Numerical Investigation of the Effect of an Axial Pre-Swirl Nozzle With a Radial Angle in a Pre-Swirl Rotor-Stator System

2021 ◽  
Author(s):  
Gang Zhao ◽  
Shuiting Ding ◽  
Tian Qiu ◽  
Shenghui Zhang
Keyword(s):  
Gas Turbines ◽  
Pressure Loss ◽  
Total Pressure ◽  
Discharge Coefficient ◽  
Pressure Ratio ◽  
Temperature Drop ◽  
Tangential Velocity ◽  
Total Pressure Loss ◽  
Adiabatic Effectiveness ◽  
Cooling Air

Abstract Pre-swirl nozzles are often used in gas turbines to deliver the cooling air to the turbine blades. The static axial nozzles swirl the cooling air in the direction of rotation of the turbine disk, thereby reducing the relative total temperature of the air. Most studies about nozzles focus on its shape, radial location, tangential angle to reduce the pressure loss and increase the temperature drop of the pre-swirl system, but few of them consider the benefit of a radial angle of nozzles. This paper investigated numerically the performance of a pre-swirl system whose pre-swirl nozzles have a radial angle. Six radial angles are selected to study the flow dynamics of a direct-transfer pre-swirl system in terms of the total pressure loss coefficient of the pre-swirl cavity, the discharge coefficient of the receiver holes, and the adiabatic effectiveness. It is shown that the nozzles with radial angles can adjust the tangential velocity and radial velocity and thus can influence the performance of a pre-swirl system. There is a lowerest value of total pressure loss in pre-swirl cavity, that is α = 90°, which can hardly be influenced by the radial angle of nozzle and pressure ratio π. For a specific swirl ratio β∞, there exists an optimal αopt where the discharge coefficient of receiver hole is maximum. Moreover, αopt decreases as pressure ratio π increases. And so is the adiabatic effectiveness Θad.

scholarly journals Impact of a Cooled Cooling Air System on the External Aerodynamics of a Gas Turbine Combustion System

2017 ◽  
Vol 139 (5) ◽  
Author(s):  
A. Duncan Walker ◽  
Bharat Koli ◽  
Liang Guo ◽  
Peter Beecroft ◽  
Marco Zedda
Keyword(s):  
Gas Turbines ◽  
Pressure Loss ◽  
Total Pressure ◽  
Outer Wall ◽  
Total Pressure Loss ◽  
Test Facility ◽  
Combustion System ◽  
Air System ◽  
External Aerodynamics ◽  
Cooling Air

To manage the increasing turbine temperatures of future gas turbines a cooled cooling air system has been proposed. In such a system some of the compressor efflux is diverted for additional cooling in a heat exchanger (HX) located in the bypass duct. The cooled air must then be returned, across the main gas path, to the engine core for use in component cooling. One option is do this within the combustor module and two methods are examined in the current paper; via simple transfer pipes within the dump region or via radial struts in the prediffuser. This paper presents an experimental investigation to examine the aerodynamic impact these have on the combustion system external aerodynamics. This included the use of a fully annular, isothermal test facility incorporating a bespoke 1.5 stage axial compressor, engine representative outlet guide vanes (OGVs), prediffuser, and combustor geometry. Area traverses of a miniature five-hole probe were conducted at various locations within the combustion system providing information on both flow uniformity and total pressure loss. The results show that, compared to a datum configuration, the addition of transfer pipes had minimal aerodynamic impact in terms of flow structure, distribution, and total pressure loss. However, the inclusion of prediffuser struts had a notable impact increasing the prediffuser loss by a third and consequently the overall system loss by an unacceptable 40%. Inclusion of a hybrid prediffuser with the cooled cooling air (CCA) bleed located on the prediffuser outer wall enabled an increase of the prediffuser area ratio with the result that the system loss could be returned to that of the datum level.

scholarly journals Optimization Design and Experimental Study of Low-Pressure Axial Fan with Forward-Skewed Blades

2007 ◽  
Vol 2007 ◽  
pp. 1-10 ◽  
Author(s):  
Li Yang ◽  
Ouyang Hua ◽  
Du Zhao-Hui
Keyword(s):  
Experimental Study ◽  
Pressure Loss ◽  
Total Pressure ◽  
Optimization Design ◽  
Pressure Ratio ◽  
Low Pressure ◽  
Broadband Noise ◽  
Total Pressure Loss ◽  
Aerodynamic Noise ◽  
Axial Fan

This paper presents an experimental study of the optimization of blade skew in low pressure axial fan. Using back propagation (BP) neural network and genetic algorithm (GA), the optimization was performed for a radial blade. An optimized blade is obtained through blade forward skew. Measurement of the two blades was carried out in aerodynamic and aeroacoustic performance. Compared to the radial blade, the optimized blade demonstrated improvements in efficiency, total pressure ratio, stable operating range, and aerodynamic noise. Detailed flow measurement was performed in outlet flow field for investigating the responsible flow mechanisms. The optimized blade can cause a spanwise redistribution of flow toward the blade midspan and reduce tip loading. This results in reduced significantly total pressure loss near hub and shroud endwall region, despite the slight increase of total pressure loss at midspan. In addition, the measured spectrums show that the broadband noise of the impeller is dominant.

Optimal Design of the Volute for a Turbocharger Radial Flow Compressor

2014 ◽  
Author(s):  
Mohammad Mojaddam ◽  
Ali Hajilouy-Benisi ◽  
Mohammad Reza Movahhedy
Keyword(s):  
Pressure Loss ◽  
Total Pressure ◽  
Radial Flow ◽  
Pressure Ratio ◽  
Radial Force ◽  
Total Pressure Loss ◽  
Effective Parameters ◽  
Flow Rates ◽  
Design Point ◽  
Flow Compressor

In this research the design methods of radial flow compressor volutes are reviewed and the main criterions in volute primary designs are recognized and most effective ones are selected. The effective parameters i.e. spiral cross section area, circumferential area distribution, exit cone and tongue area of the compressor volute are parametrically studied to identifythe optimum values. A numerical model is prepared and verified through experimental data which are obtained from the designed turbocharger test rig. Different volutes are modeled and numerically evaluated using the same impeller and vane-less diffuser. For each model, the volute total pressure ratio, static pressure recovery and total pressure loss coefficients and the radial force on the impeller are calculated for different mass flow rates at design point and off-design conditions. The volute which shows better performanceand causes lower the net radial force on the impeller, at desiredmass flow rates is selected as an optimal one. The results show the volute design approach differences at the design point and off-design conditions. Improving the pressure ratio and reducing total pressure loss at design point, may result inthe worse conditions at off-design conditions as well as increasing radial force on the impeller.

scholarly journals Investigation of Wash Fluid Preheating on the Effectiveness of Online Compressor Washing in Industrial Gas Turbines

2020 ◽  
Author(s):  
Roupa Agbadede ◽  
Biweri Kainga
Keyword(s):  
Gas Turbines ◽  
Pressure Loss ◽  
Total Pressure ◽  
Loss Coefficient ◽  
Total Pressure Loss ◽  
Compressor Cascade ◽  
Flow Angle ◽  
Pressure Loss Coefficient ◽  
Industrial Gas Turbines ◽  
Total Pressure Loss Coefficient

Abstract This study presents an investigation of wash fluid preheating on the effectiveness of online compressor washing in industrial gas turbines. Crude oil was uniformly applied on the compressor cascade blades surfaces using a roller brush, and carborundum particles were ingested into the tunnel to create accelerated fouled blades. Demineralized water was preheated to 500C using the heat coil provided in the tank. When fouled blades washed with preheated demineralized and the one without preheating were compared, it was observed that there was little or no difference in terms of total pressure loss coefficient and exit flow angle. However, when the fouled and washed cases were compared, there was a significant different in total pressure loss coefficient and exit flow angle.

The Effect of Tip Clearance and Hub Rotation on the Performance of an Axial Compressor Stator

2019 ◽  
Author(s):  
Chao Jiang ◽  
Jun Hu ◽  
Jiayu Wang ◽  
Jun Li ◽  
Rong Xu
Keyword(s):  
Pressure Loss ◽  
Total Pressure ◽  
High Speed ◽  
Pressure Ratio ◽  
Important Influence ◽  
Tip Clearance ◽  
Total Pressure Loss ◽  
Computation Method ◽  
Efficiency Curve ◽  
Total Pressure Ratio

Abstract In this paper, 1.5-stage high-speed compressor stator was studied using numerical computation method. Four gap cases were calculated under the condition of the hub being stationary or rotating, and the characteristic curves of the 1.5-stage compressor was obtained. Firstly, the influence of the change of the gap on the total pressure ratio and the efficiency curve was studied when the state of the hub is fixed. Then, the influence of the rotation of the hub on the total pressure ratio and the efficiency curve was discussed when the tip clearance is fixed. Finally, the total pressure loss of the stator channel would be analyzed. The above research would make people understand that the relative motion of the end wall has an important influence on the performance of the axial-flow compressor; when hub is stationary, the optimal gap is greater than 0, while when hub was rotating, the optimal gap was 0; and recognize that the variation of clearance and the motion state of the hub have an important influence on the distribution of total pressure loss along the span.

Numerical Study on the Effects of Blade Leading Edge Shape to the Performance of Supersonic Rotor

2003 ◽  
Author(s):  
Kicheol Park
Keyword(s):  
Rotational Speed ◽  
Pressure Loss ◽  
Total Pressure ◽  
Performance Test ◽  
Numerical Study ◽  
Pressure Ratio ◽  
Leading Edge ◽  
Total Pressure Loss ◽  
Manufacturing Cost ◽  
Blade Leading Edge

Recently, it is required to design a fan and compressor with higher stage pressure ratio while maintaining adiabatic efficiency high also. To increase the stage pressure ratio, blade rotational speed or diffusion factor should be increased. In the case of increased rotational speed, relative speed of flow at blade leading edge is well supersonic. With supersonic rotor blade, total pressure loss is mainly due to leading edge shock waves and the thickness should be thin enough to minimize this. As a result, the blade is like to be week in terms of mechanical strength and the manufacturing cost would be increased because high-precision NC machining is required. Furthermore, it is one of the biggest hurdles to maintain proper level of thickness while one making small stages. In this paper, aerodynamic performance of supersonic rotor blades with different leading edge thickness and shapes are calculated using the finite volume method. The effects of blade leading edge shape and thickness to the performance are investigated especially in terms of total pressure loss and the already known loss correlations of leading edge thickness are examined. Subsequently this will be verified by performance test on rig.

scholarly journals Accurate 2-D Modelling of Transonic Compressor Cascade Aerodynamics

Aerospace ◽  
2019 ◽  
Vol 6 (5) ◽  
pp. 57 ◽  
Author(s):  
Tommaso Piovesan ◽  
Andrea Magrini ◽  
Ernesto Benini
Keyword(s):  
Pressure Loss ◽  
Total Pressure ◽  
Static Pressure ◽  
Pressure Ratio ◽  
Loss Coefficient ◽  
Total Pressure Loss ◽  
Compressor Cascade ◽  
Transonic Compressor ◽  
Pressure Loss Coefficient ◽  
Total Pressure Loss Coefficient

Modern aeronautic fans are characterised by a transonic flow regime near the blade tip. Transonic cascades enable higher pressure ratios by a complex system of shockwaves arising across the blade passage, which has to be correctly reproduced in order to predict the performance and the operative range. In this paper, we present an accurate two-dimensional numerical modelling of the ARL-SL19 transonic compressor cascade. A large series of data from experimental tests in supersonic wind tunnel facilities has been used to validate a computational fluid dynamic model, in which the choice of turbulence closure resulted critical for an accurate reproduction of shockwave-boundary layer interaction. The model has been subsequently employed to carry out a parametric study in order to assess the influence of main flow variables (inlet Mach number, static pressure ratio) and geometric parameters (solidity) on the shockwave pattern and exit status. The main objectives of the present work are to perform a parametric study for investigating the effects of the abovementioned variables on the cascade performance, in terms of total-pressure loss coefficient, and on the shockwave pattern and to provide a quite large series of data useful for a preliminary design of a transonic compressor rotor section. After deriving the relation between inlet and exit quantities, peculiar to transonic compressors, exit Mach number, mean exit flow angle and total-pressure loss coefficient have been examined for a variety of boundary conditions and parametrically linked to inlet variables. Flow visualisation has been used to describe the shock-wave pattern as a function of the static pressure ratio. Finally, the influence of cascade solidity has been examined, showing a potential reduction of total-pressure loss coefficient by employing a higher solidity, due to a significant modification of shockwave system across the cascade.

Characteristics of a Turbine Cascade at Low Reynolds Numbers

1997 ◽  
Author(s):  
Koji Murata ◽  
Hiroyuki Abe ◽  
Yasukata Tsutsui
Keyword(s):  
Reynolds Number ◽  
Gas Turbines ◽  
Pressure Loss ◽  
Total Pressure ◽  
Separation Bubble ◽  
Reynolds Numbers ◽  
Boundary Layer Separation ◽  
Total Pressure Loss ◽  
Turbine Cascade ◽  
Low Reynolds

The aerodynamic characteristics of turbine cascades are thought to be relatively satisfactory due to the favorable pressure of the accelerating flow. But within the low Reynolds number region of 50,000 where the 300kW ceramic gas turbines which are being developed under the New-Sunshine Project of Japan operate, the characteristics such as boundary layer separation and reattachment which lead to prominent power losses cannot be easily predicted. In this research, experiments have been conducted to evaluate the performance of a linear two dimensional turbine cascade. Surface pressure distributions of the airfoil were measured for a range of blade chord Reynolds numbers from 40,000 to 160,000 (at inlet), and at 1.3% inlet turbulence intensity. In addition, the wake of the cascade was measured simultaneously using a five hole pilot tube. Traverses of the wake show that there is a drastic increase in the mean total pressure loss at the wake between the Reynolds number of 80,000 to 40,000, and in some conditions, a rise as much as 10% was confirmed. Thus, in accordance with the pressure distribution of the surface of the airfoil, a relation between the total pressure loss and the length of the laminar separation bubble formed on the airfoil could be seen.

scholarly journals Investigation of the Effects of Different Working Fluids on Compressor Cascade Performance

2021 ◽  
Vol 11 (5) ◽  
pp. 1989
Author(s):  
Zhitao Tian ◽  
Chengze Wang ◽  
Qun Zheng
Keyword(s):  
Physical Properties ◽  
Specific Heat ◽  
Pressure Loss ◽  
Total Pressure ◽  
Static Pressure ◽  
Pressure Ratio ◽  
Total Pressure Loss ◽  
Compressor Cascade ◽  
Working Fluids ◽  
Specific Heat Ratio

The compressor of closed Brayton cycle (CBC) plant operating with working fluid other than air is a vital element of the energy conversion unit. However, due to insufficient understanding of the influence of the physical properties of working fluids on the performance of the compressor, the actual working conditions and design conditions of the compressor’s performance deviate greatly. In this paper, the objective is to analyze the influence mechanism of the physical properties on the performance of the cascade of compressor (static pressure ratio and total pressure loss coefficient). Therefore, the impact of a specific heat ratio on the performance of the compressor cascade is studied utilizing carbon dioxide (γ = 1.29), air and carbon monoxide (γ = 1.4), argon and helium (γ = 1.667). Moreover, the relationships of static pressure ratio and total pressure loss coefficient with physical properties of the working fluids are analyzed in the compressor cascade. It is established that a higher specific heat ratio fluid gives a higher coefficient of total pressure loss and static pressure ratio in contrast to smaller specific heat ratio at matching inlet Reynolds number and Mach number.

scholarly journals Correlation between total pressure losses of highly loaded annular diffusers and integral stage design parameters

2018 ◽  
Vol 2 ◽  
pp. I9AB30 ◽  
Author(s):  
Dajan Mimic ◽  
Christoph Jätz ◽  
Florian Herbst
Keyword(s):  
Boundary Layer ◽  
Gas Turbines ◽  
Pressure Loss ◽  
Total Pressure ◽  
Total Pressure Loss ◽  
Pressure Recovery ◽  
Design Parameters ◽  
Pressure Losses ◽  
Tip Leakage ◽  
Highly Loaded

Diffusers convert kinetic flow energy into a rise in static pressure. This pressure recovery is the primary aerodynamic design objective for exhaust gas diffusers in power-generating steam and gas turbines. The total pressure loss is an equally important diffuser design parameter. It is strongly linked to the pressure recovery and the residual kinetic energy of the diffuser outlet flow. A reduction benefits the overall thermodynamic cycle, which requires the adjacent components of a diffuser to be included in the design process. This paper focuses on the total pressure losses in the boundary layer of a highly loaded annular diffuser. Due to its large opening angle the diffuser is susceptible to flow separation under uniform inlet conditions, which is a major source for total pressure losses. However, the unsteady tip leakage vortices of the upstream rotor, which are a source of losses, stabilise the boundary layer and prevent separation. Experiments and unsteady numerical simulation conducted show that the total pressure loss reduction caused by the delayed boundary layer separation exceed the vortex-induced losses by far. This flow interaction between the rotor and diffuser consequently decreases the overall total pressure losses. The intensity of the tip leakage vortex is linked to three rotor design parameters, namely work coefficient, flow coefficient and reduced blade-passing frequency. Based on these parameters, we propose a semi-empiric correlation to predict and evaluate the change in total pressure losses with regards to design operating conditions.

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