This includes a sufficient number of such characteristics to account for the variations in speed and machine geometry. The developed dynamic model was verified by comparison of simulation results with experimentally recorded operating parameters for a real engine. This includes the start-up sequence and the change of load. Additional simulation covers the system response to a step increase in fuel flow. The simulation is carried out faster than the real process.

Impact of Exit Duct Dynamic Response on Compressor Stability

Abrupt distortions can appear as a result of transient crosswind or during rapid aircraft maneuvers. Such distortions are known to reduce the aerodynamic stability of engines and therefore present a major concern to all aero-engine manufacturers. To assess the aerodynamic stability of fan blades due to distortions, rig tests are usually carried out to establish the loss in stall margin. In such test campaigns, an exit duct (which is followed by a nozzle) is placed downstream of the fan blade and the operating condition of the fan is controlled by this nozzle. It is shown in this paper that in such rig tests the length of duct downstream of a fan has a significant impact on fan stall margin. The key contributor for such interaction is the dynamic response of the exit duct and the aerodynamic stability of the fan is affected by the acoustic reflection from the exit nozzle. To study the underlying physics, transient response in the exit duct downstream of a transonic fan stage was studied numerically using a simplified model. Simulation results, along with calculations based on analytical theories, confirmed the generation, propagation and reflection of waves induced by the inlet distortion. A quantitative relationship concerning the lengths of the compression system is introduced which determines whether a duct setup would have beneficial or detrimental influences on compressor aerodynamic stability. The findings of this research have great implications for the stability assessment of fans as the stability margin can be affected by the waves generated in bypass ducts.

Deficiencies in Turbulence Modelling for the Prediction of the Stability Boundary in Highly Loaded Compressors

Reynolds-averaged Navier-Stokes (RANS) equations are employed for aerodynamic and aeroelastic modelling in axial compressors. Their solutions are highly dependent on the turbulence models for closure. The main objective of this work is to assess the widely used Spalart-Allmaras model’s suitability for compressor flows. For this purpose, an extensive investigation of the sources of uncertainties in a high-speed multi-stage compressor rig was carried out. The grid resolution near the casing end wall, which affects the tip leakage flow and casing boundary layer, was found to have a major effect on the stability limit prediction. Refinements in this region led to a stall margin loss prediction. It was found that this loss was exclusively due to the destruction term in the SA model.

Recognition of Structures Leading to Transition in a Low Pressure Turbine Cascade: Effect of Reduced Frequency

The boundary layer developing over the suction side of a low pressure turbine cascade operating under unsteady inflow conditions has been experimentally investigated. Time-resolved Particle Image Velocimetry (PIV) measurements have been performed in two orthogonal planes, the blade to blade and a wall parallel plane embedded within the boundary layer, for two different wake reduced frequencies. Proper Orthogonal Decomposition (POD) has been used to analyze the data and to provide an interpretation of the most significant flow structures for each phase of the wake passing cycle. To this purpose, a POD based procedure that sorts the data synchronizing the measurements of the two planes has been developed. Phase averaged data are then obtained for both cases. Moreover, once properly sorted, POD has been applied to sub-ensembles of data at the same relative phase within the wake passing cycle. Detailed information on the most energetic turbulent structures at a particular phase are obtained with this procedure (called phased POD), overcoming the limit of classical phase average that just provides a statistical representation of the turbulence field. Furthermore, the synchronization of the measurements in the two planes allows the computation of the characteristic dimension of boundary layer structures that are responsible for transition. These structures are often identified as vortical filaments parallel to the wall, typically referred to as boundary layer streaks. The largest and most energetic structures are observed when the wake centerline passes over the rear part of the suction side, and they appear practically the same for both reduced frequencies. The passing wake forces transition leading to the breakdown of the boundary layer streaks. Otherwise, the largest differences between the low and high reduced frequency are observed in the calmed region. The post-processing of these two planes further allowed us to compute the spacing of the streaks and make it non-dimensional by the boundary layer displacement thickness observed for each phase. The non-dimensional value of the streaks spacing is about constant, irrespective of the reduced frequency.

Numerical Investigation on the Effect of Bleed Port With Self-Recirculating Casing Treatment on the Stability of a 1.5-Stage Transonic Compressor

Compared to the traditional casing treatment, the self-recirculating casing treatment (SCT) can improve or not decrease the compressor efficiency while achieving the stall margin improvement. For the bleed port, the main design indicator is to reduce the flow loss caused by suction, while providing sufficient jet flow and jet pressure to the injector. In order to gain a better study of the bleed port stabilization mechanisms, the bleed configuration was parameterized with the bleed port inlet width and the bleed port axial position. Five kinds of recirculating casing treatments were applied to a 1.5-stage transonic axial compressor with the method of three-dimensional unsteady numerical simulation. Fifteen identical self-recirculating devices are uniformly mounted around the annulus. The numerical results show that the SCT can improve compressor total pressure ratio and stability, shift the stall margin towards lower mass flows. Furthermore, it has no impact on compressor efficiency. The optimal case presents that stability margin is improved by 6.7% employing 3.1% of the annulus mass flow. Expanding bleed port inlet width to an intermediate level can further enhance compressor stability, but excessive bleed port inlet width will reduce the stabilization effect. The optimal bleed port position is located in the blocked area of the low energy group at the top of the rotor. In the case of solid casing, stall inception was the tip blockage, which was mainly triggered by the interaction of the tip leakage vortex and passage shock. From radial distribution, the casing treatment predominantly affects the above 70% span. The reduction of tip reflux region by suction effect is the main reason for the extension of stable operation range. The SCT also has an obvious stability improvement in tip blockage stall, while delaying the occurrence of compressor stall.