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Film Cooling, Heat Transfer and Aerodynamic Measurements in a Three Stage Research Gas TurbineSuryanarayanan, Arun 2009 May 1900 (has links)
The existing 3-stage turbine research facility at the Turbomachinery Performance and Flow
Research Laboratory (TPFL), Texas A and M University, is re-designed and newly installed to enable coolant
gas injection on the first stage rotor platform to study the effects of rotation on film cooling and heat
transfer. Pressure and temperature sensitive paint techniques are used to measure film cooling
effectiveness and heat transfer on the rotor platform respectively. Experiments are conducted at three
turbine rotational speeds namely, 2400rpm, 2550rpm and 3000rpm. Interstage aerodynamic measurements
with miniature five hole probes are also acquired at these speeds. The aerodynamic data characterizes the
flow along the first stage rotor exit, second stage stator exit and second stage rotor exit. For each rotor
speed, film cooling effectiveness is determined on the first stage rotor platform for upstream stator-rotor
gap ejection, downstream discrete hole ejection and a combination of upstream gap and downstream hole
ejection. Upstream coolant ejection experiments are conducted for coolant to mainstream mass flow ratios
of MFR=0.5%, 1.0%, 1.5% and 2.0% and downstream discrete hole injection tests corresponding to
average hole blowing ratios of M = 0.5, 0.75, 1.0, 1.25, 1.5, 1.75 and 2.0 for each turbine speed. To
provide a complete picture of hub cooling under rotating conditions, experiments with simultaneous
injection of coolant gas through upstream and downstream injection are conducted for an of MFR=1% and
Mholes=0.75, 1.0 and 1.25 for the three turbine speeds. Heat transfer coefficients are determined on the
rotor platform for similar upstream and downstream coolant injection. Rotation is found to significantly
affect the distribution of coolant on the platform. The measured effectiveness magnitudes are lower than that obtained with numerical simulations. Coolant streams from both upstream and downstream injection
orient themselves towards the blade suction side. Passage vortex cuts-off the coolant film for the lower
MFR for upstream injection. As the MFR increases, the passage vortex effects are diminished.
Effectiveness was maximum when Mholes was closer to one as the coolant ejection velocity is
approximately equal to the mainstream relative velocity for this blowing ratio. Heat transfer coefficient
and film cooling effectiveness increase with increasing rotational speed for upstream rotor stator gap
injection while for downstream hole injection the maximum effectiveness and heat transfer coefficients
occur at the reference speed of 2550rpm.
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Experimental Study of Gas Turbine Blade Film Cooling and Heat TransferNarzary, Diganta P. 2009 August 1900 (has links)
Modern gas turbine engines require higher turbine-entry gas temperature to improve their
thermal efficiency and thereby their performance. A major accompanying concern is the heat-up
of the turbine components which are already subject to high thermal and mechanical stresses.
This heat-up can be reduced by: (i) applying thermal barrier coating (TBC) on the surface, and
(ii) providing coolant to the surface by injecting secondary air discharged from the compressor.
However, as the bleeding off of compressor discharge air exacts a penalty on engine performance,
the cooling functions must be accomplished with the smallest possible secondary air injection.
This necessitates a detailed and systematic study of the various flow and geometrical parameters
that may have a bearing on the cooling pattern.
In the present study, experiments were performed in three regions of a non-rotating gas
turbine blade cascade: blade platform, blade span, and blade tip. The blade platform and blade
span studies were carried out on a high pressure turbine rotor blade cascade in medium flow
conditions. Film-cooling effectiveness or degree of cooling was assessed in terms of cooling hole
geometry, blowing ratio, freestream turbulence, coolant-to-mainstream density ratio, purge flow
rate, upstream vortex for blade platform cooling and blowing ratio, and upstream vortex for blade
span cooling. The blade tip study was performed in a blow-down flow loop in a transonic flow
environment. The degree of cooling was assessed in terms of blowing ratio and tip clearance.
Limited heat transfer coefficient measurements were also carried out. Mainstream pressure loss
was also measured for blade platform and blade tip film-cooling with the help of pitot-static
probes. The pressure sensitive paint (PSP) and temperature sensitive paint (TSP) techniques were
used for measuring film-cooling effectiveness whereas for heat transfer coefficient measurement,
temperature sensitive paint (TSP) technique was employed.
Results indicated that the blade platform cooling requires a combination of upstream purge
flow and downstream discrete film-cooling holes to cool the entire platform. The shaped cooling
holes provided wider film coverage and higher film-cooling effectiveness than the cylindrical
holes while also creating lesser mainstream pressure losses. Higher coolant-to-mainstream density ratio resulted in higher effectiveness levels from the cooling holes. On the blade span, at
any given blowing ratio, the suction side showed better coolant coverage than the pressure side
even though the former had two fewer rows of holes. Film-cooling effectiveness increased with
blowing ratio on both sides of the blade. Whereas the pressure side effectiveness continued to
increase with blowing ratio, the increase in suction side effectiveness slowed down at higher
blowing ratios (M=0.9 and 1.2). Upstream wake had a detrimental effect on film coverage. 0%
and 25% wake phase positions significantly decreased film-cooling effectiveness magnitude.
Comparison between the compound shaped hole and the compound cylindrical hole design
showed higher effectiveness values for shaped holes on the suction side. The cylindrical holes
performed marginally better in the curved portion of the pressure side. Finally, the concept tip
proved to be better than the baseline tip in terms of reducing mainstream flow leakage and
mainstream pressure loss. The film-cooling effectiveness on the concept blade increased with
increasing blowing ratio and tip gap. However, the film-coverage on the leading tip portion was
almost negligible.
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Investigation of Film Cooling Strategies CFD versus Experiments -Potential for Using Reduced ModelsNadalina Jafabadi, Hossein January 2010 (has links)
The ability and efficiency of today’s gas turbine engines are highly dependent on development of cooling technologies, among which film cooling is one of the most important. Investigations have been conducted towards discovering different aspects of film cooling, utilizing both experiments and performing CFD simulations. Although, investigation by using CFD analysis is less expensive in general, the results obtained from CFD calculations should be validated by means of experimental results. In addition to validation, in cases like simulating a turbine vane, performing CFD simulations can be time consuming. Therefore, it is essential to find approaches that can reduce the computational cost while results are validated by experiments. This study has shown the potential for reduced models to be utilized for investigation of different aspects of film cooling by means of CFD at low turn-around time. This has been accomplished by first carrying out CFD simulations and experiments for an engine-like setting for a full vane. Then the computational domain is reduced in two steps where all results are compared with experiments including aerodynamic validation, heat transfer coefficient and film effectiveness. While the aerodynamic results are in close agreement with experiments, the heat transfer coefficient and film effectiveness results have also shown similarities within the expected range. Thus this study has shown that this approach can be very useful for e.g. early vane and film cooling design.
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Film cooling effectiveness measurements on rotating and non-rotating turbine componentsAhn, Jaeyong 25 April 2007 (has links)
Detailed film cooling effectiveness distributions were measured on the stationary
blade tip and on the leading edge region of a rotating blade using a Pressure Sensitive
Paint technique. Air and nitrogen gas were used as the film cooling gases and the
oxygen concentration distribution for each case was measured. The film cooling
effectiveness information was obtained from the difference of the oxygen concentration
between air and nitrogen gas cases by applying the mass transfer analogy. In the case of
the stationary blade tip, plane tip and squealer tip blades were used while the film
cooling holes were located (a) along the camber line on the tip or (b) along the span of
the pressure side. The average blowing ratio of the cooling gas was controlled to be 0.5,
1.0, and 2.0. Tests were conducted in a five-bladed linear cascade with a blow down
facility. The free stream Reynolds number, based on the axial chord length and the exit
velocity, was 1,100,000 and the inlet and the exit Mach number were 0.25 and 0.59,
respectively. Turbulence intensity level at the cascade inlet was 9.7%. All
measurements were made at three different tip gap clearances of 1%, 1.5%, and 2.5% of
blade span. Results show that the locations of the film cooling holes and the presence
of squealer have significant effects on surface static pressure and film-cooling effectiveness. Same technique was applied to the rotating turbine blade leading edge
region. Tests were conducted on the first stage rotor of a 3-stage axial turbine. The
Reynolds number based on the axial chord length and the exit velocity was 200,000 and
the total to exit pressure ratio was 1.12 for the first rotor. The effects of the rotational
speed and the blowing ratio were studied. The rotational speed was controlled to be
2400, 2550, and 3000 rpm and the blowing ratio was 0.5, 1.0, and 2.0. Two different
film cooling hole geometries were used; 2-row and 3-row film cooling holes. Results
show that the rotational speed changes the directions of the coolant flows. Blowing
ratio also changes the distributions of the coolant flows. The results of this study will
be helpful in understanding the physical phenomena regarding the film injection and
designing more efficient turbine blades.
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Large eddy simulation of cooling practices for improved film cooling performance of a gas turbine bladeAl-Zurfi, Nabeel January 2017 (has links)
The Large Eddy Simulation approach is employed to predict the flow physics and heat transfer characteristics of a film-cooling problem that is formed from the interaction of a coolant jet with a hot mainstream flow. The film-cooling technique is used to protect turbine blades from thermal failure, allowing the gas inlet temperature to be increased beyond the failure temperature of the turbine blade material in order to enhance the efficiency of gas turbine engines. A coolant fluid is injected into the hot mainstream through several rows of injection holes placed on the surface of a gas turbine blade in order to form a protective coolant film layer on the blade surface. However, due to the complex, unsteady and three-dimensional interactions between the coolant and the hot gases, it is difficult to achieve the desired cooling performance. Understanding of this complex flow and heat transfer process will be helpful in designing more efficiently cooled rotor blades. A comprehensive numerical investigation of a rotating film-cooling performance under different conditions is conducted in this thesis, including film-cooling on a flat surface and film-cooling on a rotating gas turbine blade. The flow-governing equations are discretised based on the finite-volumes method and then solved iteratively using the well-known SIMPLE and PISO algorithms. An in-house FORTRAN code has been developed to investigate the flat plate film-cooling configuration, while the gas turbine blade geometry has been simulated using the STAR-CCM+ CFD commercial code. The first goal of the present thesis is to investigate the physics of the flow and heat transfer, which occurs during film-cooling from a standard film hole configuration. Film-cooling performance is analysed by looking at the distribution of flow and thermal fields downstream of the film holes. The predicted mean velocity profiles and spanwise-averaged film-cooling effectiveness are compared with experimental data in order to validate the reliability of the LES technique. Comparison of adiabatic film-cooling effectiveness with experiments shows excellent agreement for the local and spanwise-averaged film-cooling effectiveness, confirming the correct prediction of the film-cooling behaviour. The film coverage and film-cooling effectiveness distributions are presented along with discussions of the influence of blowing ratio and rotation number. Overall, it was found that both rotation number and blowing ratio play significant roles in determining the film-cooling effectiveness distributions. The second goal is to investigate the impact of innovative anti-vortex holes on the film-cooling performance. The anti-vortex hole design counteracts the detrimental kidney vorticity associated with the main hole, allowing coolant to remain attached to the blade surface. Thus, the new design significantly improves the film-cooling performance compared to the standard hole arrangement, particularly at high blowing ratios. The anti-vortex hole technique is unique in that it requires only readily machinable round holes, unlike shaped film-cooling holes and other advanced concepts. The effects of blowing ratio and the positions of the anti-vortex side holes on the physics of the hot mainstream-coolant interaction in a film-cooled turbine blade are also investigated. The results also indicate that the side holes of the anti-vortex design promote the interaction between the vortical structures; therefore, the film coverage contours reveal an improvement in the lateral spreading of the coolant jet.
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COMPUTATIONAL SIMULATION AND ANALYSIS OF FILM COOLING FOR THE LEADING-EDGE MODEL OF A TURBINE BLADELITZLER, JEFFREY W. 03 July 2007 (has links)
No description available.
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Experimental characterisation of the coolant film generated by various gas turbine combustor liner geometriesChua, Khim Heng January 2005 (has links)
In modern, low emission, gas turbine combustion systems the amount of air available for cooling of the flame tube liner is limited. This has led to the development of more complex cooling systems such as cooling tiles i.e. a double skin system, as opposed to the use of more conventional cooling slots i.e. a single skin system. An isothennal experimental facility has been constructed which can incorporate 10 times full size single and double skin (cooling tile) test specimens. The specimens can be tested with or without effusion cooling and measurements have been made to characterise the flow through each cooling system along with the velocity field and cooling effectiveness distributions that subsequently develop along the length of each test section. The velocity field of the coolant film has been defined using pneumatic probes, hot-wire anemometry and PIV instrumentation, whilst gas tracing technique is used to indicate (i) the adiabatic film cooling effectiveness and (ii) mixing of the coolant film with the mainstream flow. Tests have been undertaken both with a datum low turbulence mainstream flow passing over the test section, along with various configurations in which large magnitudes and scales of turbulence were present in the mainstream flow. These high turbulence test cases simulate some of the flow conditions found within a gas turbine combustor. Results are presented relating to a variety of operating conditions for both types of cooling system. The nominal operating condition for the double skin system was at a coolant to mainstream blowing ratio of approximately 1.0. At this condition, mixing of the mainstream and coolant film was relatively small with low mainstream turbulence. However, at high mainstream turbulence levels there was rapid penetration of the mainstream flow into the coolant film. This break up of the coolant film leads to a significant reduction in the cooling effectiveness. In addition to the time-averaged characteristics, the time dependent behaviour of the .:coolantfilm was. also investigated. In particular, unsteadiness associated with large scale structures in the mainstream flow was observed within the coolant film and adjacent to the tile surface. Relative to a double skin system the single skin geometry requires a higher coolant flow rate that, along with other geometrical changes, results in typically higher coolant to mainstream velocity ratios. At low mainstream turbulence levels this difference in velocity between the coolant and mainstream promotes the generation of turbulence and mixing between the streams so leading to some reduction in cooling effectiveness. However, this higher momentum coolant fluid is more resistant to high mainstream turbulence levels and scales so that the coolant film break up is not as significant under these conditions as that observed for the double skin system. For all the configurations tested the use of effusion cooling helped restore the coolant film along the rear of the test section. For the same total coolant flow, the minimum value of cooling effectiveness observed along the test section was increased relative to the no effusion case. In addition the effectiveness of the effusion patch depends on the amount of coolant injected and the axial location of the patch. The overall experimental data suggested the importance of the initial cooling film conditions together with better understanding of the possible mechanisms that results in the rapid cooling film break-up, such as high turbulence mainstream flow and scales, and this will lead to a more effective cooling system design. This experimental data is also thought to be ideal for the validation of numerical predictions.
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Turbine blade platform film cooling with simulated stator-rotor purge flow with varied seal width and upstream wake with vortexBlake, Sarah Anne 15 May 2009 (has links)
The turbine blade platform can be protected from hot mainstream gases by injecting
cooler air through the gap between stator and rotor. The effectiveness of this film
cooling method depends on the geometry of the slot, the quantity of injected air, and the
secondary flows near the platform. The purpose of this study was to measure the effect
of the upstream vane or stator on this type of platform cooling, as well as the effect of
changes in the width of the gap.
Film cooling effectiveness distributions were obtained on a turbine blade platform within
a linear cascade with upstream slot injection. The width of the slot was varied as well as
the mass flow rate of the injected coolant. Obstacles were placed upstream to model the
effect of the upstream vane. The coolant was injected through an advanced labyrinth
seal to simulate purge flow through a stator-rotor seal. The width of the opening of this
seal was varied to simulate the effect of misalignment. Stationary rods were placed
upstream of the cascade in four phase locations to model the unsteady wake formed at
the trailing edge of the upstream vane. Delta wings were also placed in four positions to
create a vortex similar to the passage vortex at the exit of the vane. The film cooling
effectiveness distributions were measured using pressure-sensitive paint (PSP).
Reducing the width of the slot was found to decrease the area of coolant coverage,
although the film cooling effectiveness close to the slot was slightly increased. The
unsteady wake was found to have a trivial effect on platform cooling, while the passage
vortex from the upstream vane may significantly reduce the film cooling effectiveness.
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Experimental and Numerical Study of Endwall Film CoolingMahadevan, Srikrishna 01 January 2015 (has links)
This research work investigates the thermal performance of a film-cooled gas turbine endwall under two different mainstream flow conditions. In the first part of the research investigation, the effect of unsteady passing wakes on a film-cooled pitchwise-curved surface (representing an endwall without airfoils) was experimentally studied for heat transfer characteristics on a time-averaged basis. The temperature sensitive paint technique was used to obtain the local temperatures on the test surface. The required heat flux input was provided using foil heaters. Discrete film injection was implemented on the test surface using cylindrical holes with a streamwise inclination angle of 35? and no compound angle relative to the mean approach velocity vector. The passing wakes increased the heat transfer coefficients at both the wake passing frequencies that were experimented. Due to the increasing film cooling jet turbulence and strong jet-mainstream interaction at higher blowing ratios, the heat transfer coefficients were amplified. A combination of film injection and unsteady passing wakes resulted in a maximum pitch-averaged and centerline heat transfer augmentation of ? 28% and 31.7% relative to the no wake and no film injection case. The second part of the research study involves an experimental and numerical analysis of secondary flow and coolant film interaction in a high subsonic annular cascade with a maximum isentropic throat Mach number of ? 0.68. Endwall (platform) thermal protection is provided using discrete cylindrical holes with a streamwise inclination angle of 30? and no compound angle relative to the mean approach velocity vector. The surface flow visualization on the inner endwall provided the location of the saddle point and the three-dimensional separation lines. Computational predictions showed that the leading-edge horseshoe vortex was confined to approximately 1.5% of the airfoil span for the no film injection case and intensified with low momentum film injection. At the highest blowing ratio, the film cooling jet weakened the horseshoe vortex at the leading-edge plane. The passage vortex was intensified with coolant injection at all blowing ratios. It was seen that increasing average blowing ratio improved the film effectiveness on the endwall. The discharge coefficients calculated for each film cooling hole indicated significant non-uniformity in the coolant discharge at lower blowing ratios and the strong dependence of discharge coefficients on the mainstream static pressure and the location of three-dimensional separation lines. Near the airfoil suction side, a region of coalesced film cooling jets providing close to uniform film coverage was observed, indicative of the mainstream acceleration and the influence of three-dimensional separation lines.
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Experimental investigation of the performance of a fully cooled gas turbine vane with and without mainstream flow and experimental analysis supporting the redesign of a wind tunnel test sectionMosberg, Noah Avram 16 February 2015 (has links)
This study focused on experimentally determining the cooling performance of a fully cooled, scaled-up model of a C3X turbine vane. The primary objective was to determine the differences in overall effectiveness in the presence and absence of a hot mainstream flowing over the vane. Overall effectiveness was measured using a thermally scaled matched Biot number vane with an impingement plate providing the internal cooling. This is the first study focused on investigating the effect of removing the mainstream flow and comparing the contour and laterally-averaged effectiveness data in support of the development of an assembly line thermal testing method. It was found that the proposed method of factory floor testing of turbine component cooling performance did not provide comparable information to traditional overall effectiveness test methods. A second experiment was performed in which the effect of altering the angle of attack of a flow into a passive turbulence generator was investigated. Measurements in the approach flow were taken using a single wire hot-wire anemometer. This study was the first to investigate the effects such a setup would have on fluctuating flow quantitates such as turbulence intensity and integral length scale rather than simply the mean quantities. It was found that both the downstream turbulence intensity and the turbulence integral length scale increase monotonically with approach flow incidence angle at a specified distance downstream of the turbulence generator. / text
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