Numerical Investigation of a Turbine Guide Vane Exposed to Rotating Detonation Exhaust Flow

2019 ◽  
Author(s):  
Majid Asli ◽  
Cleopatra Cuciumita ◽  
Panagiotis Stathopoulos ◽  
Christian Oliver Paschereit
Keyword(s):  
Shock Wave ◽  
Flow Field ◽  
Numerical Investigation ◽  
Gas Turbines ◽  
Total Pressure ◽  
Guide Vane ◽  
The Body ◽  
Pressure Amplitude ◽  
Turbine Vane ◽  
Rotating Detonation

Abstract Detonative Pressure Gain Combustion has the potential to increase the propulsion efficiency of aero-engines and the thermal efficiency of stationary gas turbines. Important advances were made in this field, especially in the case of Rotating Detonation Combustion (RDC). Although experimental and numerical studies reported in the literature have significantly increased in number, the major open problem is a lack of efficient turbomachinery to transform the fluctuating potential energy from an RDC into power output. For this problem to be properly addressed, time resolved data at the outlet of an RDC needs to be collected. As a first step, numerical data can be used to generate a geometry for the turbine, which must be validated experimentally. To determine the performance of a turbine vane row, total pressure losses need to be measured. There are several challenges in measuring the total pressure between the outlet of an RDC and the inlet of a turbine vane row. The high temperature values, the distance of the pressure transducer from the outlet of the combustor lead to a lower time resolution of the pressure signal. The confined space is also an issue, allowing for very few options in measuring the total pressure. Another major problem is the shock wave that may form as a detached shock wave with respect to the body of the pressure probe at certain moments in the flow cycle, which leads to measuring a different value rather than the actual value of the flow field. To address these issues, the current study presents a numerical investigation of a guide vane row that was experimentally tested at the outlet of an RDC working on hydrogen and air under stoichiometric conditions. One of the vane rows was 3D printed with a geometry allowing the measurement of total pressure. Static pressure at the outlet of the RDC was also measured. It was observed that the measured pressures are average values in time. Based on these averages, the total inlet pressure and velocity variations in time were reconstructed in an exponential trend, according to the ones reported in the literature and the aforementioned experiments. These variations were set as inlet conditions for transient numerical simulations. Results show that the total pressure amplitude decreases significantly when the flow passes the annulus and the vanes as well. By looking in to the flow field detail, the presence of shock wave in front of the blade is investigated. Additionally, it is calculated that the average total pressure decreases 7.9% by the vane row.

scholarly journals Aerodynamic Investigation of Guide Vane Configurations Downstream a Rotating Detonation Combustor

2020 ◽  
Author(s):  
Majid Asli ◽  
Panagiotis Stathopoulos ◽  
Christian Oliver Paschereit
Keyword(s):  
Flow Field ◽  
Unsteady Flow ◽  
Pressure Loss ◽  
Total Pressure ◽  
Guide Vane ◽  
Turbine Blades ◽  
Total Pressure Loss ◽  
Design Parameters ◽  
Geometrical Parameters ◽  
Rotating Detonation

Abstract Pressure Gain Combustion (PGC) is considered a possible solution to increase gas turbine cycle efficiency, due to the lower entropy generation in the combustion process. However, the highly unsteady flow produced by PGC makes it more difficult to extract work from its exhaust gas. Any outlet restriction downstream of PGC, such as turbine blades, affects its flow field and may cause additional thermodynamic losses. The unsteadiness in the form of pressure, temperature and velocity vector fluctuations has a negative impact on the operation of conventional turbines. Therefore, evaluating early turbine design parameters for such applications is of great interest. Additionally, experimental measurements and data acquisition present researchers with challenges that have to do mostly with the high temperature exhaust of PGC and the high frequency of its operation. Numerical simulations can provide important insights into PGC exhaust flow and its interaction with turbine blades. In this paper, a Rotating Detonation Combustor (RDC) and a row of nozzle guide vanes have been modeled based on the data from literature and an available experimental setup at TU Berlin. Five guide vane configurations with different geometrical parameters have been modeled. URANS simulations were done for all guide vane arrangements to investigate the effect of solidity and blade type representing different outlet restrictions on the RDC exhaust flow. Total pressure loss and velocity fluctuation were computed upstream and downstream of the vanes. The results analzed the connection between total pressure loss and the vanes solidity and thickness to chord ratio. It is observed that more than 57% of the upstream velocity angle fluctuation amplitude was damped by the vanes. Furthermore, the area reduction was found to be the significant driving factor for damping the velocity angle fluctuations, whether in the form of solidity or thickness on chord ratio increment. A further study of the flow field details revealed that the vane passages act as convergent divergent nozzles in the unsteady flow field and no compression wave exists upstream. This RDC exhaust flow investigation is an important primary step from a turbomachinery standpoint, which provided details of blade behavior in such an unsteady flow field.

Computational Design and Experimental Evaluation of Using a Leading Edge Fillet on a Gas Turbine Vane

2001 ◽  
Author(s):  
G. A. Zess ◽  
K. A. Thole
Keyword(s):  
Boundary Layer ◽  
Flow Field ◽  
Pressure Gradient ◽  
Gas Turbine ◽  
Total Pressure ◽  
Guide Vane ◽  
Leading Edge ◽  
Field Measurements ◽  
Computational Design ◽  
Horseshoe Vortex

With the desire for increased power output for a gas turbine engine comes the continual push to achieve higher turbine inlet temperatures. Higher temperatures result in large thermal and mechanical stresses particularly along the nozzle guide vane. One critical region along a vane is the leading edge-endwall juncture. Based on the assumption that the approaching flow to this juncture is similar to a two-dimensional boundary layer, previous studies have shown that a horseshoe vortex forms. This vortex forms because of a radial total pressure gradient from the approaching boundary layer. This paper documents the computational design and experimental validation of a fillet placed at the leading edge-endwall juncture of a guide vane to eliminate the horseshoe vortex. The fillet design effectively accelerated the incoming boundary layer thereby mitigating the effect of the total pressure gradient. To verify the CFD studies used to design the leading edge fillet, flow field measurements were performed in a large-scale, linear, vane cascade. The flow field measurements were performed with a laser Doppler velocimeter in four planes orientated orthogonal to the vane. Good agreement between the CFD predictions and the experimental measurements verified the effectiveness of the leading edge fillet at eliminating the horseshoe vortex. The flowfield results showed that the turbulent kinetic energy levels were significantly reduced in the endwall region because of the absence of the unsteady horseshoe vortex.

Numerical Simulation of Deposition Phenomena on High-Pressure Turbine Vane Using UPACS

2019 ◽  
Author(s):  
Kenta Mizutori ◽  
Koji Fukudome ◽  
Makoto Yamamoto ◽  
Masaya Suzuki
Keyword(s):  
Numerical Simulation ◽  
High Pressure ◽  
Flow Field ◽  
Pressure Loss ◽  
Total Pressure ◽  
Total Pressure Loss ◽  
High Pressure Turbine ◽  
Pressure Loss Coefficient ◽  
Turbine Vane ◽  
Total Pressure Loss Coefficient

Abstract We performed numerical simulation to understand deposition phenomena on high-pressure turbine vane. Several deposition models were compared and the OSU model showed good adaptation to any flow field and material, so it was implemented on UPACS. After the implementation, the simulations of deposition phenomenon in several cases of the flow field were conducted. From the results, particles adhere on the leading edge and the trailing edge side of the pressure surface. Also, the calculation of the total pressure loss coefficient was conducted after computing the flow field after deposition. The total pressure loss coefficient increased after deposition and it was revealed that the deposition deteriorates aerodynamic performance.

scholarly journals Numerical Investigation of Influence of Entropy Wave on the Acoustic and Wall Heat Transfer Characteristics of a High-Pressure Turbine Guide Vane

Acoustics ◽  
2020 ◽  
Vol 2 (3) ◽  
pp. 524-538
Author(s):  
Keqi Hu ◽  
Yuanqi Fang ◽  
Yao Zheng ◽  
Gaofeng Wang ◽  
Stéphane Moreau
Keyword(s):  
Heat Transfer ◽  
Heat Transfer Coefficient ◽  
Flow Field ◽  
Transfer Coefficient ◽  
Gas Turbines ◽  
Guide Vane ◽  
Turbulence Models ◽  
Local Pressure ◽  
Maximum Heat ◽  
Turbine Guide Vane

As an indirect noise source generated in the combustion chamber, entropy waves are widely prevalent in modern gas turbines and aero-engines. In the present work, the influence of entropy waves on the downstream flow field of a turbine guide vane is investigated. The work is mainly based on a well-known experimental configuration called LS89. Two different turbulence models are used in the simulations which are the standard k-ω model and the scale-adaptive simulation (SAS) model. In order to handle the potential transition issue, Menter’s ð-Reθ transition model is coupled with both models. The baseline cases are first simulated with the two different turbulence models without any incoming perturbation. Then one forced case with an entropy wave train set at the turbine inlet at a given frequency and amplitude is simulated. Results show that the downstream maximum Mach number is rising from 0.98 to 1.16, because the entropy waves increase the local temperature of the flow field; also, the torque of the vane varies as the entropy waves go through, the magnitude of the oscillation is 7% of the unforced case. For the wall (both suction and pressure side of the vane) heat transfer, the entropy waves make the maximum heat transfer coefficient nearly twice as the large at the leading edge, while the minimum heat transfer coefficient stays at a low level. As for the averaged normalized heat transfer coefficient, a maximum difference of 30% appears between the baseline case and the forced case. Besides, during the transmission process of entropy waves, the local pressure fluctuates with the wake vortex shedding. The oscillation magnitude of the pressure wave at the throat is found to be enhanced due to the inlet entropy wave by applying the dynamic mode decomposition (DMD) method. Moreover, the transmission coefficient of the entropy waves, and the reflection and transmission coefficients of acoustic waves are calculated.

A New Test Facility to Investigate Film Cooling on a Non-Axisymmetric Contoured Turbine Endwall: Part I — Introduction and Aerodynamic Measurements

2015 ◽  
Author(s):  
Franz Puetz ◽  
Johannes Kneer ◽  
Achmed Schulz ◽  
Hans-Joerg Bauer
Keyword(s):  
Heat Transfer ◽  
Flow Field ◽  
Pressure Distribution ◽  
Film Cooling ◽  
Gas Turbines ◽  
Guide Vane ◽  
Leading Edge ◽  
Suction Side ◽  
Test Facility ◽  
Endwall Contouring

An increased demand for lower emission of stationary gas turbines as well as civil aircraft engines has led to new, low emission combustor designs with less liner cooling and a flattened temperature profile at the outlet. As a consequence, the heat load on the endwall of the first nozzle guide vane is increased. The secondary flow field dominates the endwall heat transfer, which also contributes to aerodynamic losses. A promising approach to reduce these losses is non-axisymmetric endwall contouring. The effects of non-axisymmetric endwall contouring on heat transfer and film cooling are yet to be investigated. Therefore, a new cascade test rig has been set up in order to investigate endwall heat transfer and film cooling on both a flat and a non-axisymmetric contoured endwall. Aerodynamic measurements that have been made prior to the upcoming heat transfer investigation are shown. Periodicity and detailed vane Mach number distributions ranging from 0 to 50% span together with the static pressure distribution on the endwall give detailed information about the aerodynamic behavior and influence of the endwall contouring. The aerodynamic study is backed by an oil paint study, which reveals qualitative information on the effect of the contouring on the endwall flow field. Results show that the contouring has a pronounced effect on vane and endwall pressure distribution and on the endwall flow field. The local increase and decrease of velocity and the reduced blade loading towards the endwall is the expected behavior of the 3d contouring. So are the results of the oil paint visualization, which show a strong change of flow field in the leading edge region as well as that the contouring delays the horse shoe vortex hitting the suction side.

scholarly journals Aerodynamic Investigation of Guide Vane Configurations Downstream a Rotating Detonation Combustor

2020 ◽  
Author(s):  
Majid Asli ◽  
Panagiotis Stathopoulos ◽  
Christian Oliver Paschereit
Keyword(s):  
Flow Field ◽  
Negative Impact ◽  
Guide Vane ◽  
Turbine Blades ◽  
Geometrical Parameters ◽  
Area Reduction ◽  
Flow Investigation ◽  
Pressure Gain Combustion ◽  
Nozzle Guide ◽  
Rotating Detonation

Abstract Any outlet restriction downstream of Pressure Gain Combustion (PGC), such as turbine blades, affects its flow field and may cause additional thermodynamic losses. The unsteadiness in the form of pressure, temperature and velocity vector fluctuations has a negative impact on the operation of conventional turbines. Additionally, experimental measurements and data acquisition present researchers with challenges that have to do mostly with the high temperature exhaust of PGC and the high frequency of its operation. Nevertheless, numerical simulations can provide important insights into PGC exhaust flow and its interaction with turbine blades. In this paper, a Rotating Detonation Combustor (RDC) and a row of nozzle guide vanes have been modeled based on the data from literature and an available experimental setup. URANS simulations were done for five guide vane configurations with different geometrical parameters to investigate the effect of solidity and blade type representing different outlet restrictions on the RDC exhaust flow. The results analyzed the connection between total pressure loss and the vanes solidity and thickness to chord ratio. It is observed that more than 57% of the upstream velocity angle fluctuation amplitude was damped by the vanes. Furthermore, the area reduction was found to be the significant driving factor for damping the velocity angle fluctuations, whether in the form of solidity or thickness on chord ratio increment. This RDC exhaust flow investigation is an important primary step from a turbomachinery standpoint, which provided details of blade behavior in such an unsteady flow field.

scholarly journals Effects of Tip Endwall Contouring on the Three Dimensional Flow Field in an Annular Turbine Nozzle Guide Vane: Part 2 — Numerical Investigation

1985 ◽  
Author(s):  
Tony Arts
Keyword(s):  
Flow Field ◽  
Numerical Investigation ◽  
Guide Vane ◽  
Three Dimensional ◽  
Three Dimensional Flow ◽  
Dimensional Flow ◽  
Nozzle Guide Vane ◽  
Contour Plots ◽  
Endwall Contouring ◽  
Nozzle Guide

This paper describes the numerical investigation of the three dimensional flow through a low speed, low aspect ratio, high turning annular turbine nozzle guide vane with meridional tip endwall contouring. This rotational flow field has been simulated using a finite volume discretization and a time marching technique to solve the three dimensional, time dependent Euler equations expressed in a cylindrical coordinates system. The results are presented under the form of contour plots, spanwise pitch-averaged distributions and blade static pressure distributions. Detailed comparisons with the measurements described in part I of the paper are also provided.

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