Recent/Current Research at the Gas Turbine Laboratory
A Unified Approach for Vaned Diffuser Design in Advanced Centrifugal Compressors
Jonathan Everitt
Advisor: Prof. Spakovszky
Modern internal combustion engines utilize high pressure ratio turbochargers combined with NOx control strategies to improve efficiency and reduce emissions. In such an application, the centrifugal compressor has to simultaneously achieve high efficiency, high pressure ratio and a broad operating range. To meet these requirements, the trend has been towards highly loaded compressors, utilizing high speed impellers with backward-leaning blades for extended operating range and vaned diffusers for enhanced pressure recovery with compact geometry.
The flow out of the impeller presents multiple challenges for the vaned diffuser: it is transonic, unsteady, and highly non-uniform in both axial and circumferential directions. The relative importance of each of these factors is not well understood: different researchers have reached different conclusions regarding, for example, the importance of the impeller outflow non-uniformity. Diffuser design therefore largely depends upon historical correlations, CFD simulation and careful experimentation.
The objective of the investigation is therefore to rigorously establish the links between diffuser geometry, performance, component matching and stability. The technical approach combines first principles based modeling with high-fidelity calculations and experiments using a unique swirling flow diffuser test rig at the Gas Turbine Laboratory. The goal is to develop design criteria and to define performance metrics expressed in terms of overall vane parameters and appropriately averaged inflow properties that can be applied in the preliminary design stage.
A Methodology for Centrifugal Compressor Stability Prediction
Jonathan Everitt and Anjaney Kottapalli
Advisor: Prof. Spakovszky
Although centrifugal compressors exhibit the same type of instabilities as axial compressors, rotating stall and surge are characterized by a much broader spectrum of unstable behavior. The wide variety of instability behavior, along with the inherently complicated flow in such a machine, are primary reasons that rotating stall and surge in centrifugal compressors are less well understood than similar phenomena in axial compressors. As a consequence, a general theory or a criterion for the onset of instability in centrifugal compressors does not exist. Instead, correlations are used to describe the surge point for a certain class of centrifugal compressors and to estimate the stability limit based on a priori knowledge of blade row characteristics. The major limitation of these methods is that these characteristics are only available after experimental measurements and thus the method is not of predictive nature. This research project is different from past efforts in that the prediction is purely based on centrifugal compressor geometry and does not rely on correlations or a priori knowledge of compressor characteristics. The approach is two-pronged. Previous research indicates that for certain classes of centrifugal compressors the inception of instability is in the diffuser; however the underlying fluid mechanics is not well understood. To gain insight, unsteady 3-D RANS calculations were carried out on the isolated diffuser using an inlet flow field derived from full stage calculations. The inlet conditions were perturbed with a short wavelength total pressure disturbance. It was shown that flow separation at the diffuser vane leading edge, combined with recirculating flow in the vaneless space, results in the development of vortical structures which convect at similar speed to experimentally measured spike stall precursors.
Cartoon indicating the mechanism for the onset of instability in the isolated diffuser simulations, suggested to be responsible for short-wavelength “spike” stall inception.
Unsteady pressure traces for pressure taps located at constant radius around vaneless space. Left: Response to forcing for an unstable operating point at 100% speed in the isolated diffuser simulations (from Everitt & Spakovszky, 2011). Right: Experimental data showing spike stall inception in the same compressor (from Spakovszky & Roduner, 2009).
The second prong to the approach borrows ideas from previous work on axial compressors and consists of 3-D steady RANS calculations to determine the body force distributions representing the effects of discrete blades on the flow field. The body forces are then coupled to a 3-D unsteady RANS solver, which can be run much faster than an unsteady bladed simulation. The compressor model is then forced with a short wavelength body force impulse in the vaneless space. The goal is to demonstrate that the method can accurately predict both the stall point and the type of stall inception pattern (short wavelength spikes or long wavelength modal waves) in centrifugal compressors.
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Representation of bladed centrifugal compressor with axisymmetric body forces. |
Improved Performance Return Channel Design for Multistage Centrifugal Compressors
Anne-Raphaelle Aubry
Advisor: Prof. Greitzer
High-pressure multistage centrifugal compressors are used extensively in the energy industry across a wide variety of applications from refinery processes to gas injection for carbon capture and sequestration. Centrifugal compressor manufacturers are looking towards reduced radial and axial dimension compressors to meet customer’s demands for lower cost and higher reliability. As the dimensions of the centrifugal compressors shrink, the job of the return channel—which must turn the flow by 180° and remove the tangential component of the flow—becomes more difficult.
MIT, in collaboration with Mitsubishi Heavy Industries (MHI), is developing a novel return channel design for these multistage compressors with the objective of improving efficiency, while meeting geometry constraints.
Opportunities to improve “traditional” return channel design were identified in a previous investigation and qualitative best practices established. A quantitative assessment of these best practices is being undertaken, and use of an adjoint method to optimize the return channel shape is also under consideration. Candidate designs obtained with this adjoint method would then be refined to develop a design that addresses the desired performance improvements. Performance of the candidate design are to be assessed in a full-scale stage test at the MHI single-stage test facility.
Contours of radial velocity (left) and entropy production rate (right) in the baseline return vane show a region of reduced velocity flow on the vane suction side, surrounded by a region of high entropy production. |
Investigation of Real Gas Effects in Supercritical CO2 Compression Systems
Nikola Baltadjiev, and Dr. Claudio Lettieri
Advisor: Prof. Spakovszky
Reduction of harmful CO2 emissions from power plants is becoming a major concern in industry and a high priority for research, development, and deployment projects. Carbon capture and sequestration in underground wells requires the fluid to be compressed to high pressures (exceeding 100 bar) reaching supercritical conditions. At this state the fluid has density similar to a liquid, but at the same time expends to fill up a volume just like a gas does. Although the properties of these working fluids can be characterized, little is known about their behavior in turbomachinery.
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Contours of compressibility factor Z in a Temperature-Entropy diagram |
This project aims to develop a fundamental understanding of the changes in flow behavior associated with the supercritical state of the fluid and determine the root cause for performance and stability issues observed in compressor stages. The goal is to achieve a step change in performance and stability margin via aerodynamic redesign of the individual stage components.
Compressor systems working with dense gases usually operate at low flow rates, using low flow coefficient stages. This requires relatively narrow passages characterized by high friction losses. At the same time leakage and windage losses become more prominent. Careful aerodynamic re-design of individual stage components is necessary to limit the performance penalty associated with these applications.
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Kelvin-Helmholtz instability from 3D Unsteady Real Gas CFD simulation |
Unlike ideal fluids, thermodynamic properties of real fluids can vary significantly, especially near the critical point, altering the gas dynamic
behavior. Investigating the impact on basic compressible flow relations can help gain insight into the operation of the compressor stage. Furthermore, flow acceleration around the impeller leading edge can lead to the working fluid locally entering the two-phase region, suggesting the possibility of condensation.
CFD simulations are carried out on selected design concepts to characterize supercritical fluid behavior in realistic turbomachinery environments and to quantify potential performance gains. Future work may involve laboratory scale experiments on canonical flow situations with supercritical CO2, and performance tests on new stage designs to validate key findings and proposed improvement concepts.
Modeling Instabilities in High-Pressure Pumping Systems
Dr. Jeff Defoe
Advisor: Prof. Spakovszky
Pumps are used in the process industry and in power plant applications. They must be able to operate over a wide range of flow rates resulting from variations in load. The particular pumping system under consideration is unstable near its best efficiency point. The challenges are to determine the physical mechanism leading to this system-wide instability and how to redesign the system to ensure stable operation at all relevant flow rates. The system is modeled using a previously-established dynamic modeling framework. This is combined with appropriate boundary conditions to obtain eigenvalues, which are natural frequencies of the system, and their associated growth rates. This determines the stability characteristics of the system. The pumping system model is schematically shown below and includes a plenum containing a gas spring, various area changes, and piping both upstream and downstream of the pump.
Some system components dissipate energy, while others act as energy sources. The pump is an active element, which can dissipate or add energy to the system. The working hypothesis is that the pump is the cause of the observed instabilities. The pump is comprised of an impeller (rotating part) and a volute (stationary part). The impeller dynamics are known. The volute dynamics, however, are not known. It is therefore necessary to obtain a dynamic model of the volute to complete the system stability model. To obtain such a model and to identify the source of the unsteadiness, unsteady RANS CFD calculations of the flow in the volute are conducted. These revealed that the key mechanism governing the self-excited unsteady flow in the volute is bluff body flow separation in the return channel. At the best efficiency operating point, the dynamic behavior in the volute due to this bluff body separation is such that energy is fed into the system. Overall there is insufficient damping, leading to dynamic instability. The figure below illustrates the mechanism which leads to the unsteady behavior in a simplified planar diffuser geometry and relates it to the flow in the volute.
The next step is to conduct a numerical parametric study of the effects of velocity and pressure perturbations on the flow in the volute. This system identification study will result in a dynamic model for use in the existing modeling framework. The framework can then be used to predict at which flow rates unstable operation will occur. Such a prediction tool, coupled with the insight gained into the unsteady flow in the volute, will be used to guide design changes to render the system stable.
Aeromechanic Response in High Performance Centrifugal Compressor Stage
Christopher Lusardi
Advisor: Dr. Tan
Impeller blades in centrifugal compressors are exposed to unsteady forces that can increase stress levels in the part, leading to premature structural failure. These unsteady forces may arise from different sources, a significant one of which is the unsteady pressure field from the impeller interaction with the downstream diffuser. The resulting time-varying loads can induce vibratory stresses in the blades that could be significantly higher than the steady-state stresses. There have been experimental/test/field observations of impeller blades breaking at trailing edge as well as leading edge that are traceable to unsteadiness associated with impeller-diffuser interactions. Furthermore, test data shows that not only does the unsteady impeller-diffuser interaction impact the impeller forced response characteristic but that it is also highly sensitive to impeller-diffuser gap variation; the situation with impeller blade leading edge exhibiting high response that is linked to impeller-diffuser interactions constitutes an upstream manifestation of a downstream stimulus. The overall goal is to characterize and identify the unsteady flow process in impeller-diffuser interaction on the observed impeller aeromechanic difficulty such as those described above.
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Ported Shroud Operation in Turbochargers
George Christou
Advisor: Dr. Tan
In recent years, due to environmental regulations, automotive turbochargers have been increasingly implemented to accomplish high powering and downsizing of internal combustion engines. The operability of the compressor is bound at low mass flow rate by the surge line. Surge is characterized as a breakdown of the flow with large pressure fluctuations that can cause rapid deterioration and in some cases failure of the compressor and the bearing system. A technique used to control the development of surge is by implementing a ported shroud at the inlet of the compressor. The ported shroud configuration is used to improve both the choke and surge lines on the compressor performance map.
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Garret by Honeywell, Turbo Tech 103 (Expert), 2006 |
The overall goal of this research project, in collaboration with Honeywell Turbo Technologies, is to improve the performance of ported shroud centrifugal turbochargers. Specific goals include: providing an explanation of changes in the flow processes with and without ported shroud relative to compressor operation; identifying and quantifying loss mechanisms present in ported shroud centrifugal compressors; increasing the effective operating range by increasing surge and choke margins; and increasing the efficiency at off-design operating points.
Manifestation of Forced Response in High Performance Centrifugal Compressor Stage for Aerospace Applications
Edward Walton
Advisor: Dr. Tan
Impeller blades in a centrifugal compressor stage operate in an unsteady pressure field due to the presence of a downstream diffuser. This unsteady pressure or loading is more generally the result of what is referred to as the “impeller-diffuser” interaction. The primary result of this interaction is to set up pressure waves which traverse and decay from the trailing edge to leading edge of the blades. It is these unsteady pressure waves which are thought to be the primary driver of whether an indicated resonance on a Campbell diagram will achieve resonance or not. Both frequency and shape of the forcing are important in determining whether a blade will encounter aeromechanic difficulty. A goal of this research is to delineate (design and operating) parameters that set impeller blade aerodynamic and structural response; this is then to be followed by defining what constitutes an adequate characterization of impeller blade system response so that it can be used to develop guidelines for avoidance of aeromechanic difficulties in centrifugal compressor stages.
Centrifugal compressor stage for computational research
Modal displacement of splitter vane
Unsteady pressure field around splitter vane expressed as a phase and magnitude:
Return Channel Design Optimization Using Adjoint Method for Multistage Centrifugal Compressors
Wei Guo
Advisor: Prof. QiQi Wang
Multistage centrifugal compressors are widely used across industries and demand is growing in radial and axial compactness to reduce cost and increase reliability. Optimized design is therefore needed to reduce the loss caused by the innate 180 degree change in flow direction in the return bend. Previous quantitative evaluation by Ms. Anne-Raphaelle Aubry has shown opportunities for performance improvements. However, the parametric study based on empirically selected geometric parameters meets difficulty in establishing a systematic and automated algorithm for design optimization. Adjoint method has been successfully applied to design optimization in external flows for about two decades. This approach is now being introduced to return bend design optimization in internal flows. The basic idea is evaluating the sensitivity of a desired objective function regarding to geometry perturbation. The specific form of the objective function is carried by the solution of adjoint equations, and the geometry perturbation is converted into the residuals of the primal flow field equations. Then a continuous sensitivity distribution can be obtained by integrating the product of adjoint solution and residuals. Using the idea of control theory, the evolution of geometry perturbation will be automated based on the feedback of sensitivity distribution. Meanwhile, since the sensitivity distribution provides continuous information on design modification, the optimization process would be ideally more comprehensive than optimizing with a limited number of discrete parameters.
Residual field obtained by mapping flow field solution on the unperturbed mesh to a perturbed mesh. From left to right: non-dimensionalized residual field of pressure, axial velocity, radial velocity and circumferential velocity.
Two Engine Integrated Propulsion System
Alex Espitia, and Dr. Alejandra Uranga
Advisor: Prof. Greitzer
In 2008, NASA awarded four research contracts to define advanced concepts and enabling technologies for subsonic aircraft, in the 2035 timeframe, that could address the challenges posed by the increased demand while significantly reducing fuel consumption. The research was part of the NASA N+3 program, where N+3 refers to aircraft three generations beyond those currently flying. MIT, in collaboration with industrial partners Aurora Flight Sciences and Pratt & Whitney, is developing the D8 series aircraft to meet future demands. The D8 aircraft fields a “double bubble” fuselage and has two engines flush-mounted at the top-rear of the fuselage. This new engine configuration for commercial aircraft is being further evaluated. A parametric study of various separation distances between the two engines using high fidelity simulations is being performed. Currently, a simplified study based on two-dimensional simulations has shown that a merged double engine-model (no separation) yields the highest thrust performance due to a reduction in total drag on the engine’s nacelles and the elimination of flow separation that occurred for models with engines that are close together. Three-dimensional simulations are now being performed in order to determine the engine design that will be used to power the D8 aircraft. In addition to engine separation, the shape of the fuselage aft of the engines inlet is to be determined so as to provide the flow diffusion necessary for optimal performance.
Inlet Swirl Distortion Effects on the Generation and Propagation of Fan Rotor Shock Noise
Dr. Jeff Defoe
Advisor: Prof. Spakovszky
Reducing emissions, fuel burn, and noise are the main drivers for innovative aircraft design. Embedded propulsion systems, such as those used in hybrid wing-body aircraft, can offer fuel burn and noise reduction benefits but one of the major challenges in high-speed fan stages used in these embedded propulsion systems is inlet distortion noise, in particular fan rotor shock noise.
A new approach was developed to solve this problem based on a body force description of the fan blade row. The body force field not only represents the overall rotor characteristics, capable of capturing the distortion transfer effects, but for the first time is also used as the fan noise source. An unsteady perturbation force field generates the rotor-locked shock and expansion fan system which gives rise to rotor shock noise. The approach has been validated on NASA's Source Diagnostic Test fan and inlet. The generated shock Mach numbers are in good agreement with experimental results, with the peak values predicted within 6%. An assessment of the far-field acoustics against experimental data showed agreement of 8 dB on average for the blade-passing tone.
This approach is employed in a parametric study to assess the effects of inlet geometry parameters (offset-to-diameter ratio and downstream-to-upstream area ratio) on flow distortion and rotor shock noise. Mechanisms related to the vortical inlet structures were found to govern changes in the rotor shock noise generation and propagation. The vortex whose circulation is in the opposite direction to the fan rotation (counter-swirling vortex) increases incidence angles on the fan blades near the tip, enhancing noise generation. The vortex with circulation in the direction of fan rotation (co-swirling vortex) creates a region of subsonic relative flow near the blade tip radius which decreases the sound power propagated to the far-field.
The parametric study revealed that the overall sound power level at the fan leading edge is set by the ingested streamwise circulation, and that for inlet designs in which the streamwise vortices are displaced away from the duct wall, the sound power at the upstream inlet plane increased by as much as 9 dB. By comparing the far-field noise results obtained to those for a conventional inlet, it is deduced that the changes in rotor shock noise are predominantly due to the ingestion of streamwise vorticity. The far-field spectra are also altered by inlet distortion. Tones at up to 3 times the blade-passing frequency are amplified and tones above one-half of the blade-passing frequency are attenuated and appear to be cut-off.
Future work might focus on broadening the applicability of the present approach to other types of fan noise, such as rotor-stator interaction tones and/or fan broadband noise. The challenge lies in accurately modeling the viscous effects such as blade wakes with body forces. Additional studies could also be undertaken using the current approach. These might involve varying other inlet duct parameters and broadening the parameter space under consideration. The details of the non-uniform flow entering the inlet duct could also be varied. These types of studies could provide additional insight into the mechanisms already discovered.
Propulsion System Integration and Noise Assessment of a Hybrid Wing-Body Aircraft
Dorian Colas, Dr. Elena De la Rosa Blanco, (former students: Leo Ng, Phil Weed)
Advisor: Prof. Spakovszky
Reducing the environmental impact of air travel is a major impetus to current research in aeronautics. A potential configuration that could enable step changes in fuel consumption, noise and emissions is a hybrid wing-body aircraft where a lifting fuselage is blended with the wings. Building on previous work from the Silent Aircraft Initiative, this project aims to develop a set of advanced predictive methods that will enable the design of a hybrid wing-body aircraft to meet NASA’s N+2 goals: (i) 25% less fuel burn, (ii) 80% less emissions, and (iii) 52 dB less noise compared to current aircrafts in service. MIT, in collaboration with Boeing, NASA, and UC Irvine, is defining the aircraft configuration and propulsion system to meet such goals.
One approach reducing propulsion system noise is to mount the engines above the airframe, utilizing the large planform area to shield the noise generated by the turbomachinery. A fast algorithm of medium-fidelity was developed based on Kirchoff’s diffraction theory to compute the shielding effect of the airframe using directivity compact sources. The method includes flight effects and is applicable to any kind of aircraft configuration.
An alternative configuration uses engines embedded in the airframe where the airframe boundary layer is ingested by a distributed propulsion system. In such configurations thrust and drag cannot be simply separated and instead the overall aircraft performance is assessed using a previously established power balance analysis. The design of an S-shaped inlet and distortion tolerant fan stage is also being pursued.
The approach is based on high-fidelity simulations of the coupled airframe, inlet and fan system using a body force based representation of the fan stage. Various design concepts will be explored with the goal to improve power savings and to mitigate inlet flow distortion and fan performance penalties.
Fan-Inlet Integration for Low FPR Propulsors
Andreas Peters
Advisor: Prof. Spakovszky
Aircraft engine design trends tend towards higher bypass ratio, lower pressure ratio fan designs for improved fuel burn, reduced emissions and noise. Low-pressure ratio fans offer increased propulsive efficiency and, besides enabling thermodynamic cycle changes for improved fuel efficiency, significant acoustic benefits can be achieved. Fan diameters increase as fan pressure ratios (FPR) are reduced, and the design of innovative nacelle concepts becomes critical to limit the impact of larger diameter fans on nacelle weight and drag. The proposed work addresses the uncharted design space of low FPR propulsors and their nacelles and will provide new inlet and nacelle design guidelines to minimize nacelle drag and maximize fuel burn benefits in low FPR propulsors without jeopardizing operability.
Since low-pressure ratio fans and their nacelles are more closely coupled than current turbofan engines, inlet-fan interaction and inlet flow distortion at the fan face are increased. Consequently, a coupled fan-nacelle approach capable of capturing inlet-fan and fanexhaust interactions is required to evaluate the performance of low FPR propulsors. In this work, a fast and reliable body force based approach was developed to assess the performance of innovative nacelle concepts. In this approach, rotor and stator blade rows are replaced by body force fields determined from steady single-passage RANS simulations. Steady full-annulus simulations are carried out to determine the performance of fan stage and nacelle in the presence of non-uniform inflow and back pressure distortion due to pylon and bifurcation. As illustrated in the figure below, the developed method was demonstrated to capture the coupling of internal and external flows and the distortion transfer through the fan stage and reduces the computational cost by up to two orders of magnitude compared to full 3D unsteady RANS simulations.
The next step is to use the body force based approach to conduct a parametric study of candidate inlet and nacelle geometries with the objective to improve the propulsor performance by reducing nacelle drag and weight.
Aerodynamics and Heat Transfer in Gas Turbine Tip Shroud Cavity Flow
Timothy Palmer, Department of Mechanical Engineering
Advisor: Dr. Tan
Past research effort on gas turbine technology has focused on reducing loss generation and cooling flow requirements in the main flow path. To further improve turbine efficiency and durability, the secondary air flow system, critical to operation of these engines, needs to be investigated and its associated loss mechanisms reduced. This project aims to determine the specific drivers that set the loss generating mechanisms and heat transfer in the secondary flow system. Understanding of these drivers would allow the formulation of strategies for turbine performance and durability enhancement to benefit the next generation of large industrial gas turbines for power generation. The project seeks to address, on a quantitative basis, the following: 1) the effects of the cavity on the aerodynamics of and characteristic turbine operating parameters in the blade-tip region; 2) response of the blade tip shroud cavity flow to injected cooling and seal leakage flows and turbine tip configurations; 3) the role of unsteadiness on the tip shroud cavity flow and the associated loss generation; and 4) the impact of (1) and (3) on overall multistage axial turbine performance including the downstream diffuser. Once the aerodynamic loss generation mechanisms have been isolated, heat transfer will be incorporated to determine its effect on the turbine tip shroud cavity flow.
Turbine cutaway with inset showing an example of the turbine tip shroud cavity geometry.
Generic 4th Stage low-pressure turbine for CFD
Secondary Air Interactions with Main Flow in Axial Turbines
Research Assistant TBA
Advisor: Dr. Tan
In the past decade, industrial gas turbines have by far become the most popular type of plant for power generation due to their compactness, low emissions and potential for power-heat cogeneration. In the effort to increase energy conversion efficiency, engineers have raised turbine inlet temperatures to well above the metal melting point. Turbine blades are generally protected by expensive thermal barrier coatings and various forms of internal and film cooling. However, in order to prevent hot gasses from being ingested into the unprotected cavities between rotating and stationary components, cool air bled from the compressor is used to purge the gaps at the endwalls. MIT, in collaboration with Siemens Energy and Siemens Corporate Research, is developing a computational approach to identify and understand loss generating flow processes of purge air interacting with mainstream flow in axial turbines.
Contours of change in volumetric entropy generation rate relative to a baseline case with no purge flow bring out the regions in a rotor blade passage that have modified losses as a consequence of purge flow injection from the hub gap upstream of the rotor. We have identified a number of effects that result in these changes: mixing out of the velocity difference between purge and mainstream flows, the generation of radial velocity gradients as a consequence of purge flow interacting with the passage vortex structures, and increased wetted and tip clearance flow losses due to a change of reaction. There is also a positive effect of reduced tip clearance losses when purge flow is injected from the shroud. These effects have been rigorously quantified, and their drivers have been pinpointed. This new knowledge provides clear guidelines for better turbine designs.
Compressor Aerodynamics in Large Industrial Gas Turbines for Power Generation
Sitanun Sakulkaew
Advisor: Dr. Tan
The overall goal of the research is to improve the efficiency of large industrial gas turbines through improvement of compressor performance. Specifically, the research focuses on two important aspects of compressor science and technology. The first aspect addresses loss and flow blockage generation in high-speed multistage axial compressors to establish a design philosophy for high efficiency and for broadening the island of peak efficiency. The second aspect seeks to quantify the variation of efficiency as blade (rotor tip and stator hub) clearance approaches zero and its implication on peak efficiency in a multistage environment. The overall framework of the approach consists of using computational analyses to first establish the traceability of flow features as they impact compressor performance changes; this is then to be followed by experimental assessments.
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A representative large industrial gas turbine for power generation. |
Turbine Tip Clearance Loss Mechanisms
Steven Mazur
Advisors: Prof. Greitzer, Dr. Tan
One of the large loss sources in a turbine stage arises from the flow through the gap between the rotor tip and the shroud. The pressure difference across the tip drives the flow through the gap, and this leakage flow subsequently rolls up into a vortex on the suction side of the blade and convects downstream. As the vortex mixes out and decays, entropy is generated. Previous work by Arthur Huang has identified the pressure gradient external to the vortex as a major mechanism for determining the loss generated by the tip vortex. The current project aims to consider new influencing factors on the vortex evolution and associated loss. 3D computational simulations are being used to study the influence of several classes of effects. Downstream influence of the transition duct at the exit of the high pressure turbine can have an impact on the external conditions the tip leakage vortex is subjected to. A parametric study is underway to illustrate how the governing design parameters influence the tip clearance loss. Future project goals are to discover how upstream and unsteady effects change the loss created by turbine tip gap flows.
Flow and Heat Transfer in Modern Turbine Rim Seal Cavity
Peter Catalfamo, and Rachel Berg
Advisor: Dr. Tan
Sponsor: General Electric
Ingestion of hot gas from the flowpath into the gaps between the rotor and stator can cause turbine components to overheat and lead to deterioration in component life. To prevent this, modern gas turbines, both industrial and aerospace, use compressor bleed air to provide positive outflow through the rim seal (known as “purge” flow). This purge flow can be a substantial fraction of the total flow bled off of the compressor and as such it represents a substantial performance penalty. Past efforts have focused primarily on generating correlative orifice models using experimental data. These results are limited in their applicability by the geometry and conditions tested. In this research MIT, in collaboration with GE Energy and GE Aviation, seeks to investigate the fundamental flow physics in the turbine rim cavity region. Of particular interest is the response of the wheelspace and rim cavity to external stimuli set up by the main annulus flow such as flow unsteadiness due to rotor stator interactions. Rig data being collected by GE will be used to assess the analysis and to guide the investigation. Understanding these mechanisms is fundamental to optimizing seal design and minimizing the purge flow requirements, thus minimizing the associated performance penalty.
