Preparation and characterization of LPPS NiCoCrAlYTa coatings for gas turbine engine

NiCoCrAlYTa coatings have been deposited onto an aircraft gas turbine engine blade using a LPPS unit equipped with a computerized robot. Optimal processing conditions, including spray parameters, the trajectory of the robot, and the synchronized movements between the torch and the blade, have been developed for superior coating properties. Transferred arc treatment, providing a preheating and a cleaning of the substrate surface, enhances the adherence of the coatings to the substrate. The resulting LPPS coatings show dense and uniform characteristics with ideal hardness, and good corrosion resistance to cycle oxidation.

Fault Identification and Health Monitoring of Gas Turbine Engines Using Hybrid Machine Learning-based Strategies

Ahealth monitoring scheme is developed in this work by using hybrid machine learning strategies to iden-tify the fault severity and assess the health status of the aircraft gas turbine engine that is subject to component degrada-tions that are caused by fouling and erosion.The proposed hybrid framework involves integrating both supervised recur-rent neural networks and unsupervised self-organizing maps methodologies,where the former is developed to extract ef-fective features that can be associated with the engine health condition and the latter is constructed for fault severity modeling and tracking of each considered degradation mode.Advantages of our proposed methodology are that it ac-complishes fault identification and health monitoring objectives by only discovering inherent health information that are available in the system I/O data at each operating point.The effectiveness of our approach is validated and justified with engine data under various degradation modes in compressors and turbines.

A DEMATEL Approach Based on Fuzzy Sets for Evaluating Critical Factors of Gas Turbine in Marine Engineering

In power production,gas turbines are commonly used components that generate high amount of energy depending on size and weight.They function as integral parts of helicopters,aircrafts,trains,ships,electrical generators,and tanks.Notably,many researchers are focusing on the design,operation,and maintenance of gas turbines.The focal point of this paper is a DEMATEL approach based on fuzzy sets,with the attempt to use these fuzzy sets explicitly.Using this approach,the cause–effect diagram of gas turbine failures expressed in the literature is generated and aimed to create a perspective for operators.The results of the study show that,"connecting shaft has been broken between turbine and gear box"selected the most important cause factor and"sufficient pressure fuel does not come for fuel pump"is selected the most important effect factor,according to the experts.

Application of ultrasonic fatigue technology in very-high-cycle fatigue testing of aviation gas turbine engine blade materials:A review

The need for very-high-cycle fatigue(VHCF)testing up to 1010cycles of aviation gas turbine engine blade materials under combined mechanical loads and complex environments has encouraged the development of VHCF testing instrumentation and technology.This article begins with a comprehensive review of the existing available techniques that enable VHCF testing.Recent advances in ultrasonic fatigue testing(UFT)techniques are highlighted,containing their new capabilities and methods for single load,multiaxial load,variable amplitude fatigue,and combined cycle fatigue.New techniques for conducting UFT in high-temperature,humid environments,and corrosive environments are summarized.These developments in mechanical loading and environmental building techniques provide the possibility of laboratory construction for real service conditions of blade materials.New techniques that can be used for in situ monitoring of VHCF damage are summarized.Key issues in the UFT field are presented,and countermeasures are collated.Finally,the existing problems and future trends in the field are briefly described.

Auto-updating model-based control for thrust variation mitigation and acceleration performance enhancement of gas turbine aero-engines

Model-based control shows promising potential for engine performance improve-ment and future aero-propulsion requirements.In this paper,an auto-updating thrust variation mitigation(AuTVM)control approach using on-board model strategies is proposed for gas tur-bine aero-engines under in-service degradation effects,which aims at active thrust regulation and acceleration protection in a simultaneous way.The AuTVM control is integrated with an on-line block,based on a reliable on-board engine model,and an off-line part for the periodical update of control parameters via post-flight engine monitoring data.The core feature of the AuTVM control is a set of auto-updating loops within the on-line part,including thrust regu-lation loop,surge margin loop,turbine entry temperature loop,and the steady loop,whose con-trol parameters are periodically adjusted with increasingflight cycles.Meanwhile,an industrial sensor-based baseline controller and two tailored model-based controllers,i.e.,a thrust variation mitigation(TVM)controller withfixed gains and a self-enhancing active transient protection(SeATP)controller with pro-active transient protection and passive thrust control,are also developed as comparison bases.Numerical simulations for idle to full-power acceleration tests are carried on a validated aero-thermal turbofan engine model using publicly available degra-dation data.Simulation results demonstrate that both new engines and severely degraded en-gines regulated by the AuTVM controller show significant thrust response enhancement,compared to the baseline controller.Moreover,thrust variation at the maximum steady state of degraded engines,which exists within the SeATP controller and the baseline controller,is suppressed by the proposed AuTVM controller.Robustness analysis against degradation uncer-tainties and sensor accuracy confirms that the AuTVM controller owns a closer maximum steady-state thrust distribution to the desired value than those of the SeATP and the baseline controller while utilizing transient margins of controlled engines more effectively.Hence,the control performance of the AuTVM controller for in-service engines is guaranteed.

Analysis of Emissions Profiles of Hydraulic Fracturing Engine Technologies

Today, the oil and gas industry, and in particular hydraulic fracturing operations, have come under increasing pressure from regulators and the public to reduce emissions. As the industry evolves, oil and gas producers are in the position of evaluating alternative technologies which will support their objectives of reducing their overall emissions profile and carbon footprint. As a response, the deployment of technology and solutions to reduce emissions related to hydraulic fracturing applications has recently accelerated, creating various options to address these industry challenges. BJ Energy Solutions and West Virginia University have been working on the application and emissions characterization of various hydraulic fracturing technologies. A study was conducted to evaluate the efficiency and resultant emissions from various technologies, including natural gas reciprocating engines, diesel-natural gas dual-fuel engines, large (>24 MW) gas turbines, and direct drive turbines. The study involved the development of an emissions model with the purpose of estimating total emissions of carbon dioxide (CO2), nitrous oxide (N2O) and exhaust methane (CH4) slip, all Greenhouse Gases (GHGs), and converted to tons of CO2 equivalent emissions per day of operation. The model inputs are the required Hydraulic Horsepower (HHP) based on pumping rate and pressure for various shale play scenarios. The model calculates emissions from the TITAN, which is a direct-drive turbine model fielded by BJ, using data collected following U.S. Environmental Protection Agency (EPA) testing protocols. The model also calculates and compares other hydraulic fracturing technologies utilizing published Original Equipment Manufacturer (OEM) data. Relevant EPA-regulated criteria emissions of oxides of nitrogen (NOx), Carbon Monoxide (CO) and Particulate Matter (PM) are also reported. Modeling results demonstrated that in most cases, the TITAN gas turbine system has lower total GHG emissions than conventional diesel and other next-generation technologies, and also has lower criteria emissions. The benefits of the TITAN gas turbine system compared to the other technologies stems from significantly lower methane slip, and the high-power transfer efficiency resulting from directly connecting a turbine to a reciprocating pump, despite the comparatively lower thermal efficiency.

Numerical simulation method of surge experiments on gas turbine engines

In order to obtain the surge margin of an aero-engine during its operation,an engine surge experiment is required.A multi-dimensional simulation method for an aero-engine is established in this paper.The simulation of a surge experiment using high-pressure air-injection is then carried out on a turbo-shaft engine to obtain the surge boundary using this method.More specifically,firstly,a body-force model is employed to calculate the compressor performance owing to its capability of capturing the main three-dimensional features of compressor surge and avoiding excessive simulation time required by the traditional fully-three-dimensional Reynolds Averaged Navier-Stokes(RANS)method.Then,a one-dimensional model combining a lumped-parameter plenum model is used for the combustor to account for the propagation of pressure waves and the heat-release process,and a zero-dimensional throttle model is used to mimic the choking effect at the turbine nozzle.Finally,the air-injection system is modeled by imposing an injection boundary condition,which can be used conveniently in changing injection parameters.Based on the established method,the influences of different test parameters,such as the air-injection location,the pressure,the orifice size,the number of injection orifices,and the injection time duration on the surge characteristics and boundary are further studied,which offer effective guidance to optimize an actual experimental design.

Endwall aerodynamic losses from turbine components within gas turbine engines

A survey of research on aerodynamic loss investigations for turbine components of gas tuibine engines is presented.Experimental and numerically predicted results are presented from investigations undertaken over the past 65 plus years.Of particular interest are losses from the development of secondary flows from airfoil/endwall interactions.The most important of the airfoilAmdwall secondary flows are passage vortices,counter voitices,and corner vortices.The structure and development of these secondaiy flows are described as they affect aerodynamic perfonnance within and downstream of turbine passage flows in compressible,high speed flows with either subsonic or transonic Mach number distributions,as well as within low-speed,incompressible flows.Also discussed are methods of endwall contouring,and its consequences in regard to airfoil/endwall secondary flows.

A novel model-based multivariable framework for aircraft gas turbine engine limit protection control

Control technologies are innovated to satisfy increasingly complicated control demands of gas turbine engines.In terms of limit protection control,a novel model-based multivariable limit protection control method,which is achieved by adaptive command reconstruction and multiplecontrol loop selection and switch logic,is proposed in this paper to address the problem of balancing smaller thrust loss and safe operations by comparing with widely-used Min-Max logic.Five different combination modes of control loops,which represent the online control loop of last time instant and that of current time instant,is analyzed.Different command reconstructions are designed for these modes,which is based on static gain conversion of amplitude beyond limits by using an onboard model.The double-prediction based control loop selection and switch logic is developed to choose a control loop appropriately by comparing converted amplitude beyond limits regardless of one or more parameters tending to exceed limits.The proposed method is implemented in a twin-spool turbofan engine to achieve limit protection with direct thrust control,and the loss of thrust is improved by about 30% in comparison with the loss of thrust caused by Min-Max logic when limit protection control is activated,which demonstrates the effectiveness of the proposed method.

Performance of small-scale aero-derivative industrial gas turbines derived from helicopter engines

This paper considers comparative assessment of simple and advanced cycle small-scale aero-derivative industrial gas turbines derived from helicopter engines.More particularly,investigation was made of technical performance of the small-scale aero-derivative engine cycles based on existing and projected cycles for applications in industrial power generation,combined heat and power concept,rotating equipment driving,and/or allied processes.The investigation was done by carrying out preliminary design and performance simulation of a simple cycle(baseline)two-spool small-scale aero-derivative turboshaft engine model,and some advanced counterpart aero-derivative configurations.The advanced configurations consist of recuperated and intercooled/recuperated engine cycles of same nominal power rating of 1.567 MW.The baseline model was derived from the conversion of an existing helicopter engine model.In doing so,design point and off-design point performances of the engine models were established.In comparing their performances,it was observed that to a large extent,the advanced engine cycles showed superior performance in terms of thermal efficiency,and specific fuel consumption.In numerical terms,thermal efficiencies of recuperated engine cycle,and intercooled/recuperated engine cycles,over the simple cycle at DP increased by 13.5%,and 14.5%respectively,whereas specific fuel consumption of these cycles over simple cycle at DP decreased by 12.5%,and 13%respectively.This research relied on open access public literature for data.

Performance characterization of different configurations of gas turbine engines

This paper investigates the performance of different configurations of gas turbine engines.A full numerical model for the engine is built.This model takes into account the variations in specific heat and the effects of turbine cooling flow.A lso,the model considers the efficiencies of all component,effectiveness of heat exchangers and the pressure drop in relevant components.The model is employed to compare the engine performances in cases of employing intercooler,recuperation and reheat on a single spool gas turbine engine.A comparison is made between single-spool engine and two-spool engine with free power turbine.Also,the performance of the eng ine with inter-stage turbine burner is investigated and compared with engine employing the nominal reheat concept.The engine employing inter-stage turbine bumers produces superior improvements in both net work and efficiency over all other configurations.The effects of ignoring the variations on specific heat of gases and turbine cooling flow on engine performance are estimated.Ignoring the variation in specific heat can cause up to 30%difference in net specific work.The optimum locations of the intercooler and the reheat combustor are detemined using the numerical model of the engine.The maximum net specific work is obtained if the reheat combustor is placed at 40%of the expansion section.On the other hand,to get maximum efficiency the reheat combustor has to be placed at nearly 10%-20%of the expansion section.The optimum location of the intercooler is almost at 50%of the compression section for both maximum net specific work and efficiency.

Experimental study on impacts of fuel type on thermo-acoustic instability in a gas turbine model combustor

Effects of liquid fuel composition variations on characteristics of self-excited thermo-acoustic instabilities in a lean premixed,pre-vaporized gas turbine model combustor were experimentally studied.Test fuels included practical RP-3 jet fuel and its blending with iso-octane and n-dodecane,which were branched and linear alkanes respectively.Under the test conditions,dynamic pressure measurements indicated that the dominant instability frequency was highest for RP-3 flame,while RP-3/ndodecane flame exhibited the strongest instability strength.A further analysis showed that the instability frequency correlated well with the profiles of adiabatic flame temperature,and the strength of the instability highly depended on the ignition delay times of the fuels.Measurements of the flame structure and flow field with OH*chemiluminescence (CL) imaging and twodimensional particle image velocimetry (PIV) techniques indicated that changes in the fuel composition did not alter the unstable modes and general sequences of flame-flow structure oscillations.Further power spectra and proper orthogonal decomposition(POD) analysis suggested that axial oscillations along with precessing vortex core (PVC) induced helical motion predominated periodic flame structure and flow field oscillations.

涡轴发动机细长柔性转子动力特性及高速动平衡技术研究

建立涡轴发动机动力涡轮转子动力特性有限元分析模型,计算得到转子的前三阶临界转速值,在高速旋转试验器上完成全转速范围内的动力特性和高速动平衡试验,实测得到转子在额定工作转速范围内高速动平衡前后等状态下振动幅值随转速的变化曲线,研究结果表明:建立的计算模型反映了转子的真实情况,高速动平衡效果良好,解决了型号研制过程中的一项关键技术。

航空发动机数值仿真试验台建立中几个关键技术问题的讨论

本文主要讨论航空发动机数值仿真试验台建立中几个关键技术问题 :发动机数值仿真试验台建立的目标与基本要求 ;发动机三维多功能数值仿真数学模型、精度可靠性与并行算法 ;部件与发动机数值仿真试验台集成软件包 ;发动机仿真软件平台的框架 ;发动机三维数值仿真硬件系统。只有系统解决上述六个问题才能顺利建立这一仿真试验台。

涡轮发动机三维多场耦合数值仿真的数学模型

给出了用于发动机多场耦合数值仿真的统一的数学模型,这一模型由于采用了广义控制体系,在用来求解非定常气、热、弹耦合的变域差分问题时很方便. 讨论了统一的数学模型在不同场的演化,分析了多场耦合的4种类型及发动机多种层次的物理模型与数学模型.

局部进气条件下空气涡轮火箭发动机掺混燃烧研究

为揭示空气涡轮火箭发动机燃烧室中富燃燃气与空气在涡轮局部进气条件下的混合增强机制和燃烧反应机理,对富燃燃气与空气的湍流混合及燃烧过程进行了数值模拟,并结合试验结果定量分析了两类燃烧组织方案的掺混和燃烧效率。研究结果表明:涡轮局部进气条件下波瓣混流器强化掺混的主导因素是大尺度流向涡的对流型混合,涡轮局部进气对涡系的初始空间分布及涡量强度具有显著影响,其对下游掺混质量的影响与波瓣型面相关;肼分解燃气与空气的燃烧是一种分支链锁反应,其主要反应历程是氢气的氧化反应和氨气的分解,热混合效率可作为掺混燃烧效率预测的重要参考量。

基于一维模型的涡轮过渡流道优化设计

降低涡轮过渡流道的损失,缩短其长度,增大其径向偏移,是提高高涵道比涡扇发动机经济性和环境友好性的重要手段之一。目前研究基于所发展的一维涡轮过渡流道性能预测方法,提出了涡轮过渡流道设计的优化问题,融合单纯性优化方法,建立了一套有自主知识产权的涡轮过渡流道初步优化设计体系,用于涡轮过渡流道的快速设计。文中简要讨论了涡轮过渡流道的一维性能预测方法,给出了一维优化过程中所涉及诸多因素的确定,如几何的参数化、目标函数的确定,优化方法的选取、初步设计的建立以及优化的流程等。最后,以某实际涡轮过渡流道的优化设计为研究对象,对所发展的方法进行了验证,结果表明优化后的流道在总压损失和压升方面都较原型有了提高。通过目前的设计算例,充分确认了方法的鲁棒性和有效性。

涡轴发动机涡轮部件改型气动分析

首先对涡轴发动机涡轮部件的变工况特性进行了数值研究,计算结果与实验测量吻合良好。在此基础上,引入拓扑分析手段,对有、无中间承力支板的涡轮内部流场进行了详细分析。结果表明:支板对效率、流量等总体性能参数影响不大。最后对涡轮进行了改型优化,去掉支板,缩小尺寸并由动力涡轮静叶栅作为承力框架。计算表明:通过适当的气动优化,无支板的涡轮气动性能可以进一步提高。

大涵道比涡扇发动机涡轮过渡段的数值研究

采用CFD数值模拟方法对大涵道比涡扇发动机高压涡轮与低压涡轮之间的过渡流道进行了气动性能计算。通过与实验数据的对比,验证了CFD计算模型预测涡轮过渡段气动性能的可靠性以及计算精度。分析了径向偏移较大的涡轮过渡段流动特点和损失机理,为进一步提高过渡段气动性能提供研究方向。高性能大涵道比发动机中,涡轮过渡流道流动机理复杂,设计难度大,其可供挖掘潜力较大。

预冷却吸气式涡轮冲压膨胀循环发动机发展简介

分析了吸气式涡轮冲压膨胀循环(ATREX)发动机模型和工作原理,介绍了其主要部件的研究和发展情况,与常规发动机相比说明了其发展的优势,总结归纳了这种发动机技术特点和发展的关键技术。研究认为ATREX发动机采用液态氢预冷却方式,大大扩展了常规涡轮发动机的工作范围,可满足高速飞行器或两级入轨航空航天飞机第一级动力装置的要求。