Aero-engine Design Research Articles

Analysis of the effect of squeeze film damper on the bending-torsional coupling vibration characteristics of dual-rotor system

The study of the vibration reduction characteristics of SFD to the bending-torsional coupled vibration of the dual-rotor system can provide support for the structural design of aero-engines. Based on the overall eccentric disc model, the bending-torsional coupling dynamic equation of the dual-rotor system is established in this paper. The amplitude and nonlinear periodic characteristics of the dual-rotor system are obtained by combining Newmark-β method and Newton-Raphson method. The results show that the vibration reduction effect of SFD on bending-torsional coupling vibration is closely related to the speed. The vibration reduction effect of SFD is weak at the first-order critical speed range and below, and reduces the torsional amplitude by 70 % at the second-order critical speed range and above. SFD kept the bending-torsional coupling vibration of the dual-rotor system stable in a stable 5-period motion state.

Investigation on hybrid thermal features of aero- engines from combustor to turbine

The thermal interplay between the combustor and the turbine exerts remarkable effects on the performance of an aero-engine. The non-uniform distributions of the temperature and thermal radiation exist at the combustor outlet due to the operation characteristics of combustor and turbine. These uneven thermal features may cause hot spots on the blade surface, and even lead to material damage in regions such as the blade tip. Generally, the combustor and the turbine were separately treated and the thermal interconnection of each other were seldom involved except for the predesigned inlet conditions, which cannot be completely consistent with the actual turbine inlet conditions. To resolve such a problem, the present study proposes an integrated numerical simulation approach on the combustor-guide vane-rotating blade combined structure operating at engine-representative parameters (pressure ratios, temperature, Mach number and so on) with taking the effect of thermal radiation into account. The detailed analysis reveals that the distributions of temperature distributions of turbine inlet are closely related to the combustion gas swirl flow, dilution hole injection flow and combustor liner cooling air. The secondary flow at the combustor exit increases the mixing of the cold fluid and the hot streaks (HS) and reduces the temperature unevenness of the rotor blades. This elaborate investigation provides a deeper insight into the complex aero-thermal physics in turbine passage relevant to the aero-engine design and analysis.

Fatigue Reliability Analysis System for Key Components of Aero-Engine

Aero-engine is known as the heart of an aircraft. Fatigue is one of the main causes of aero-engine failure, therefore, it is essential to predict the fatigue life in the aero-engine design process. Due to the uncertainty of influencing factors, it is necessary to further analyze the fatigue reliability. First, the fatigue life should be predicted on the basic of finite element analysis. The steps include parametric modeling, stress-strain analysis, load spectrum acquisition, and selection of fatigue life prediction model. Then, the reliability estimation of fatigue life should be employed, including the statistical analysis of influencing factors, reliability analysis method, and reliability estimation of fatigue life. Taking turbine blade and test probe of aero-engine as the research objects, the fatigue reliability analysis system is developed based on the ABAQUS-MATLAB platform. Statistical analysis shows that fatigue life approximately obeys lognormal distribution, and the distribution parameters estimated by MCS and Kriging are coincide, while Kriging only needs dozens of training samples. Under different reliability indexes, the design fatigue life error between Kriging and MCS is less than 1%, which meets the accuracy requirements and can effectively guide the fault detection and maintenance of aero-engine.

A State Space Modeling Method for Aero-Engine Based on AFOS-ELM

State space models (SSMs) are important for multi-variable performance analysis and controller design of aero-engines. In order to solve the problems of the traditional state space modeling methods that rely on component-level models (CLMs) and cannot be carried out in real time, an aero-engine state space modeling method based on adaptive forgetting factor online sequential extreme learning machine (AFOS-ELM) is proposed in this paper. The structure of the extreme learning machine (ELM) is determined according to the form of the state space model, and the inverse-free ELM algorithm is used to automatically select the appropriate number of hidden nodes to improve the efficiency of offline initialization. The focus of the ELM on current operation performance is enhanced by the adaptive renewed forgetting factor, which reduces the impact of aero-engine history and deviated data on the current output and improves the accuracy of the model. Then, according to the analytical equation of the ELM model, the state space model of an aero-engine at each sampling time is obtained by using the partial derivative method. The simulation results based on engine test data show that the real-time performance and accuracy of the state space model established online in this paper can meet the needs of aero-engine control system requirement.

Aeroelastic Stability of Combined Plunge-Pitch Mode Shapes in a Linear Compressor Cascade

Modern aeroengine designs strive for peak specific fuel and thermal efficiency. To achieve these goals, engines have more highly loaded compressor stages, thinner aerofoils, and blended titanium integrated disks (blisks) to reduce weight. These configurations promote the occurrence of aeroelastic phenomena such as flutter. Two important parameters known to influence flutter stability are the reduced frequency and the ratio of plunge and pitch components in a combined flap mode shape. These are used as design criteria in the engine development process. However, the limit of these criteria is not fully understood. The following research aims to bridge the gap between semi-analytical models and modern compressors by systematically investigating the flutter stability of a linear compressor cascade. This paper introduces the plunge-to-pitch incidence ratio, which is defined as a function of reduced frequency and pitch axis setback for a first flap (1F) mode shape. Using numerical simulations, in addition to experimental validation, aerodynamic damping is computed for many modes to build stability maps. The results confirm the importance of these two parameters in compressor aeroelastic stability as well as demonstrate the significance of the plunge-to-pitch incidence ratio for predicting the flutter limit.

Research on disruptive design pre-identification in the preliminary design phase of aero engine flow pass multidisciplinary optimization

In aero engine design, determining whether the preliminary design will have disruptive effects on the detailed design is the key to multidisciplinary design optimization in the preliminary design stage. In order to adapt to the non-orthogonal parameter value range caused by the self-constrained parametric modeling method, a non-orthogonal space mapping method that maps the optimal Latin hypercube sampling points of the traditional orthogonal design space to the non-orthogonal design space is proposed. Based on the logical regression method in machine learning field, a kind of feasible domain boundary identification method is employed to identify whether the sample spatial response meets the relevant criteria. The method proposed in this paper is used to identify and analyze the key technologies of the high-pressure turbine mortise joint structure. It is found that the preliminary design of the aero engine may lead to the failure to obtain a mortise joint structure meeting the design requirements in the detailed design stage. The mortise joint structure needs to be pre-optimized in the preliminary design stage.

Leakage and wear characteristics of carbon seals for aero-engines

Experimental investigation has been performed to study the leakage and wear characteristics of carbon seal working at high circumferential speed. To expand the scope of application, two newly designed carbon seals were compared: #1 Carbon Seal (CS1) with the inner diameter of 136 mm and including 4 segments, #2 Carbon Seal (CS2) with the inner diameter of 212 mm and including 6 segments. Air leakage tests were firstly conducted in the Medium-speed Seal Test Rig. The pressure ratio changed from 1.04 to 2.02 with the rotating speed varying from 0 to 18300 r/min. Of paramount concern was the durability test, including 300 h running time accumulated by three different working conditions, which was separately implemented on each carbon seal. The morphology variation of the friction surface, wear and leakage were recorded. Results indicated that the leakage monotonously increases with the pressure ratio and decreases with the rotating speed. Comparing with CS1, more typical features exist on the friction surface of CS2, which are generated by more severe wear. Continually, leakage characteristics deteriorate. Furthermore, fitted formula has been educed for the life prediction of carbon seal, which could provide some supports for aero-engine design.

Impact of installation on a civil large turbofan exhaust system at idle descent conditions

Recent trends in civil aero-engine design aim at lowering specific thrust and improving propulsive efficiency by increasing the bypass ratio and therefore, usually also the fan diameter. The integration of these larger diameter engines with the airframe is critical to exhaust performance, and it is important to include these effects in engine performance analysis. The discharge coefficient of the bypass and core nozzles of a high-bypass ratio aero-engine at idle descent conditions is investigated numerically for an aero-engine in isolation and installed on an airframe. The discharge coefficients influence the engine operating conditions and turbomachinery re-matching at these off-design conditions. The maximum difference in the bypass nozzle discharge coefficient between the installed and isolated aero-engine across the descent phase is ≃1.6%. The differences in the core nozzle discharge coefficient between the installed and the reference isolated configuration are ≃43% and ≃−5.4% at the start and the end of the descent phase, respectively. The nozzle discharge coefficients depend on flight Mach number, incidence angle, and the nozzle pressure ratios of the fan and core nozzles. Multiple competing flow mechanisms govern the static pressure on the core nozzle base, which influences the core nozzle discharge coefficient. A novel reduced-order model is developed to estimate the core nozzle discharge coefficient for the installed configuration in idle descent conditions. This approach is based on the effective nozzle pressure ratio and can be implemented in engine performance simulations.

The Data-Driven Surrogate Model-Based Dynamic Design of Aeroengine Fan Systems

Abstract High-cycle fatigue failures of fan blade systems due to vibrational loads are of great concern in the design of aeroengines, where energy dissipation by the relative frictional motion in the dovetail joints provides the main damping to mitigate the vibrations. The performance of such a frictional damping can be enhanced by suitable coatings. However, the analysis and design of coated joint roots of gas turbine fan blades are computationally expensive due to strong contact friction nonlinearities and also complex physics involved in the dovetail. In this study, a data-driven surrogate model, known as the Nonlinear in Parameter AutoRegressive with eXegenous input (NP-ARX) model, is introduced to circumvent the difficulties in the analysis and design of fan systems. The NP-ARX model is a linear input–output model, where the model coefficients are nonlinear functions of the design parameters of interest, such that the Frequency Response Function (FRF) can be directly obtained and used in the system analysis and design. A simplified fan-bladed disc system is considered as the test case. The results show that using the data-driven surrogate model, an efficient and accurate design of aeroengine fan systems can be achieved. The approach is expected to be extended to solve the analysis and design problems of many other complex systems.

Single-Pulse Shock Tube Experimental and Kinetic Modeling Study on Pyrolysis of a Direct Coal Liquefaction-Derived Jet Fuel and Its Blends with the Traditional RP-3 Jet Fuel.

A basic understanding of the high-temperature pyrolysis process of jet fuels is not only valuable for the development of combustion kinetic models but also critical to the design of advanced aeroengines. The development and utilization of alternative jet fuels are of crucial importance in both military and civil aviation. A direct coal liquefaction (DCL) derived liquid fuel is an important alternative jet fuel, yet fundamental pyrolysis studies on this category of jet fuels are lacking. In the present work, high-temperature pyrolysis studies on a DCL-derived jet fuel and its blend with the traditional RP-3 jet fuel are carried out by using a single-pulse shock tube (SPST) facility. The SPST experiments are performed at averaged pressures of 5.0 and 10.0 bar in the temperature range around 900–1800 K for 0.05% fuel diluted by argon. Major intermediates are obtained and quantified using gas chromatography analysis. A flame-ionization detector and a thermal conductivity detector are used for species identification and quantification. Ethylene is the most abundant product for the two fuels in the pyrolysis process. Other important intermediates such as methane, ethane, propyne, acetylene, and 1,3-butadiene are also identified and quantified. The pyrolysis product distributions of the pure RP-3 jet fuel are also performed. Kinetic modeling is performed by using a modern detailed mechanism for the DCL-derived jet fuel and its blends with the RP-3 jet fuel. Rate-of-production analysis and sensitivity analysis are conducted to compare the differences of the chemical kinetics of the pyrolysis process of the two jet fuels. The present work is not only valuable for the validation and development of detailed combustion mechanisms for alternative jet fuels but also improves our understanding of the pyrolysis characteristics of alternative jet fuels.

Application and Development of Electronic Computers in Aero Engine Design and Manufacture

This paper has introduced the development of CAD and CAM at home and abroad, especially the application in aero engine; The development goal, implementation steps and assumption of computer system configuration of CAD / CAM technology for aero engine in China are put forward combined with the current situation of CAD / CAM technology in China.

A Comparison of Single-Entry and Multiple-Entry Casing Impingement Manifolds for Active Thermal Tip Clearance Control

In this paper, we compare using computational fluid dynamics the aero-thermal performance of two candidate casing manifolds for supplying an impingement-actuated active tip clearance control system for an aero-engine high-pressure turbine. The two geometries are (a) single-entry: an annular manifold fed at one circumferential location; (b) multiple-entry: a casing manifold split into four annular sectors, with each sector supplied separately from an annular ring main. Both the single-entry and multiple-entry systems analysed in this paper are idealised versions of active clearance control systems in current production engines. Aero-thermal performance is quantitatively assessed on the basis of the heat transfer coefficient distribution, driving temperature difference for heat transfer between the jet and casing wall and total pressure loss within the high-pressure turbine active clearance control system. We predict that the mean heat transfer coefficient (defined with respect to the inlet temperature and local wall temperature) of the single-entry active clearance control system is 77% greater than the multiple-entry system, primarily because the coolant in the multiple-entry case picks up approximately 40 K of temperature from the ring main walls, and secondarily because the average jet Reynolds number of impingement holes in the single-entry system is 1.2 times greater than in the multiple-entry system. The multiple-entry system exhibits many local hot and cold spots, depending on the position of the transfer boxes, while the single-entry case has a more predictable aero-thermal field across the system. The multiple-entry feed system uses an average of 20% of the total available pressure drop, while the feed system for the single-entry geometry uses only 2% of the total available pressure drop. From the aero-thermal results of this computational study, and in consideration of holistic aero-engine design factors, we conclude that a single-entry system is closer to an optimal solution than a multiple-entry system.

Novel Aero-Engine Multi-Disciplinary Preliminary Design Optimization Framework Accounting for Dynamic System Operation and Aircraft Mission Performance

This paper presents a modular, flexible, extendable and fast-computational framework that implements a multidisciplinary, varying fidelity, multi-system approach for the conceptual and preliminary design of novel aero-engines. In its current status, the framework includes modules for multi-point steady-state engine design, aerodynamic design, engine geometry and weight, aircraft mission analysis, Nitrogen Oxide (NOx) emissions, control system design and integrated controller-engine transient-performance analysis. All the modules have been developed in the same software environment, ensuring consistent and transparent modeling while facilitating code maintainability, extendibility and integration at modeling and simulation levels. Any simulation workflow can be defined by appropriately combining the relevant modules. Different types of analysis can be specified such as sensitivity, design of experiment and optimization. Any combination of engine parameters can be selected as design variables, and multi-disciplinary requirements and constraints at different operating points in the flight envelope can be specified. The framework implementation is exemplified through the optimization of an ultra-high bypass ratio geared turbofan engine with a variable area fan nozzle, for which specific aircraft requirements and technology limits apply. Although the optimum design resulted in double-digit fuel-burn benefits compared to current technology engines, it did not meet engine-response requirements, highlighting the need to include transient-performance assessments as early as possible in the preliminary engine design phase.

Systematic design and implementation of a semantic assistance system for aero-engine design and manufacturing

Systematic design and implementation of a semantic assistance system for aero-engine design and manufacturing

Systematic design and implementation of a semantic assistance system for aero-engine design and manufacturing

Systematic design and implementation of a semantic assistance system for aero-engine design and manufacturing

Multi-Objective Optimization of Cascade Blade Profile Based on Reinforcement Learning

The multi-objective optimization of compressor cascade rotor blade is important for aero engine design. Many conventional approaches are thus proposed; however, they lack a methodology for utilizing existing design data/experiences to guide actual design. Therefore, the conventional methods require and consume large computational resources due to their need for large numbers of stochastic cases for determining optimization direction in the design space of problem. This paper proposed a Reinforcement Learning method as a new approach for compressor blade multi-objective optimization. By using Deep Deterministic Policy Gradient (DDPG), the approach modifies the blade profile as an intelligent designer according to the design policy: it learns the design experience of cascade blade as accumulated knowledge from interaction with the computation-based environment; the design policy can thus be updated. The accumulated computational data is therefore transformed into design experience and policies, which are directly applied to the cascade optimization, and the good-performance profiles can be thus approached. In a case study provided in this paper, the proposed approach is applied on a blade profile, which is thus optimized in terms of total pressure loss and laminar flow area. Compared with the initial profile, the total pressure loss coefficient is reduced by 3.59%, and the relative laminar flow area at the suction surface is improved by 25.4%.

Spinning Wave Scattering From a Flow Pipe With Serrations

Abstract It is well known that trailing-edge serrations, which are also known as chevrons, are able to reduce the turbulent mixing noise from an aeroengine. The study of the associated control capability of the scattering of incident waves from the rotor–stator assembly is rare. To address this issue, a theoretical model is proposed to predict sound wave scattering from a cylindrical pipe with trailing-edge serrations in the presence of plug flows. The model incorporates Fourier series expansion into the Wiener–Hopf method and, therefore, is a natural extension of the previous aerofoil work by Huang (2017, “Theoretical Model of Acoustic Scattering From a Flat Plate With Serrations,” J. Fluid Mech., 819, pp. 228–257). The nature of the flow duct problem, however, leads to a much more complicated matrix kernel, and the associated factorization method is given in this article. The proposed model is validated by comparison with the numerical simulations at certain representative setups, which show the overall agreements to be satisfactory. The comparisons also show that the proposed model is so efficient that it can enable rapid predictions. A series of parametric studies are performed to study the two mechanisms behind the noise reduction of a serrated flow duct. One is the redistribution of acoustic energy to new higher cutoff modes. The other is the destructive interference due to multiple scattering from serrations. Overall, the proposed model should be helpful in offering deep physical insights and would be able to assist the aeroacoustic design and optimization of new low-noise aeroengines and flow duct systems after considering the tradeoff with aerodynamic impacts.

Artificial intelligence aided design of film cooling scheme on turbine guide vane

In recent years, artificial intelligence (AI) technologies have been widely applied in many different fields including in the design, maintenance, and control of aero-engines. The air-cooled turbine vane is one of the most complex components in aero-engine design. Therefore, it is interesting to adopt the existing AI technologies in the design of the cooling passages. Given that the application of AI relies on a large amount of data, the primary task of this paper is to realize massive simulation automation in order to generate data for machine learning. It includes the parameterized three-dimensional (3-D) geometrical modeling, automatic meshing and computational fluid dynamics (CFD) batch automatic simulation of different film cooling structures through customized developments of UG, ICEM and Fluent. It is demonstrated that the trained artificial neural network (ANN) can predict the cooling effectiveness on the external surface of the turbine vane. The results also show that the design of the ANN architecture and the hyper-parameters have an impact on the prediction precision of the trained model. Using this established method, a multi-output model is constructed based on which a simple tool can be developed. It is able to make an instantaneous prediction of the temperature distribution along the vane surface once the arrangements of the film holes are adjusted.

Usage of Data Provenance Models in Collaborative Multidisciplinary Aero-Engine Design

AbstractThe collaborative multidisciplinary design of aircraft engines is a complex and highly iterative process. An essential characteristic of this design process is the involvement of a large number of experts from different disciplines, as well as the usage of numerous tools and workflows. Large amounts of data are produced and need to be exchanged via a multitude of interfaces. Furthermore, the data undergo various transformations in the course of the design process. Understanding where a certain piece of data originates from and how it is connected to other datasets becomes therefore progressively essential. The purpose of this paper is to present a methodology to apply data provenance models in collaborative multidisciplinary aero-engine design, supported by an approach for data standardization and identification. Besides the methodology, the software implementation to support this approach is presented in detail, including automated capturing and storage of provenance data, as well as methods for data investigation. In addition, the presented methodology is evaluated by means of practical examples from the field of preliminary aero-engine design.

A conformable high temperature nitride coating for Ti alloys

There are many applications including aeroengine design where one would like to operate Ti or its alloys at higher temperatures, but the threat of oxidation or fire remains a longstanding challenge. Here, we have designed a bilayer nitride coating for Ti and its alloys produced by magnetron sputter deposition of a SiAlN coating (1.2 μm thick) along with a Mo interlayer. We have taken advantage of interdiffusion and inter-reaction at the interface during cyclic oxidation at 800 °C to form a layered nitride coating system comprising: a SiAlN top layer, a TiN0.26 and Ti5Si3 mixed phase interlayer, and a Ti-Mo solid solution. The novel TiN0.26 interlayer exhibits adaptive conformability via mechanical twinning, thereby accommodating the thermal mismatch strain between the coating and substrate. This, along with high adhesion, confers excellent thermal cycling life with no cracking, spallation and oxidation of the coating, evident after hundreds of hours of cyclic oxidation (>40 cycles) in air at 800 °C. This work provides a design pathway for a new family of coatings displaying excellent adhesion, adaptive conformability and superior environmental protection for Ti alloys at high temperature.