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

2018 ◽  
Vol 140 (7) ◽  
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 three-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.

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.

Flow Measurements in a Fishtail Diffuser With Strong Curvature

1999 ◽  
Vol 121 (2) ◽  
pp. 410-417 ◽  
Author(s):  
M. I. Yaras
Keyword(s):  
Boundary Layer ◽  
Pressure Loss ◽  
Total Pressure ◽  
Static Pressure ◽  
Secondary Flows ◽  
Total Pressure Loss ◽  
Cross Sectional ◽  
Inlet Flow ◽  
Inlet Conditions ◽  
Strong Curvature

The paper presents detailed measurements of the incompressible flow development in a large-scale 90 deg curved diffuser with strong curvature and significant streamwise variation in cross-sectional aspect ratio. The flow path approximates the so-called fishtail diffuser utilized on small gas turbine engines for the transition between the centrifugal impeller and the combustion chamber. Two variations of the inlet flow, differing in boundary layer thickness and turbulence intensity, are considered. Measurements consist of three components of velocity, static pressure and total pressure distributions at several cross-sectional planes throughout the diffusing bend. The development and mutual interaction of multiple pairs of streamwise vortices, redistribution of the streamwise flow under the influence of these vortices, the resultant streamwise variations in mass-averaged total-pressure and static pressure, and the effect of inlet conditions on these aspects of the flow are examined. The strengths of the vortical structures are found to be sensitive to the inlet flow conditions, with the inlet flow comprising a thinner boundary layer and lower turbulence intensity yielding stronger secondary flows. For both inlet conditions a pair of streamwise vortices develop rapidly within the bend, reaching their peak strength at about 30 deg into the bend. The development of a second pair of vortices commences downstream of this location and continues for the remainder of the bend. Little evidence of the first vortex pair remains at the exit of the diffusing bend. The mass-averaged total pressure loss is found to be insensitive to the range of inlet-flow variations considered herein. However, the rate of generation of this loss along the length of the diffusing bend differs between the two test cases. For the case with the thinner inlet boundary layer, stronger secondary flows result in larger distortion of the streamwise velocity field. Consequently, the static pressure recovery is somewhat lower for this test case. The difference between the streamwise distributions of measured and ideal static pressure is found to be primarily due to total pressure loss in the case of the thick inlet boundary layer. For the thin inlet boundary layer case, however, total pressure loss and flow distortion are observed to influence the pressure recovery by comparable amounts.

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.

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.

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.

Location Effect of Boundary Layer Suction on Compressor Hub-Corner Separation

2014 ◽  
Author(s):  
Ping-Ping Chen ◽  
Wei-Yang Qiao ◽  
Karsten Liesner ◽  
Robert Meyer
Keyword(s):  
Boundary Layer ◽  
Pressure Loss ◽  
Total Pressure ◽  
Three Dimensional ◽  
Base Flow ◽  
Total Pressure Loss ◽  
Compressor Cascade ◽  
Suction Surface ◽  
Corner Separation ◽  
Boundary Layer Suction

The large secondary flow area in the compressor hub-corner region usually leads to three-dimensional separation in the passage with large amounts of total pressure loss. In this paper numerical simulations of a linear high-speed compressor cascade, consisting of five NACA 65-K48 stator profiles, were performed to analyze the flow mechanism of hub-corner separation for the base flow. Experimental validation is used to verify the numerical results. Active control of the hub-corner separation was investigated by using boundary layer suction. The influence of the selected locations of the endwall suction slot was investigated in an effort to quantify the gains of the compressor cascade performance. The results show that the optimal chordwise location should contain the development section of the three-dimensional corner separation downstream of the 3D corner separation onset. The best pitchwise location should be close enough to the vanes’ suction surface. Therefore the optimal endwall suction location is the MTE slot, the one from 50% to 75% chord at the hub, close to the blade suction surface. By use of the MTE slot with 1% suction flow ratio, the total-pressure loss is substantially decreased by about 15.2% in the CFD calculations and 9.7% in the measurement at the design operating condition.

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.

Numerical Investigation of the Influence of Real World Blade Profile Variations on the Aerodynamic Performance of Transonic Nozzle Guide Vanes

2011 ◽  
Vol 134 (2) ◽  
Author(s):  
R. Edwards ◽  
A. Asghar ◽  
R. Woodason ◽  
M. LaViolette ◽  
K. Goni Boulama ◽  
...  
Keyword(s):  
Boundary Layer ◽  
Flow Field ◽  
Flow Pattern ◽  
Pressure Loss ◽  
Total Pressure ◽  
Aerodynamic Performance ◽  
Guide Vanes ◽  
Trailing Edges ◽  
Nozzle Guide ◽  
Nozzle Guide Vanes

This paper addresses the issue of aerodynamic consequences of small variations in airfoil profile. A numerical comparison of flow field and cascade pressure losses for two representative repaired profiles and a reference new vane were made. Coordinates for the three airfoil profiles were obtained from the nozzle guide vanes of refurbished turboshaft engines using 3D optical scanning and digital modeling. The repaired profiles showed differences in geometry in comparison with the new vane, particularly near the leading and trailing edges. A numerical simulation was conducted using a commercial CFD code, which uses the finite volume approach for solving the governing equations. The computational predictions of the aerodynamic performance were compared with experimental results obtained from a cascade consisting of blades with the same airfoil profiles. The CFD analysis was performed for the cascade at subsonic inlet and transonic exit conditions. Boundary layer growth, wake formation, and shock boundary layer interactions were observed in the two-dimensional computations. The flow field showed the presence of shock waves downstream of the passage throat and near the trailing edges of the blades. A conspicuous change in flow pattern due to subtle variation in airfoil profile was observed. The calculated flow field was compared with the flow pattern visualized in the experimental test rig using the schlieren method. The total pressure calculation for the cascade exit showed an increase in pressure loss for one of the off-design profiles. The pressure loss calculations were also compared with the multihole total pressure probe measurement in the transonic cascade rig.

Accurate Estimation of Profile Losses and Analysis of Loss Generation Mechanisms in a Turbine Cascade

2017 ◽  
Vol 139 (12) ◽  
Author(s):  
D. Lengani ◽  
D. Simoni ◽  
M. Ubaldi ◽  
P. Zunino ◽  
F. Bertini ◽  
...  
Keyword(s):  
Boundary Layer ◽  
Pressure Loss ◽  
Total Pressure ◽  
Loss Coefficient ◽  
Total Pressure Loss ◽  
Accurate Estimation ◽  
Loss Mechanism ◽  
Pressure Loss Coefficient ◽  
Generation Mechanisms ◽  
Total Pressure Loss Coefficient

The paper analyzes losses and the loss generation mechanisms in a low-pressure turbine (LPT) cascade by proper orthogonal decomposition (POD) applied to measurements. Total pressure probes and time-resolved particle image velocimetry (TR-PIV) are used to determine the flow field and performance of the blade with steady and unsteady inflow conditions varying the flow incidence. The total pressure loss coefficient is computed by traversing two Kiel probes upstream and downstream of the cascade simultaneously. This procedure allows a very accurate estimation of the total pressure loss coefficient also in the potential flow region affected by incoming wake migration. The TR-PIV investigation concentrates on the aft portion of the suction side boundary layer downstream of peak suction. In this adverse pressure gradient region, the interaction between the wake and the boundary layer is the strongest, and it leads to the largest deviation from a steady loss mechanism. POD applied to this portion of the domain provides a statistical representation of the flow oscillations by splitting the effects induced by the different dynamics. The paper also describes how POD can dissect the loss generation mechanisms by separating the contributions to the Reynolds stress tensor from the different modes. The steady condition loss generation, driven by boundary layer streaks and separation, is augmented in the presence of incoming wakes by the wake–boundary layer interaction and by the wake dilation mechanism. Wake migration losses have been found to be almost insensitive to incidence variation between nominal and negative (up to −9 deg) while at positive incidence, the losses have a steep increase due to the alteration of the wake path induced by the different loading distribution.

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