Experimental Investigation of the Clocking Effect in a 1.5-Stage Axial Turbine—Part I: Time-Averaged Results

2009 ◽  
Vol 131 (2) ◽  
Author(s):  
Sven König ◽  
Bernd Stoffel ◽  
M. Taher Schobeiri
Keyword(s):  
Pressure Loss ◽  
Total Pressure ◽  
Separation Bubble ◽  
Measurement Techniques ◽  
Suction Side ◽  
Total Pressure Loss ◽  
Experimental Investigations ◽  
Flow Angle ◽  
Factors Affecting ◽  
Lp Turbine

Comprehensive experimental investigations were conducted to get deeper insight into the physics of stator clocking in turbomachines. Different measurement techniques were used to investigate the influence of varying clocking positions on the highly unsteady flow field in a 1.5-stage axial low-pressure (LP) turbine. A Reynolds number typical for LP turbines as well as a two-dimensional blade design were chosen. Stator 2 was developed as a high-lift profile with a separation bubble on the suction side. This paper presents the results that were obtained by means of static pressure tappings and five-hole probes as well as the time-averaged results of unsteady x-wire measurements. The probes were traversed in different measuring planes for ten clocking positions. Depending on the clocking position, a variation in total pressure loss for Stator 2, a change of the rotor exit flow angle, and a dependency of the Stator 2 exit flow angle were found. The influence of these parameters on turbine efficiency was studied. Three main factors affecting the total pressure loss could be separated: the size of the separation bubble, the production of turbulent kinetic energy, and the strength of the periodic fluctuations downstream of Stator 2.

Experimental Investigation of the Clocking Effect in a 1.5-Stage Axial Turbine—Part II: Unsteady Results and Boundary Layer Behavior

2009 ◽  
Vol 131 (2) ◽  
Author(s):  
Sven König ◽  
Bernd Stoffel ◽  
M. Taher Schobeiri
Keyword(s):  
Boundary Layer ◽  
Separation Bubble ◽  
Measurement Techniques ◽  
Leading Edge ◽  
Suction Side ◽  
Experimental Investigations ◽  
Stress Signal ◽  
Upstream Direction ◽  
Loss Parameter ◽  
Lp Turbine

Comprehensive experimental investigations were conducted to get deeper insight into the physics of stator clocking in turbomachines. Different measurement techniques were used to investigate the influence of varying clocking positions on the highly unsteady flow field in a 1.5-stage axial low-pressure (LP) turbine. A Reynolds number typical for LP turbines as well as a two-dimensional blade design were chosen. Stator 2 was developed as a high-lift profile with a separation bubble on the suction side. This paper presents the results that were obtained by means of unsteady x-wire measurements upstream and downstream of Stator 2 and surface mounted hot-film measurements on the Stator 2 suction side. It was found that for the case when the Stator 1 wakes impinge close to the leading edge of Stator 2 the interaction between the Stator 1 and the rotor vortical structures takes place in proximity of the Stator 2 boundary layer, which leads to a shift of the transition point in the upstream direction. The major loss parameter concerning the Stator 2 aerodynamic performance could be attributed to the strength of the periodic fluctuations within the Stator 2 suction side boundary layer. A phase shift in the quasiwall shear stress signal in the front region of the Stator 2 vane was observed for different clocking positions.

scholarly journals Numerical and Experimental Investigations on Optimized 3D Compressor Airfoils

2018 ◽  
Author(s):  
Jan Mihalyovics ◽  
Christian Brück ◽  
Dieter Peitsch ◽  
Ilias Vasilopoulos ◽  
Marcus Meyer
Keyword(s):  
Pressure Loss ◽  
Total Pressure ◽  
Experimental Studies ◽  
Total Pressure Loss ◽  
Design Parameters ◽  
Flow Conditions ◽  
Flow Angle ◽  
Flow Phenomena ◽  
The Impact ◽  
Stator Design

The objective of the presented work is to perform numerical and experimental studies on compressor stators. This paper presents the modification of a baseline stator design using numerical optimization resulting in a new 3D stator. The Rolls Royce in-house compressible flow solver HYDRA was employed to predict the 3D flow, solving the steady RANS equations with the Spalart-Allmaras turbulence model, and its corresponding discrete adjoint solver. The performance gradients with respect to the input design parameters were used to optimize the stator blade with respect to the total pressure loss over a prescribed incidence range, while additionally minimizing the flow deviation from the axial direction at the stator exit. Non-uniform profile boundary conditions, being derived from the experimental measurements, have been defined at the inlet of the CFD domain. The presented results show a remarkable decrease in the axial exit flow angle deviation and a minor decrease in the total pressure loss. Experiments were conducted on two compressor blade sets investigating the three-dimensional flow in an annular compressor stator cascade. Comparing the baseline flow of the 42° turning stator shows that the optimized stator design minimizes the secondary flow phenomena. The experimental investigation discusses the impact of steady flow conditions on each stator design while focusing on the comparison of the 3D optimized design to the baseline case. The flow conditions were investigated using five-hole probe pressure measurements in the wake of the blades. Furthermore, oil-flow visualization was applied to characterize flow phenomena. These experimental results are compared with the CFD calculations.

Aerodynamic optimisation of a winglet-cavity tip in a high-pressure axial turbine cascade

2016 ◽  
Vol 232 (4) ◽  
pp. 649-663 ◽  
Author(s):  
Zhihua Zhou ◽  
Shaowen Chen ◽  
Songtao Wang
Keyword(s):  
High Pressure ◽  
Mass Flow ◽  
Pressure Loss ◽  
Total Pressure ◽  
Suction Side ◽  
Tip Clearance ◽  
Total Pressure Loss ◽  
Pressure Side ◽  
Tip Clearance Flow ◽  
Clearance Flow

Tip clearance flow between rotating blades and the stationary casing in high-pressure turbines is very complex and is one of the most important factors influencing turbine performance. The rotor with a winglet-cavity tip is often used as an effective method to improve the loss resulting from the tip clearance flow. In this study, an aerodynamic geometric optimisation of a winglet-cavity tip was carried out in a linear unshrouded high-pressure axial turbine cascade. For the purpose of shaping the efficient winglet geometry of the rotor tip, a novel parameterisation method has been introduced in the optimisation procedure based on the computational fluid dynamics simulation and analysis. The reliability of a commercial computational fluid dynamics code with different turbulence models was first validated by contrasting with the experimental results, and the numerical total pressure loss and flow angle using the Baseline k-omega Model (BSL κ-ω model) shows a better agreement with the test data. Geometric parameterisation of blade tips along the pressure side and suction side was adopted to optimise the tip clearance flow, and an optimal winglet-cavity tip was proven to achieve lower tip leakage mass flow rate and total pressure loss than the flat tip and cavity tip. Compared to the numerical results of flat tip and cavity tip, the optimised winglet-cavity design, with the winglet along the pressure side and suction side, had lower tip leakage mass flow rate and total pressure loss. It offered a 35.7% reduction in the change ratio [Formula: see text]. In addition, the optimised winglet along pressure side and suction side, respectively, by using the parameterisation method was studied for investigating the individual effect of the pressure-side winglet and suction-side winglet on the tip clearance flow. It was found that the suction-side extension of the optimal winglet resulted in a greater reduction of aerodynamic loss and leakage mass flow than the pressure-side extension of the optimal winglet. Moreover, with the analysis based on the tip flow pattern, the numerical results show that the pressure-side winglet reduced the contraction coefficient, and the suction-side winglet reduced the aerodynamic loss effectively by decreasing the driving pressure difference near the blade tips, the leakage flow velocity, and the interaction between the leakage flow and the main flow. Overall, a better aerodynamic performance can be obtained by adopting the pressure-side and suction-side winglet-cavity simultaneously.

Aerodynamic Performance of Suction-Side Gill Region Film Cooling

2010 ◽  
Vol 132 (3) ◽  
Author(s):  
Justin Chappell ◽  
Phil Ligrani ◽  
Sri Sreekanth ◽  
Terry Lucas ◽  
Edward Vlasic
Keyword(s):  
Mach Number ◽  
Film Cooling ◽  
Pressure Loss ◽  
Total Pressure ◽  
Suction Side ◽  
Total Pressure Loss ◽  
Pressure Loss Coefficient ◽  
Blowing Ratio ◽  
Exit Mach Number ◽  
Total Pressure Loss Coefficient

The performance of suction-side gill region film cooling is investigated using the University of Utah transonic wind tunnel and a simulated turbine vane in a two-dimensional cascade. The effects of film cooling hole orientation, shape, and number of rows, and their resulting effects on the aerodynamic losses, are considered for four different hole configurations: round axial (RA), shaped axial (SA), round radial (RR), and round compound (RC). The mainstream Reynolds number based on axial chord is 500,000, exit Mach number is 0.35, and the tests are conducted using the first row of holes, or both rows of holes at blowing ratios of 0.6 and 1.2. Carbon dioxide is used as the injectant to achieve density ratios of 1.77–1.99 similar to values present in operating gas turbine engines. Presented are the local distributions of total pressure loss coefficient, local normalized exit Mach number, and local normalized exit kinetic energy. Integrated aerodynamic losses (IAL) increase anywhere from 4% to 45% compared with a smooth blade with no film injection. The performance of each hole type depends on the airfoil configuration, film cooling configuration, mainstream flow Mach number, number of rows of holes, density ratio, and blowing ratio, but the general trend is an increase in IAL as either the blowing ratio or the number of rows of holes increase. In general, the largest total pressure loss coefficient Cp magnitudes and the largest IAL are generally present at any particular wake location for the RR or SA configurations, regardless of the film cooling blowing ratio and number of holes. The SA holes also generally produce the highest local peak Cp magnitudes. IAL magnitudes are generally lowest with the RA hole configuration. A one-dimensional mixing loss correlation for normalized IAL values is also presented, which matches most of the both rows data for RA, SA, RR, and RC hole configurations. The equation also provides good representation of the RA, RC, and RR first row data sets.

The Design, Analysis and Test of a Unique Gas Turbine Scroll for the F-35, Joint Strike Fighter

2006 ◽  
Author(s):  
Hakan Aksoy ◽  
Stony W. Kujala ◽  
Craig W. McKeever ◽  
Ly D. Nguyen
Keyword(s):  
Mach Number ◽  
Pressure Loss ◽  
Total Pressure ◽  
Total Pressure Loss ◽  
Flow Distortion ◽  
Cfd Simulations ◽  
Flow Angle ◽  
Auxiliary Power Unit ◽  
Auxiliary Power ◽  
Stringent Performance

The design of the APU (Auxiliary Power Unit) for the F-35 JSF (Joint Strike Fighter) focused on minimizing size and weight while meeting stringent performance goals. To help realize that goal, a unique turbine scroll was designed. The scroll design delivers air from the combustor to the turbine inlet with minimal loss and flow distortion while minimizing design space. CFD (Computational Fluid Dynamics) results of scroll total pressure loss and exit peripheral distribution of total pressure, Mach number, and flow angle are presented. Rig tests were utilized for measuring and validating the computed total pressure and Mach number distributions around the periphery of the scroll exit. Comparisons of the CFD simulations and test data indicate strong correlation in values of average total pressure loss, local total pressure loss and Mach number around the exit periphery.

Different Effects of Cantilevered and Shrouded Stators on Axial Compressor Performance

2017 ◽  
Author(s):  
Xiayi Si ◽  
Jinfang Teng ◽  
Xiaoqing Qiang ◽  
Jinzhang Feng
Keyword(s):  
Pressure Loss ◽  
Total Pressure ◽  
Loss Coefficient ◽  
Total Pressure Loss ◽  
Axial Position ◽  
Leakage Flow ◽  
Flow Angle ◽  
Design Point ◽  
Pressure Loss Coefficient ◽  
Total Pressure Loss Coefficient

Numerical simulations with the steady 3D RANS were performed on the rear stage of a modern high pressure compressor. The labyrinth seal cavity model of the shrouded stator was simplified according to the actual stator structure, which the seal cavity gap is 1% of blade height. Several typical configurations (shrouded stator, idealized stator and cantilevered stators) were designed and carried out, and cantilevered stators contained no gap, small gap (CS1%), design gap (CS2.5%) and large gap (CS4%/CS5%). The results indicate due to the effect of leakage flow from 1% span seal cavity gap, the total pressure loss of SS is larger than IS, while IS instead of SS in the process of the compressor design, the stall margin will be enlarged nearly 6% numerically. At the design point, when the hub gap is 3.5% span clearance CS has the same loss with IS, and when the hub gap is 4.5% span clearance CS has almost the same loss with SS. Among all operation range, the total pressure loss of S1 increases with the increase of the hub clearance. When the hub gap is 0 (CS0), there is no leakage flow and the loss is the least. At the design point, comparing with SS, the total pressure loss coefficient of CS0 decreases 18.34%, CS2.5% decreases 8.46% and IS decreases 6.45%. It means if the cantilevered stator with 2.5% span hub clearance were adopted in the HPC, the performance would be better than the shrouded stator. However, because of the matching condition, the rotor that follows after cantilevered stator should be redesigned according to blade loading and inlet flow angle changed. The performance of cantilevered stator is impacted of various hub clearance, the loss below 25% span increases significantly with hub clearance, the maximum value of outlet flow angle deviation is 2.3 degree. The stator hub peak loading is shifted upstream toward the leading edge when hub clearance size is increased. The total pressure loss coefficient and pressure coefficient at different axial position had the function relation. When the hub clearance increases, the position of double leakage flow start backwards, in the rear part of stator the secondary flow becomes stronger leading to more mixing loss and lower total pressure.

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.

Aerodynamic Optimization of a Winglet-Shroud Tip Geometry for a Linear Turbine Cascade

2017 ◽  
Vol 139 (10) ◽  
Author(s):  
Min Zhang ◽  
Yan Liu ◽  
Tianlong Zhang ◽  
Mengchao Zhang ◽  
Ying He
Keyword(s):  
Pressure Loss ◽  
Total Pressure ◽  
Separation Bubble ◽  
Aerodynamic Performance ◽  
Total Pressure Loss ◽  
Turbine Cascade ◽  
Aerodynamic Optimization ◽  
Tip Leakage ◽  
Viscous Loss ◽  
Partial Shroud

This paper presents a continued study on a previously investigated novel winglet-shroud (WS) (or partial shroud) geometry for a linear turbine cascade. Various widths of double-side winglets (DSW) and different locations of a partial shroud are considered. In addition, both a plain tip and a full shroud tip are applied as the datum cases which were examined experimentally and numerically. Total pressure loss and viscous loss coefficients are comparatively employed to execute a quantitative analysis of aerodynamic performance. The effectiveness of various widths (w) of DSW set at 3%, 5%, 7%, and 9% of the blade pitch (p) is numerically investigated. Skin-friction lines on the tip surface indicate that different DSW cases do not alter flow field features including the separation bubble and reattachment flow within the tip gap region, even for the case with the broadest width (w/p = 9%). However, the pressure side extension of the DSW exhibits the formation of separation bubble, while the suction side platform of the DSW turns the tip leakage vortex (TLV) away from the suction surface (SS). Meanwhile, the horse-shoe vortex (HV) near the casing is not generated even for the case with the smallest width (w/p = 3%). As a result, both the tip leakage and the upper passage vortices are weakened and further dissipated with wider w/p in the DSW cases. Larger width of the DSW geometry is indeed able to improve the aerodynamic performance, but only to a slight degree. With the w/p increasing from 3% to 9%, the mass-averaged total pressure loss coefficient over an exit plane is reduced by only 2.61%. Therefore, considering both the enlarged (or reduced) tip area and the enhanced (or deteriorated) performance compared to the datum cases, a favorable width of w/p = 5% is chosen to design the WS structure. Three locations for the partial shroud (linkage segment) are devised, locating them near the leading edge, in the middle and close to the trailing edge, respectively. Results demonstrate that all three cases of the WS design have advantages over the DSW arrangement in lessening the aerodynamic loss, with the middle linkage segment location producing the optimal effect. This conclusion verifies the feasibility of the previously studied WS configuration.

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.

Numerical Research on the Mixing Mechanism of Lobed Mixer With New De-Swirling Structure

2016 ◽  
Author(s):  
Zhijun Lei ◽  
Jianbo Gong ◽  
Yanfeng Zhang ◽  
Shangmei Su ◽  
Chunyan Hu
Keyword(s):  
Numerical Simulation ◽  
Pressure Loss ◽  
Total Pressure ◽  
Recirculation Zone ◽  
Swirling Flow ◽  
Separation Bubble ◽  
Total Pressure Loss ◽  
Mixing Efficiency ◽  
Flow Mixing ◽  
Inlet Swirl

A detailed numerical simulation is presented to investigate the new de-swirling methods and their effect on the mixing mechanisms of a turbofan mixer with 12 lobes. The numerical simulation employed a commercial solver, ANSYS CFX, using k-ω SST model. The core-to-bypass temperature ratio and pressure ratio were set to 2.59, and 0.97 respectively, giving the Mach number of 0.66 and bypass ratio of 2.65 at mixing nozzle outlet. The inlet swirl typically accelerates the jet-flow mixing by enhancing the vortices intensity and interaction, but leakage swirling flow can cause a three-dimensional separation bubble and the recirculation zone resulting in the dramatic increasing the total pressure loss and thrust loss. Removal of the leakage swirling flow between the lobes’ trough and centre-body was the key to limit the negative influence of inlet swirl. Two IGV design were investigated, DS1 and DS2. DS1 was installed at the upstream of the lobed mixer, could remove the negative effect of inlet swirl properly, but also inhibited the active role of the inlet swirl. The total pressure and thrust loss reduced by 0.31% and 3.8%, respectively, but the mixing efficiency also decreased by 1.72%. DS2, an integrated strut with the lobed mixer design, not only ensured the structure strength of the lobed mixer, but also reduced the length and weight of the exhaust system. This method suppressed the flow separation bubble on centre-body to some extent, and eliminated the recirculation zone downstream of the cenrebody, resulting in the total pressure loss decrease of 0.31% and thrust gain of 3.63%. On the other hand, the method DS2 also made full use of the inlet swirl to enhance the jet-flow mixing, resulting in the mixing efficiency increased 1.54% compared with that of the DS1 case. Under the off-design conditions with the incidence angle of ±10°, the aerodynamic performance of the DS2 cases didn’t changed too much such as the DS1 cases.

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