A pragmatic and versatile airfoil design and fabrication system, named the Arbitrary Surface Flank Milling (ASFM) system, has been continuously developed and applied to the design and manufacturing of fan and axial compressor airfoils — in the form of integrally bladed rotors (IBR) and centrifugal impeller blades — for jet engines at Pratt & Whitney over the past 28 years. In a previous paper 18 years ago, we gave an account of the general concept and methodology of the system, illustrated with a number of actual production engine applications. We have since then greatly expanded the domain of applicability of the system to successfully handle blades of unusual geometric complexity. We have made strides in the area of design for flank mill-ability with conventional cutters, and cutters shaped according to blade geometry. Based on the same concept and software methodology, we have also successfully demonstrated the superior quality and cost effectiveness of Flank Super Abrasive Machining (Flank SAM) of blades in nickel-based alloy IBRs.

The heat transfer occurring within a turbocharger is closely coupled to the internal combustion exhaust. This process is highly complex and it is not taken account either in the choice of components nor is it reflected in the efficiency maps. In particular the flow within the compressor is usually treated as adiabatic. This is not the case since the heat fluxes from both the turbine and the engine must be taken into account in the compression process. Tests at Imperial College enabled the heat fluxes on a turbocharger installed on a 2.0 litre diesel engine to be quantified. The main aim was to evaluate the heat transfer impact on the compressor efficiency. In addition to this a simplified 1-D model was then developed and validated with experimental measurements. Based on the data generated by the model, a correlation for the compressor non-adiabatic efficiency is presented in the paper. A multiple regression procedure was developed; the work is based on a statistical description of the different parameters that affect the heat transfer model. The regression analysis showed that the compressor non-adiabatic efficiency can be evaluated with good degree of confidence by three main independent parameters: Mach number, pressure ratio and “temperature parameter”. The Mach number had the largest effect on the compressor efficiency, while the temperature factor was revealed to be significant only at low speeds. The impact of the turbocharger size on efficiency was also assessed. The length of the bearing housing and the compressor diameter turned out to be significant within a few percentage points of the compressor efficiency.

The processes occurring in turbo machinery applications are frequently treated as adiabatic. However, in a turbocharger significant heat transfer occurs, leading to a deficit of turbocharger performance. The overall objective of this experimental work is to improve the understanding of the heat transfer process taking place in a turbocharger when installed on an internal combustion engine. In order to do this, beyond the standard set of measurements needed to define the turbo operating point, a large number of thermocouples were installed on the turbocharger. The tests results allow the quantification of the temperatures within the turbocharger and revealed that a nonuniform temperature distribution exists on the compressor and turbine casings. This is partly attributed to the proximity of the turbocharger to the engine. This process plays a role on the deterioration of the compressor efficiency when compared to the corresponding adiabatic efficiency. A correlation that allows the calculation of the compressor exit temperature is proposed. The method uses the surface temperature of the bearing housing; it was validated against experimental data with deviations no larger than 3%. A simplified 1-dimensional heat transfer model was also developed and compared with experimental measurements. The algorithms calculate the heat transferred through the turbocharger, from the hot end to the cold end by means of lump masses. The compressor performance deterioration from the adiabatic map is predicted.

Taking into account the increasing availability of computational power at an affordable cost, two-dimensional through-flow calculation methods are gaining more and more attention, given the fact that the required time for convergence is continuously reducing. Consequently, several application fields (i.e. whole engine performance simulation), that were traditionally dominated by simpler and faster zero-dimensional or one-dimensional methods, purely because of computational power restrictions, gradually move towards two-dimensional analyses. These tend to offer more information about the flow-field at a greater accuracy. The Radial Equilibrium Equation (REE), in its either simple or full version, has been the basis of several two-dimensional and quasi-three-dimensional through-flow techniques that are being used for the flow analysis within ducts, compressors and turbines. The aim of this paper is to provoke a thorough discussion on the actual solution of the full REE for the determination of the meridional velocity profile. More precisely, this manuscript discusses in detail the implications on the solution of the full REE when the blade lean angle related terms are included in the equation. This issue has only been superficially addressed in the existing literature up to this stage. The expressions for radial equilibrium addressed in the context of this paper, mainly consist the basis of a particular streamline curvature code ( 2D SLC Compressor Software ), developed as a performance investigation and design tool of axial flow compressors. This code has been through a number of ‘improvement cycles’ over its several years of existence. One such cycle included the elaborate study of several final versions of the full REE, in order to reassure a stable and fast convergence for the final solution, while maintaining the highest possible level of accuracy. Firstly, this manuscript presents the final version of the full REE, commenting on each individual term in the equation, as well as on the various assumptions made during its derivation process. The two different solutions of the equation are given for zero and non-zero blade lean angle values. Moreover, the implications of the solution of the non-zero blade lean angle equation on the stability, convergence time and accuracy of the final results are pointed out. Finally, some conclusions are expressed as far as the effects of the blade lean angle on a compressor blade row performance and the actual applicability of the two forms of the REE are concerned. These conclusions were drawn from personal experience applying the equations but also from an extensive literature review conducted.

Distributed generation emerged as a new philosophy for electricity generation in our time, since then, it has been possible to see new concepts of technology following the idea of energy production away from the main producers or in remote areas, mainly in the countryside. Distributed generation technologies include small gas turbine engines, internal combustion reciprocating engines, photovoltaic panels, fuel cells, solar thermal conversion and Stirling engines using fossil and biofuels. Among them, the small gas turbine engine that generates electricity and heat working with fossil or renewable fuels is a promising technology for the near future. The aim of this work is the cycle analysis and preliminary compressor design of a 600kW simple cycle gas turbine engine that has been developed in Brazil. The 600kW engine will be the first prototype of its class in Brazil. A cycle performance calculation for different pressure ratios and turbine inlet temperature was carried out for fixed component efficiencies and losses. A selection of the design point was discussed and compared with the existing commercial engines of the same class. A compressor design point calculation was carried out with a mean line calculation CODE developed in FORTAN language. A CFD simulation was used for flow field analyses and design refinement.

The accurate prediction of centrifugal compressor stability continues to be an important area of interest in the oil and gas industry. Ensuring stability is critical to the cost-effective installation and operation of these machines in remote environments where field stability problems are much more expensive to diagnose and correct. Current industry standards and tools for the prediction of impeller destabilizing forces are based on empirical methods that, to date, have served fairly well for systems with reasonable stability margins. However, as stability margins are decreased, use of a modeling method that is more physics based that can better represent the observed trends in machine behavior at low stability margins is needed. Furthermore, the development of mega-class Liquefied Natural Gas (LNG) compressors and ultra-high pressure re-injection compressors provide further motivation to improve accuracy. In this paper, a new physics based expression for the prediction of impeller cross-coupling, previously introduced by Moore et al. [1] is further investigated by analyzing several classes and scale factors of impellers ranging from 2-D designs used in re-injection up to full 3-D impellers typically used in LNG. The new expression is based on both numerical simulation (CFD) and experimental test data from a known instability. Comparisons are made for overall stability prediction as well as sensitivity to system changes. Conclusions are made regarding the applicability and limits of this new approach.

In order to reduce the amount of carbon dioxide (CO 2 ) greenhouse gases released into the atmosphere, significant consideration has been given to the sequestration of CO 2 from power plants and other major producers of greenhouse gas emissions. Integrated Gasification Combined Cycle (IGCC) power plants offer an alternative to pulverized coal plants because the carbon dioxide may be separated from the process gas stream prior to combustion. The compression of the captured carbon dioxide stream requires a sizeable amount of power, which impacts plant availability, capital expenditures and operational cost. Preliminary analysis has estimated that the CO 2 compression process reduces the plant efficiency by 8% to 12% for a typical IGCC plant. The detailed thermodynamic analysis presented here examines methods to minimize the power penalty to the producer through integrated, low-power compression concepts. The goal of the present research is to reduce this penalty through novel compression concepts and integration with existing IGCC processes. The research supports the U.S. Department of Energy (DOE) National Energy Technology Laboratory (NETL) objectives of reducing the energy requirements for carbon capture and sequestration in electrical power production. The primary objective of this study is to boost the pressure of CO 2 to pipeline pressures with the minimal amount of energy required. Fundamental thermodynamic analysis methods related to the compression of CO 2 are used in the following paper to explore pressure and enthalpy rise in both liquid and gaseous states.

Development of wet gas compressors is challenging due to the liquid phase impact on performance. Experimental investigation of airfoil behavior in wet condition contributes to a revised compressor design and increased understanding of multiphase flow mechanisms. The performance of an airfoil was investigated in wet gas flow. An air-water mixture was used as the experimental fluid. The influence of wet gas flow on airfoil performance was investigated at different angles of incidence and gas volume fractions. A qualitative description of the complex physical process observed when liquid is introduced into the flow field is given. Airfoil performance was degraded at increased liquid mass flow rate owing to premature boundary layer separation. The initiation of separation was observed as a local film thickening, followed by increased liquid film fluctuations. A continuity wave was observed surrounding the airfoil, forming a U shape of increased liquid concentration. The wave was initiated by deposited droplets and the formation of secondary droplets. The investigation reveals that compressor operating range, surge and stall margins are affected by the wet gas fluid. Reviewed literature and experiments confirm a reduced stall and surge margin when a compressor is exposed to wet gas. Further investigation will involve sub-scale impeller tests to determine the effects on the performance and stability ranges.

The tandem airfoil has potential to do more work as a compressor blade than a single airfoil without incurring higher losses. The goal of this work is to evaluate the fluid mechanics of a tandem rotor in the rear stages of a core compressor. As such, the results are constrained to shock-free, fully turbulent flow with thick endwall boundary layers at the inlet. A high hub-to-tip ratio 3D blade geometry was developed based upon the best-case tandem airfoil configuration from a previous 2D study. The 3D tandem rotor was simulated in isolation in order to scrutinize the fluid mechanisms of the rotor, which had not previously been well documented. A geometrically similar single blade rotor was also simulated under the same conditions for a baseline comparison. The tandem rotor was found to outperform its single blade counterpart by attaining a higher work coefficient, polytropic efficiency and numerical stall margin. An examination of the tandem rotor fluid mechanics revealed that the forward blade acts in a similar manner to a conventional rotor. The aft blade is strongly dependent upon the flow it receives from the forward blade, and tends to be more three-dimensional and non-uniform than the forward blade.

Given the current level of computational resources that are readily available, three dimensional (3-D) gas turbine engine performance simulation remains extremely time consuming. The current paper presents a synthesis of existing flow simulation methods coupled together in the form of a new software package. The software is able to assess the impact of a 3-D flow profile at the intake inlet on engine performance, demanding relatively low computational resources. More precisely four flow simulation techniques are employed, represented respectively by four individual stand alone software sub-modules. 3-D Vortex Lattice Method (VLM) is used to simulate the intake flow. Subsequently the intake outlet 3-D flow profile is decomposed into a radial and a circumferential component. For the compressor performance simulation, that receives those components as inlet boundary conditions, a two dimensional (2-D) Streamline Curvature (SLC) simulation method coupled with an extended parallel compressor model is used. SLC addresses the impact of the radial flow distortion, whereas the extended parallel compressor model examines the impact of circumferential flow distortion on engine performance. The results of the above analysis are stored into an intake-compressor performance characteristic map, which is then fed into a zero dimensional (0-D) performance simulation tool in order to evaluate the overall impact of the intake inlet distorted flow on engine performance. The paper is divided into two major sections. The first one presents the individual flow simulation techniques, together with the corresponding software modules. A short summary of each method is given first and then the software module is described, followed by brief comments on the validation results that have been already published. The section in concluded by the description of the synthesized software. The second major section deals with the application of the synthesized simulation method on a turbojet engine. A generic turbojet engine has been chosen mounted behind a generic intake, given the lack of relevant experimental results. The engine has a four stage axial flow compressor driven by a single stage axial flow turbine, followed by a converging nozzle. 3-D total pressure profiles were imposed at the intake inlet and several comparative graphs of engine’s performance parameters between “clean” and distorted inlet flow conditions are given. The paper is concluded with a discussion on software’s abilities and weaknesses as well as on its potential future expansion.

The various incidence, deviation and loss models used in through-flow analysis methods, such as Streamline Curvature, are nothing more than statistical curve fits. A closer look at public domain data reveals that these statistical correlations and curve fits are usually based on experimental cascade data that actually display a fairly large scatter, resulting in a relatively high degree of uncertainty. This usually leads to substantial differences between the calculated and actual performances of a given gas turbine engine component. Typically, matching calculated results from a throughflow analysis against experimental data requires the combination of various correlations available in the public domain, through a very tedious, complex and time consuming ‘trial and error’ process. This particular study supports the view that it might actually be much more time-effective to “adopt” a given loss model against experimental data through an iterative, physics-based approach, rather than try to identify the best combination of available correlations. For example, the well-established “Swan’s model” for calculating the blade profile loss factor in subsonic and transonic axial flow compressors depends strongly on approximate correlations for calculating the blade wake momentum thickness, and therefore represents such a case. This study demonstrates this by looking into an iterative approach to blade profile loss model adaptation that can provide a relatively simple and quick, but also physics-based way of ‘calibrating’ profile loss models against available experimental data for subsonic applications. This paper presents in detail all the analysis necessary to support the above concept and discusses Swan’s model in particular as an example. Finally, the paper discusses the performance comparison of a two-dimensional, Streamline Curvature compressor model against experimental data before and after the adaptation of that particular loss model.

One of the effective methods for improvement of compressor stage operation is a flow control within the compressor passages. An application of boundary layer control on the suction surface of a blade profile may be useful in tandem cascade configuration. Total tandem guide vane consists of a series of two vanes: guide vane (the first segment) and straightening vane (the rear segment). These vanes have a definite shift along the axial and peripheral direction, providing a perfect re-distributing of energy within the blade passages between different flow zones. There are several parameters influencing mainly on loss in a tandem cascade. Experimental results of the parametric test are used for the creation of a statistical model of loss determination of compressor tandem cascade using the Group Method of Data Handling (GMDH). The paper focuses on the statistical-based correlation that encompasses most pertinent parameters featuring tandem configuration and operating condition.

Major challenges are related to compressor and driver integration during run down. In order to understand these challenges, the pipeline compressor facility at Troll Kollsnes gas treatment plant, Norway, has been subjected to detailed trip testing and dynamic simulation analysis. The plant includes five pipeline compressors and is utilised as a pilot for analysing the transient response of a 40 MW compressor driven by a variable speed electric motor. The compressor control and protection system include an anti-surge and a hot gas bypass system. Vibration records have shown that under power outage the compressors were exposed to violent vibrations. Further investigation revealed that during a short power outage, the compressor enters the surge- and rotating stall area under certain operating scenarios. The rotating stall response resulted in reduced operating range and flexibility for the pipeline compressors. Specific precautions had to be taken to prevent the compressor from running into the low flow operating area of the performance envelope. Dynamic simulations cover important aspects related to the transient scenario analyses performed to reveal the root cause of the compressor problems. The simulation system enables sophisticated plant models to be configured from high quality standard model algorithm building blocks. Verification of the model blocks have been performed against plant records in order to validate the transient predictions. The paper reports experience from testing and verification of compressor and driver integration with reference to transient behaviour during run down. This includes the validation of the dynamic models, which apply both to the design and commissioning phase where actual plant trip tests should be used to verify the design and stability margins.

A streamline curvature method with improvements to key loss models is applied to a two-stage, low aspect ratio, transonic fan with design tip relative Mach number of approximately 1.65. Central to the improvements is the incorporation of a physics-based shock model. The attempt here is to capture the effects of key flow phenomena relative to the off-design performance of the fan. A quantitative analysis regarding solution sensitivities to model parameters that influence the key phenomena over a wide range of operating conditions is presented. Predictions are compared to performance determined from overall and interstage measurements, as well as from a three-dimensional, steady, Reynolds-averaged Navier-Stokes method applied across the first rotor. Overall and spanwise comparisons demonstrate that the improved model gives reasonable performance trending and generally accurate results. The method can be used to provide boundary conditions to higher-order solvers, or implemented within novel approaches using the streamline curvature method to explore complex engine-inlet integration issues, such as time-variant distortion.

This paper describes the introduction of 3D blade designs into the core compressors for the Rolls-Royce Trent engine with particular emphasis on the use of sweep and dihedral in the rotor designs. It follows the development of the basic ideas in a university research project, through multistage low-speed model testing, to the application to high pressure engine compressors. An essential element of the project was the use of multistage CFD and some of the development of the method to allow the designs to take place is also discussed. The first part of the paper concentrates on the university-based research and the methods development. The second part describes additional low-speed multistage design and testing and the high-speed engine compressor design and test.

A streamline curvature throughflow numerical approach is assessed and modified to better approximate the flow fields of transonic axial compression systems. Improvements in total pressure loss modeling are implemented, central to which is a physics-based shock model, to ensure accurate and reliable off-design performance prediction. The new model accounts for shock geometry changes, with shock loss estimated as a function of inlet relative Mach number, blade section loading (flow turning), solidity, leading edge radius, and suction surface profile. Data from a single-stage, isolated rotor provide the basis for experimental comparisons. Improved performance prediction is shown. The importance of properly accounting for shock geometry and loss changes with operating conditions is demonstrated.

This paper describes the introduction of 3D blade designs into the core compressors of the Rolls-Royce Trent engine series with particular emphasis on the use of sweep and dihedral in the rotor designs. It follows the development of the basic ideas in a university research project, through multistage low-speed model testing, to their application to the high pressure engine compressor. An essential element of the project was the use of multistage CFD and some of the development of the method to allow the designs to take place is also discussed. Part I of the paper concentrated on the fundamental university-based research and the methods development. Part II describes additional low-speed multistage design and testing and the high-speed engine compressor designs and tests.

The present study is concerned with wake-induced unsteady effects in axial-compressor blade rows. The goal is to exploit these effects in order to design high-lift blades without increasing the profile loss, as has been achieved for low-pressure turbine blades. In the first part of this paper, the experimental means and the computational fluid dynamics tools are described. The rig features a flat plate that can be subjected to different velocity distributions representative of the suction side of a real compressor blade. Cylindrical bars mounted on a moving system simulate the incoming wakes from the upstream blade row in the compressor. Results are presented for steady flow and for unsteady compressor-like conditions. In all cases, the separation bubble of the steady flow is suppressed by the turbulence that is induced in the boundary layer by the wakes at approximately 10% of the suction side. Its reappearance is then delayed by a region of stable laminar-like flow and low loss, the so-called calming effect that follows the wake-induced patch. The paper describes these phenomena for an example of configuration. It is then shown that it should be possible to increase the lift by 35% while keeping the same level of loss as the initial conventional pressure distribution of the study.

Thermodynamic model of wet compression process is established in this paper. The topics of ideal wet compression process, actual wet compression process, water droplet evaporative rate, wet compression work, inlet evaporative cooling, concept of wet compression efficiency, aerodynamic breaking of water droplets etc. are investigated and discussed in this paper.

A micro-scale, high-speed compressor impeller (12mm diameter, 800,000 rpm) was tested for feasibility in regard to aerodynamic performance. The compressor was designed for application in a fist-sized gas-turbine-generator. To survive high stresses at such high temperatures, the rotor was manufactured as a single turbine/compressor/shaft unit in silicon nitride, by the Mold SDM process. Performance testing was conducted in a cold-flow rig at reduced speed of 420,000 rpm. Results from a CFD code compared favorably to measured data at this speed. Extrapolation from test conditions to full design speed was accomplished by application of CFD applied at both speeds.