Probabilistic and Numerical Modelling of a Lobed Mixer at Windmilling Conditions

2013 ◽  
Author(s):  
Maxime Lecoq ◽  
Nicholas Grech ◽  
Pavlos K. Zachos ◽  
Vassilios Pachidis
Keyword(s):  
Flow Field ◽  
Response Surface ◽  
Gas Turbine ◽  
Pressure Loss ◽  
Total Pressure ◽  
Engine Performance ◽  
Operating Conditions ◽  
Total Pressure Loss ◽  
Operating Point ◽  
Mixing Losses

Aero-gas turbine engines with a mixed exhaust configuration offer significant benefits to the cycle efficiency relative to separate exhaust systems, such as increase in gross thrust and a reduction in fan pressure ratio required. A number of military and civil engines have a single mixed exhaust system designed to mix out the bypass and core streams. To reduce mixing losses, the two streams are designed to have similar total pressures. In design point whole engine performance solvers, a mixed exhaust is modelled using simple assumptions; momentum balance and a percentage total pressure loss. However at far off-design conditions such as windmilling and altitude relights, the bypass and core streams have very dissimilar total pressures and momentum, with the flow preferring to pass through the bypass duct, increasing drastically the bypass ratio. Mixing of highly dissimilar coaxial streams leads to complex turbulent flow fields for which the simple assumptions and models used in current performance solvers cease to be valid. The effect on simulation results is significant since the nozzle pressure affects critical aspects such as the fan operating point, and therefore the windmilling shaft speeds and air mass flow rates. This paper presents a numerical study on the performance of a lobed mixer under windmilling conditions. An analysis of the flow field is carried out at various total mixer pressure ratios, identifying the onset and nature of recirculation, the flow field characteristics, and the total pressure loss along the mixer as a function of the operating conditions. The data generated from the numerical simulations is used together with a probabilistic approach to generate a response surface in terms of the mass averaged percentage total pressure loss across the mixer, as a function of the engine operating point. This study offers an improved understanding on the complex flows that arise from mixing of highly dissimilar coaxial flows within an aero-gas turbine mixer environment. The total pressure response surface generated using this approach can be used as look-up data for the engine performance solver to include the effects of such turbulent mixing losses.

Numerical and Experimental Analysis of Inlet Non-Uniformity Influence on Intercooler Performance

2012 ◽  
Author(s):  
Wei Dong ◽  
Chun Mao ◽  
Jian-Jun Zhu ◽  
Yong Chen
Keyword(s):  
Heat Transfer ◽  
Flow Field ◽  
Gas Turbine ◽  
Pressure Loss ◽  
Total Pressure ◽  
Flow Path ◽  
Total Pressure Loss ◽  
Design Calculation ◽  
Flow Area ◽  
Inlet Flow

The inlet flow of intercooler will be not uniform when the air flows through diffuser behind the low-pressure compressor. With the aid of the CFD technology, the flow field and pressure drop of gas turbine intercooler are analyzed. The equivalent flow area and the equivalent heat transfer methods are proposed in numerical simulations, hence the modeling problem of entire intercooler flow path is solved effectively. Three kinds of scheme of flow path are computed and the flow fields and pressure drops are given in this paper. The influence of inlet flow non-uniformity on the performance of intercooler is analyzed. The numerical computation results indicate that the CFD technology is valuable for analyzing the details of the flow field and improving the intercooler flow path design. The combination of the CFD technology and the intercooler design technique can improve the design of the flow path of gas turbine intercooler. The total pressure loss of the intercooler can be effectively reduced by improving the inlet flow non-uniformity. By comparing the experiment results and computation results of intercooler flow and heat transfer with the plate-fin heat exchanger design calculation, inlet flow non-uniformity has less influence on the heat transfer, but has more influence on the total pressure loss. When the intercooler with special flow structure and layout for marine gas turbine is designed, inlet flow non-uniformity should be fully considered and the total pressure loss should be corrected based on experiments.

Experimental Study of Combustion on a Micro Gas Turbine Combustor

2008 ◽  
Author(s):  
Feng-Shan Wang ◽  
Wen-Jun Kong ◽  
Bao-Rui Wang
Keyword(s):  
Gas Turbine ◽  
Pressure Loss ◽  
Loss Factor ◽  
Total Pressure ◽  
Combustion Efficiency ◽  
Total Pressure Loss ◽  
Test Facility ◽  
Inlet Temperature ◽  
Micro Gas Turbine ◽  
Oil Fuel

A research program is in development in China as a demonstrator of combined cooling, heating and power system (CCHP). In this program, a micro gas turbine with net electrical output around 100kW is designed and developed. The combustor is designed for natural gas operation and oil fuel operation, respectively. In this paper, a prototype can combustor for the oil fuel was studied by the experiments. In this paper, the combustor was tested using the ambient pressure combustor test facility. The sensors were equipped to measure the combustion performance; the exhaust gas was sampled and analyzed by a gas analyzer device. From the tests and experiments, combustion efficiency, pattern factor at the exit, the surface temperature profile of the outer liner wall, the total pressure loss factor of the combustion chamber with and without burning, and the pollutants emission fraction at the combustor exit were obtained. It is also found that with increasing of the inlet temperature, the combustion efficiency and the total pressure loss factor increased, while the exit pattern factor coefficient reduced. The emissions of CO and unburned hydrogen carbon (UHC) significantly reduced, but the emission of NOx significantly increased.

Film Cooling and Aerodynamic Performance on Multi-Cavity Squealer Tip of a Turbine Blade

2021 ◽  
Author(s):  
Feng Li ◽  
Zhao Liu ◽  
Zhenping Feng
Keyword(s):  
Gas Turbine ◽  
Film Cooling ◽  
Pressure Loss ◽  
Total Pressure ◽  
Aerodynamic Performance ◽  
Total Pressure Loss ◽  
Cooling Performance ◽  
Pressure Loss Coefficient ◽  
Blade Tip ◽  
Total Pressure Loss Coefficient

Abstract The blade tip region of the shroud-less high-pressure gas turbine is exposed to an extremely operating condition with combined high temperature and high heat transfer coefficient. It is critical to design new tip structures and apply effective cooling method to protect the blade tip. Multi-cavity squealer tip has the potential to reduce the huge thermal loads and improve the aerodynamic performance of the blade tip region. In this paper, numerical simulations were performed to predict the aerothermal performance of the multi-cavity squealer tip in a heavy-duty gas turbine cascade. Different turbulence models were validated by comparing to the experimental data. It was found that results predicted by the shear-stress transport with the γ-Reθ transition model have the best precision. Then, the film cooling performance, the flow field in the tip gap and the leakage losses were presented with several different multi-cavity squealer tip structures, under various coolant to mainstream mass flow ratios (MFR) from 0.05% to 0.15%. The results show that the ribs in the multi-cavity squealer tip could change the flow structure in the tip gap for that they would block the coolant and the leakage flow. In this study, the case with one-cavity (1C) achieves the best film cooling performance under a lower MFR. However, the cases with multi-cavity (2C, 3C, 4C) show higher film cooling effectiveness under a higher MFR of 0.15%, which are 32.6%%, 34.2%% and 41.0% higher than that of the 1C case. For the aerodynamic performance, the case with single-cavity has the largest total pressure loss coefficient in all MFR studied, whereas the case with two-cavity obtains the smallest total pressure loss coefficient, which is 7.6% lower than that of the 1C case.

Numerical Simulation of Cold Swirl Field for Solid Fuel Ramjet with NACA Airfoil

2014 ◽  
Vol 716-717 ◽  
pp. 711-716
Author(s):  
Jie Yu ◽  
Xiong Chen ◽  
Hong Wen Li
Keyword(s):  
Numerical Simulation ◽  
Flow Field ◽  
Radial Velocity ◽  
Solid Fuel ◽  
Velocity Gradient ◽  
Pressure Loss ◽  
Total Pressure ◽  
Swirl Flow ◽  
Total Pressure Loss ◽  
Swirl Number

In order to study the swirl flow characteristics in the solid fuel ramjet chamber, a new type of annular vane swirler with NACA airfoil is designed. The cold swirl flow field in the chamber is numerically simulated with different camber and t attack angle, while the swirl number , swirl flow field structure, total pressure recovery coefficient were studied. According to numerical simulation result, the main factors in swirl number are camber and angle of attack, the greater angle of attack, the greater the camber ,the stronger swirl will be. Results show that the total pressure loss is mainly concentrated in the inlet section, the total pressure loss cause by vane swirler is small. Radial velocity gradient exists in swirling flow, and increases with the swirl number. With the influence of centrifugal force and combustion chamber structure, the radial velocity gradient increases.

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.

Experimental Investigation of Total Pressure Loss Development in a Highly Loaded Low-Pressure Turbine Cascade

2017 ◽  
Vol 140 (3) ◽  
Author(s):  
Philip Bear ◽  
Mitch Wolff ◽  
Andreas Gross ◽  
Christopher R. Marks ◽  
Rolf Sondergaard
Keyword(s):  
Flow Field ◽  
Secondary Flow ◽  
Pressure Loss ◽  
Total Pressure ◽  
Low Pressure ◽  
Total Pressure Loss ◽  
Low Pressure Turbine ◽  
Pressure Transport ◽  
Secondary Flow Field ◽  
Highly Loaded

Improvements in turbine design methods have resulted in the development of blade profiles with both high lift and good Reynolds lapse characteristics. An increase in aerodynamic loading of blades in the low-pressure turbine (LPT) section of aircraft gas turbine engines has the potential to reduce engine weight or increase power extraction. Increased blade loading means larger pressure gradients and increased secondary losses near the endwall. Prior work has emphasized the importance of reducing these losses if highly loaded blades are to be utilized. The present study analyzes the secondary flow field of the front-loaded low-pressure turbine blade designated L2F with and without blade profile contouring at the junction of the blade and endwall. The current work explores the loss production mechanisms inside the LPT cascade. Stereoscopic particle image velocimetry (SPIV) data and total pressure loss data are used to describe the secondary flow field. The flow is analyzed in terms of total pressure loss, vorticity, Q-Criterion, turbulent kinetic energy, and turbulence production. The flow description is then expanded upon using an implicit large eddy simulation (ILES) of the flow field. The Reynolds-averaged Navier–Stokes (RANS) momentum equations contain terms with pressure derivatives. With some manipulation, these equations can be rearranged to form an equation for the change in total pressure along a streamline as a function of velocity only. After simplifying for the flow field in question, the equation can be interpreted as the total pressure transport along a streamline. A comparison of the total pressure transport calculated from the velocity components and the total pressure loss is presented and discussed. Peak values of total pressure transport overlap peak values of total pressure loss through and downstream of the passage suggesting that the total pressure transport is a useful tool for localizing and predicting loss origins and loss development using velocity data which can be obtained nonintrusively.

scholarly journals Experimental Prediction of Lean Blowout Limits for 3kW Micro Gas Turbine Combustor fuelled with LPG

2021 ◽  
Vol 13 (1) ◽  
pp. 89-95
Author(s):  
V. KIRUBAKARAN ◽  
David BHATT
Keyword(s):  
Pressure Drop ◽  
Gas Turbine ◽  
Pressure Loss ◽  
Total Pressure ◽  
Total Pressure Loss ◽  
Gas Turbine Combustor ◽  
Inlet Velocity ◽  
Lean Blowout ◽  
Micro Gas Turbine ◽  
Experimental Prediction

The Lean Blowout Limit of the combustor is one of the important performance parameters for a gas turbine combustor design. This study aims to predict the total pressure loss and Lean Blowout (LBO) limits of an in-house designed swirl stabilized 3kW can-type micro gas turbine combustor. The experimental prediction of total pressure loss and LBO limits was performed on a designed combustor fuelled with Liquefied Petroleum Gas (LPG) for the combustor inlet velocity ranging from 1.70 m/s to 11 m/s. The results show that the predicted total pressure drop increases with increasing combustor inlet velocity, whereas the LBO equivalence ratio decreases gradually with an increase in combustor inlet velocity. The combustor total pressure drop was found to be negligible; being in the range of 0.002 % to 0.065 % for the measured inlet velocity conditions. These LBO limits predictions will be used to fix the operating boundary conditions of the gas turbine combustor.

scholarly journals Gas turbine outlets: test results

2020 ◽  
Vol 4 (394) ◽  
pp. 121-128
Author(s):  
Nikolay N. Ponomarev
Keyword(s):  
Experimental Study ◽  
Gas Turbine ◽  
Pressure Loss ◽  
Total Pressure ◽  
Mutual Influence ◽  
Engineering Calculation ◽  
Total Pressure Loss ◽  
Flow Section ◽  
Pressure Losses ◽  
Flow Parameters

Object and purpose of research. The object of this work is gas turbine outlet consisting of axial-radial diffuser with the struts and the volute. The purpose is to create a methodology for engineering calculations, taking into account the mutual influence of the diffuser and the volute. Materials and methods. Experimental study of the flow in the models of outlets by measuring total and static pressure in characteristic sections. Calculation of integral and averaged flow parameters in measurement sections. Visualization of boundary flows. Based on the experimental results, development of regression models for the correction factors to be applied in the theoretical model, with selection of relevant factors. Main results. An experimental study of 23 variants of models with a total volume of 112 experimental points (modes) was carried out. On the basis of the experiment, methodology and program for engineering calculation of total pressure losses in the outlets were developed. It was found that the installation of guide blades and radial ribs in the diffuser in order to reduce local expansion angles with the ultimate purpose of mitigating total pressure losses actually does not lead to this result due to the because the flow in the diffuser becomes asymmetric due to its interaction with the volute. Visualization of boundary flows in the diffusers and the volutes has been performed, which makes it possible to identify the locations of separations causing increased pressure losses. Conclusion. An engineering method for calculating the total pressure loss in gas turbine outlet has been developed. The technique makes it possible, taking size restrictions into account, to select the geometry of the flow section that ensures minimum total pressure loss.

Influence of Pulsating Flow on Turbine Performance Investigated by DES and PIV

2018 ◽  
Author(s):  
Yohei Nakamura ◽  
Manato Chinen ◽  
Masamichi Sakakibara ◽  
Kazuyoshi Miyagawa
Keyword(s):  
Pressure Loss ◽  
Total Pressure ◽  
Internal Flow ◽  
Pulsating Flow ◽  
Total Pressure Loss ◽  
Operating Point ◽  
Pulsation Frequency ◽  
Flow Conditions ◽  
Experimental Apparatus ◽  
Loss Mechanisms

Recently, the downsizing of engine using turbocharger attracts more and more attention. Generally speaking, a turbocharger is usually designed based on its steady performance curve. However, the operating point of a turbocharger turbine does not match the steady operating point: instead it shows hysteresis behavior because of the pulsating flow generated by the engine valves. Unfortunately, turbine efficiency drops under pulsating flow conditions, but the loss mechanisms of the turbine under these conditions are not understood. Internal flow measurements under pulsating flow are actually very difficult. In this study, the internal flow under pulsating conditions was measured using a high speed PIV (Particle Image Velocimetry) system. The loss mechanisms were investigated by experimental investigation and computational fluid dynamics (CFD). The instantaneous pressure, velocity and torque were measured using a turbine experimental apparatus at WASEDA University. To generate the pulsating flow, a pulse generator was placed upstream of the turbine: a rotational disk with holes that only lets the flow through periodically. The pulsating frequency could be changed freely by changing the rotational speed of the disk. The visualization using PIV was performed at a frequency of 1 kHz at the turbine outlet. Many fine vortices which rotate in various directions were observed under pulsating flow. Such vortices mix in the exhaust diffuser and under low frequency flow, mixing of vortices took a long time. It was observed that one loss mechanism under unsteady conditions is the mixing of vortices at the turbine outlet. CFD was performed using ANSYS-CFX, with approximately 10 million nodes. Turbulent flows were treated by using the Reynolds-averaged Navier-Stokes (RANS) and Detached Eddy Simulation (DES) with the SST k-ω turbulence model. It was confirmed that the wheel and exhaust diffuser total pressure loss under pulsating flow was higher under steady flow conditions. In addition, the total pressure loss is proportional to the flow pulsation frequency. The analysis with DES agreed with the PIV results qualitatively. On the other hand, the analysis with RANS could not simulate the flow pattern at the turbine outlet.

A Comprehensive Investigation of Blade Row Interaction Effects on Stator Loss Utilizing Vane Clocking

2017 ◽  
Author(s):  
Natalie R. Smith ◽  
Nicole L. Key
Keyword(s):  
Boundary Layer ◽  
Flow Field ◽  
Pressure Loss ◽  
Total Pressure ◽  
Boundary Layer Transition ◽  
Total Pressure Loss ◽  
Tip Leakage Flow ◽  
Inlet Flow ◽  
Blade Row ◽  
Vane Clocking

Blade row interactions drive the unsteady performance of high pressure compressors. Vane clocking is the relative circumferential positioning of consecutive stationary vane rows with the same vane count. By altering the upstream vane wake’s path with respect to the downstream vane, vane clocking changes the blade row interactions and results in a change in steady total pressure loss on the downstream vane. The open literature lacks a conclusive discussion of the flow physics governing these interactions in compressors. This paper presents the details of a comprehensive vane clocking study on the embedded stage of the Purdue 3-stage axial compressor. The steady loss results, including radial total pressure profiles and surface flow visualization, suggest a shift in the Stator 2 corner separations occurs between clocking configurations associated with the maximum and minimum total pressure loss. To better understand the flow mechanisms driving the vane clocking effects on the steady Stator 2 performance, time-resolved interrogations of the Stator 2 inlet flow field, surface pressure unsteadiness, and boundary layer response were conducted. The Stator 2 surface flows, both pressure unsteadiness and boundary layer transition, are influenced by vane clocking and interactions between Rotor 1 and Rotor 2, but neither of these results indicate a cause for the change in steady total pressure loss. Moreover, they are a result of upstream changes in the flow field: the interaction between the Stator 1 wake and Rotor 2 results in a circumferentially varying pattern which alters the inlet flow field for the downstream row, including the unsteadiness and frequency content in the tip and hub regions. Therefore, under different clocking configurations, Stator 2 experiences significantly different inlet blockage and unsteadiness from the Rotor 2 tip leakage flow and hub corner separation, which, in turn, shifts the radial blade loading distribution and subsequent loss development of Stator 2.

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