Aero Gas Turbine Engine Research Articles

Wax Patterns for Integrally Cast Rotors/Stators of Aeroengine Gas Turbines

Modern turbine engines for aerospace applications need precision investment cast components such as integrally cast rotors and stators, for their hot end turbine stages. Traditionally, these turbines are used as starter engines. In recent times, such engines are also used for strategic missile applications. The rotor/stator castings consist of a central hub (shrouded in some designs) over which a number of aerofoil shaped blades are located. Since these components cannot be machined, investment casting is the only available route for manufacture and hence stringent dimensional aerospace quality has to be in-built in the casting process itself. In the process of investment casting, pattern generation by injection of wax into dedicated dies/moulds is the first critical step. Traditional approach deals in producing individual blades with hub/shroud features through wax injection and assembly of a set of such injected patterns onto a dedicated and precisely manufactured fixture to wax-weld and generate an integral wax pattern, a process known as the ‘segmental approach’. It is possible to design a single-injection die with retractable metallic inserts in the case of untwisted blades of stator patterns without the shroud. Such an approach is also possible for twisted blades of rotors with highly complex design of inter-blade inserts and retraction mechanisms. DMRL has for long established methods and procedures for the above to successfully supply precision castings for various defence related projects. In recent times, urea based soluble insert approach has also been successfully applied to overcome the need to design and manufacture a precision assembly fixture, leading to substantial reduction in component development times. Present paper deals in length various approaches tried and established at DMRL to generate precision wax patterns for aerospace quality turbine rotors and stators. In addition to this, the importance of simulation in solving issues related to wax injection is also touched upon.

Impact of Foreign Object Damage on an Aero Gas Turbine Engine

An aero gas turbine engine is investigated for foreign object damage. Extensive damage in the low pressure compressor and abnormal noise forced the engine to shut down. Metallic screw type foreign object debris was retrieved from air intake fairing. Chemical analysis of the metallic debris and failed surface of LP compressor blade show the smearing of blade material on the debris and debris material on compressor blade. This confirms that impact of metallic foreign debris has caused extensive damage to LP compressor blades. The paper presents the details of investigation carried out to assess the extent of damage and its root cause. Effective foreign object prevention and elimination program has also been suggested in the paper to address the FOD in aircraft engines.

Design of the gas turbine engine rotor wheel with ceramic blades

The use of ceramic parts in the hot section of a gas turbine engine will make it possible to increase theengine efficiency due to the increased operating gas temperature at the combustion chamber outlet and reducedair consumption for blade cooling, as well as significantly reduce the weight of the parts due to the low densityof ceramics. The main disadvantage of ceramic parts is the brittleness of the material. Mechanical properties ofadvanced ceramic materials are insufficient for making a whole wheel, however, they are sufficient for makingblades. The main problem in making a structure of this kind is ensuring the strength of the interlock between ametal disk and ceramic blades. Previous works of the authors presented analysis of various types of interlocksbetween ceramic blades and a metal disk. This paper is devoted to the design of an aero engine gas turbine rotorwheel using a high-strength silicon nitride ceramic material. A software code in the APDL Ansys environment iscreated that makes three-dimensional design definition possible. A rotor wheel with ceramic blades is designedon the basis of a typical aircraft high-pressure gas turbine engine. A commercial heat-resistant nickel alloy isconsidered as the disk material, while silicon nitride is used for the production of blades. Evaluation of thestress-strain state and strength analysis are conducted, the properties of the ceramic material are analyzed. Theresults of the research show that it is necessary to make a special cooling system for a metal disk and that theproperties of the ceramics presented are to be improved. It is reasonable to apply similar constructions on stationarygas turbine plants.

Development and Integration of Rain Ingestion Effects in Engine Performance Simulations

Rain ingestion can significantly affect the performance and operability of gas turbine aero-engines. In order to study and understand rain ingestion phenomena at engine level, a performance model is required that integrates component models capable of simulating the physics of rain ingestion. The current work provides, for the first time in the open literature, information about the setup of a mixed-fidelity engine model suitable for rain ingestion simulation and corresponding overall engine performance results. Such a model can initially support an analysis of rain ingestion during the predesign phase of engine development. Once components and engine models are validated and calibrated versus experimental data, they can then be used to support certification tests, the extrapolation of ground test results to altitude conditions, the evaluation of control or engine hardware improvements and eventually the investigation of in-flight events. In the present paper, component models of various levels of fidelity are first described. These models account for the scoop effect at engine inlet, the fan effect and the effects of water presence in the operation and performance of the compressors and the combustor. Phenomena such as velocity slip between the liquid and gaseous phases, droplet breakup, droplet–surface interaction, droplet and film evaporation as well as compressor stages rematching due to evaporation are included in the calculations. Water ingestion influences the operation of the components and their matching, so in order to simulate rain ingestion at engine level, a suitable multifidelity engine model has been developed in the Proosis simulation platform. The engine model's architecture is discussed, and a generic high bypass turbofan is selected as a demonstration test case engine. The analysis of rain ingestion effects on engine performance and operability is performed for the worst case scenario, with respect to the water quantity entering the engine. The results indicate that rain ingestion has a strong negative effect on high-pressure compressor surge margin, fuel consumption, and combustor efficiency, while more than half of the water entering the core is expected to remain unevaporated and reach the combustor in the form of film.

Failure Analysis of Nozzle Guide Vane of a Low Pressure Turbine in an Aero Gas Turbine Engine

Failure of low pressure turbine nozzle guide vane (NGV) in an aero gas turbine engine is analyzed to determine its root cause. Forensic and metallurgical investigations are carried out on the NGV failed as well as the failed components of the downstream modules. Failure in the cooling system in the NGV due to crack in the sealing plate was found to be the cause of NGV getting burnt. This failure has caused extensive damages in low pressure turbine modules. Remedial measures are also suggested to prevent such failures.

The aerodynamic challenges of aeroengine gas-turbine combustion systems

AbstractThe components of an aeroengine gas-turbine combustor have to perform multiple tasks – control of external and internal air distribution, fuel injector feed, fuel/air atomisation, evaporation, and mixing, flame stabilisation, wall cooling, etc. The ‘rich-burn’ concept has achieved great success in optimising combustion efficiency, combustor life, and operational stability over the whole engine cycle. This paper first illustrates the crucial role of aerodynamic processes in achieving these performance goals. Next, the extra aerodynamic challenges of the ‘lean-burn’ injectors required to meet the ever more stringent NOx emissions regulations are introduced, demonstrating that a new multi-disciplinary and ‘whole system’ approach is required. For example, high swirl causes complex unsteady injector aerodynamics; the threat of thermo-acoustic instabilities means both aerodynamic and aeroacoustic characteristics of injectors and other air admission features must be considered; and high injector mass flow means potentially strong compressor/combustor and combustor/turbine coupling. The paper illustrates how research at Loughborough University, based on complementary use of advanced experimental and computational methods, and applied to both isolated sub-components and fully annular combustion systems, has improved understanding and identified novel ideas for combustion system design.

Fatigue Failure of LP Compressor Blade in an Aero Gas Turbine Engine

Failure of low-pressure compressor rotor blade in an aero gas turbine engine is analyzed to determine its root cause. Forensic and metallurgical investigations are carried out on the blade failed. The failure of the first stage rotor blade is found to be the first in the chain of events that led to the engine failure. The mode of failure in the blade is found to be fatigue and has originated from the mounting lug fillet region due to high stress concentrations. The failure has caused extensive damages in low-pressure compressor module and also in downstream modules as a secondary effect. Remedial measures are also suggested to prevent such failures.

A Model for Creep and Creep Damage in the γ-Titanium Aluminide Ti-45Al-2Mn-2Nb.

Gamma titanium aluminides (γ-TiAl) display significantly improved high temperature mechanical properties over conventional titanium alloys. Due to their low densities, these alloys are increasingly becoming strong candidates to replace nickel-base superalloys in future gas turbine aeroengine components. To determine the safe operating life of such components, a good understanding of their creep properties is essential. Of particular importance to gas turbine component design is the ability to accurately predict the rate of accumulation of creep strain to ensure that excessive deformation does not occur during the component’s service life and to quantify the effects of creep on fatigue life. The theta (θ) projection technique is an illustrative example of a creep curve method which has, in this paper, been utilised to accurately represent the creep behaviour of the γ-TiAl alloy Ti -45Al-2Mn-2Nb. Furthermore, a continuum damage approach based on the θ-projection method has also been used to represent tertiary creep damage and accurately predict creep rupture.

A comparative study between combustion performances of turbine inter-guide-vane burner and trapped vortex combustor

To improve the performance of aero gas turbine engines, more and more interests have been shown on turbine inter-guide-vane burner based on the ultra-compact combustion concept. To make a universal turbine inter-guide-vane burner, a new concept is proposed using a trapped vortex cavity to replace the high swirling circumferential cavity combustor to address the need to scale the configuration for a larger turbine spool. Three models, including trapped vortex combustor, transition model, and turbine inter-guide-vane burner, are designed. Comparative analysis between combustion performances of three models by using numerical simulation method is carried out. The scale-adaptive simulation turbulence model is used in the simulation process, aiming to reduce the deviation between numerical simulation value and actual value. Finally, the turbine inter-guide-vane burner model is found to be the superior design proposal for turbine inter-guide-vane combustion technology, compared with the other two models.

SIMULATION AND EXPERIMENTAL STUDIES ON GRAIN SELECTION BEHAVIOR OF SINGLE CRYSTAL SUPERALLOY I. Starter Block

从实验及模拟角度对比研究了引晶段的晶粒密度和取向随晶粒生长高度（研究位置距试样底面的高度）的变化规律，并给出了引晶段参数的设计准则．采用EBSD晶体取向成像技术获得了引晶段截面的晶粒形貌和取向极图；采用CA—FD方法，针对单晶定向凝固过程进行数理建模与仿真，实现了凝固过程的三维宏观温度场与微观组织生长的模拟计算．从宏、微观角度解释了定向凝固过程的晶粒竞争演化行为，为引晶段设计提供理论支持．

Modelling and simulation of manufacturing process chains

Modelling and simulation of manufacturing process chains are important for decreasing the defects induced by the manufacturing processes and increasing the life of the components during production. A life-based methodology is proposed for aero-engine gas turbine components where the main aim is to reduce the level of tensile stresses in order to improve the life. A new Finite Element (FE) software FEDES (FE Data Exchange System) is presented and used to simulate three manufacturing process chains where the final residual stress state and distortions are predicted to estimate the life of the components. Also, the paper reviews the current developments in manufacturing process chains and highlights the future challenges in the field.

Determination of optimum specific thrust for civil aero gas turbine engines: a multidisciplinary design synthesis and optimisation

A systematic methodology for the determination of optimum specific thrust of civil turbofan engines is presented. The optimum is defined as the specific thrust at which the engine direct operating cost is minimised. Based on publicly available data and clean‐sheet analysis, the developed method provides a rationale for the values of specific thrust found in existing engines and the basis for future designs. The optimum specific thrust determined here complements the thermodynamic optimisation strategy for engine parameters developed in a previous publication by Guha. The process of optimisation thus involves a complex multidisciplinary methodology involving aerodynamics, thermodynamics, structures, system integration, economics, legislations, and other design and operational issues. This study utilises two models: the fixed aircraft rubber engine model which simulates the retrofitting of new engines to a current airframe, and the rubber aircraft rubber engine model which simulates the thermodynamic optimisation of the engines alongside the optimisation of the airframe dimensions. Both models demonstrate that as the design range of the aircraft increases, the optimum value of specific thrust decreases. As fuel prices rise, the optimum value of specific thrust reduces further, taking it to levels where radical changes in engine design and integration would be necessary.

Effect of Combustor Geometry on Performance of Airblast Atomizer Under Sub-Atmospheric Conditions

One of the certification requirements that a jet engine has to fulfil is its altitude relight capability. Relighting an aero gas turbine engine at high altitudes is more challenging than at sea level conditions. The pressure, air velocity, and temperature within the combustor at such conditions are very low, hindering the fuel atomization and evaporation process. After ignition, combustion efficiency can be relatively low due to the poor fuel atomization quality, leading to slow shaft acceleration rates. Further studies in this field can help determine and predict the fuel spray characteristics, which limit the relight and pull-away capability of the engine at these sub-idle conditions. Reported in this paper is the CFD analysis of a typical airblast atomizer, simulated at different sub-idle operating conditions. Two sets of models were used; one with a simple liner-only combustor with co-flow air, and the other with a more detailed geometry, including wall cooling slots, primary and secondary dilution holes, and co-flow air. For the simpler model, three different inner and outer liner wall spacing were modelled to examine the effect of the chamber volume on the fuel spray behaviour. The CFD models were then run at a chamber pressure of 101, 41 and 3 1 kPa, typical of sub-idle conditions. The effect of such conditions on the atomization quality of the fuel spray was analysed. The study carried out indicates how the chamber pressure, chamber volume and AFR (through amount of co-flow air introduced) significantly affect the resulting spray characteristics. A parametric analysis was performed to extract a correlation between the spray SMD (Sauter Mean Diameter) and fuel flow rate.

ZONAL RANS-LES MODELING FOR TURBINES IN AEROENGINES

The cost of large-eddy simulation (LES) modeling in various zones of gas turbine aeroengines is outlined. This high cost clearly demonstrates the need to perform hybrid Reynolds-averaged Navier-Stokes-LES (RANS-LES) over the majority of engine zones because the Reynolds number is too high for pure LES. The RANS layer is used to cover over the fine streaks found in the inner part of the boundary layer. The hybrid strategy is applied to various engine zones, which is shown to typically give much greater predictive accuracy than pure RANS simulations. However, the cost estimates show that the RANS layer should be disposed within the low-pressure turbine zone. Also, the nature of the flow physics in this zone makes LES most sensible. © 2012 by Begell House, Inc.

Real-time diagnostics of a jet engine exhaust using an intra-pulse quantum cascade laser spectrometer

It has been demonstrated that an intra-pulse scanned quantum cascade laser spectrometer may be used to obtain real-time diagnostics of the amounts of carbon monoxide, carbon dioxide, and water, in the exhaust of an aero gas turbine (turbojet) engine operated in a sea level test cell. Measurements have been made of the rapid changes in composition following ignition, the composition under steady state operating conditions, and the composition changes across the exhaust plume. The minimum detection limit for CO in a double pass through a typical gas turbine plume of 50 cm in diameter, with 0.4 seconds integration time, is approximately 2 ppm.

Evolutionary Optimization for Gain Tuning of Jet Engine Min-Max Fuel Controller

AS turbine aero engines (GTEs) have played a significant role in the expansion of the flight capabilities of modern aircraft. For a GTE, stable and safe operation of the engine components must be ensured in order to operate at the operability and performance level for which it is designed. The thrust response is also of great importance for an aircraft, both in terms of the speed of response and accomplishment of the adequate thrust [1]. Hence, an appropriate control system has always been an essential part of the GTEs to provide both regulation and management. The main task of a GTE control system is to provide the required performance while maintaining safe and stable operation [2]. The safetyconsiderationsincludethe flowinstability,highheatloads,and high cyclic stresses. In other words, the GTE control requirements include amainfuel controlforsetting andholdingsteady-state thrust (steady-state control mode) and fuel acceleration and deceleration schedules to provide limit protection (physical limitation control mode) and rapid time response for gas turbine engine (transient control mode) [3]. AsurveyonGTEcontrolstrategiesshowsthatvariousfuelcontrol methods have been used for GTE fuel control in the last two decades [4–16]. These studies show that the complexity of the engine control comes from the need to operate the engine as close as possible to its limits. In other words, the design of the control system for a gas turbine aero engine results in a series of single input single output controllers. Many methods such as PID, LQ, LQG, and H1 are used as one of the single feedback control loops of the GTE control system. In this case, the design method is based on a selection approach between various control loops. This strategy is named “min-max selection strategy,” as the selection of the control loops is based on a min-max approach between transient control loops. Minmax approach is an industrial method for the design of the GTE fuel control system. However, in order to provide an improved engine performance as well as the engine protection against the physical limitations, one of the most common complexities of the min-max selection strategy is controller gain tuning. Taking the nonlinearity and switching nature of this control strategy into account, gradient-based optimization methods have weak performance for gain tuning. As a result, this problem requires a non gradient optimization technique on the basis of the evolutionary algorithms. In this paper, application of an evolutionary algorithm for optimization of the min-max fuel controller parameters is presented. For this purpose, the control requirements and constraints for a gas turbine aero engine are first explained. The min-max fuel control strategy is then described and an initial min-max fuel controller is designed. Subsequently, using the genetic algorithm (GA), the parameters of the initial min-max fuel controller is tuned so that the engine requirements and constraints are satisfied. In optimization process, the fitness function is defined to minimize the settling time andfuelconsumption,wheretheweightfactorsaresetwithrespectto the importance of each function. In addition, a computer simulation programisdevelopedtoinvestigatetheeffectivenessoftheapproach for a single-spool jet engine. The results of simulation are validated by experimental data in both steady-state and transient modes to support the simulation model. Finally, the results are provided to investigate the performance of the optimized controller and to evaluate the effectiveness of the approach.

Crack growth in the creep-fatigue regime under constrained loading of thin sheet combustor alloys

Damage tolerant lifing methodologies are an essential requirement to support the safe and reliable operation of combustors in the modern gas turbine aero-engine. These static components experience a wide range of temperature superimposed upon a complex stress field resulting from their geometrical form. Features such as seam welds, injector ports and cooling hole arrays could potentially initiate and subsequently interact with fatigue cracks, while the thin sheet structure of the combustion annulus locally imparts plane stress dominant conditions. However, on the macro-scale the stress system is highly constrained, ensuring that cracks usually arrest before achieving a critical size, thereby avoiding failures. These service observations must now be supplemented by a detailed scientific understanding of fatigue behaviour in combustor alloys under representative conditions.The technical challenge of evaluating crack growth rates in thin sheet laboratory scale specimens under strain controlled fatigue at high temperatures is addressed. The alloy system of interest was the nickel based superalloy Haynes 230. A matrix of novel fatigue crack growth tests is reported at temperatures ranging from 816°C to 982°C under a variety of strain ratios (εmin/εmax) with crack growth monitored via an automated pulsed DCPD system. Contemporaneous measurements of constitutive behaviour will demonstrate the significant degree of stress relaxation that occurs under these constrained conditions, continually modifying the crack tip driving force. This was best quantified by the corresponding stress ratio, which generally fell throughout the period of the test and under dwell loading cycles often resulted in a highly negative value (e.g. stress ratio lower than minus fifteen) incorporating a minimal tensile stress contribution. Crack growth can clearly continue under such highly compressive regimes. The contribution from creep damage mechanisms will be illustrated via post test fractographic and microstructural examination.

Advanced Ni base superalloys for small gas turbines

Ni base superalloy cast materials provide an outstanding balance of high temperature strength, fatigue resistance, oxidation resistance and coating performance, and can be produced to very tight tolerances in extremely complex configurations, such as axial and centrifugal integral cast turbine wheels. As a result, cast superalloys are used in the most demanding applications of aero and industrial gas turbine engines. Use of these materials is expanding to smaller microturbine, turbojet, turbocharger and missile engine applications due to this unique combination of desirable properties. This paper will present an overview of the application of investment cast Ni base superalloys and process capability to small turbine and missile engines.Les matériaux moulés de superalliage à base de Ni fournissent une balance exceptionnelle de résistance à haute température, de résistance à la fatigue, de résistance à l’oxydation et de performance du revêtement. On peut également les produire à des tolérances très serrées dans des configurations extrêmement complexes, comme des roues de turbine à coulée axiale et centrifuge d’une seule pièce. Conséquemment, les superalliages moulés sont utilisés dans les applications les plus exigeantes des turbines à gaz aéro- et industrielles. L’utilisation de ces matériaux s’étend à des applications de plus petite micro turbine, de turboréacteur, de turbocompresseur et de moteur de missile grâce à cette combinaison unique de propriétés désirables. Ce document présente une revue de l’application des superalliages à moulage de précision à base de Ni et de la capacité opérationnelle du processus pour les petites turbines et les moteurs de missile.

Nonlinear Multiple Points Gas Turbine Off-Design Performance Adaptation Using a Genetic Algorithm

Accurate gas turbine performance models are crucial in many gas turbine performance analysis and gas path diagnostic applications. With current thermodynamic performance modeling techniques, the accuracy of gas turbine performance models at off-design conditions is determined by engine component characteristic maps obtained in rig tests and these maps may not be available to gas turbine users or may not be accurate for individual engines. In this paper, a nonlinear multiple point performance adaptation approach using a genetic algorithm is introduced with the aim to improve the performance prediction accuracy of gas turbine engines at different off-design conditions by calibrating the engine performance models against available test data. Such calibration is carried out with introduced nonlinear map scaling factor functions by “modifying” initially implemented component characteristic maps in the gas turbine thermodynamic performance models. A genetic algorithm is used to search for an optimal set of nonlinear scaling factor functions for the maps via an objective function that measures the difference between the simulated and actual gas path measurements. The developed off-design performance adaptation approach has been applied to a model single spool turbo-shaft aero gas turbine engine and has demonstrated a significant improvement in the performance model accuracy at off-design operating conditions.

Robust Design Optimization of Gas Turbine Compression Systems

Gas turbine compression systems are required to perform adequately over a range of operating conditions. Complexity has encouraged the conventional design process for compressors to focus initially on one operating point, usually the most commonor arduous, to draw up an outline design. Generally, only as this initial design is refined is its offdesign performance assessed in detail. Not only does this necessarily introduce a potentially costly and timeconsuming extra loop in the design process, but it also may result in a design whose offdesign behavior is suboptimal. Aversion of nonintrusive polynomial chaos was previously developed in which a set of orthonormal polynomials was generated to facilitate a rapid analysis of robustness in the presence of generic uncertainties with good accuracy. In this paper, this analysis method is incorporated in real time into the design process for the compression system of a three-shaft gas turbine aeroengine. This approach to robust optimization is shown to lead to designs that exhibit consistently improved system performance with reduced sensitivity to offdesign operation.