High Pressure Gas Turbine Stage Research Articles

Effect of clocking on entropy noise generation within an aeronautical high pressure turbine stage

This article reports a detailed experimental and numerical study of the clocking effect between entropy wave spots and stator leading edges on the evolution of the entropy wave and on indirect noise generation in a high pressure turbine stage. Experimental campaigns considering burner-representative temperature fluctuations injected upstream of an uncooled high-pressure gas turbine stage have been performed in the high-speed closed-loop test rig of the Fluid Machine Laboratory (LFM) of the Politecnico di Milano. Acoustic measurements focused on entropy noise generation have been carried out by the German Aerospace Center, DLR with two flush mounted in-duct microphone arrays. URANS CFD simulations with and without entropy fluctuations imposed at the stage inlet were performed with the TRAF code, developed by the Università degli Studi di Firenze. A numerical post-processing procedure, based on the time and space discrete Fourier transform (DFT) of the conservative variables, has been implemented to extract the low-frequency content connected to the entropy fluctuations and indirect noise emissions. Simulations show an overall good agreement with the experiments, especially at the stage outlet thus confirming the possibility to predict entropy noise by means of URANS simulations with time varying inlet conditions. By exploiting the combination of experiments and simulations, the evolution of the entropy wave, the identification of entropy noise generation areas, and the evaluation of indirect noise emissions, for the different clocking positions were properly assessed. This study also demonstrates that this type of simulations can be used for the geometric design optimization of combustor/turbine coupling performing detailed parametric studies concerning the clocking position of the perturbation source in order to reduce the turbine-combustor interactions and the indirect noise emissions.

Influence of the Rotor-Driven Perturbation on the Stator-Exit Flow within a High-Pressure Gas Turbine Stage

In stator–rotor interaction studies on axial turbines, the attention is commonly focused on the unsteady rotor aerodynamics resulting from the periodic perturbations induced by the stator flow structures. Conversely, less interest has been historically attracted regarding the influence of the rotor on the flow released by the stator, correlated to propagation of the blade potential field upstream of the rotor leading edge. In this paper, experiments in the research high-pressure turbine of the Laboratory of Fluid-Machines of the Politecnico di Milano, performed by applying a fast-response aerodynamic pressure probe, alongside fully-3D time-accurate CFD simulations of the flow, are combined with the aim of discussing the rotor-to-stator interaction. While rotating, the rotor induces periodic perturbations on the pressure and velocity field in the stator–rotor gap, altering the evolution of the total quantities and the flow rate discharged by each stator channel and eventually triggering energy-separation effects which result in total pressure and total temperature oscillations in the stator-exit flow. Such oscillations were found to rise up to almost ±10% of the stage total temperature drop.

Computational and Experimental Study of the Unsteady Convection of Entropy Waves Within a High-Pressure Turbine Stage

Abstract This article describes the transport and the interaction of pulsating entropy waves generated by combustor burners within a high-pressure turbine stage for aeronautical application. Experiments and computational fluid dynamics (CFD) simulations were carried out in the context of the European Research Project RECORD. Experimental campaigns considering burner-representative temperature fluctuations (in terms of spot shape, fluctuation frequency, and total temperature variation percentage) injected upstream of an un-cooled high-pressure gas turbine stage have been performed in the high-speed closed-loop test rig of the Fluid Machine Laboratory (LFM) of the Politecnico di Milano (Italy). The pulsating entropy waves are injected at the stage inlet in streamwise direction at four different azimuthal positions featuring a 7% overtemperature with respect to the main flow with a frequency of 90 Hz. Detailed time-resolved temperature measurements (in the range of 0–200 Hz) upstream and downstream of the stage, as well as in the stator–rotor axial gap, were performed. Time-accurate CFD simulations with and without entropy fluctuations imposed at the stage inlet were performed with the TRAF code, developed by the Università degli Studi di Firenze (Italy). A numerical postprocessing procedure, based on the discrete Fourier transform (DFT) of the conservative variables, has been implemented to extract the low-frequency content connected to the entropy fluctuations. Measurements highlighted a significant attenuation of the entropy wave spot throughout their transport within the stator channel and their interaction with the rotor blade rows, highly depending on their injection azimuthal position. Simulations show an overall good agreement with the experiments on the measurement traverses, especially at the stage outlet. By exploiting the combination of experiments and simulations, the aerodynamic and thermal implications of the temperature fluctuation injected upstream of the stage were properly assessed, thus allowing suggest useful information to the designer. The comparison with the experiments confirms the accuracy of the CFD method to solve the periodic, but characterized by a low-frequency content event, associated with the entropy wave fluctuation.

가스터빈 연소기 출구 열·유동장 분포가 터빈에 미치는 효과의 수치해석 연구

Effects of temperature and flow fields at a combustor outlet on a high pressure gas turbine stage are investigated by using numerical methods. Numerical simulations of the annular type combustor are carried out by using large eddy simulation (LES) and unsteady Reynolds-averaged Navier-Stokes equations (URANS) techniques. The temperature distribution is found to be significantly affected by vortical structures at the primary zone and dilution jets. Numerical simulations of high pressure turbine stage are also conducted by using the Reynolds-averaged Navier-Stokes equations (RANS). Temperature distributions and aerodynamic loads at blade and end wall surfaces are drastically altered by changes in temperature, momentum, turbulent kinetic energy and turbulence dissipation rate at the exit of the combustor which are imposed as turbine inlet boundary conditions.

Investigations Into Heat Transfer and Aerodynamic Performance of a Worn Squealer Tipped Turbine Stage

Abstract Heat transfer and aerodynamic performance in worn squealer tip gap of a high-pressure gas turbine stage were numerically investigated. Effects of the starting location of wear and wear depth on tip heat transfer coefficient distributions and stage efficiency were analyzed to evaluate the aero-thermal performance degradations in the gas turbine stage after wear. At three starting locations of wear and five wear depths, flow patterns in worn squealer tip gap of the turbine stage were visualized and compared with the original design case. The results show that the counter-rotating vortex systems in tip cavity, as well as the interactions between leakage vortex and passage vortex, are significantly affected by the degree of wear damage. The starting location of wear and wear depth have pronounced influences on heat transfer and aerodynamic performance in squealer tip gap. After wear, the stage efficiency is decreased by about 0.3–1%, as the wear depth is equal to clearance gap size. In the serious worn case, thermal load on tip cavity floor is increased by about 60%, while the heat transfer on rims is reduced by about 20%. However, compared with the original design case, the area-averaged heat transfer coefficient on shroud is reduced by 5% at most.

Computational and Experimental Study of Hot Streak Transport Within the First Stage of a Gas Turbine

Abstract The paper discusses the migration, the interaction with the blades, and the attenuation of hot streaks generated by combustor burners, during their propagation within the first turbine stage of aero-engines. Experiments and computational fluid dynamics (CFD) simulations were carried out in the framework of the European Project RECORD and on its follow-up. Measurements considering burner-representative temperature perturbations injected upstream of an un-cooled high-pressure gas turbine stage were performed in the high-speed closed-loop test-rig of the Politecnico di Milano (Italy). The hot streaks were injected in a streamwise direction at the stage inlet in four different circumferential positions with respect to the stator blade. They feature a 20% over-temperature with respect to the main flow. Detailed temperature measurements as well as unsteady aerodynamic measurements upstream and downstream of the blade rows were performed. Time-accurate CFD simulations of the flow upstream and within the turbine stage were performed with the TRAF code, developed by the Università degli Studi di Firenze (Italy). Measurements show a relevant attenuation of hot streaks throughout their transport within the stator and the rotor blade rows, highly depending on the injection azimuthal position. The perturbations were observed to lose their spatial coherence, especially in the transport within the rotor, and to undergo severe spanwise migration. Simulations exhibit a good agreement with the experiments on the measurement planes and allow tracking the complex flow phenomena occurring within the blade rows. Finally, the aerodynamic and thermal implications of the inlet temperature perturbations are properly highlighted and discussed.

Experimental and Analytical Assessment of Cavity Modes in a Gas Turbine Wheelspace

Fast response pressure data acquired in a high-speed 1.5-stage turbine hot gas ingestion rig (HGIR) show the existence of pressure oscillation modes in the rim-seal-wheelspace cavity of a high pressure gas turbine stage with purge flow. The experimental results and observations are complemented by computational assessments of pressure oscillation modes associated with the flow in canonical cavity configurations. The cavity modes identified include shallow cavity modes and Helmholtz resonance. The response of the cavity modes to variation in design and operating parameters are assessed. These parameters include cavity aspect ratio (AR), purge flow ratio, and flow direction defined by the ratio of primary tangential to axial velocity. Scaling the cavity modal response based on computational results and available experimental data in terms of the appropriate reduced frequencies appears to indicate the potential presence of a deep cavity mode as well. While the role of cavity modes on hot gas ingestion cannot be clarified based on the current set of data, the unsteady pressure field associated with turbine rim cavity modal response can be expected to drive ingress/egress.

Tip Leakage Flow and Heat Transfer on Turbine Blade Tip and Casing, Part 1: Effect of Tip Clearance Height and Rotation Speed

Steady simulations were performed to investigate tip leakage flow and heat transfer characteristics on the rotor blade tip and casing in a single-stage gas turbine engine. A typical high-pressure gas turbine stage was modeled with a pressure ratio of 3.2. The predicted isentropic Mach number and adiabatic wall temperature on the casing showed good agreement with available experimental data under similar operating condition. The present numerical study focuses extensively on the effects of tip clearance heights and rotor rotational speeds on the blade tip and casing heat transfer characteristics. It was observed that the tip leakage flow structure is highly dependent on the height of the tip gap and the speed of the rotor. In all cases, the tip leakage flow was seen to separate and recirculate just around the corner of the pressure side of the blade tip. This region of re-circulating flow enlarges with increasing clearance heights. The separated leakage flow reattaches afterwards on the tip surface. Leakage flow reattachment was shown to enhance surface heat transfer at the tip. The interaction between tip leakage flow and secondary flows that is induced by the relative casing motion is found to significantly influence the blade tip and casing heat transfer distribution. A region of critical heat transfer exists on the casing near the blade tip leading edge and along the pressure-side edge for all the clearance heights that were investigated. At high rotation speed, the region of critical heat transfer tends to move towards the trailing edge due to the change in inflow angle.

The Effect of Coolant Injection on the Endwall Flow of a High Pressure Turbine

This paper presents a study of the effects of two types of hub coolant injection on the rotor of a high pressure gas turbine stage. The first involves the leakage flow from the hub cavity into the mainstream. The second involves a deliberate injection of coolant from a row of angled holes from the edge of the stator hub. The aim of this study is to improve the distribution of the injected coolant on the rotor hub wall. To achieve this, it is necessary to understand how the coolant and leakage flows interact with the rotor secondary flows. The first part of the paper shows that the hub leakage flow is entrained into the rotor hub secondary flow and the negative incidence of the leakage strengthens the secondary flow and increases its penetration depth. Three-dimensional unsteady calculations were found to agree with fast response pressure probe measurements at the rotor exit of a low speed test turbine. The second part of the paper shows that increasing the injected coolant swirl angle reduced the secondary flow penetration depth, improves the coolant distribution on the rotor hub, and improves stage efficiency. Most of the coolant however, was still found to be entrained into the rotor secondary flow.

Desensitization of the Flowfield from Rotor Tip-Gap Height by Casing-Air Injection

This paper presents an investigation that aims to desensitize the flowfield of an unshrouded high-pressure gas-turbine stage from the radial clearance between the rotor tip and the casing. A novel method of cooling-air injection from the rotor casing into the rotor tip region is applied. The injection opposes the tip-leakage flow and affects the development of the rotor secondary flow. In a previous investigation, an increase of stage efficiency of 0.55 percentage points was achieved with an injection of 0.7% of the core mass flow. The present study investigates the effect of the injection on the flowfield for two different rotor tip clearances. Measurements conducted with a two-sensor fast-response pressure probe provide data describing the time-resolved behavior of flow angles and pressures as well as turbulence intensity. The experimental investigation is done on the tip-gap heights of 0.65 and 1.00% span and injection mass flow rates representing 0, 0.7, and 1% of the turbine core mass flow. The results show that an increasing injection rate reduces the difference of rotor exit flow angle and flow unsteadiness between the two different tip-gap heights. The stage efficiency increases with injection for both tip-gap cases, whereas identical values are obtained at the injection rate of 1 % of the turbine mass flow.

Control of Rotor Tip Leakage Through Cooling Injection From the Casing in a High-Work Turbine

This paper presents an experimental investigation of a novel approach for controlling the rotor tip leakage and secondary flow by injecting cooling air from the stationary casing onto the rotor tip. It contains a detailed analysis of the unsteady flow interaction between the injected air and the flow in the rotor tip region and its impact on the rotor secondary flow structures. The experimental investigation has been conducted on a one-and-1/2-stage, unshrouded turbine, which has been especially designed and built for the current investigation. The turbine test case models a highly loaded, high pressure gas turbine stage. Measurements conducted with a two-sensor fast-response aerodynamic probe have provided data describing the time-resolved behavior of flow angles and pressures, as well as turbulence intensity in the exit plane of the rotor. Cooling air has been injected in the circumferential direction at a 30 deg angle from the casing tangent, opposing the rotor turning direction through a circumferential array of ten equidistant holes per rotor pitch. Different cooling air injection configurations have been tested. Injection parameters such as mass flow, axial position, and size of the holes have been varied to see the effect on the rotor tip secondary flows. The results of the current investigation show that with the injection, the size and the turbulence intensity of the rotor tip leakage vortex and the rotor tip passage vortex reduce. Both vortices move toward the tip suction side corner of the rotor passage. With an appropriate combination of injection mass flow rate and axial injection position, the isentropic efficiency of the stage was improved by 0.55 percentage points.