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 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.

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.

Techniques for Aerodynamic Loss Measurement of Transonic Turbine Cascades With Trailing-Edge Region Coolant Ejection

1992 ◽  
Author(s):  
D. J. Mee
Keyword(s):  
Total Pressure ◽  
Free Stream ◽  
Turbine Blades ◽  
Flux Ratio ◽  
Edge Region ◽  
Trailing Edge ◽  
Aerodynamic Loss ◽  
Transonic Turbine ◽  
Turbine Cascades ◽  
Density Ratios

Experimental techniques associated with the measurement of loss of transonic turbine blades with trailing-edge region coolant ejection are considered. Results from experiments with different coolant to free stream gas density ratios indicate that it is not always adequate to simulate only the coolant blowing rate. However, for the measurement of loss, the present experimental results indicate that it may be adequate to simulate the momentum flux ratio. In loss calculations the value used for the total pressure of the coolant gas is discussed and shown to influence a comparison of different cooling geometries.

The Trailing Edge Loss of Transonic Turbine Blades

1990 ◽  
Vol 112 (2) ◽  
pp. 277-285 ◽  
Author(s):  
J. D. Denton ◽  
L. Xu
Keyword(s):  
Surface Pressure ◽  
Turbine Blades ◽  
Back Pressure ◽  
Base Pressure ◽  
Mass Energy ◽  
Suction Surface ◽  
Trailing Edge ◽  
Transonic Turbine ◽  
Energy And Momentum ◽  
Unique Relationship

Trailing edge loss is one of the main sources of loss for transonic turbine blades, contributing typically 1/3 of their total loss. Transonic trailing edge flow is extremely complex, the basic flow pattern is understood but methods of predicting the loss are currently based on empirical correlations for the base pressure. These correlations are of limited accuracy. Recent findings that the base pressure and loss can be reasonably well predicted by inviscid Euler calculations are justified and explained in this paper. For unstaggered choked blading, it is shown that there is a unique relationship between the back pressure and the base pressure and any calculation that conserves mass, energy and momentum should predict this relationship and the associated loss exactly. For realistic staggered blading, which operates choked but with subsonic axial velocity, there is also a unique relationship between the back pressure and the base pressure (and hence loss) but the relationship cannot be quantified without knowing a further relationship between the base pressure and the average suction surface pressure downstream of the throat. Any calculation that conserves mass, energy and momentum and also predicts this average suction surface pressure correctly will again predict the base pressure and loss. Two-dimensional Euler solutions do not predict the suction surface pressure exactly because of shock smearing but nevertheless seem to give reasonably accurate results. The effects of boundary layer thickness and trailing edge coolant ejection are considered briefly. Coolant ejection acts to reduce the mainstream loss. It is shown that suction surface curvature downstream of the throat may be highly beneficial in reducing the loss of blades with thick trailing edges operating at high subsonic or low supersonic outlet Mach numbers.

scholarly journals The Aerodynamics of Trailing-Edge-Cooled Transonic Turbine Blades: Part 1 — Experimental Approach

1997 ◽  
Author(s):  
Mathias Deckers ◽  
John D. Denton
Keyword(s):  
Turbine Blades ◽  
Base Pressure ◽  
Stagnation Temperature ◽  
Trailing Edge ◽  
Static Pressure Distribution ◽  
Discernible Effect ◽  
Two Dimensional Flow ◽  
Shock System ◽  
Flow Through ◽  
Transonic Turbine

The research presented in this part of the paper involved a detailed experimental study of the flow through transonic turbine blading with trailing edge coolant ejection. The tests were carried out on (nearly) flat plate models representing the region of uncovered turning downstream of the throat. The investigation focused on the aerodynamic aspects associated with trailing edge ejection in steady two-dimensional flow over a range of exit Mach numbers, coolant pressure ratios and temperature ratios. The experiments showed that the simple existence of the coolant cavity leads to a substantial change of the flow field in the near wake. Consequently, the slotted unblown base was found to have considerably less loss than the solid one. The effect of coolant ejection is shown to cause a substantial increase in base pressure and reduction in overall loss. The surface static pressure distribution and boundary layers were affected by the coolant in two ways: directly from downstream, via the base pressure, and indirectly through a changed trailing edge shock system. However, the coolant stagnation temperature ratio was found to have no discernible effect on the base pressure and loss.

Model Tests for the Detailed Investigation of the Trailing Edge Flow in Convergent Transonic Turbine Cascades

1976 ◽  
Author(s):  
C. Sieverding ◽  
M. Decuyper ◽  
J. Colpin ◽  
O. Amana
Keyword(s):  
Base Pressure ◽  
Test Series ◽  
Base Region ◽  
Flow Conditions ◽  
Trailing Edge ◽  
Separation Shock ◽  
Cascade Flow ◽  
Transonic Turbine ◽  
Turbine Cascades ◽  
Edge Flow

The paper presents some experiments concerning the trailing edge flow of transonic turbine bladings. The experiments are conducted on a model simulating very closely the flow in the overhang section of convergent turbine cascades. The first test series simulates cascade flow conditions with shock free flow in the overhang section (limit load conditions), while the second test series simulates cascade flow conditions characterized by trailing edge shock interferences with the flow in the overhang section. Particular attention is given to the trailing edge separation shock and the length of the isobaric trailing edge base region. The test results indicate that the strength of the separation shock for the rounded trailing edge is of the same order as the strength of the lip shock found by Merzkirch for sharp-edged wedges. The length of the isobaric base region is only 60 to 80 percent of the trailing edge thickness, while the pressure recovery region is about twice as long as the constant pressure region. From an analysis of the flow parameters influencing the base pressure, it is concluded that the boundary-layer shape factor plays a significant role in establishing the base pressure.

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 Trailing Edge Loss of Transonic Turbine Blades

1989 ◽  
Author(s):  
J. D. Denton ◽  
L. Xu
Keyword(s):  
Surface Pressure ◽  
Turbine Blades ◽  
Back Pressure ◽  
Base Pressure ◽  
Mass Energy ◽  
Suction Surface ◽  
Trailing Edge ◽  
Transonic Turbine ◽  
Energy And Momentum ◽  
Unique Relationship

Trailing edge loss is one of the main sources of loss for transonic turbine blades, contributing typically 1/3 of their total loss. Transonic trailing edge flow is extremely complex, the basic flow pattern is understood but methods of predicting the loss are currently based on empirical correlations for the base pressure. These correlations are of limited accuracy. Recent findings that the base pressure and loss can be reasonably well predicted by inviscid Euler calculations are justified and explained in this paper. For unstaggered choked blading it is shown that there is is a unique relationship between the back pressure and the base pressure and any calculation that conserves mass, energy and momentum should predict this relationship and the associated loss exactly. For realistic staggered blading which operates choked but with subsonic axial velocity there is also a unique relationship between the back pressure and the base pressure (and hence loss) but the relationship cannot be quantified without knowing a further relationship between the base pressure and the average suction surface pressure downstream of the throat. Any calculation that conserves mass, energy and momentum and also predicts this average suction surface pressure correctly will again predict the base pressure and loss. Two dimensional Euler solutions do not predict the suction surface pressure exactly because of shock smearing but nevertheless seem to give reasonably accurate results.

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.

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