Performance Improvement of a Turbocharger Twin Scroll Type Turbine Stage
The need for internal combustion engines with high efficiency and less CO2 exhaust rises as the government regulation over the environment tightened. One traditional solution to this problem is to implement a turbocharger, which compresses the inlet flow into the internal combustion engine by means of redundant energy from the engine exhaust gas. These turbo charged engines tend to have high fuel efficiency as well as low CO2 production rate while the size of the engine remains compact. Generally, the quasi-steady state assumption has been adopted for estimating the performance of turbine of the turbocharger. While this is reasonable approximation for single scroll turbocharger as long as the volute storage effect has been accounted. However, such may not be the case for twin-scroll turbocharger where the unsteadiness effect can become more prevailing.
An objective here is on assessing the role of flow unsteadiness induced by the pulsating flow environment that the twin-scroll turbine is subjected to. This is then to be followed by determining the attributes of turbocharger twin scroll turbine stage that would substantially improve efficient extraction of flow energy by the turbine. Thus another objective is to identify the efficiency limiting flow processes when the turbine is subjected to the pulsating flow environment representative of turbocharger twin scroll turbine operation. Accomplishment of this objective would provide a physical basis to formulate means, passive as well as active, to achieve a step improvement in the operating efficiency of the turbine stage. While the focus here is on turbocharger-engine, the outcomes and ideas conceived might be of utility to assessing turbine operations in pulsating flow environment induced by pulse detonation combustors.
Compressor Aerodynamics in Large Industrial Gas Turbines for Power Generation
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.
Aerodynamics and Heat Transfer in Gas Turbine Tip Shroud Cavity Flow
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.
Forced Response of a Centrifugal Compressor Stage Due to the Impeller-Diffuser Interaction
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.
Fan-Inlet Integration for Low FPR Propulsors
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.
A Methodology for Centrifugal Compressor Stability Prediction
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.
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.
Improved Performance Return Channel Design for Multistage Centrifugal Compressors
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.
Return Channel Design Optimization Using Adjoint Method for Multistage Centrifugal Compressors
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
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.
Aeromechanic Response in High Performance Centrifugal Compressor Stage
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.
Inlet Swirl Distortion Effects on the Generation and Propagation of Fan Rotor Shock Noise
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.
Loss Modeling of Turbine Tip Leakage Flows
Formation of tip leakage vortex (Mischo, Behr, Abhari, 2008). DS1 is the dividing streamline between incidence-driven flow and pressure-driven flow. DS2 is the dividing streamline between flow ending up in the passage vortex and flow ending up in the leakage vortex. A major source of inefficiency in a turbine results from pressure-driven flow leaking across the rotor tip from the pressure side to the suction side. The flow emerges from the tip gap in a jet, which rolls up into a vortex near the shroud/suction side corner of the blade passage. Entropy is generated as the leakage flow mixes with the mainstream flow. In addition to creating aerodynamic losses, tip leakage flows also transfer heat to the rotor tip so that an uncooled rotor tip may be damaged. Because of this, turbine designers introduce cooling flows, which bring with them their own mixing losses, as well as lower total work due to the cooling flow bypassing the combustor. This project aims to model the losses associated with turbine tip leakage in order to better design the rotor tip. Schematic of tip leakage flow (Krishnababu et al., 2009) Currently used models for aerodynamic tip leakage losses are correlations based on rotor tip lift coefficients, blade geometries, or simply an efficiency penalty proportional to the gap height. We have modeled the tip gap region as a series of 2D planes in the leakage streamline and radial directions. In each of these 2D planes, the flow is viewed as a 1D sudden expansion over a vena contracta. Mass and momentum control volume equations are solved to determine leakage mass flows and velocities, and hence entropy due to mixing, which is the efficiency loss. The next steps in this project are to conduct CFD analysis of tip leakage flow to determine whether the assumptions used in the modeling are reasonable and to develop and test models for losses from required cooling associated with the tip leakage.
The "Swirl Tube" - an Aircraft Drag Management Device to Reduce Noise and Fuel Burn
Aircraft on approach in high-drag and high-lift configuration create unsteady flow structures which inherently generate noise. For devices such as flaps, spoilers and the undercarriage there is a strong correlation between overall noise and drag such that, in the quest for quieter aircraft, one challenge is to generate drag at low noise levels. The invention is a novel aircraft drag management concept to reduce aircraft noise during approach and to improve fuel burn in cruise. The idea is based on a swirling exhaust flow emanating, for example, from a jet engine nacelle (see figure) or a wing-tip mounted duct. A novel application is to exploit the low pressure in the vortex core of the swirling exhaust flow to generate drag. The idea is that in a steady streamwise vortex the centripetal acceleration of fluid particles is balanced by a radial pressure gradient. The very low pressure near the vortex core at the exit of the duct generates pressure drag. This streamwise vortex is in essence steady, yielding low noise levels and a quiet acoustic signature. To see a Quicktime movie of the swirl tube in action, click here (this is a large file so please be patient while it loads).
Effects of Impeller-Diffuser Interaction on Aerodynamic Performance of Centrifugal Compressors
This work is aimed at a key problem of modern centrifugal compressor stages: the impact of unsteady interaction between the rotating impeller blades and stationary diffuser vanes on performance and aeromechanics. Three important technical issues are of engineering interest:
- The effect of these unsteady interactions on the time-averaged performance of a modern centrifugal compressor stage;
- Initiation and development of aerodynamic instabilities in a centrifugal compressor; and
- Use of (1) to propose design guidelines and potential innovative designs for achieving incremental improvements in performance.
The current focus is on technical issues (1) and (3), however future project phases will use the initial results to infer their potential impact on issue (2). In addition to addressing the technical aerodynamic issues, assessment and validation of new analytical approaches and computational tools for use in centrifugal compressor stage design and analysis will be carried out during the course of the research program. This research used MSU TURBO, developed by Mississippi State University.
Blade Row Interactions in High-Speed Axial Compressors
The purpose of this project is to understand and quantitatively assess the role of blade row interactions in the performance of highly-loaded, high Mach number (HLHM) axial compressors. These interactions include that of the rotor shock on the upstream blade row, as well as the influence of blade wakes on downstream blade rows. Using this knowledge, guidelines for the design of efficient HLHM compressors can be recommended, ultimately resulting in smaller, lighter, and less complex compressor cores.
Inlet-Engine Integration for High Performance Aircraft
Current practice in aircraft inlet computations does not appear to account for, in a rigorous manner, the impact of flow redistribution due to the presence of engine compressors. This can be of import in the determination of inlet-engine dynamic behavior, e.g., ASTOVL vehicles at takeoff and landing. The research activity represents an effort to develop a consistent computational methodology for inlet-engine integration aerodynamics on high performance aircraft.
Performance and Flow Phenomena in Vaned Diffusers
Computational studies have been initiated to examine the impact of unsteady impeller-diffuser interactions on performance, specifically: (i) time-averaged effects of unsteady impeller-diffuser interactions and (ii) initiation and development of aerodynamic instabilities in a centrifugal compressor.
Analysis of Aerodynamically Mistuned Bladed Disks
In typical analyses of bladed disks, the problem is assumed to be tuned, that is all blades are assumed to have identical geometries, mass and stiffness characteristics. In reality, both the manufacturing process and engine wear create a situation where the blades differ slightly from one another. These blade-to-blade variations are known as mistuning and can significantly impact the operation of bladed disks. In particular, mistuning causes the forced response amplitudes of individual blades to be much larger than that predicted by a tuned analysis, which has serious implications for high cycle fatigue. The purpose of this research is to develop high-fidelity, low-order aerodynamic models suitable for use in a mistuning context. While significant progress has been made is the area of structural mistuning, the effects of blade geometric variations is not well understood. This research uses systematic model order reduction techniques (such as the proper orthogonal decomposition) to develop models suitable for the aerodynamic mistuning problem.
Modeling of Turbomachinery Aerodynamic Instabilities
The central focus of this research is the initiation and development of compressor instabilities in situations where the three-dimensionally of the flow is significant. Models have been developed to assess effects of: radial blade loading distribution on stability, inlet distortion on stall inception in three dimensional flows, and three-dimensional phenomena due to mismatching in multistage compressors. Procedures are also being developed to define the nonlinear evolution of compressor flow disturbances in situations in which compressibility has a strong role. Work has also been initiated to examine the effects of rotor stator interaction on blade row stalling behavior. The goal is to establish, in a rigorous manner, causal links between design characteristics and the stability of compressor flows .
Impeller-Diffuser Interaction On Aerodynamic Performance of Centrifugal Compressors
The work is aimed at a key problem on the aerodynamics of modern centrifugal compressor stages, the impact of unsteady impeller-diffuser interactions on performance and design. Three important technical issues are of engineering interest: (1) time-averaged effects of unsteady impeller-diffuser interactions on the performance of a modern centrifugal compressor stage; (2) initiation and development of aerodynamic instabilities in a centrifugal compressor; and (3) use of (1) to suggest design guidelines and potential innovative design for achieving step changes in performance. The current focus will be on technical issue (1) and (3); however the results will be used to infer their potential impact on issue (2). In addition to addressing the technical issues, assessment and validation of analytical/computational tool/tools for use in centrifugal compressor stage design and analysis will be carried out during the course of the research program.
Characterization of Aeromechanics Response and Instability in
High Performance Centrifugal Compressor Stage /Rocket Centrifugal Pump
The work here is address the role of impeller diffuser interactions on the forced response of impeller blades and as such is somewhat complementary to and synergistic with the effort on impeller-diffuser interaction on aero performance of centrifugal compressors.
Impact of End-Wall Flows and Wakes in Multistage Axial Compressor Performance
The effort proposes to examine the unsteady response of endwall flow, specifically the rotor tip leakage flow, to the flow conditions associated with the upstream stator (wakes and flow non-uniformity as set by design of upstream stator) and downstream stator (mostly unsteadiness due to potential interaction), and the potential time-average impact on multistage compressor performance.
Hydrodynamics of Centrifugal Pumps for Space Propulsion
While much useful work has been done on turbopumps for earth-to-space rocket propulsion, it has been mostly of a developmental nature. To raise the level of understanding and the design capability for turbopumps to the desired level, we see a need to address several specific areas of centrifugal pump fluid dynamics. One of these is the effect of tip clearance flows on performance and stability; others are the unsteady effects associated with fluid-structure interaction and with impeller-diffuser interaction. It is the first of these that we target here.
Aeromechanic Response of High Speed Compressor Stage to Inlet Distortion
The focus is on using data (both experimental from CRF at AFRL and computations) to delineate the response of high speed compressor stage to inlet distortion in terms of the flow processes/drivers responsible for the observed unsteady blade loads and response. Also effort will be undertaken to assess the observations/understanding from CRF test rigs together with the accompanying numerical simulations against those from engine in flight test situations; this is needed to answer the question of what constitute an adequate ground/model rig test for representing the forced response on engine compressor encountering inlet distorted flows in flight. The intellectual challenge here is how to organize the results so that physical insight can be extracted to aid future design of fan/compressors with fewer aeromechanics difficulties, including HCF (High Cycle Fatigue) failure.
Characterization of Wakes/Tip Vortex/Secondary Vortex Induced Blade Excitations
in Multi-Stage Environment
The overall goal here is to understand and predict two- and three-dimensional sources of unsteady loads on fans, compressors and turbine blades under representative operating environments and to show their impact on forced vibration and component flutter. The key objectives are to: (1) define key links between unsteady blade load and its sources which include rotor tip vortices, discrete secondary flow vortices, wakes, blade surface unsteady boundary layer separation, and blade motion over the operating range of interest; (2) quantify the effects of adjacent blade rows, or adjacent stages (i.e. multi-stage influence); and (3) develop a predictive physical model that includes aero-structural coupling effects and hence can provide data to predict dynamic stress and strain. In the larger picture, it is important to quantify parametric trends which fix aeromechanics response and instability (flutter) onset of turbomachinery blades, and to assess the feasibility of tailoring the aerodynamic and structural design of blades
Computational Model for Multistage Compressor Aerodynamics:
Performance Map Generation and Stability Characterization
A framework for an effective computational methodology was developed for multi-stage compressor map generation. The methodology consists of using a few isolated-blade row Navier-Stokes solutions for each blade row to construct a body force database. The purpose of the body force database is to replace each blade row in a multi-stage compressor by a body force distribution to produce same pressure rise and flow turning. To do this, each body force database is generated in such a way that it can respond to the changes in local flow conditions. Once the database is generated, no further Navier-Stokes computations are necessary. The process is repeated for every blade row in the multi-stage compressor. The body forces are then embedded as source terms in an Euler solver. The method is developed to have the capability to compute the performance in a flow that has radial as well as circumferential non-uniformity with a length scale larger than a blade pitch; thus it can potentially be used to characterize the stability of a compressor under design. It is these two latter features as well as the procedure to obtain the body force representation that distinguish the present methodology from the streamline curvature method.
Spectral Computations of Three-Dimensional Flow for Axial Turbomachinery
A three-dimensional spectral code has been developed for computation of 3-D flow through an axial turbomachinery blade row. It is currently being used to examine the response of the blade passage flow field to the incoming moving wakes and streamwise vortices.
Computational Studies of Flutter and Forced Response in Compressors/Fans
Research is underway to explore the capability and the role of computational fluid dynamic tools for: (i) addressing blade flutter phenomena and associated controlling mechanisms as well as the parametric trends; and (ii) assessing the effect of density nonuniformities on the unsteady response of turbomachinery blading.
Experimental Investigations of Axial Turbine Fluid Physics
Transient (blowdown) testing of advanced performance turbines is being conducted at realistic Mach and Reynolds numbers, ratios of turbine inlet temperature to metal and coolant temperature, and inlet temperature profiles. A current topic under investigation is the effect film cooling on aerodynamic performance and heat transfer.
Aspirated compressors employ suction on the blading and endwalls to realize greatly increased pressure rise and efficiency from axial turbomachinery. Single stage compressors scheduled for test include a 700 fps design with a pressure rise of 1.5:1 and a 1500 fps design with a pressure rise of 3.5:1. This work is in partnership with NASA GRC, PW, and Honeywell.
Evaporative Cooling for Gas Turbines
A new approach to the problem of using two phase flow to internally cool gas turbine hot parts is under investigation. Experiments on a rotating turbine blade simulator have demonstrated the cooling rates needed for the engine environment. Work is proceeding toward demonstrating this concept in the engine environment.
Modeling and Integration of Active Internal Flow Control in Aero-Engine Compressors
This project constitutes a research program on the key technical issues relating to the modeling and integration of flow control devices in aero-engine compressors. The project focuses on the physical understanding of the governing flow phenomena, their implementation and experimental application with the goal to further enhance the performance and operability of axial-flow compressor stages beyond the current loading and stability limits. The theme of the technical approach is a unique analytical, reduced-order dynamic system modeling methodology, which incorporates a system rather than a component view of this multi-disciplinary problem. This work is in collaboration with the NASA Glenn Research Center where a proof-of-concept experiment with synthetic jets will be conducted in the Low Speed Axial Compressor test facility.
Quiet Supersonic Platform (QSP) Propulsion
We are examining the propulsion requirements for very quiet long range, MACH 2+ cruise vehicles. As part of this effort we are designing a 2 stage aspirated compressor with a pressure ratio of 10:1.
Carbon Nanotube Bearing
Carbon nanotube rotor [not to scale] Rotating Micro Electro-Mechanical Systems (MEMS) require rotary bearings, but current MEMS bearing technologies have drawbacks. Silicon rubbing on silicon wears out quickly. Gas bearings require a gas source, and are relatively low stiffness. A promising alternative, proposed and being pursued by the Charles Stark Draper Laboratory in collaboration with the GTL, is to use Carbon Nanotubes (CNTs). Multi-walled CNTs have a concentric-tube structure that lends itself to bearings. Each tube is strong, but there is little or no bonding between tubes, allowing them to slide relative to each other. However, the friction characteristics of these bearings are not precisely quantified. This project’s goal is to construct a simple CNT bearing rotary device, demonstrating MEMS and CNT compatible fabrication techniques, and allowing some data on the friction characteristics to be gathered. Applications of such a bearing technology could include microscale turbomachinery, as well as gyroscopes, pumps, and other rotating devices.
Small-scale Gas Turbine Engines
Small scale gas turbine engines can provide much higher power densities than conventional batteries, and show promise as a portable and enduring power source. A 1kW class small scale gas turbine generator is being developed for this purpose. Experimentally identified key issues for making the engine work are thermal management and rotordynamic stability. The heat generated by the engine through windage losses in the bearings and in the generator need to be removed and the rotor has to be effectively shielded from high temperature sources to ensure the mechanical integrity of the generator. The challenge is set by the small scale architecture. The goal is to establish an appropriate thermal management scheme. The technical approach is based on a parametric thermal resistance network that is calibrated with experimental data. The high rotational speeds required to reach acceptable aerodynamic efficiency call for gas lubricated bearings, which are known to have a low threshold of stability and are prone to large amplitude sub-synchronous vibration. Hence, another goal is to investigate these phenomena and develop design guidelines for high-speed gas bearings. Our approach is based on high level modelling of the bearings and to perform a sensitivity analysis to identify significant scaling effects. Due to the small scale of these engines the effect of the interactions between the different components on the efficiency is significant. Hence a future goal is to apply an integrated design and optimization approach to investigate alternative engine configuration.
Single Crystal Silicon as a Macro-World Structural Material: Design of Compact, Lightweight High Pressure Vessels
Exploration of the use of micromachined single crystal silicon as a macro-world structural material, understanding its advantages and limitations. Single crystal silicon is a material with theoretical strengths higher than steel and with a lower density than aluminum. This high strength, light weight nature of silicon make it an ideal structural material. Silicon has shown favorable performance as a structural material in micro-scale applications but the brittle nature of this material makes it difficult to reliably achieve high usable strengths on a macro-scale. Exploration on how microfabrication techniques affect silicon strength and how high strength macro-scale silicon structures can be made. This research had an engineering objective to find better ways of building compact, lightweight high pressure vessels for demanding applications such as spacecraft. The advantages of a silicon based pressure vessel is the potential to integrate the vessel regulator and control circuit on chip, the large inherent strength of silicon, the low density and thus lightweight characteristic of silicon and the ability to machine silicon into unique, compact geometries. This work was part of a joint effort of MIT and Ventions, LLC, to design a silicon-based microlaunch vehicle system for small satellite applications. This launch system was complete with chamber and nozzle, valve, regulator, power supply and tanks. The motivation of this endeavor was to expand the definition of low cost access to space with a cost per mission versus cost per payload pound.
A Fully-Integrated Permanent Magnet Turbine Generator
There is a need for compact, high-performance power sources that can outperform the energy density of modern batteries for use in portable electronics, autonomous sensors, robotics, and other applications. The current research aims to produce a fully-integrated, synchronous permanent magnet microturbogenerator capable of generating 10W DC output power using compressed air as its energy source. Presently, all the silicon die fabrication is complete, and the magnetic components are being integrated onto the die in preparation for power generation testing.
While the magnetic integration is in progress, efforts are underway to separately test and qualify the gas-lubricated bearings that will support the magnetic rotor to very high speeds. To make the tests relevant, they are conducted on silicon dies similar to the final generator dies, with the only differences being the lack of surface windings and a laminated magnetic stator. Figure 1 shows a bearing rig die enclosed in an acrylic package, as well as the metal tubulations and O-rings used to bring nitrogen into the die. The circular hole on the top of the package, together with additional holes on the backside, serves as an air vent.
Three sets of bearing rig tests are currently planned. All of them involve the same bearing rig die shown in Figure 1, but different rotors – light silicon rotor, heavy silicon rotor, and magnetic rotor – will be spun. A light rotor made purely of silicon and shown in Figure 2 will be used to assess the nominal imbalance, defined as the distance between the geometric and mass center of the rotor, introduced by the fabrication process. This rotor has approximately half the mass of the magnetic rotor, so a solder-filled rotor twice as heavy will be tested next to determine whether the bearings perform well with a massive rotor. After these two sets of experiments are complete, the magnetic rotor, which has permanent magnets and a soft magnetic back iron embedded, will be characterized. Because the silicon die can be easily opened along its eutectic interface, it is anticipated that the magnetic rotor can be removed from the die after testing and reused for the generator die.
Microengine Materials, Structures and Packaging
Includes tasks of materials characterization and constitutive modeling and the overall thermal and structural design of the microengine and associated devices. Packaging includes the design and fabrication of the interfaces of these micro-devices with the macro-world, including fuel supplies, air intakes and electrical contacts.
Micro Bearing Rig Rotordynamics
A fundamental enabling technology needed for all the high-power-density micro devices is the ability to spin silicon rotating elements, supporting the turbomachinery and electrical components, at peripheral speeds of several hundred meters per second, over two orders of magnitude faster than silicon rotors have previously achieved. This research focuses on experimental studies on the rotordynamic and hydrostatic journal gas bearings and thrust bearings of the micro devices.
MicroEngines (MEMS Gas Turbines, Generators, & Rocket Engines)
A multidisciplinary effort, in cooperation with the MIT Micro Technology Laboratory, is underway to develop gas turbine and rocket engines a few millimeters in diameter spinning at 2-5 x 106 rpm. The devices would be capable of producing 10-50 watts of power or 10-30 grams of thrust. Applications include battery replacement and micro-airplane propulsion. A subset of this effort is a program to build micro electric motor driven compressors of similar size. Also, a bipropellant, turbopump equipped, centimeter sized micro rocket engine, producing 3-5 lb. of thrust, is under development. There efforts encompasses all aspects of gas turbine and rocket propulsion engineering, including the fluid mechanics of turbomachinery, mechanical design, structures and materials, combustion, bearings, electric generators, materials, and microfabrication.
MEMS in Turbomachinery
An analytical and experimental program which is evaluating the use of large arrays of high frequency response micro-fabricated flow valves arranged on the tip casing of a compressor to alter the tip flowfield in an advantageous manner.
Assessment of Propfan Propulsion Systems for Reduced Environmental Impact
Baseline CRP blade-tip vortex system: Front rotor tip-vortices and viscous wakes interacting with rear rotor contribute to interaction tone noise. Current aircraft engine design studies tend towards higher bypass ratio, low-speed fan configurations in order to attain reductions in fuel consumption, emissions, and noise.
Propfan (advanced turboprop) engine concepts investigated in the past by American, European, and Russian aircraft manufacturers have demonstrated significant benefits in these areas. However, considerable concern remains about the potential noise generated by propfan engines, including both inflight cabin noise and community noise during takeoff and approach. The overall goal of this project is to define an advanced CRP configuration with improved noise characteristics while maintaining the required aerodynamic performance for a given aircraft mission. An aircraft performance, weight and balance, and mission analysis is conducted on a candidate CRP-powered aircraft configuration and a detailed aerodynamic design of a pusher CRP is carried out. Full wheel unsteady 3-D RANS simulations are then used to determine the time-varying blade surface pressures and unsteady flow features necessary to define the acoustic source terms.
Polar directivity at first interaction tone frequency: Implementing advanced source mitigation concepts in re-designed CRP significantly reduces interaction tone noise compared to baseline CRP design. A frequency domain approach based on Goldstein’s formulation of the acoustic analogy for moving media and an existing single rotor noise method is extended to counter-rotating configurations. Using the developed CRP noise estimation method, the underlying noise mechanisms front-rotor wake interaction, aft-rotor upstream influence, hub-endwall secondary flows, and front-rotor tip-vortices to interaction tone noise are dissected and quantified. Based on this investigation, the CRP is re-designed for reduced noise incorporating a clipped rear-rotor and increased rotor-rotor spacing to reduce upstream influence, tip-vortex, and wake interaction effects.
Maintaining the thrust and propulsive efficiency at takeoff, the noise is calculated for both designs. On the engine/aircraft system level, the re-designed CRP demonstrates significant noise reductions and the results suggest that advanced open rotor designs can possibly meet Stage 4 noise requirements. Re-designed CRP for low noise: Clipping rear rotor reduces interaction of front rotor tip-vortices with rear rotor, thereby decreasing interaction tone noise.
A Functionally Silent Aircraft to Transform Commercial Air Transportation
Aircraft noise is a major inhibitor of the growth of air transport. This project focuses on revolutionary enabling technologies for a functionally silent aircraft. Silent in this context means sufficiently quiet that the aircraft noise is less than that of the background noise in a typical well-populated environment. In order to make aircraft operations quiet enough such that the noise is not perceived as annoying by the community, airframe and propulsion system noise have to be reduced dramatically and beyond the current noise reduction goals. The proposed work introduces revolutionary concepts for a functionally silent aircraft and focuses on feasibility and quantitative assessment of the system integration of these enabling technologies.
Post-Combustion Flow Field Effects on Engine Exhaust Composition
Aircraft emissions are implicated in a diverse range of local and global atmospheric effects. This ongoing effort endeavors to make significant contributions towards understanding the role of engine hot section aerodynamics and kinetics in determining the composition of engine exhaust, specifically as these processes influence trace species constituents. A combined chemistry/flow model has been developed in collaboration with Aerodyne Research, Inc., that for the first time has allowed detailed 1-D, 2-D, and 3-D simulations of intra-engine trace species evolution. This program is currently pursuing detailed investigations of sulfur emissions, model development, and direct validation using planned experiments.
Gas Turbine Combustor Research
Some of the most complex flows within a gas turbine engine can be found in the combustor. As a result, design methods have historically been based largely on empiricism. We are working in collaboration with Pratt and Whitney and United Technologies Research Center to develop reduced-order models of various phenomena in gas turbine combustors. These models are being developed to provide insight into flow physics and chemistry, and to provide a basis for design tools that would serve as complements to empirical data and 3-D CFD.
High Fuel-Air Ratio Combustor and Turbine Research
As temperatures and overall fuel-to-air ratios increase in commercial and military aircraft engines, the potential for significant heat release due to post-combustor oxidation of partially reacted fuel is increased. Film-cooling flows can be the site of significant heat release, potentially altering surface heat transfer characteristics and damaging sections of the internal gas path. A shock tunnel test facility, numerical simulations and analytical modeling are being exercised to understand and provide design options to alleviate adverse effects that high fuel-air ratio combustion may have on hot-section durability, turbine performance, and pollutant emissions. The work is funded by Pratt and Whitney.
Low Greenhouse Gas Emissions Aircraft
We are performing systematic assessments of cost and emissions impacts of future aircraft technologies designed to reduce greenhouse gas emissions. Tools are being developed for use in a global, multiple transport mode context to conduct inter-modal comparisons of relative cost per unit emissions reduction potential for various technologies under different emissions regulation and demand scenarios. This work is jointly funded by MIT's Center for Environmental Initiatives and the MIT Cooperative Mobility Program.
System for Assessing Global Emissions of Aviation
Working with a team of researchers from the MIT Flight Transportation Laboratory and the DOT Volpe National Transportation Systems Center, we are developing what will become FAA's primary tool for assessing various policy options for regulating pollutant emissions from aircraft.
The Economic Value of Silence
Rising congestion and delays throughout the air transport system are in part associated with noise-related operational restrictions and airport expansion delays. The full societal costs of the constraints that these restrictions impose on service to the flying public, the quality of life for local communities, regional economic growth, and expansion of aerospace-related industries have never been rigorously documented. As a result, policy decisions regarding both regulatory strategies and federal expenditures to address aviation noise have not been adequately informed. We are collaborating with researchers from Cambridge University to articulate and provide a survey of the explicit and implicit costs of aircraft noise. Through this account, the potential benefits that follow reduced external, direct, investment, and opportunity costs associated with aircraft noise can be assessed. The work is thus a step towards enabling a rational account of the comparative utility of operational or technologically-oriented regulatory strategies, federal expenditures on noise reduction research, and local noise abatement.
Dynamic Control of Compressor and Compression System Aerodynamic Instability
A multi-disciplinary research area is "smart engines," in which the components (inlet, compressor, turbine, etc.) are under local feedback control. A three-stage axial compressor, two helicopter engines, a subsonic and a supersonic inlet diffuser are instrumented to support this research.
The Active Rotor
The active rotor is a composite transonic fan with embedded actuators capable of altering both the static shape and dynamic response of the fan blades. It is designed to be a research tool for flutter and forced response, aerodynamic performance, noise, and compressor stability. The program is currently focused on the very challenging rotor mechanical design and construction.
Active Control of Shocks in Supersonic Inlets
A major design driver for supersonic diffusers is the requirement to prevent unwanted shock formation and shock blow-out (supersonic unstart). As part of the Quiet Supersonic Platform (QSP) program to create a new efficient supersonic vehicle in the business jet size range, efficient supersonic inlets are being designed that rely on feedback control, rather than traditional design methods, to meet the unstart requirement. Small-scale experiments at MIT will use high-speed schlieren imaging to study the dynamics and control of shock formation and movement.
Diffuser Separation Control
Serpentine inlets common in tactical aircraft introduce significant levels of distortion to the flow into the compressor. Open loop and feedback methods to reduce this distortion, mitigate associated unsteadiness, and improve the pressure recovery of the diffuser are being investigated. A 1/6th scale Unmanned Combat Aerial Vehicle (UCAV) inlet is being tested at MIT, with plans for large scale testing at NASA Glenn.
Modeling and Integration of Active Internal Flow Control in Aero-Engine Compressors
Active flow control is one possible strategy to enhance both the performance and the stability of compressor and turbine flow systems. Synthetic jets are a means of injection actuation and consist of an orifice or neck that is driven by an ocillating wall in a cavity. This research program focuses on the key technical issues relating to the modeling and integration of flow control devices in aero-engine compressors with the goal to further enhance the performance and operability beyond the current loading and stability limits. The theme of the technical approach is a unique analytical, reduced-order dynamic system modeling methodology, which incorporates a system rather than a component view of this multi-disciplinary problem. Preliminary test will be conducted in MIT GTL's linear compressor cascade wind tunnel and final experiments are planned to be carried out in NASA GRC's Low Speed Axial Compressor (LSAC) test facility.
The Robust Jet Engine Project
Variability in gas turbine engine performance due to manufacture and in-service wear is a key factor in the competitiveness in the gas turbine industry. Aircraft engines are subject to a number of competing requirements and current designs often represent a compromise between efficiency and many of the other "ilities" (affordability, reliability, operability, etc). Further, while a given design may be highly efficient, if small perturbations from in geometry or operating state lead to large variations in engine performance, then the engine is not performing robustly. A continuing need exists for (1) engines that are robust to variability with greater reliability and efficiency yet at lower costs and (2) next generation design tools to develop these improved engines. The Robust Jet Engine project is a response to these needs with a particular focus on aerothermal robustness. The goals of this program are: Identification and quantification of key drivers for uncertainty and engine-to-engine variability in aerothermal quality including validation against data; Definition of criteria for the design of engines with a commercially-significant reduction in sensitivity to uncertainty and variability including analysis of cost trade-offs; Development of improved processes for monitoring and controlling the effects of variability on aerothermal quality; Implementation of one or more of the above elements in an industrial setting.