Influence of Mach Number on Profile Loss of Axial-Flow Gas Turbine Blades

2016 ◽  
Author(s):  
M. D. Kibsey ◽  
S. A. Sjolander
Keyword(s):  
Mach Number ◽  
Turbine Blades ◽  
Base Pressure ◽  
Current Profile ◽  
Trailing Edge ◽  
Complex Behaviour ◽  
Number Range ◽  
Profile Losses ◽  
Mach Numbers ◽  
Profile Loss

The current profile loss prediction methods for axial turbine blades usually predict a monotonic increase in profile losses at outlet Mach numbers above 1.0, while linear cascade testing in the literature has revealed a more complex behaviour. An objective of this investigation was to help clarify the flow features that are most influential on the profile losses in the transonic and supersonic regimes. Four linear cascades of turbine blades were investigated both experimentally and computationally, at design incidence. Measurements were carried out over an outlet Mach number range of roughly 0.5 to 1.4, and a Reynolds number range of about 5 × 105 to 1.4×106. It was found that the profile losses of the four cascades exhibited a loss “plateau”, where the total pressure loss coefficient became approximately constant over a range of outlet Mach numbers spanning the low supersonic range. Cascades of different geometries exhibited different extents of this loss plateau, and a commonly used Mach number correction for profile losses did not capture the behaviour. In the literature, a relationship has been observed between the base pressure and the profile losses. The base pressure was linked to the losses in the trailing edge wake and in the trailing edge shock system. For this reason, base pressure data were obtained from blades instrumented with a static tap at the trailing edge, and also from computational fluid dynamics (CFD). The results provided insight into the role of the base pressure in the profile losses through the transonic regime. It was concluded from this study that an accurate prediction of the base pressure may serve as a basis for a revised Mach number correction to be applied to the profile loss correlation in the transonic and supersonic flow regimes.

Auto and Cross Correlation Measurements in a Turbine Cascade Using a Digital Correlator

1982 ◽  
Author(s):  
P. J. Bryanston-Cross ◽  
J. J. Camus
Keyword(s):  
Mach Number ◽  
High Speed ◽  
Cross Correlation ◽  
Turbine Blades ◽  
Image Plane ◽  
Strongly Correlated ◽  
Trailing Edge ◽  
Number Range ◽  
Cross Correlations ◽  
Digital Correlator

A simple technique has been developed which samples the dynamic image plane information of a schlieren system using a digital correlator. Measurements have been made in the passages and in the wakes of transonic turbine blades in a linear cascade. The wind tunnel runs continuously and has independently variable Reynolds and Mach number. As expected, strongly correlated vortices were found in the wake and trailing edge region at 50 KHz. Although these are strongly coherent we show that there is only limited cross-correlation from wake to wake over a Mach no. range M = 0.5 to 1.25 and variation of Reynolds number from 3 × 105 to 106. The trailing edge fluctuation cross correlations were extended both upstream and downstream and preliminary measurements indicate that this technique can be used to obtain information on wake velocity. The vortex frequency has also been measured over the same Mach number range for two different cascades. The results have been compared with high speed schlieren photographs.

scholarly journals The Influence of Trailing Edge Ejection on the Base Pressure in Transonic Turbine Cascades

1982 ◽  
Author(s):  
C. H. Sieverding
Keyword(s):  
Turbine Blades ◽  
Density Ratio ◽  
Main Flow ◽  
Coolant Flow ◽  
Base Pressure ◽  
Trailing Edge ◽  
Number Range ◽  
Transonic Turbine ◽  
Turbine Cascades ◽  
Transonic Cascade

This paper summarizes the results of base pressure studies on transonic turbine blades in presence of an ejection of coolant flow from a slot in the trailing edge. The first part of the paper reports on tests carried out on a enlarged model of the overhang section of a typical transonic cascade. This model provides valuable information about the detailed trailing edge pressure distribution and points to an asymmetric evolution of the base pressure on both sides of the slot in presence of a bleed. The second part of the paper presents experimental results from cascade tests covering an outlet Mach number range from M2, is = 0.5 to 1.35. These experiments underline the importance of the coolant flow impact on the base pressure and confirm the asymmetry of the base pressure with respect to the cooling slot. Tests with different coolant flow gases point to the significance of a proper simulation of the density ratio between coolant flow and main flow.

The Base Pressure and Loss of a Family of Four Turbine Blades

1988 ◽  
Vol 110 (1) ◽  
pp. 9-17 ◽  
Author(s):  
L. Xu ◽  
J. D. Denton
Keyword(s):  
Dominant Role ◽  
Turbine Blades ◽  
Base Pressure ◽  
Trailing Edge ◽  
Edge Geometry ◽  
Trailing Edges ◽  
Loss Prediction ◽  
Turbine Cascades ◽  
Profile Loss ◽  
Made In

Measurements of the effect of trailing edge geometry on the base pressure and loss of a family of four turbine cascades are presented. The measurements were made in the transonic range of Mach number from 0.8 to 1.2. It is found that, for blades with typical trailing edge thickness, the trailing edge loss is the major source of profile loss at these speeds and that the base pressure plays a dominant role in determining the loss. For blades with thick trailing edges an accurate prediction of base pressure is crucial to loss prediction. However, it is found that current methods of base pressure prediction are unable to give reliable predictions.

The Influence of Trailing Edge Ejection on the Base Pressure in Transonic Turbine Cascades

1983 ◽  
Vol 105 (2) ◽  
pp. 215-222 ◽  
Author(s):  
C. H. Sieverding
Keyword(s):  
Turbine Blades ◽  
Density Ratio ◽  
Main Flow ◽  
Coolant Flow ◽  
Base Pressure ◽  
Trailing Edge ◽  
Number Range ◽  
Transonic Turbine ◽  
Turbine Cascades ◽  
Transonic Cascade

This paper summarizes the results of base pressure studies on transonic turbine blades in the presence of an ejection of coolant flow from a slot in the trailing edge. The first part of the paper reports on tests carried out on a enlarged model of the overhang section of a typical transonic cascade. This model provides valuable information about the detailed trailing edge pressure distribution and points to an asymmetric evolution of the base pressure on both sides of the slot in the presence of a bleed. The second part of the paper presents experimental results from cascade tests covering an outlet Mach number range M2, is = 0.5 to 1.35. These experiments underline the importance of the coolant flow impact on the base pressure and confirm the asymmetry of the base pressure with respect to the cooling slot. Tests with different coolant flow gases point to the significance of a proper simulation of the density ratio between coolant flow and main flow.

The Effects of Trailing Edge Thickness on the Losses of Ultrahigh Lift Low Pressure Turbine Blades

2014 ◽  
Vol 136 (8) ◽  
Author(s):  
Chao Zhou ◽  
Howard Hodson ◽  
Christoph Himmel
Keyword(s):  
Turbine Blades ◽  
Reynolds Numbers ◽  
Low Pressure ◽  
Base Pressure ◽  
Manufacturing Cost ◽  
Surface Boundary ◽  
Trailing Edge ◽  
Low Pressure Turbine ◽  
Sst Model ◽  
Profile Loss

Experimental, numerical and analytical methods were used to investigate the effects of the blade trailing edge thickness on the profile loss of ultrahigh-lift low-pressure turbine blades. Two cascades, the T106C and the T2, were studied. The loss obtained based on the data at the blade trailing edge plane and the plane 0.3 Chord downstream of the trailing edge agree with each other for T106C blade with and without upstream wakes at different Reynolds numbers. The blade profile losses were broken down as the suction surface boundary loss, the pressure side boundary loss and the mixing loss downstream of the trailing edge for six Reynolds numbers. Trailing edge thicknesses varying from 1.4% to 4.7% pitch were investigated at a Reynolds number of 210,000. It was found that the flow distributions across the passage at the trailing edge planes were highly nonuniform. In particular, and as a result, the trailing edge base pressure was higher than the mixed-out static pressure, so the contribution of the base pressure to the mixing loss downstream of the trailing edge plane was to reduce the loss. When the trailing edge thickness increases, there are three main effects: (1) the area with high base pressure region increases, which tends to reduce the downstream mixing loss; (2) the base pressure reduces, which tends to increase the loss; and (3) the flow diffusion downstream of the trailing edge, which tends to increase the loss. The overall result is the combined effect. For the T106C cascade, increasing the trailing edge thickness from 1.9% pitch to 2.8% pitch has a small effect on the loss. Further increasing the trailing edge thickness increases the loss. The T2 blade has a higher lift than the T106C blade, so the effects of the base pressure in reducing the mixing loss downstream of the trailing edge is more evident. The experimental results show that the profile loss first decreases and then increases as the trailing edge thickness increases. CFD, using the transition k-ω SST model and the k-ω SST model, provides good predictions of the aerodynamic performance. It was used to study the cases with trailing edge thicknesses of 1.4% pitch and 2.9% pitch. The profile loss is almost the same for these two trailing edge thickness. The results show that it is possible to use thicker blade trailing edges in low pressure turbines without aerodynamic penalty. This can lead to benefits in terms of mechanical integrity and manufacturing cost reductions.

The Effects of Trailing Edge Thickness on the Losses of Ultra-High Lift LP Turbine Blades

2013 ◽  
Author(s):  
Chao Zhou ◽  
Howard Hodson ◽  
Christoph Himmel
Keyword(s):  
Turbine Blades ◽  
Reynolds Numbers ◽  
Low Pressure ◽  
Base Pressure ◽  
Manufacturing Cost ◽  
Surface Boundary ◽  
Trailing Edge ◽  
High Lift ◽  
Sst Model ◽  
Profile Loss

Experimental, numerical and analytical methods were used to investigate the effects of the blade trailing edge thickness on the profile loss of ultra-high lift low pressure turbine blades. Two cascades, the T106C and the T2, were studied. The loss obtained based on the data at the blade trailing edge plane and the plane 0.3 Chord downstream of the trailing edge agree with each other for T106C blade with and without upstream wakes at different Reynolds numbers. The blade profile losses were broken down as the suction surface boundary loss, the pressure side boundary loss and the mixing loss downstream of the trailing edge for six Reynolds numbers. Trailing edge thicknesses varying from 1.4% to 4.7% Pitch were investigated at a Reynolds number of 210000. It was found that the flow distributions across the passage at the trailing edge planes were highly non-uniform. In particular, and as a result, the trailing edge base pressure was higher than the mixed-out static pressure, so the contribution of the base pressure to the mixing loss downstream of the trailing edge plane was to reduce the loss. When the trailing edge thickness increases, there are three main effects: 1) The area with high base pressure region increase, which tends to reduce the downstream mixing loss; 2) The base pressure reduces, which tends to increase the loss; 3) The flow diffusion downstream of the trailing edge, which tends to increase the loss. The overall result is the combined effect. For the T106C cascade, increasing the trailing edge thickness from 1.9% Pitch to 2.8% Pitch has a small effect on the loss. Further increasing the trailing edge thickness increases the loss. The T2 blade has a higher lift than the T106C blade, so the effects of the base pressure in reducing the mixing loss downstream of the trailing edge is more evident. The experimental results show that the profile loss first decreases and then increases as the trailing edge thickness increases. CFD, using the transition k-ω SST model and the k-ω SST model, provides good predictions of the aerodynamic performance. It was used to study the cases with trailing edge thicknesses of 1.4% Pitch and 2.9% Pitch. The profile loss is almost the same for these two trailing edge thickness. The results show that it is possible to use thicker blade trailing edges in low pressure turbines without aerodynamic penalty. This can lead to benefits in terms of mechanical integrity and manufacturing cost reductions.

scholarly journals Numerical Study on the Biomimetic Trailing Edge of a Turbine Blade Under a Wide Range of Outlet Mach Numbers

2021 ◽  
Vol 9 ◽  
Author(s):  
Fengbo Wen ◽  
Yuxi Luo ◽  
Shuai Wang ◽  
Songtao Wang ◽  
Zhongqi Wang
Keyword(s):  
Mach Number ◽  
Energy Loss ◽  
Numerical Study ◽  
Turbine Blades ◽  
Loss Coefficient ◽  
Detached Eddy Simulation ◽  
Trailing Edge ◽  
Loss Mechanism ◽  
Wide Range ◽  
Mach Numbers

This study was carried out to investigate the loss mechanism of a blade with a harbor seal whisker structure on the trailing edge under different Mach numbers. The loss of high-pressure turbine blades with four different trailing edge geometries, including a prototype, an elliptical trailing edge (ETE), a sinusoidal trailing edge (STE), and a biomimetic trailing edge (BTE) at Mach numbers of 0.38–1.21 is studied. The delayed detached-eddy simulation method is used to predict the detailed flow of the four cascades. The result shows that, when the Mach number is less than 0.9, the BTE can effectively reduce the energy loss coefficient compared with the other three cases. As the Mach number increases, the three-dimensional characteristics of the wake behind the BTE weaken. The energy loss coefficient of the blade with the BTE is close to that of the blade with the ETE and STE when the Mach number is greater than 0.9. Besides this, by controlling the wake, the BTE can effectively suppress the dynamic movement of shock waves in the cascade at high Mach numbers.

Experimental and Numerical Investigation on the Performance of a Family of Three HP Transonic Turbine Blades

2004 ◽  
Author(s):  
D. Corriveau ◽  
S. A. Sjolander
Keyword(s):  
Flow Velocity ◽  
Turbine Blades ◽  
Suction Side ◽  
Experimental Results ◽  
Base Pressure ◽  
Suction Surface ◽  
Trailing Edge ◽  
Lower Base ◽  
Shock System ◽  
Transonic Turbine

Experimental results concerning the performance of three high-pressure (HP) transonic turbine blades having fore-, aft- and mid-loadings have been presented previously by Corriveau and Sjolander [1]. Results from that study indicated that by shifting the loading towards the rear of the airfoil, improvements in loss performance of the order of 20% could be obtained near the design Mach number. In order to gain a better understanding of the underlying reasons for the improved loss performance of the aft-loaded blade, additional measurements were performed on the three cascades. Furthermore, 2-D numerical simulations of the cascade flow were performed in order to help in the interpretation of the experimental results. Based on the analysis of additional wake traverse data and base pressure measurements combined with the numerical results, it was found that the poorer loss performance of the baseline mid-loaded profile compared to the aft-loaded blade could be traced back to the former’s higher rear suction side curvature. The presence of higher rear suction surface curvature resulted in higher flow velocity in that region. Higher flow velocity at the trailing edge in turn contributed to reducing the base pressure. The lower base pressure at the trailing edge resulted in a stronger trailing edge shock system for the mid-loaded blade. This shock system increased the losses for the mid-loaded baseline profile when compared to the aft-loaded profile.

CFD Investigation of the Flow of Trailing Edge Cooling Slots

2018 ◽  
Author(s):  
Y. Jiang ◽  
N. Gurram ◽  
E. Romero ◽  
P. T. Ireland ◽  
L. di Mare
Keyword(s):  
Mach Number ◽  
Kinetic Energy ◽  
Flow Separation ◽  
Film Cooling ◽  
High Speed ◽  
Turbine Blades ◽  
Cross Flow ◽  
Turbulent Energy ◽  
Trailing Edge ◽  
Flow Interactions

Slot film cooling is a popular choice for trailing edge cooling in high pressure (HP) turbine blades because it can provide more uniform film coverage compared to discrete film cooling holes. The slot geometry consists of a cut back in the blade pressure side connected through rectangular openings to the internal coolant feed passage. The numerical simulation of this kind of film cooling flows is challenging due to the presence of flow interactions like step flow separation, coolant-mainstream mixing and heat transfer. The geometry under consideration is a cutback surface at the trailing edge of a constant cross-section aerofoil. The cutback surface is divided into three sections separated by narrow lands. The experiments are conducted in a high speed cascade in Oxford Osney Thermo-Fluids Laboratory at Reynolds and Mach number distributions representative of engine conditions. The capability of CFD methods to capture these flow phenomena is investigated in this paper. The isentropic Mach number and film effectiveness are compared between CFD and pressure sensitive paint (PSP) data. Compared to steady k–ω SST method, Scale Adaptive Simulation (SAS) can agree better with the measurement. Furthermore, the profiles of kinetic energy, production and shear stress obtained by the steady and SAS methods are compared to identify the main source of inaccuracy in RANS simulations. The SAS method is better to capture the unsteady coolant-hot gas mixing and vortex shedding at the slot lip. The cross flow is found to affect the film significantly as it triggers flow separation near the lands and reduces the effectiveness. The film is non-symmetric with respect to the half-span plane and different flow features are present in each slot. The effect of mass flow ratio (MFR) on flow pattern and coolant distribution is also studied. The profiles of velocity, kinetic energy and production of turbulent energy are compared among the slots in detail. The MFR not only affects the magnitude but also changes the sign of production.

Numerical investigation of critical lip thickness of subsonic co-flowing jet

2021 ◽  
Vol ahead-of-print (ahead-of-print) ◽  
Author(s):  
Naren Shankar R. ◽  
Ganesan V.G. ◽  
Dilip Raja N. ◽  
Sathish Kumar K. ◽  
Vijayaraja K.
Keyword(s):  
Mach Number ◽  
Total Pressure ◽  
Static Pressure ◽  
Velocity Ratio ◽  
Critical Limit ◽  
Number Range ◽  
Content Type ◽  
Jet Mixing ◽  
Pressure Decay ◽  
Mach Numbers

Purpose The effect of increasing lip thickness (LT) and Mach number on subsonic co-flowing Jet (CFJ) decay at subsonic and correctly expanded sonic Mach numbers has been analysed experimentally and numerically in this study. This study aims to a critical LT below which mixing enhances and above which mixing inhibits. Design/methodology/approach LT is the distance, separating the primary nozzle and the secondary duct, present in the co-flowing nozzle. The CFJ with LT ranging from 2 mm to 150 mm at jet exit Mach numbers of 0.6, 0.8 and 1.0 were studied in detail. The CFJ with 2 mm LT is used for comparison. Centreline total pressure decay, centreline static pressure decay and near field flow behaviour were analysed. Findings The result shows that the mixing enhances until a critical limit and a further increase in the LT does not show any variation in the jet mixing. Beyond this critical limit, the secondary jet has a detrimental effect on the primary jet, which deteriorates the process of mixing. The CFJ within the critical limit experiences a significantly higher mixing. The effect of the increase in the Mach number has marginal variation in the total pressure and significant variation in static pressure along the jet axis. Practical implications In this study, the velocity ratio (VR) is maintained constant and the bypass ratio (BR) was varied from low value to very high values for subsonic and correctly expanded sonic. Presently, commercial aircraft engine operates under these Mach numbers and low to ultra-high BR. Hence, the present study becomes essential. Originality/value This is the first effort to find the critical value of LT for a constant VR for a Mach number range of 0.6 to 1.0, compressible CFJ. The CFJs with constant VR of unity and varying LT, in these Mach number range, have not been studied in the past.

Sign in / Sign up

Export Citation Format

Share Document