Download or read book Heat Transfer Due to Unsteady Effects as Investigated in a High speed Full scale Fully cooled Turbine Vane and Rotor Stage written by Jonathan R. Mason and published by . This book was released on 2008 with total page 236 pages. Available in PDF, EPUB and Kindle. Book excerpt:

Download or read book The Influence of Film Cooling and Inlet Temperature Profile on Heat Transfer for the Vane Row of a 1 1 2 Stage Transonic High pressure Turbine written by Harika Senem Kahveci and published by . This book was released on 2010 with total page 269 pages. Available in PDF, EPUB and Kindle. Book excerpt: Abstract: The goal of this research was to establish an extensive database for typical engine hardware with a film-cooled first stage vane, which represents the foundation for future turbomachinery film cooling modeling and component heat transfer studies. Until this time, such a database was not available within the gas turbine industry. Accordingly, the study focuses on determination of the local heat flux for the airfoil and endwall surfaces of the vane row of a fully-cooled turbine stage. The measurements were performed at the Ohio State University Gas Turbine Laboratory using the Turbine Test Facility. The full-scale rotating 1 and 1/2 turbine stage is operated at the proper corrected engine design conditions: Flow Function (FF), corrected speed, stage Pressure Ratio (PR), and temperature ratios of gas to wall and gas to coolant. The primary measurements of temperature, pressure, and heat flux are repeated for different vane inlet temperature profiles and different vane cooling flows to establish an understanding of the influence of film cooling on local heat transfer. Double-sided Kapton heat-flux gauges are used for heat-flux measurements at different span locations along the airfoil surfaces and along the inner endwall. The cooling scheme consists of numerous cooling holes located on the endwalls, at the airfoil leading edge, on the airfoil pressure and suction surfaces, and at the trailing edge, resulting in a fully cooled first stage vane. The unique film-cooled endwall heat transfer data demonstrated in contour plots reveals insight to the complex flow behavior that is dominant in this region, which becomes even more complicated with the addition of coolant. Varying profile shapes resulted in significant heat transfer variations in a growing fashion towards the trailing edge region, which increased in magnitude when there is no coolant supply. The largest cooling effect is observed on 5% span pressure surface and at the inner endwall region. Heat transfer decreases from tip towards hub with addition of cooling. However, a similar decrease is not observed at the inner endwall region by doing so, which suggests excess coolant once beyond an optimum blowing ratio. Cooling flow rate and temperature profile shape affect the distributions on the airfoil surface very similarly, the latter observed more clearly at the endwall region. The vane outer cooling effect is comparable to the combined coolant effect at all surfaces, while no impact of purge flow is observed. Aligning the hot streaks with the vane leading edge lowered heat transfer compared to mid-passage alignment at the mid-span suction surface and through the endwall passage, and increased it at the endwall exit, while the pressure surface is found to be insensitive to this switch. Comparison with a previous research program with the un-cooled version of the vane gave good agreement on the pressure surface and at the endwall, but significantly lower heat transfer on the suction surface due to ingestion of the hot flow through the cooling holes when there is no cooling.

Download or read book Effects of Hot Streak and Phantom Cooling on Heat Transfer in a Cooled Turbine Stage Including Particulate Deposition written by and published by . This book was released on 2016 with total page 83 pages. Available in PDF, EPUB and Kindle. Book excerpt: The objective of this research effort was to develop a validated computational modeling capability for the characterization of the effects of hot streaks and particulate deposition on the heat load of modern gas turbines. This was accomplished with a multi-faceted approach including analytical, experimental, and computational components. A 1-year no cost extension request was approved for this effort, so the total duration was 4 years. The research effort succeeded in its ultimate objective by leveraging extensive experimental deposition studies complemented by computational modeling. Experiments were conducted with hot streaks, vane cooling, and combinations of hot streaks with vane cooling. These studies contributed to a significant body of corporate knowledge of deposition, in combination with particle rebound and deposition studies funded by other agencies, to provide suitable conditions for the development of a new model. The model includes the following physical phenomena: elastic deformation, plastic deformation, adhesion, and shear removal. It also incorporates material property sensitivity to temperature and tangential-normal velocity rebound cross-dependencies observed in experiments. The model is well-suited for incorporation in CFD simulations of complex gas turbine flows due to its algebraic (explicit) formulation. This report contains model predictions compared to coefficient of restitution data available in the open literature as well as deposition results from two different high temperature turbine deposition facilities. While the model comparisons with experiments are in many cases promising, several key aspects of particle deposition remain elusive. The simple phenomenological nature of the model allows for parametric dependencies to be evaluated in a straightforward manner. This effort also included the first-ever full turbine stage deposition model published in the open literature. The simulations included hot streaks and simulated vane cooling. The new deposition model was implemented into the CFD model as a wall boundary condition, with various particle sizes investigated in the simulation. Simulations utilizing a steady mixing plane formulation and an unsteady sliding mesh were conducted and the flow solution of each was validated against experimental data. Results from each of these simulations, including impact and capture distributions and efficiencies, were compared and potential reasons for differences discussed in detail. The inclusion of a large range of particle sizes allowed investigation of trends with particle size, such as increased radial migration and reduced sticking efficiency at the larger particle sizes. The unsteady simulation predicted lower sticking efficiencies on the rotor blades than the mixing plane simulation for the majority of particle sizes. This is postulated to be due to the preservation of the hot streak and cool vane wake through the vane-rotor interface (which are smeared out circumferentially in the mixing-plane simulation). The results reported here represent the successful implementation of a novel deposition model into validated vane-rotor flow solutions that include a non-uniform inlet temperature profile and simulated vane cooling.

Download or read book Applied mechanics reviews written by and published by . This book was released on 1948 with total page 400 pages. Available in PDF, EPUB and Kindle. Book excerpt:

Download or read book Blade to coolant Heat transfer Results and Operating Data from a Natural convection Water cooled Single stage Turbine written by Anthony J. Diaguila and published by . This book was released on 1951 with total page 28 pages. Available in PDF, EPUB and Kindle. Book excerpt:

Download or read book Determination of Blade to coolant Heat transfer Coefficients on a Forced convection Water cooled Single stage Turbine written by John C. Freche and published by . This book was released on 1951 with total page 30 pages. Available in PDF, EPUB and Kindle. Book excerpt:

Download or read book Investigation of Hot Streak Migration and Film Cooling Effects on Heat Transfer in Rotor Stator Interacting Flows Report 1 written by and published by . This book was released on 1992 with total page 226 pages. Available in PDF, EPUB and Kindle. Book excerpt: Experimental data taken from turbine engines has shown that hot streaks exiting combustors can have a significant impact upon the secondary flow and will temperature of the first stage turbine rotor. Understanding the secondary flow and heat transfer effects due to combustor hot streaks is essential to turbine designers attempting to optimize turbine cooling systems. A numerical investigation has been performed which addresses the issues of multi-blade count ratio and three-dimensionality effects on the prediction of combustor hot streak migration in a turbine stage. The two- and three-dimensional Navier-Stokes analyses developed by Rai et al are used to predict unsteady viscous rotor-stator interacting flow in the presence of a combustor hot streak with heat transfer and film cooling. Predicted results are presented for a two-dimensional 3-stator/4-rotor and a three-dimensional 1-stator/1-rotor simulations of streak migration through a turbine stage. Comparison of these results with experimental data demonstrates the capability of the three- dimensional procedure to capture most of the flow physics associated with hot streak migration including the effects of combustor hot streaks on turbine rotor surface temperatures.

Download or read book Investigation of Hot Streak Migration and Film Cooling Effects on the Heat Transfer in Rotor Stator Interacting Flows written by Daniel J. Dorney and published by . This book was released on 1991 with total page 22 pages. Available in PDF, EPUB and Kindle. Book excerpt: This effort extended the three-dimensional unsteady rotor/stator interaction code, ROTOR3, to investigate hot streak migration and film cooling effects on the passage flow and blade surface heat transfer for an axial flow turbine stage. These objectives are part of an overall plan to extend the capabilities of this numerical procedure and to determine the potential of this technique to impact the design of future rotating turbomachinery components. The principal benefits that will result from this effort are: A diagnostic analysis and code which can be used to help design turbine blades with increased efficiency and reduced cooling requirements, An open literature demonstration on the use of Computational Simulation and Scientific Visualization for gaining insight into complex turbomachinery flows, Acceleration of the transition of large-scale computational analyses to the turbomachinery design process through joint involvement between UTRC and Pratt and Whitney turbine engineers, and A useful and proven design tool which can be used with existing and future engine design procedures.

Download or read book Heat transfer measurements for a film cooled turbine vane cascade written by Douglas R. Thurman and published by . This book was released on 2008 with total page 20 pages. Available in PDF, EPUB and Kindle. Book excerpt:

Download or read book Unsteady Analysis of Hot Streak Migration in a Turbine Stage written by Nateri K. Madavan and published by . This book was released on 1990 with total page 15 pages. Available in PDF, EPUB and Kindle. Book excerpt: Experimental data taken from gas turbine engines has shown that hot streaks exiting combustors can have a significant impact upon the secondary flow and wall temperature of the first stage turbine rotor. Understanding the secondary flow and heat transfer effects due to combustor hot streaks is essential to turbine designers attempting to optimize turbine cooling systems. A numerical investigation is presented which addresses the issues of multi-blade count ratio and three-dimensionality effects on the prediction of combustor hot streak migration in a turbine stage. Keywords: Gas generator engines, Gas turbine rotors, Unsteady flow. (CP).

Download or read book An Investigation of the Effects of Nozzle Guide Vane Trailing Edge Cooling on Rotor Heat Transfer in a Transonic Turbine Stage written by Aaron James Gleixner and published by . This book was released on 1992 with total page 330 pages. Available in PDF, EPUB and Kindle. Book excerpt:

Download or read book Heat Transfer in Gas Turbines written by Bengt Sundén and published by Witpress. This book was released on 2001 with total page 544 pages. Available in PDF, EPUB and Kindle. Book excerpt: This title presents and reflects current active research on various heat transfer topics and related phenomena in gas turbine systems. It begins with a general introduction to gas turbine heat transfer, before moving on to specific areas.

Download or read book Heat Transfer in the Blade Row and Tip Region of a Modern Transonic High Pressure Turbine with and Without Forward Cavity Purge Flow written by Stephen M. Molter and published by . This book was released on 2006 with total page 322 pages. Available in PDF, EPUB and Kindle. Book excerpt: Abstract: A full scale rotating turbine rig operated at design corrected conditions has been used to study the heat transfer mechanisms affecting the flowpath surfaces within a modem single stage high-pressure (HP) turbine. The experimental rig was first run completely un-cooled and is currently being re-constructed to accommodate HP vane airfoil and endwall cooling and inner-stage cavity purge flow injection. Heat flux and pressure data were measured for both a flat and recessed, or squealer, HP blade tip and the stationary shroud above. The measurements indicate that the recessed tip, used in the majority of modem turbines to minimize blade damage from rubs, increases the blade heat load overall, and creates several hot spots on the floor of the recess for an un-cooled airfoil. The tip data also showed there were significant unsteady variations in the heat load at the vane passing frequency. Steady state CFD calculations were completed for both flat and squealer tip configurations to examine if the analysis could capture the details that were measured. The CFD, while not capable of estimating the unsteady heat load component and generally over predicting the overall heat flux by 10-25%, did capture the measured heat flux trends in the recessed tip. The steady state CFD prediction did show good agreement with the time-accurate data along the stationary shroud. These results show that steady-state CFD analysis can be useful in predicting the complex flow field and heat load distribution in turbine blade tips to help guide future blade designs. Pre-test CFD predictions were also performed for the upcoming series of experiments that include replacing the un-cooled vane row with a fully cooled HP vane row and the introduction of HP blade forward cavity purge flow, while leaving the HP blade un-cooled. The focus of the steady state predictions for the HP blade row was two fold; to assist in guiding the placement of new heat flux and pressure instrumentation and to study the cold flow migration through the HP blade row. Adiabatic wall temperature and Nusselt number predictions along the blade surface showed large radial, or spanwise, gradients, mostly along the suction side of the blade. Surface visualization on the suction side of the blade revealed two bands of cooler regions located at the upper and lower spans, with the middle spans being hotter, comparable to the pressure side. The lower band of cool flow is a result of the forward cavity purge flow, which mostly migrates to the suction side of the blade passage. By the trailing edge of the blade the purge flow has migrated upwards to an extent of approximately 20% of the blade span surface. The upper band of cool flow is a result of the cooling flow from the HP vane outer endwall. Blade tip secondary flows and the tip leakage vortex act to entrain this cool flow into the tip gap, resulting in its migration to the suction side of the blade. Due to a downward migration of the tip vortex along the suction side through the blade row, this cool band along the airfoil surface affects the upper 20% of the blade span. Results from the pre-test CFD predictions were also analyzed along the blade platform and rim seal surfaces. Migration of the forward cavity purge flow towards the suction side increases the Nusselt number along the rim seal and platform in these areas. Adiabatic wall temperatures on the platform surface were reasonably constant and lower than those on the blade surface, an effect of a portion of the purge flow being entrained into the platform boundary layer. Pressure asymmetries along the rim seal created circumferential variations in the local purge mass flow rates, with the leading edge of blade being the location of highest pressure and thus lower purge injection. Results also indicate the possibility of hot gas ingestion into the upper region of the rim seal at the leading edge. When data is available from the updated turbine rig, the comparisons will help to further validate this code as a useful design tool.

Download or read book Film Cooling Heat Transfer and Aerodynamic Measurements in a Three Stage Research Gas Turbine written by Arun Suryanarayanan and published by . This book was released on 2010 with total page pages. Available in PDF, EPUB and Kindle. Book excerpt: 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.

Download or read book Effects of Shocks on the Unsteady Heat Transfer in a Film Cooled Transonic Turbine Cascade written by and published by . This book was released on 2001 with total page 0 pages. Available in PDF, EPUB and Kindle. Book excerpt: An experimental program was performed to study the effects of high strength shocks on the film cooling heat transfer in a transonic turbine blade. Shocks stronger than seen in actual engines were generated by a shock tube and directed to pass into the cascade upstream of a linear set of blades. The results indicate that the increased strength of the shocks do not disrupt the boundary layer on the blade. Therefore, the heat transfer is still predictable by the methods determined for lower strength shocks, as reported in the attached two papers (presented in Munich during this research program). As a transition to work on high free-stream turbulence effects, initial results for a blown grid were obtained in low speed flow. The corresponding report is attached, but has not yet been published.

Download or read book Effects of unsteady inflow conditions on the heat transfer to turbine blades written by S. Wittig and published by . This book was released on 1992 with total page 141 pages. Available in PDF, EPUB and Kindle. Book excerpt: Convective heat transfer, blade cooling, experimental investigation, numerical investigation, unsteady wakes, boundary layer, unsteady interaction, wake induced transition, boundary layer code, intermittency model.

Download or read book The Effects of Leading Edge and Downstream Film Cooling on Turbine Vane Heat Transfer written by National Aeronautics and Space Administration (NASA) and published by Createspace Independent Publishing Platform. This book was released on 2018-07-23 with total page 178 pages. Available in PDF, EPUB and Kindle. Book excerpt: The progress under contract NAS3-24619 toward the goal of establishing a relevant data base for use in improving the predictive design capabilities for external heat transfer to turbine vanes, including the effect of downstream film cooling with and without leading edge showerhead film cooling. Experimental measurements were made in a two-dimensional cascade previously used to obtain vane surface heat transfer distributions on nonfilm cooled airfoils under contract NAS3-22761 and leading edge showerhead film cooled airfoils under contract NAS3-23695. The principal independent parameters (Mach number, Reynolds number, turbulence, wall-to-gas temperature ratio, coolant-to-gas temperature ratio, and coolant-to-gas pressure ratio) were maintained over ranges consistent with actual engine conditions and the test matrix was structured to provide an assessment of the independent influence of parameters of interest, namely, exit Mach number, exit Reynolds number, coolant-to-gas temperature ratio, and coolant-to-gas pressure ratio. Data provide a data base for downstream film cooled turbine vanes and extends the data bases generated in the two previous studies. The vane external heat transfer obtained indicate that considerable cooling benefits can be achieved by utilizing downstream film cooling. The data obtained and presented illustrate the interaction of the variables and should provide the airfoil designer and computational analyst the information required to improve heat transfer design capabilities for film cooled turbine airfoils. Hylton, L. D. and Nirmalan, V. and Sultanian, B. K. and Kaufman, R. M. Unspecified Center EQUIPMENT SPECIFICATIONS; FILM COOLING; HEAT TRANSFER; LEADING EDGES; STRUCTURAL DESIGN; VANES; AIRCRAFT ENGINES; CASCADE FLOW; DATA PROCESSING; GAS TURBINES; HIGH TEMPERATURE; PARAMETERIZATION; TWO DIMENSIONAL FLOW...