Influence of Purge Flow Injection Angle on the Aerothermal Performance of a Rotor Blade Cascade

2013 ◽  
Vol 136 (4) ◽  
Author(s):  
G. Barigozzi ◽  
G. Franchini ◽  
A. Perdichizzi ◽  
M. Maritano ◽  
R. Abram
Keyword(s):  
Flow Injection ◽  
Rotor Blade ◽  
Thermal Protection ◽  
Secondary Flows ◽  
Pressure Probe ◽  
Film Cooling Effectiveness ◽  
Injection Angle ◽  
Blade Cascade ◽  
Purge Flow ◽  
End Wall

This paper is focused on the influence of stator-rotor purge flow injection angle on the aerodynamic and thermal performance of a rotor blade cascade. Tests were performed in a seven-blade cascade of a high-pressure gas turbine rotor at low Mach number (Ma2is = 0.3) under different blowing conditions. A number of fins were installed inside the upstream slot to simulate the effect of rotation on the seal flow exiting the gap in a linear cascade environment. The resulting coolant flow is ejected with the correct angle in the tangential direction. Purge flow injection angle and blowing conditions were changed in order to identify the best configuration in terms of end wall thermal protection and secondary flows reduction. The 3D flow field was surveyed by traversing a five-hole miniaturized pressure probe in a downstream plane. Secondary flow velocities, loss coefficient, and vorticity distributions are presented for the most significant test conditions. Film cooling effectiveness distributions on the platform were obtained by thermochromic liquid crystals (TLC) technique. Results show that purge flow injection angle has an impact on secondary flows development and, thus, on the end wall thermal protection, especially at high injection rates. Passage vortex is enhanced by a negative injection angle, which simulates the real counter rotating purge flow direction.

Influence of Purge Flow Injection Angle on the Aero-Thermal Performance of a Rotor Blade Cascade

2013 ◽  
Author(s):  
G. Barigozzi ◽  
G. Franchini ◽  
A. Perdichizzi ◽  
M. Maritano ◽  
R. Abram
Keyword(s):  
Thermal Performance ◽  
Flow Injection ◽  
Rotor Blade ◽  
Thermal Protection ◽  
Secondary Flows ◽  
Pressure Probe ◽  
Injection Angle ◽  
Blade Cascade ◽  
Purge Flow ◽  
End Wall

This paper is focused on the influence of stator-rotor purge flow injection angle on the aerodynamic and thermal performance of a rotor blade cascade. Tests were performed in a seven-blade cascade of a high-pressure gas turbine rotor at low Mach number (Ma2is = 0.3) under different blowing conditions. A number of fins were installed inside the upstream slot to simulate the effect of rotation on the seal flow exiting the gap in a linear cascade environment. The resulting coolant flow is ejected with the correct angle in the tangential direction. Purge flow injection angle and blowing conditions were changed in order to identify the best configuration in terms of end wall thermal protection and secondary flows reduction. The 3D flow field was surveyed by traversing a 5-hole miniaturized pressure probe in a downstream plane. Secondary flow velocities, loss coefficient and vorticity distributions are presented for the most significant test conditions. Film cooling effectiveness distributions on the platform were obtained by Thermochromic Liquid Crystals technique. Results show that purge flow injection angle has an impact on secondary flows development and thus on the end wall thermal protection, especially at high injection rates. Passage vortex is enhanced by a negative injection angle, which simulates the real counter rotating purge flow direction.

Aero-Thermal Performance of a Rotor Blade Cascade With Stator-Rotor Seal Purge Flow

2012 ◽  
Author(s):  
G. Barigozzi ◽  
F. Fontaneto ◽  
G. Franchini ◽  
A. Perdichizzi ◽  
M. Maritano ◽  
...  
Keyword(s):  
Secondary Flow ◽  
Thermal Performance ◽  
Flow Injection ◽  
Rotor Blade ◽  
Thermal Protection ◽  
Leading Edge ◽  
Pressure Probe ◽  
Blade Cascade ◽  
Purge Flow ◽  
End Wall

The present paper investigates the effects of purge flow from a stator-rotor seal gap on the aerodynamic and thermal performance of a rotor blade cascade. Particular attention is paid to thermal results in the leading edge area that is typically difficult to protect. Experimental tests have been performed on a seven-blade cascade of a high-pressure rotor stage of a real gas turbine at low Mach number (Ma2is = 0.3). To simulate the rotational effect in a linear cascade environment, a number of inclined fins have been installed inside the stator-rotor gap, making the coolant flow to exit with the right tangential velocity component. Tests have been carried out at different blowing conditions, with mass flow rate ratios up to 2.0%. Aerodynamic effects of purge flow on secondary flow structures were surveyed by traversing a 5-hole miniaturized pressure probe in a plane 0.08cax downstream of the trailing edge. Film cooling effectiveness distributions on the end wall platform were obtained by using Thermochromic Liquid Crystals technique. Results allowed to investigate the effect of purge flow injection from the upstream gap on the secondary flows development and on the thermal protection capability. Purge flow injection of 1.0% reduced secondary flow losses and was found to effectively protect the front end wall region, up to about 0.5cax downstream of the leading edge. Increasing the purge flow up to 1.5%–2.0% provided a better thermal protection not only stream wise, but also in the region close to the leading edge because of the weakened washing activity of the horseshoe vortex.

scholarly journals Improving the Film Cooling of a Rotor Blade Platform

2017 ◽  
Vol 140 (2) ◽  
Author(s):  
Giovanna Barigozzi ◽  
Antonio Perdichizzi ◽  
Roberto Abram
Keyword(s):  
Film Cooling ◽  
Rotor Blade ◽  
Thermal Protection ◽  
Secondary Flows ◽  
Thermal Tests ◽  
Film Cooling Effectiveness ◽  
Blowing Ratio ◽  
Purge Flow ◽  
Cylindrical Holes ◽  
The Impact

This paper shows the results of an experimental activity developed in cooperation between Ansaldo Energia and the Department of Engineering and Applied Science of Bergamo University with the aim of assessing the impact of newly designed holes on the thermal protection of a rotor blade platform. The original rotor blade platform featured ten cylindrical holes located along the blade pressure side (PS). Moreover, the channel front side was cooled exploiting the seal purge flow exiting the stator to rotor interface gap. The front midchannel, and particularly the region around the interplatform gap, remained uncooled. To protect this region, two sets of cylindrical holes were designed and manufactured on a seven blade cascade model for experimental verification. Aerodynamic and thermal tests were carried out at low Mach number. To evaluate the interaction of injected flow with secondary flows a five hole probe was traversed downstream of the trailing edge plane. The thermal behavior was analyzed by using thermochromic liquid crystals technique, so to obtain film cooling effectiveness distributions. The seven-hole configuration coupled with a low blowing ratio of about 1.0 provided the best thermal protection without any impact on the aerodynamic performance.

Influence of Coolant Flow Rate on Aero-Thermal Performance of a Rotor Blade Cascade With Endwall Film Cooling

2012 ◽  
Vol 134 (5) ◽  
Author(s):  
G. Barigozzi ◽  
F. Fontaneto ◽  
G. Franchini ◽  
A. Perdichizzi ◽  
M. Maritano ◽  
...  
Keyword(s):  
High Pressure ◽  
Thermal Performance ◽  
Film Cooling ◽  
Rotor Blade ◽  
Secondary Flows ◽  
Film Cooling Effectiveness ◽  
Blade Cascade ◽  
Comprehensive Picture ◽  
Cylindrical Holes ◽  
Endwall Film Cooling

This paper investigates the influence of coolant injection on the aerodynamic and thermal performance of a rotor blade cascade with endwall film cooling. A seven blade cascade of a high-pressure-rotor stage of a real gas turbine has been tested in a low speed wind tunnel for linear cascades. Coolant is injected through 10 cylindrical holes distributed along the blade pressure side. Tests have been preliminarily carried out at low Mach number (Ma2is = 0.3). Coolant-to-mainstream mass flow ratio has been varied in a range of values corresponding to inlet blowing ratios M1 = 0–4.0. Secondary flows have been surveyed by traversing a five-hole miniaturized aerodynamic probe in two downstream planes. Local and overall mixed-out secondary loss coefficient and vorticity distributions have been calculated from measured data. The thermal behavior has been also analyzed by using thermochromic liquid crystals technique to obtain film cooling effectiveness distributions. All this information, including overall loss production for variable injection conditions, allows us to draw a comprehensive picture of the aero-thermal flow field in the endwall region of a high pressure rotor blade cascade.

Influence of Coolant Flow Rate on Aero-Thermal Performance of a Rotor Blade Cascade With Endwall Film Cooling

2011 ◽  
Author(s):  
G. Barigozzi ◽  
F. Fontaneto ◽  
G. Franchini ◽  
A. Perdichizzi ◽  
M. Maritano ◽  
...  
Keyword(s):  
High Pressure ◽  
Thermal Performance ◽  
Film Cooling ◽  
Rotor Blade ◽  
Secondary Flows ◽  
Film Cooling Effectiveness ◽  
Blade Cascade ◽  
Comprehensive Picture ◽  
Cylindrical Holes ◽  
Endwall Film Cooling

This paper investigates the influence of coolant injection on the aerodynamic and thermal performance of a rotor blade cascade with endwall film cooling. A 7 blade cascade of a high-pressure-rotor stage of a real gas turbine has been tested in a low speed wind tunnel for linear cascades. Coolant is injected through ten cylindrical holes distributed along the blade pressure side. Tests have been preliminarily carried out at low Mach number (Ma2is = 0.3). Coolant-to-mainstream mass flow ratio has been varied in a range of values corresponding to inlet blowing ratios M1 = 0–4.0. Secondary flows have been surveyed by traversing a 5-hole miniaturized aerodynamic probe in two downstream planes. Local and overall mixed-out secondary loss coefficient and vorticity distributions have been calculated from measured data. The thermal behaviour has been also analysed by using Thermochromic Liquid Crystals technique, so to obtain film cooling effectiveness distributions. All this information, including overall loss production for variable injection conditions, allow to draw a comprehensive picture of the aero-thermal flow field in the endwall region of a high pressure rotor blade cascade.

Coolant Gas Injection on a Blunt-Nosed Re-Entry Vehicle

2013 ◽  
Author(s):  
Jeswin Joseph ◽  
S. R. Shine
Keyword(s):  
Mach Number ◽  
Film Cooling ◽  
Thermal Protection ◽  
Thermal Protection System ◽  
Protection System ◽  
Aerodynamic Heating ◽  
Cooling Effectiveness ◽  
Film Cooling Effectiveness ◽  
Injection Angle ◽  
Mach Numbers

Very high thermal loads are expected in re-entry vehicles traveling at hypersonic Mach numbers due to severe aerodynamic heating. In the present study, numerical investigations are carried out to analyze the use of film cooling technology for a fully reusable and active thermal protection system of the re-entry vehicle. Simulations are done to examine the fundamental flow phenomenon and the performance of blunt body film cooling in hypersonic flows. Simulations are conducted for a blunt -nosed spacecraft flying at Mach numbers varying from 4 to 8 and 40 deg angle of attack. Film cooling holes are provided on the bottom of the blunt-nosed body. Standard values at an altitude of 30 km are used as in flow boundary conditions. The dependency of blowing ratios, stream-wise injection angle and inlet Mach number on the film cooling effectiveness are investigated. It is observed that the film cooling effectiveness reduces with increase in coolant injection angle. The film cooling performance is found to be decreasing with increase in Mach number. The results could provide useful inputs for optimization of an active thermal protection system of re-entry vehicles.

Numerical Investigation of the Effect of Purge Flow on Aerodynamic Performance and Film Cooling Effectiveness on a Rotating Turbine With Non-Axisymmetric Endwall Contouring

2012 ◽  
Author(s):  
M. T. Schobeiri ◽  
K. Lu ◽  
J. C. Han
Keyword(s):  
Numerical Investigation ◽  
Mass Flow ◽  
Flow Injection ◽  
Film Cooling ◽  
Preceding Paper ◽  
Cooling Effectiveness ◽  
Film Cooling Effectiveness ◽  
Purge Flow ◽  
Endwall Contouring ◽  
The Impact

The impact of the purge flow injection on aerodynamics and film cooling effectiveness of a high pressure turbine with non-axisymmetric endwall contouring has been numerically investigated. For this purpose, the geometry and boundary condition of a three-stage turbine at the Turbomachinery Performance and Flow Research Laboratory (TPFL), Texas A&M University is utilized. The turbine is being prepared to experimentally verify the results of the current numerical investigations. Its rotor includes non-axisymmetric endwall contouring on the first and second rotor row. In the preceding paper [1] it was shown that the endwall contouring of the second rotor contouring was able to substantially increase the turbine efficiency. To investigate the film cooling in conjunction with a purge flow injection, the first turbine rotor hub was contoured. Applying the same contouring method, however, different aerodynamic behavior of the first rotor was observed due to its immediate exposure to the purge flow injection. Consequently, the endwall design of the first rotor row required particular attention. The purge flow investigation involves the reference case without endwall contouring followed by the investigation with endwall contouring. The turbine used for this numerical investigation has two independent cooling loops. The first loop supplies coolant air to the stator-rotor gap, while the second loop provides cooling air to the downstream discrete film-cooling holes and blade tip cooling injection holes. For the current investigations the second loop is closed. Film cooling effectiveness is numerically simulated for rotor frequency of 2400 rpm. Efficiency, pressure, temperature and film cooling effectiveness distributions are determined for purge mass flow ratios of MFR = 0.5%, 1.0% and 2.0%. The small amount of the injected mass flow drastically changes the development of the secondary flow structure of the contoured first turbine row partially reversing the improvement tendency obtained from the endwall contouring.

Investigation of a Novel Secondary Flow Feature in a Turbine Cascade With End Wall Profiling

2005 ◽  
Vol 127 (1) ◽  
pp. 209-214 ◽  
Author(s):  
Grant Ingram ◽  
David Gregory-Smith ◽  
Neil Harvey
Keyword(s):  
Secondary Flow ◽  
Secondary Flows ◽  
Pressure Probe ◽  
Unexpected Result ◽  
Full Potential ◽  
Flow Angle ◽  
Passage Vortex ◽  
Flow Feature ◽  
New Feature ◽  
End Wall

A novel secondary flow feature, previously unreported for turbine blading as far as the authors are aware, has been discovered. It has been found that it is possible to separate part of the inlet boundary layer on the blade row end wall as it is being over-turned and rolled up into the passage vortex. This flow feature has been discovered during a continuing investigation into the aerodynamic effects of non-axisymmetric end wall profiling. Previous work, using the low speed linear cascade at Durham University, has shown the potential of end wall profiling for reducing secondary losses. The latest study, the results of which are described here, was undertaken to determine the limits of what end wall profiling can achieve. The flow has been investigated in detail with pressure probe traversing and surface flow visualization. This has found that the inlet boundary locally separates, on the early suction side of the passage, generating significant extra loss which feeds directly into the core of the passage vortex. The presence of this new feature gives rise to the unexpected result that the secondary flow, as determined by the exit flow angle deviations and levels of secondary kinetic energy, can be reduced while at the same time the loss is increased. CFD was found to calculate the secondary flows moderately well compared with measurements. However, CFD did not predict this new feature, nor the increase in loss it caused. It is concluded that the application of non-axisymmetric end wall profiling, although it has been shown to be highly beneficial, can give rise to adverse features that current CFD tools are unable to predict. Improvements to CFD capability are required in order to be able to avoid such features, and obtain the full potential of end wall profiling.

Aero-Thermal Performance of a Rotor Blade Cascade With Film Cooling in Passage Endwall

2018 ◽  
Author(s):  
Jinglun Fu ◽  
Jahed Hossain ◽  
Jayanta Kapat
Keyword(s):  
Thermal Performance ◽  
Film Cooling ◽  
Rotor Blade ◽  
Cfd Simulations ◽  
Cooling Effectiveness ◽  
Film Cooling Effectiveness ◽  
Blade Cascade ◽  
Mainstream Flow ◽  
Endwall Film Cooling ◽  
Annular Cascade

This paper describes the numerical investigations on the aerodynamic and thermal performance of a rotor blade cascade with multiple film cooling rows in the passage. First, the experimental data on an annular cascade with upstream film cooling was compared with the numerical results to validate the numerical method. The CFD simulations of the models with a row of film holes at four different locations on the hub endwall were performed respectively. The aerodynamic and thermal performance under the interaction of the secondary flow and endwall film cooling are analyzed based on the CFD predicted streamlines of mainstream flow and film injection, the contours of total pressure loss on the sections located in the passage and at the blade exit, the pitch-averaged film cooling effectiveness and film cooling effectiveness contours. The results show that film holes placed at low level of iso-Mach line tends to provide a better cooling with a smaller amount of coolant.

Investigation of a Novel Secondary Flow Feature in a Turbine Cascade With End Wall Profiling

2004 ◽  
Author(s):  
Grant Ingram ◽  
David Gregory-Smith ◽  
Neil Harvey
Keyword(s):  
Secondary Flow ◽  
Secondary Flows ◽  
Pressure Probe ◽  
Unexpected Result ◽  
Full Potential ◽  
Flow Angle ◽  
Passage Vortex ◽  
Flow Feature ◽  
New Feature ◽  
End Wall

A novel secondary flow feature, previously unreported for turbine blading as far as the authors are aware, has been discovered. It has been found that it is possible to separate part of the inlet boundary layer on the blade row end wall as it is being over-turned and rolled up into the passage vortex. This flow feature has been discovered during a continuing investigation into the aerodynamic effects of non-axisymmetric end wall profiling. Previous work, using the low speed linear cascade at Durham University, has shown the potential of end wall profiling for reducing secondary losses. The latest study, the results of which are described here, was undertaken to determine the limits of what end wall profiling can achieve. The flow has been investigated in detail with pressure probe traversing and surface flow visualization. This has found that the inlet boundary locally separates, on the early suction side of the passage, generating significant extra loss which feeds directly into the core of the passage vortex. The presence of this new feature gives rise to the unexpected result that the secondary flow, as determined by the exit flow angle deviations and levels of secondary kinetic energy, can be reduced while at the same time the loss is increased. CFD was found to calculate the secondary flows moderately well compared with measurements. However, CFD did not predict this new feature, nor the increase in loss it caused. It is concluded that the application of non-axisymmetric end wall profiling, although it has been shown to be highly beneficial, can give rise to adverse features that current CFD tools are unable to predict. Improvements to CFD capability are required in order to be able to avoid such features, and obtain the full potential of end wall profiling.

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