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.

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.

Computational Fluid Dynamics Investigation of the Flow of Trailing Edge Cooling Slots

2019 ◽  
Vol 141 (7) ◽  
Author(s):  
Y. Jiang ◽  
N. Gurram ◽  
E. Romero ◽  
P. T. Ireland ◽  
L. di Mare
Keyword(s):  
Fluid Dynamics ◽  
Computational Fluid Dynamics ◽  
Shear Stress ◽  
Mach Number ◽  
Kinetic Energy ◽  
Flow Separation ◽  
Film Cooling ◽  
High Speed ◽  
Turbine Blades ◽  
Trailing Edge

Slot film cooling is a popular choice for trailing edge (TE) 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 cutback 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 such as 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 computational fluid dynamics (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. When compared with the steady k − ω shear stress transport (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 nonsymmetric 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.

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.

Loss Generation in Transonic Turbine Blading

2017 ◽  
Author(s):  
Penghao Duan ◽  
Choon S. Tan ◽  
Andrew Scribner ◽  
Anthony Malandra
Keyword(s):  
Experimental Data ◽  
Boundary Layer ◽  
Mach Number ◽  
High Speed ◽  
Turbine Blades ◽  
Surface Boundary ◽  
Loss Model ◽  
Exit Mach Number ◽  
Loss Characteristic ◽  
Shock Loss

The measured loss characteristic in a high-speed cascade tunnel of two turbine blades of different designs showed distinctly different trend with exit Mach number ranging from 0.8 to 1.4. Assessments using steady RANS computation of the flow in the two turbine blades, complemented with control volume analyses and loss modelling, elucidate why the measured loss characteristic looks the way it is. The loss model categorizes the total loss in terms of boundary layer loss, trailing edge loss and shock loss; it yields results in good agreement with the experimental data as well as steady RANS computed results. Thus RANS is an adequate tool for determining the loss variations with exit isentropic Mach number and the loss model serves as an effective tool to interpret both the computational and experimental data. The measured loss plateau in Blade 1 for exit Mach number of 1 to 1.4 is due to a balance between a decrease of blade surface boundary layer loss and an increase in the attendant shock loss with Mach number; this plateau is absent in Blade 2 due to a greater rate in shock loss increase than the corresponding decrease in boundary layer loss. For exit Mach number from 0.85 to 1, the higher loss associated with shock system in Blade 1 is due to the larger divergent angle downstream of the throat than that in Blade 2. However when exit Mach number is between 1.00 and 1.30, Blade 2 has higher shock loss. For exit Mach number above around 1.4, the shock loss for the two blades is similar as the flow downstream of the throat is completely supersonic. In the transonic to supersonic flow regime, the turbine design can be tailored to yield a shock pattern the loss of which can be mitigated in near equal amount of that from the boundary layer with increasing exit Mach number, hence yielding a loss plateau in transonic-supersonic regime.

Shock–boundary layer interaction and energetics in transonic flutter

2017 ◽  
Vol 832 ◽  
pp. 212-240 ◽  
Author(s):  
Pradeepa T. Karnick ◽  
Kartik Venkatraman
Keyword(s):  
Boundary Layer ◽  
Mach Number ◽  
Turbulence Model ◽  
Viscous Effects ◽  
Trailing Edge ◽  
Aeroelastic System ◽  
Number Range ◽  
Shock Reversal ◽  
Transonic Flutter ◽  
Flutter Boundary

We study the influence of shock and boundary layer interactions in transonic flutter of an aeroelastic system using a Reynolds-averaged Navier–Stokes (RANS) solver together with the Spalart–Allmaras turbulence model. We show that the transonic flutter boundary computed using a viscous flow solver can be divided into three distinct regimes: a low transonic Mach number range wherein viscosity mimics increasing airfoil thickness thereby mildly influencing the flutter boundary; an intermediate region of drastic change in the flutter boundary due to shock-induced separation; and a high transonic Mach number zone of no viscous effects when the shock moves close to the trailing edge. Inviscid transonic flutter simulations are a very good approximation of the aeroelastic system in predicting flutter in the first and third regions: that is when the shock is not strong enough to cause separation, and in regions where the shock-induced separated region is confined to a small region near the trailing edge of the airfoil. However, in the second interval of intermediate transonic Mach numbers, the power distribution on the airfoil surface is significantly influenced by shock-induced flow separation on the upper and lower surfaces leading to oscillations about a new equilibrium position. Though power contribution by viscous forces are three orders of magnitude less than the power due to pressure forces, these viscous effects manipulate the flow by influencing the strength and location of the shock such that the power contribution by pressure forces change significantly. Multiple flutter points that were part of the inviscid solution in this regime are now eliminated by viscous effects. Shock motion on the airfoil, shock reversal due to separation, and separation and reattachment of flow on the airfoil upper surface, also lead to multiple aerodynamic forcing frequencies. These flow features make the flutter boundary quantitatively sensitive to the turbulence model and numerical method adopted, but qualitatively they capture the essence of the physical phenomena.

Study of Secondary Flows in Blade Cascades of Turbomachines

1982 ◽  
Vol 104 (2) ◽  
pp. 497-509 ◽  
Author(s):  
F. Bario ◽  
F. Leboeuf ◽  
K. D. Papailiou
Keyword(s):  
Mach Number ◽  
Turbine Blades ◽  
Theoretical Method ◽  
Secondary Flows ◽  
Compressor Blades ◽  
Number Range ◽  
Mean Values ◽  
Method Of Calculation ◽  
Passage Vortex ◽  
Highly Loaded

Experiments have been performed with two cascades of turbomachines. The first cascade is composed of highly loaded turbine blades, and has been used in the low subsonic Mach number range. The second cascade consists of highly loaded compressor blades, of the DCA type. The Mach number was then in the high subsonic range. The experimental results are presented in the form of mean values in the pitch direction. Detailed local values are also described. The growth of a passage vortex and a corner effect are presented in the compressor case. Their interactions with the whole flow are analyzed. In the turbine case, the passage vortex is found to be a dominant effect. Results obtained with a theoretical method of calculation of the flow in the blade passage are used to complete the analysis.

Development of Large Intermittent Wind Tunnels

1955 ◽  
Vol 59 (532) ◽  
pp. 259-278 ◽  
Author(s):  
J. Lukasiewicz
Keyword(s):  
Mach Number ◽  
Wind Tunnel ◽  
Mass Flow ◽  
High Speed ◽  
Pressure Ratio ◽  
Flow Characteristics ◽  
Wind Tunnels ◽  
Number Range ◽  
Pressure Storage ◽  
Vacuum Storage

SummaryTwo types of intermittent wind tunnel drives, the pressure storage drive(with atmospheric exhaust) and the vacuum storage drive (with atmospheric inlet), are examined and found to match well the tunnel pressure ratio-mass flow characteristics over a wide Mach number range (0 to 4). The design of components of intermittent wind tunnel installations, their operation and instrumentation are then considered in some detail. In order to increase the output of intermittent wind tunnels to a level comparable to that of continuously running tunnels, it is proposed to drive the models during each tunnel run through a range of incidence. This technique is at present under development in the National Aeronautical Establishment's High Speed Aerodynamics Laboratory and results so far obtained are discussed. Two tunnels are considered as examples of large intermittent installations: a 4 ft. square pressure-driven tunnel and a 6 ft. square vacuum-driven tunnel. The former is found to be a more compact and economical installation. Relative merits of continuous and intermittent installations are discussed.

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.

Investigation of CFD Prediction Capabilities for Low Reynolds Turbine Aerodynamics

2009 ◽  
Author(s):  
Bastian Muth ◽  
Marco Schwarze ◽  
Reinhard Niehuis ◽  
Matthias Franke
Keyword(s):  
Mach Number ◽  
Wind Tunnel ◽  
Flow Separation ◽  
High Speed ◽  
Operating Conditions ◽  
Engine Operation ◽  
Number Range ◽  
Operation Conditions ◽  
Transition Models ◽  
Low Reynolds

The objective of this work is to study the performance of low pressure turbines operating at low Reynolds numbers by extensive experiments and to validate numerical simulation results with the experimental data. Particular attention is payed to the prediction capabilities of current numerical turbulence and transition models in order to be able to benchmark the performance of future turbine airfoil profiles and to optimise their aero design. The LPT-Cascade under consideration has been investigated at the High Speed Cascade Wind Tunnel of the Institute of Jet Propulsion to gather information about the performance of turbine airfoils under low Reynolds operating conditions. The experiments were executed in the range of Re = 40′000 to 400′000 with steady state inflow conditions at different Mach number levels. The main focus of the investigation thereby was on the range of Re = 40′000 to 70′000. The high speed cascade wind tunnel of the University of Federal Armed Forces Munich allows for an independent Reynolds and Mach number variation such that an extensive database can be generated for realistic engine operation conditions. One major test objective was related to flow separation phenomena on the suction surface and its influence on the performance of the turbine profile. For this purpose both the loss behaviour and the pressure distribution on suction and pressure surface of the blade were measured and analysed. In addition to the experiments numerical flow simulations were conducted for the same turbine profile. In order to achieve more information on the influence of different turbulence and transition models on the flow separation, transition, and reattachment behaviour, two different CFD codes were used for comparison purposes. On the one hand the CFD code TRACE, which is developed by the German Aerospace Center (DLR) and MTU Aero Engines and on the other hand the general purpose code ANSYS CFX were applied. The aim is to assess the prediction capabilities of the different codes especially in the low Reynolds number range.

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