Recent/Current Research at the Gas Turbine Laboratory
A Unified Approach for Vaned Diffuser Design in Advanced Centrifugal Compressors
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
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.
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.
Representation of bladed centrifugal compressor with axisymmetric body forces.
Improved Performance Return Channel Design for Multistage Centrifugal Compressors
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.
Investigation of Real Gas Effects in Supercritical CO2 Compression Systems
David Yang, 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.
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.
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.
A New Modeling Approach for Rotating Cavitation Instabilities in Rocket Engine Turbopumps
William Sorensen, and Dr. Claudio Lettieri
Advisor: Prof. Spakovszky
Axial inducers are often used in high performance turbopumps for rocket engines in order to reduce system weight and cost. The inducer allows operation at high rotational speeds and low inlet pressures, which in return can lead to cavitation. The dynamic cavitation behavior is often unsteady and periodic, leading to several distinct cavitation instabilities.
Of particular concern among the various instabilities is rotating cavitation, in which cavities propagate circumferentially at frequencies up to 5 times the rotor frequency. In some cases, the propagation frequency can approach the rotor shaft natural frequency, which can cause severe vibrations that present a significant risk to the turbopump. Unfortunately, the physical mechanisms of rotating cavitation are not yet well understood, and as such no general design guidelines exist to mitigate its onset. Existing low-fidelity modeling efforts, which are largely two dimensional, have no predictive capability and fail to capture the three-dimensional nature of the phenomenon.
The goal of this project is to establish a new capability to guide the design of the inducer and related casing treatment to suppress rotating cavitation instabilities. The approach exploits a recently developed body-force based modeling approach that has been used successfully in aerodynamic instability modeling of jet engine axial and centrifugal compressors and in acoustic descriptions of non-uniform flow in transonic fans. The feasibility of employing this method to determine the frequencies and mode shapes of cavitation instabilities will be examined. Extensive data obtained in testing NASA’s Space Shuttle Main Engine Low-Pressure Oxidizer Pump (SSME LPOP) will be used for validation of the approach. The technical objectives are to (1) identify the required inducer blade loading distributions and endwall flow forcing requirements to mitigate rotating cavitation, (2) determine the casing treatment type and geometry that can render the required forcing, and (3) assess the effect of inlet distortion on cavitation dynamics and investigate the linkages between rotordynamics and cavitation.
Modeling Instabilities in High-Pressure Pumping Systems
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.
Ported Shroud Operation in Turbochargers
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.
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
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.
Return Channel Design Optimization Using Adjoint Method for Multistage Centrifugal Compressors
Advisor: Prof. QiQi Wang
Multi-stage centrifugal compressors are widely used across industries and the demand is growing in the 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. Conventional gradient-based optimization becomes computationally expensive when computing gradients in such a high dimensional problem, requiring a number of simulation runs equal to the number of design variables.
In contrast, adjoint method uses linear approximation to construct adjoint equations. By solving only once the flow equations and the adjoint equations, the sensitivity is obtained for an objective function, i.e. performance metric, with regard to any number of design variables.
In practice, a generalized free-form deformation algorithm has been developed for the geometry perturbation which is free from the traditional control point configuration constraints. The perturbation converts into residuals of the primal flow equations. Then the sensitivity is computed by integrating the product of adjoint equation solution and the residuals over the computational domain. The adjoint-based sensitivity is verified against that obtained using finite-difference method using a low Reynolds number, laminar flow case. The evolution of return bend geometry deformation is then automated based on the feedback of sensitivity, using a Quasi-Newton method, until an optimal design is reached within given constraints.
The adjoint-based optimization would ideally explore the design space more comprehensively, and cost-effectively. Future work could include implementing turbulence model and adaptive meshing.
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.
Propulsor Design for Exploitation of Boundary Layer Ingestion
Advisors: Prof. Greitzer, Dr. Tan
Boundary Layer Ingestion (BLI) -- passing a portion of an airframe's wake through the engine -- has been suggested as a means of reducing civil aircraft energy consumption. An aerodynamic performance benefit results from the reduction of viscous dissipation in the airframe wake and in the propulsor jet downstream of the aircraft. A number of challenges exist in the design and performance assessment of BLI aircraft configurations. One is that definition of the propulsion system requirements becomes more difficult because the concepts of thrust and drag, conventionally associated with the engines and airframe, respectively, become ambiguous with a tightly integrated propulsion system. Further, the engine performance itself may be adversely affected by the presence of inlet distortions arising from the ingested airframe boundary layer. The goal of this computational and experimental project is to establish a strategy for the design of propulsors for operation in BLI aircraft configurations.
Propulsion system requirements are determined using the power balance method. Thus, rather than considering the forces on the aircraft, the performance of the airframe and engine are characterized in terms of sources and sinks of mechanical energy. Quantities analagous to thrust, drag, and propulsive efficiency are defined, and it is seen that the system-level benefit of BLI can be explained as the combined effect of a decrease in effective airframe drag and an increase in propulsive efficiency. Viewing the system performance in this way also provides insight into appropriate propulsion system scaling for powered wind tunnel models and for comparison between integrated and conventional podded propulsion system configurations.
The impact of BLI inlet distortion on fan stage turbomachinery performance is assessed using a non-axisymmetric throughflow model. A parametric study to determine the sensitivity of the distortion response and fan efficiency to various design parameters is being carried out to illuminate design features for propulsors more tolerant of BLI type inlet distortions. The performance of selected designs will be assessed both computationally and experimentally, as part of wind tunnel tests of a scale model BLI aircraft configuration to be conducted in the MIT Wright Brothers WindTunnel and in the NASA Langley Research Center 14'x22' Wind Tunnel.
Noise Assessment of a Hybrid Wing-Body Aircraft
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.
The method is being further developed to improve the attenuation predictions by taking into account the geometry of the airframe and propagating the diffracted noise accordingly.
Fan-Inlet Integration for Low FPR Propulsors
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.
Centrifugal Compressor Science and Technology: Multi-parameter Control for Compressor Performance Optimization
Advisor: Dr. Tan
Centrifugal compressor systems deployed in the industry are required to operate 24/7 with minimum down time. As such their operation at high thermodynamic efficiency across a wide operating range is of paramount importance to the customers.
Extending the compressor limits to meet the needs of a specific engineering mission is one of the most important aspects of compressor engineering. Strategies to quantitatively assess the potential for extending compressor performance and operating range must be developed. This can be done by determining the drivers that set the requirements for the broadest operable range with high efficiency retention.
The goal of this research is to first identify what are the parameters of high leverage that affect centrifugal compressor performance followed by establishing potential means of achieving near matching of centrifugal compressor components at all desirable operating points required for its mission. Some of the key parameters that are thought to have a high leverage on compressor performance characteristics are compressor speed, guide vane setting and diffuser vane angle setting; however there could be others that are to be identified during the course of the research. In light of this, formulating an effective control strategy (passive, active or a combination of both) for achieving desirable compressor performance requirements for its specified mission (as alluded to above) at an optimal cost would also be another goal for this research.
The general approach consists of leveraging on the technical capability and thinking at the MIT Gas Turbine Laboratory and of working collaboratively with Siemens Technologists. It will also be necessary to define physical experiments at Siemens Facility in Germany or at MIT Gas Turbine Laboratory for assessing ideas and concepts formulated during the course of the research.
Secondary Air Interactions with Main Flow in Axial Turbines
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
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.
A representative large industrial gas turbine for power generation.
Turbine Tip Clearance Loss Mechanisms
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.