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• An integrated airframe-propulsion model is developed to simulate internal/external flow. • Effects of aerodynamic and geometric parameters on combustion and performance. • Underlying mechanisms lie in variations of combustion efficiency and total pressure. • Comprehensive understanding of the thrust-drag trade-off in engine and aircraft. The oblique detonation engine (ODE)-based aircraft employs the fast oblique detonation wave (ODW) combustion for airbreathing propulsion, delivering significant technical advantages in structure and efficiency under hypersonic flight conditions. However, achieving overall optimization by harmoniously integrating the ODE internal flows with the aircraft external flows remains a crucial yet unexplored aspect. Utilizing a multi-species reactive Reynolds-Averaged Navier-Stokes (RANS) model along with a detailed hydrogen-air combustion mechanism, an integrated airframe-propulsion ODE aircraft model is developed to investigate the internal/external flow and thrust-drag characteristics of an ODE aircraft operating at a flight Mach number of 10. Simulation results demonstrate the two key aerodynamic and geometric design parameters, i.e., inlet exit Mach numbers Ma 1 and combustor wedge angles θ 1 , exhibit inverted U-shaped curve effects on ODE specific impulse I sp. The optimal configurations with Ma 1 = 5 and θ 1 = 20° are identified as the most effective balance for ODE performance, achieving the maximum I sp of about 1200s. The underlying mechanisms are elucidated from the viewpoints of variations in both combustion efficiency and total pressure. Furthermore, a slightly different inlet exit Mach number of Ma 1 = 5.5 is found to achieve the maximum flight efficiency by reducing the flight drag, indicting the thrust-drag trade-off in ODE aircraft optimal design. This study provides a comprehensive understanding of the significance of balancing propulsion and aerodynamic performance for ODE aircraft to achieve superior overall performance, which holds the key to future engineering applications. [ABSTRACT FROM AUTHOR]

The aim of this research is to analyze the flow generated by the interaction between a simplified linear blade cascade adapted from an industrial aircraft engine demonstrator and an ingested vortex with the aerodynamic characteristics of a ground vortex observed in experiments. Due to the lack of availability of experimental and high-fidelity numerical studies of this phenomenon in the literature, a representative academic test case with a high-fidelity hybrid RANS/LES resolution of turbulence, namely the Zonal Detached Eddy Simulation (ZDES) technique, is performed. After the description of the computational setup and the numerical parameters used, the instantaneous topology of this complex flow is discussed, especially the interaction of the compressor cascade with the dynamics of the vortex. The simulation shows the chopping of the vortex, its destabilization and its dislocation into turbulent structures downstream the blade cascade. Turbulence is finely resolved in the blade wake and in the chopped vortex by ZDES technique at a moderate computational cost thanks to the RANS treatment of attached boundary layers with an applicative value of Reynolds number (5 ⋅ 10 6). A URANS simulation is also performed and allows a good representation of the averaged vortical structures inside the vortex close to the phase-averaged reference simulation. Spectral analyses are performed on unsteady velocity signals from probes located inside and outside the vortex chopped by the blades, highlighting more intense axial velocity fluctuations from turbulent structures inside the transformed vortex than in the wake of the blades, these structures being convected downstream in the direction of the Outlet Guide Vane (OGV). • Problem inspired by the real-life vortex ingestion by an aircraft turbofan engine. • Simplified test case of interaction between a linear blade cascade and a vortex. • High-fidelity simulation of the problem by the ZDES hybrid RANS/LES approach. • Turbulence finely resolved in the transformed vortex at applicative Reynolds number. • Circulation, enstrophy and spectral analysis of the turbulent transformed vortex. [ABSTRACT FROM AUTHOR]

With uncrewed aircraft systems (UASs) becoming more popular in recent years there is an increasing risk of a midair collision with aircraft for which there are currently no regulations in place. Bird impacts have been thoroughly studied for years and are considered soft body impacts. The bird impact research cannot be directly applied to UAS ingestions since they are hard body impacts. Some recent research has looked into understanding the differences between soft and hard body impacts on aircraft engines. However, this research used primarily unvalidated models for the harsh conditions of an ingestion event, which creates uncertainty in the results. This work intends to fill in this gap in understanding by developing a validated computational model of a UAS for engine ingestion simulations. Part I of this work presented the experimental test plan and experimental results of UAS critical components and full UAS being launched at test articles with representative features of a fan blade. This work discusses the computational modeling completed in LS-DYNA to simulate the physical tests and to update and validate the UAS models. [ABSTRACT FROM AUTHOR]

In reliability-based topology optimization (RBTO), both quantification and maximization of structure reliability considering uncertainty of various quantities are essential. In this study, A novel gradient online learning and prediction (GoLap) method is proposed to accelerate the double-looped RBTO and reduce total computational time by using a one-hidden-layer neural networks (NN) and gradient and non-gradient based reliability quantification. The total computational time saving is achieved by (i) skipping the finite element analysis (FEA) runs and sensitivity analyses and (ii) updating the design and random variables using the gradients predicted by the NN model in selected inner and outer loops. 2D and 3D numerical examples, including an aircraft engine pylon, are presented to demonstrate the total computational time saving of up to 84.4%, and the prediction accuracy of the reliability of optimum structures for the GoLap implemented RBTO methods. [ABSTRACT FROM AUTHOR]

As a crucial component of aircraft engine combustors, the external-mixing prefilming atomizer exhibits significant advantages in enhancing combustion efficiency and reducing pollutant emissions. However, the intricate relationship between spray characteristics, droplet breakup mechanisms, flow fields, and geometric parameters of the external-mixing prefilming atomizer remains poorly understood. This paper presents a novel external-mixing prefilming atomizer, aiming to address this knowledge gap. Additionally, particle image velocimetry (PIV), high-speed Schlieren photography, and particle image analysis (PIA) measuring systems were employed to gain insights into the spray characteristics and droplet breakup mechanism of the external-mixing prefilming atomizer under various swirl flow fields. The obtained results reveal the formation of two distinct regions, namely a low-density area and a high-density area. Notably, the number of droplets in the low-density area, located within 0–2 cm from the central axis, is significantly lower compared to the high-density area situated further away from the central axis. The breakup of droplets in the low-density region is primarily attributed to the inner swirl airflow, while the outer swirl airflow induces droplet breakup in the high-density region. As the Reynolds number of the incoming flow increases, the air-liquid ratio (ALR) and Weber number also increase, thereby enhancing the fuel breakup effect. The Schlieren results reveal a two-stage spray development process. The initial stage is characterized by the dominant momentum of the dense spray, with the spray tip penetration (STP) exhibiting a relatively slow increase. The aerodynamic force of airflow plays a crucial role in governing the breakup and transport of droplets, leading to a linear variation of STP over time. As the swirl number increases, there is an enhanced momentum exchange between droplets and airflow, resulting in accelerated spray penetration and droplet breakup. [ABSTRACT FROM AUTHOR]

Centrifugal compressors are key parts of medium and small aircraft engines, the flow instabilities of centrifugal compressors have become the bottleneck of the developments of engines. It is vital to investigate the dynamic mode characteristics of flow instabilities in centrifugal compressors for solving the issues. In this study, unsteady full annular simulations of the centrifugal compressor flow are firstly implemented. It is found that the flow instability of the compressor primarily arises in the impeller. To capture the dominant transient modes in process of flow instabilities and accurately describe their temporal evolutions, a modified dynamic mode decomposition method (DMD-TC) is adopted to analyze the dynamic mode characteristics, which was proposed in our previous study. The results show that perturbations in the impeller can be mainly classified into three categories: impeller-diffuser interactions, self-excited perturbations, and vortex-induced perturbations. In detail, the modes induced by the impeller-diffuser interactions and self-excited perturbations are intrinsic modes, and they reflect the inherent flow unsteadiness in the impeller. On the one hand, the vortex-induced perturbations couple with the self-excited ones, which affect the spatial structures of modes induced by self-excited perturbations. The coupling effects enhance with the reduced mass flow rate, and the "resonance" effect might be excited. As a result, the mode of 4 RF (rotor frequency) grows during rotating stall process (RSP), which is a typical characteristic of rotating stall in the impeller. On the other hand, the vortex-induced perturbations emerge as different characteristic modes, due to the variant vortex structures with the mass flow rate. During the RSP, the growth of low-frequency characteristic modes is another typical feature of rotating stall in the impeller. [ABSTRACT FROM AUTHOR]

The performance and reliability of aircraft engine are seriously affected by multiple failures induced by multi-physical loads. Multi-failure probabilistic design is an effective measure to estimate the multi-failure response traits and quantify the multi-failure risk for the improvement of component reliability. In this paper, we propose a neural network regression-distributed collaborative strategy (NNR-DCS) based on a developed two-step error control technique, to improve the efficiency and accuracy of multi-failure probabilistic analysis. We firstly mathematically model NNR-DCS and then introduce the corresponding multi-failure probabilistic framework. With respect to various failure modes such as deformation failure, stress failure and strain failure, the multi-failure probabilistic analysis of a turbine bladed disk is conducted to evaluate the proposed method. From this simulation, we gain the probabilistic distribution features, reliability degree and sensitivity degree of each failure mode and overall failure modes on turbine bladed disk, which provides a useful reference for improving the reliability and performance of aircraft engine. The comparison of methods (Monte Carlo method, RSM, DCRSM, DCFRM, NNR and NNR-DCS) shows that the proposed NNR-DCS holds high efficiency and accuracy for multi-failure probabilistic analysis. The efforts of this study offer an effective way for multi-failure evaluation from a probabilistic perspective and shed light on the multi-objective reliability-based design optimization of complex structures besides turbine bladed disk. [ABSTRACT FROM AUTHOR]

The objective of an aircraft engine control system is to provide required thrust as well as protection against the physical and operational limits. Model predictive control (MPC) technique is an attractive approach that incorporates input/output constraints in its optimization process to fulfill the control requirements of an aircraft engine. However, due to heavy computational burden of MPC, the real-time implementation of this algorithm is challenging and selection of MPC design parameters is crucial. This paper presents the design and hardware implementation of MPC algorithm as well as its HIL testing for turbofan engine control. In addition, a feedback correction technique is employed to compensate the effect of plant-model mismatch. For this purpose, a thermodynamic nonlinear engine model is firstly developed. A multivariable model predictive controller is then designed based on a discrete-time linearized state-space model where the horizons are obtained through a genetic algorithm optimization procedure. Moreover, this controller is implemented on an appropriate hardware taking the real-time aspects into account. Finally, an HIL platform is developed for testing the turbofan engine electronic control unit (ECU). For this purpose, a multi-level throttle command is applied to the ECU and the performance of the engine is evaluated. The results indicate that the controller satisfies all the engine constraints, and confirm the successful software and hardware implementation of the control algorithm in real-time. [ABSTRACT FROM AUTHOR]

Numerical results from a three-dimensional (3D) computational fluid dynamics (CFD) model of the Jetstream 31 aircraft in conditions of one engine inoperative are presented. The objective of this work is to analyse the performance of the Jetstream 31 aircraft and provide transient data using an unsteady Detached-Eddy Simulation (DES) CFD approach and a numerical propeller model to compare computational results with a single engine flight test experiment. The propeller modelling approach has been implemented through User-Defined Functions using C programming language to replicate the propeller effect. Different angles of attack and sideslip are studied, based on records from flight test data, both with unsteady (DES) and steady-state Reynolds-Averaged Navier-Stokes (RANS) models. An error analysis on the flight test data provides an error band from 2% to 16% among all cases, high values due to the lack of many data samples. Across the RANS approach, an average deviation of 6.9% and 3.8% for respectively lift and drag coefficients is achieved. By applying the DES turbulence modelling approach, a better lift prediction is achieved (5.4%) despite a slightly worse drag (4.5%). It has also been found that 80% of the numerical results are within the error band defined. A close agreement has been found within moment coefficients, with average percentage deviations from 3.3% to 7.0%. Overall, an analysis has been carried out in the present work, both on the flight test and computational sides to provide reliable numerical results of these aerodynamic properties. [ABSTRACT FROM AUTHOR]

When extreme learning machine (ELM) is used to cope with classification problems, the ±1 is commonly used to construct the label vector. Since ELM adopts the square loss function, this means that it tends to force the margins of all the training samples exactly equaling one from the perspective of margin learning theory, which is unreasonable to some extent. To overcome this hard margin flaw, in this paper a soft extreme learning machine (SELM) is proposed, which flexibly sets a soft target margin for each training sample. Through solving a series of regularized ELMs (RELMs), SELM can be computed efficiently. Based on SELM, an improved SELM (ISELM) is proposed to deal with imbalanced classification problems, which can keep the same computational efficiency as SELM via solving a series of weighted RELMs. From the experimental results on benchmark data sets, the effectiveness and feasibility of SELM and ISELM are confirmed. More importantly, when they are applied to fault detection of aircraft engine, they are promising to be developed as the candidate techniques for it, and ISELM is especially in favor. [ABSTRACT FROM AUTHOR]

A summary is given of the validation of an integrated aircraft environmental simulation software, with focus on acoustic and engine exhaust emissions. First, an experimental campaign was carried out to gather noise data of several commercial aircraft on departure from Manchester airport (ICAO code: EGCC). Field measurements were taken with microphones placed at a point 8,500 m (4.6 n-miles) from the estimated brake-release point. The departure measuring point was chosen due to the variability in flight trajectories. Comparisons between measurements and numerical predictions are shown for 12 commercial airplanes: A319-111, A320-214, A321-211, A330-243, A330-343, A350-900, A380-861, B737-800, B757-200, B787-800 and -900 versions, and finally the ERJ-195. The validated airplane models are used with a multi-objective, multi-parameter trajectory optimisation at the same airport to determine optimal departure profiles for airplanes in three weight classes: commuter jet, medium range transport, and long-range transport. It is demonstrated how different or concurrent cost functions (noise, fuel consumption, environmental price) lead to considerably different flight profiles and environmental emissions. [ABSTRACT FROM AUTHOR]

Abstract The lean blowout (LBO) limit is a crucial performance for aircraft engine combustors. It is essential to obtain the LBO limit during the design stage of the aircraft engine combustors. The semi-empirical correlation is an important tool for quick prediction of the LBO limits. Among all the semi-empirical correlations for the prediction of the LBO limits, Lefebvre's LBO model is widely used for the swirl stabilized combustors. The Flame Volume (FV) model was proposed based on Lefebvre's LBO model to accommodate the effects of the geometry of the flame tube on the LBO. Meanwhile, the multi-point lean direct injection (MPLDI) combustor whose geometry of the dome is different from the traditional combustors is a promising low NOx emission combustor. Up to now, there are few existing semi-empirical correlations to predict the LBO limit for the MPLDI combustors although the prediction of the LBO limit is critical for them. Based on the FV concept and new physics-based analysis, the FV-MP (Flame Volume for the Multi-Point) model is derived to predict the LBO limits for the MPLDI combustors. The FV-MP model could accommodate the effects of the fuel staging and recessed pilot stage, in addition to the operating conditions, on the LBO limits of the MPLDI combustors and achieve better prediction accuracy than both the FV and Lefebvre's LBO models within the range of corresponding validation experiments. Compared with Lefebvre's LBO model, the FV model could double the prediction accuracy. Compared with the FV model, the FV-MP model could further double the prediction accuracy. [ABSTRACT FROM AUTHOR]

Abstract Gaining a full understanding of the aerothermodynamic (AT) process in the high-pressure turbine of an aircraft engine is a complex task. The data in the literature on this topic are very limited, particularly that on tridimensional simulation. We report herein, for the first time, the designs, computer simulations, tridimensional calculations of the interactions of AT parameters under various operation conditions (takeoff, cruise) of an aircraft engine by using a refined mesh system containing approximately 6 million polyhedrons and more than 1 million interactions to solve equations in a high-pressure turbine (HPT). Our research provides, for the first time, an overview of high-resolution topographic images of the distribution of AT parameters in multi-row turbomachinery. Our calculations indicate that the relationship between these parameters was convoluted and depended on each operating condition. For example, our findings show that the thermal boundary conditions and rotor speeds strongly affected the flow temperatures (14%) and flow velocity (31%). On the other hand, the cooling system appeared not to affect AT parameters. The temperature nonuniformities in the turbine in the axial and radial directions were also observed. [ABSTRACT FROM AUTHOR]

Abstract Annular combustors are traditional, while hollow combustors without centerbodies are becoming popular for rotating detonation engines. Both schemes are studied experimentally in this paper to determine which combustor demonstrates a better detonation performance and pressure gain performance. The hydrogen and air sources are at room temperatures of 280–285 K. The annular and hollow combustors have the same scheme, although the centerbody is removed from the hollow combustor. The combustors are made optically accessible by embedding a piece of quartz glass in the outerbody. The hollow combustor channel has an outer diameter of 100 mm, and the detonation channel of the annular combustor has an inner diameter of 60 mm. Fuel and air are injected into the combustor from 150 cylindrical orifices with a diameter of 0.8 mm and a circular channel with a throat width of 1 mm, respectively. The results show that the hollow combustor demonstrates a better detonation performance and a higher pressure gain than the annular combustor. High-speed images show the rotating detonation is propagating with a lap time which is in good agreement with that obtained from pressure traces. [ABSTRACT FROM AUTHOR]

Abstract The paper proposes a method to minimize stage-by-stage initial unbalance in the aero engine assembly of multistage rotors based on the connective assembly model. The process of the mass eccentric deviation propagation in the assembly is analyzed. The initial unbalance of the final assembly is improved by properly selecting the assembly orientations of multistage rotors. A constrained nonlinear programming model is extracted from the optimal assembly strategy by choosing the initial unbalance as the objective function, and choosing the assembly orientations as the nonlinear constraint. The globally optimized solution of the constrained nonlinear programming model is solved using a genetic algorithm. The proposed method is used on an experimental set-up with the horizontal balancing machine and the effectiveness of the proposed method is verified by the assembly of the multistage rotors using the optimal assembly strategy. Compared to the direct assembly strategy, which the assembly orientations without consideration, the initial unbalance of final assembly using the optimal assembly strategy are reduced by 45.8%, 63.1% and 72.7% for two, three and four rotors assembly, respectively. The proposed method can improve the assembly quality of multistage rotors in the aero engine assembly and be used for assembly guidance, tolerance allocation, and so on, especially for the assembly with a large number of rotors. [ABSTRACT FROM AUTHOR]

Abstract This paper is concerned with state estimation approach to track aircraft engine gas-path health condition in an advanced distributed architecture. The sensor measurements are divided into several subsets by installation position along gas path, and they are integrated to estimate engine health state changes with sensor fusion uncertainty. The uncertain sensor fusion is characterized by time delay and packet dropout in the fusion behavior of sensor measurements, and the delay steps occur randomly. A novel distributed extended Kalman filter with the data buffer bank (DEKF) is developed, and self-tuning buffer strategy of recursive fusion estimation is combined to the DEKF to form the self-tuning DEKF (SDEKF) algorithm for improving state estimation performance. The lengths of data buffer bank related to the local filters of SDEKF are different, and they are independently adaptive to the information loss level and local estimation accuracy. Local states are calculated using the measurements collected at the latest steps in self-tuning buffer banks, and then sent to master filter to yield global state and covariance by fusion estimation. The contribution of this study is to propose a novel EKF algorithm for state estimation in the distributed framework with sensor fusion uncertainty, and it achieves better trade-off between the estimation accuracy and computational efforts. The systematical comparisons of basic EKF, constant buffer DEKF and SDEKF algorithms are carried out for aircraft engine gas-path health estimation with sensor fusion uncertainty. The simulation results show the superiority of the SDEKF, and it confirms our viewpoints in this paper. [ABSTRACT FROM AUTHOR]

Abstract The unsteady distorted flow fields generated within convoluted intakes can have a detrimental effect on the stability of an aero-engine. The frequency signature in the distorted flow field is of key importance to the engine's response. In this work, time-resolved particle image velocimetry is used to obtain the three-component velocity field at the outlet plane of two S-duct intake configurations for a range of inlet Mach numbers. Proper orthogonal decomposition of the time-resolved velocity data allows the identification of the main frequencies and coherent structures in the flow. The most energetic unsteady structures comprise an in-plane vortex switching mode, associated with a lateral oscillation of the main loss region, and a vertical oscillation of the main loss region. The switching structure occurs at a frequency of St = 0.42 and 0.32 for the high and low offset ducts, respectively. The vertical perturbation is associated with a more broadband spectrum between approximately St = 0.6 – 1.0 and St = 0.26 – 1.0 for the high and low offset configurations, respectively. The determined frequencies for these main unsteady flow structures are within the range, which is expected to be detrimental to the operating stability of an aero-engine. The results provide a novel, time-resolved dataset of synchronous velocity measurements of high spatial resolution that enables analysis of the unsteady flows at the exit of complex aero-engine intakes. [ABSTRACT FROM AUTHOR]

Abstract Most response surface methods typically work on isotropically sampled data to predict a single variable and fitted with the aim of minimizing overall error. This study develops a metamodel for application in preliminary design of aircraft engine nacelles which is fitted to full-factorial data on two of the eight independent variables, and a Latin hypercube sampling on the other six. The specific set of accuracy requirements for the key nacelle aerodynamic performance metrics demand faithful reproduction of parts of the data to allow accurate prediction of gradients of the dependent variable, but permit less accuracy on other parts. The model is used to predict not just the independent variable but also its derivatives, and the Mach number, an independent variable, at which a certain condition is met. A simple Gaussian process model is shown to be unsuitable for this task. The new response surface method meets the requirements by normalizing the input data to exploit self-similarities in the data. It then decomposes the input data to interpolate orthogonal aerodynamic properties of nacelles independently of each other, and uses a set of filters and transformations to focus accuracy on predictions at relevant operating conditions. The new method meets all the requirements and presents a marked improvement over published preliminary nacelle design methods. [ABSTRACT FROM AUTHOR]

Abstract Imbalanced problems often occur when the size of majority class is bigger than that of the minority one. The Least squares support vector machine (LSSVM) is an effective method for solving classification problem on balanced datasets. However, LSSVM has bad performance on minority class facing with class imbalance learning for the classification boundary skewing toward the majority class. In order to overcome the drawback, LSSVM for class imbalance learning (LSSVM-CIL) is proposed. LSSVM-CIL utilizes two different regularization parameters C + and C − that evaluate different misclassification costs. Furthermore, a method of combining reduced technique and recursive strategy is proposed to reduce the size of support vectors and retain representative samples. In addition, decomposition of the matrices via Cholesky factorization is employed as a solution to enhance the computational stability. Furthermore, the effectiveness of the two algorithms presented in this paper is confirmed with experimental results on various real-world imbalanced datasets. Fault detection of aircraft engine can be regarded as a CIL problem and has the demand for the real time. Finally, experiments on aircraft engine indicate that the two algorithms can be selected as candidate techniques for fault detection of aircraft engine. [ABSTRACT FROM AUTHOR]

This paper is concerned with nonlinear Kalman filtering approach to aircraft engine gas path analysis with measurement uncertainty. The uncertain measurements are characterized by time delay and packet dropout. The delay step of physical parameters occurs randomly, and its probability is regulated by a set of uncorrelated variables following Poisson distribution and uniform distribution. Packet dropout is caused as the data are not collected in time or data buffer overflows. The novel nonlinear Kalman filters (KFs) are developed using a multistep recursive estimation strategy with self-tuning buffer in the presence of gas path measurement uncertainty. The data buffers are introduced in the state estimator, the length of which is adaptive to the information loss level. The algorithms run recursively using the new arrival data and buffer position information. With a more effective arrangement of the collected measurements in real time, the better estimation accuracy of gas path health status is expected. Simulations involving abrupt fault and degradation datasets of aircraft engines were carried out to numerically evaluate and compare the performance of the improved nonlinear KFs with their existing KFs in the context of health estimation with time delay and packet dropout. The test results demonstrate that the proposed methodology not only reduces the computational time but also obtains a satisfactory accuracy for state estimation in the cases of engine gas path measurement uncertainty. [ABSTRACT FROM AUTHOR]

The Blended Wing Body (BWB) is a type of innovative aircraft, based on the flying wing concept. For this aircraft, the literature has reported performance improvements compared to conventional aircraft: economy of fuel, reduction of the weight of the structure, increased payload capacity and less impact on the environment. However, most BWB studies have focused on large aircraft and it is not sure whether the gains are the same for smaller aircraft. The main objective of this study is to perform the conceptual design of a 200 passengers BWB and compare its performance against an equivalent conventional A320 aircraft in terms of payload and range. Moreover, an emphasis will be placed on obtaining a stable aircraft, with the analysis of static and dynamic stability over its flight envelope. This kind of aircraft has a lack of stability due to the absence of vertical tail. Most studies of stability were already realized on reduced size models of BWB, but there is no study on a 200 passengers BWB. The design of the BWB was realized with the platform called Computerized Environment for Aircraft Synthesis and Integrated Optimization Methods (CEASIOM). The airplane, the engines and the control surfaces were obtained in the geometrical module AcBuilder. This design platform, suitable for conventional aircraft design, has been modified and additional tools have been integrated in order to achieve the aerodynamic, performance and stability analysis of the BWB aircraft. The aerodynamic coefficients are calculated from Tornado program. The BWB flight envelope was created based on aeronautical data of A320 aircraft. From this flight envelope, we have got back several thousand possible points of flight. The static and dynamic stability was studied using the longitudinal and lateral matrices of stability and the Flying Qualities Requirements for every flight point. [ABSTRACT FROM AUTHOR]

The inlet is a critical component of the air-breathing system of an aircraft engine and an appropriate inlet design can address the issue of aerodynamic contradiction between low and high Mach numbers, restricting aircraft flight over a wide speed range. Therefore, in this study, an axisymmetric variable-geometry inlet with a Mach 5 design point was designed. A multi-point multi-objective optimization design, based on the Kriging surrogate model and non-dominated sorting genetic algorithm-II algorithm, was performed. Wind tunnel tests and numerical simulations were conducted to investigate the inlet performance and flow field details. The steady-state pressure distribution was recorded, and shadowgraph images were obtained simultaneously to comprehensively evaluate the inlet performance. The experimental and numerical results indicate that the design and optimization methods of the inlet are effective. The inlet operated appropriately and stably over a wide range of speeds, and the total pressure recovery (TPR) coefficient reaches 0.421 at the design point. The optimization results show that a 21.36% gain in the TPR coefficient at Ma = 3.0 was achieved at the cost of a 2.44% loss of total pressure recovery at Ma = 5.0. It is also beneficial in terms of the total pressure distortion at Ma = 5.0, which is reduced by 6.54%. This indicates that the inlet design and optimization framework developed in this study is promising for wide use in practical engineering applications. [ABSTRACT FROM AUTHOR]

Current aircraft engine control logic employs a Min–Max selector structure. This structure must provide desired thrust and prevent the engine from exceeding any safety or operational limits. Careful analysis shows that there is no assurance for traditional Min–Max algorithm with linear compensators to protect output limits in transient regime, while limit violation can cause serious damages and even lead to loss of engine. In this paper, according to analysis outcomes, a strategy is presented to design linear regulators of Min–Max selector control to improve transient limit protection. For this objective, the isolated limitation loops are designed overshoot/undershoot-free to reduce the possibility of limit violation during activeness of the individual limit regulator. Despite the success of this approach, it is observed that some outputs influenced by engine acceleration or deceleration exceed their limits when other loop regulators are active. To overcome this problem, the limit values of these outputs are considered in controller design process of other loops using state feedback method and a set of linear matrix inequalities (LMIs). Simulation results show that the proposed approach effectively decreases the possibility of transient limit violation and can improve Min–Max limit protection in aircraft engine control. [ABSTRACT FROM AUTHOR]

Against the background of the probabilistic analysis for High Pressure Turbine (HPT) Blade-tip Radial Running Clearance (BRRC) to achieve the high-performance and high-reliability of aeroengine, Distributed Collaborative Extremum Response Surface Method (DCERSM) was proposed for the dynamic probabilistic analysis of complex turbomachinery on the foundation of quadratic polynomials response surface model. On the basis of deeply investigating Extremum Response Surface Method (ERSM), the mathematical model of DCERSM was established based on quadratic polynomial function. As illustrated in BRRC transient probabilistic analysis with multiple components and multi-physics fields based on DCERSM, blade-tip radial static clearance δ = 1.82 mm is advisable synthetically considering the reliability and working efficiency of gas turbine. The reliability, distribution characteristics and failure probability of BRRC are obtained. Besides, rotational speed ω and gas temperature T are the most important factors and expansivity coefficients and surface coefficients of heat transfer show also important influence on BRRC variation. Through the comparison of three methods (DCERSM, ERSM, Monte Carlo method), it is demonstrated that DCERSM reshapes the possibility of complex turbomachinery probabilistic analysis and improves computing efficiency while preserving the accuracy. DCERSM offers a useful insight for BRRC dynamic reliability design and optimization with multi-object and multi-discipline. The efforts of this study also enrich the theory and method of mechanical reliability design. [ABSTRACT FROM AUTHOR]

A variable cycle engine (VCE) is a potential power system for future advanced military and civil aircraft. The involvement of numerous adjustable parameters enables a VCE to meet multiple flight requirements simultaneously; however, it also increases the difficulty of engine design. Global algorithms can be adopted to maximally exploit the performance potential of a VCE while satisfying constraints. However, the optimization of a VCE is a highly constrained and nonlinear problem with high computational cost. In this article, a novel acceleration technique for the global optimization of VCEs is presented. The initial guesses for the iterations of the VCE model were optimized along with the optimization variables to reduce the model call number (MCN). The differential evolution (DE) algorithm was chosen to demonstrate and verify the effectiveness of this acceleration technique. The results show that the technique can significantly reduce the MCN under most conditions without affecting the optimization process. The more rapidly the DE converges, the more significantly the MCN is reduced. In some cases, this technique can help improve the optimization results. Moreover, the technique can be applied to various DE improvement strategies. This technique can also be employed in the performance optimization of other aero engines or similar systems. [ABSTRACT FROM AUTHOR]

Flight Environment Simulation System (FESS) is an important ground test facility, which can simulate the flight environment to test the performance of the aircraft engine. Focusing on the anti-windup control of FESS in the presence of actuator (magnitude and rate) saturations, an anti-windup design method with local sector and H 2 / H ∞ optimization is proposed. Firstly, the saturation-type static nonlinearities are separated from the dynamic nonlinear model of the actuator, and then an anti-windup architecture with static compensator is constructed. Subsequently, the integrated H 2 performance and H ∞ performance are used to deal with the saturation nonlinearities to reduce the system performance degradation caused by the magnitude and rate saturations. Meanwhile, the internal stability of the closed-loop system is guaranteed via local sector condition. Finally, an optimization problem based on Linear Matrix Inequalities (LMI) is developed for anti-windup design, in which internal stability, H 2 performance and H ∞ performance are considered. The designed anti-windup controller is applied to FESS, and the numerical simulation results verify the effectiveness of the proposed method. Furthermore, some key issues in anti-windup design are discussed. [ABSTRACT FROM AUTHOR]

Based on the available experimental data, dependences of true stress on true strain under dynamic loading for titanium alloys VT6, OT4, OT4-0 are constructed, and parameters for Johnson–Cook model are determined. Simulation results for deformation and fracture of fan case of aircraft engine at high-speed collision with the fan blade simulator in LS-DYNA software package are presented. [ABSTRACT FROM AUTHOR]

The latticework duct plays an important role in turbine blade cooling for aircraft engines. The present numerical research provided information about the thermal performance in latticework ducts with different endwall shapes. Each latticework duct analyzed in this study had a crossing angle of 90° and 11 sub channels on both the two sides of the endwall. A numerical investigation was conducted at five Reynolds numbers (11,000 to 55,000) and six endwall shapes. The local Nu number and flow structures were analyzed in detail. For a rectangular latticework duct, a jet region generated considerable Nusselt numbers, and the helical flow leaded to a uniform Nu number. Moreover, the turning and jet effects caused a high friction factor. The results also showed that an appropriate endwall shape could increase the thermal performance in a latticework duct. The different endwall shapes mainly altered the flow structure nearby the sidewall of the latticework, especially for the spiral flow. Moreover, the endwall shape also influenced the heat transfer and flow behavior. This phenomenon occurred because the different endwall shapes increased the turbulence of the cooling air to different extents. [ABSTRACT FROM AUTHOR]

Data-driven fault diagnosis for aircraft engine has recently been widely studied, most of which utilized traditional machine learning tools. However, these approaches often assumed that the data used for training and testing come from the identical distribution, which is unpractical in practice, because even for the same type of engine, there may be some discrepancies in the data collected between different engine individuals. This idealized assumption may confine these promising data-driven technologies to well-design experimental environments instead of applying them to practical applications. Meanwhile, the research on data-driven engine fault diagnosis methods based on transfer learning is relatively limited. To overcome this issue, this paper proposes a methodology using unilateral alignment strategy to address the issue of cross domain engine gas path fault diagnosis, where the training data extracted from source domain and the testing data extracted from target domain are assumed to come from different distributions. More importantly, the proposed methodology is capable of maintaining the inter-class relationship of source domain, instead of aligning the discrepancies for source and target domain by mapping both of their feature spaces into an unknown intermediate space, because the label space between two domains may be different, and the label set of target domain might be only a subset of that of source domain. Meanwhile, both the marginal and conditional distribution discrepancies are considered. Finally, the proposed approaches are evaluated by extensive experiments on engine cross domain fault diagnosis, including hybrid transfer experiments, complete transfer experiments and missing class experiments, and the overall results demonstrate the feasibility of the proposed methodology. [ABSTRACT FROM AUTHOR]

The development of advanced control systems is motivated to achieve the complicated aircraft engine control, so a novel semi-alternative optimization strategy based model predictive control is proposed. In this proposed controller, a nonlinear state-space model is generated as the predictive model on the basis of a baseline engine model that estimates the current operating state of engine with an extended Kalman filter every sampling instant, and a quadratic constrained problem rather than a higher order nonlinear problem is constructed based on this model. Moreover, a novel control sequence to be optimized based on semi-alternative optimization strategy is proposed, which includes two parts. The first part represents that all the control variables of the engine are optimized for the current sampling instant while the second part represents that different control variables are optimized alternatively for different future sampling instants over the control horizon. The proposed model predictive control method is implemented to a twin-spool turbofan engine. Simulation results demonstrate not only that the proposed method can achieve a smaller control error than the standard model predictive control algorithm does, but also that the proposed controller is more time-saving than the standard model predictive control, which can save up to about 61% optimization time when control horizon increases to nine. [ABSTRACT FROM AUTHOR]

A turbine is the essential component of aircraft engines. The empirical models of turbine performance, including the characteristics of isentropic efficiency and mass flow rate, are important for the control- and diagnostics-oriented dynamic simulation of aircraft engines. They are also useful for the energy analysis and performance simulation of aircraft engines. A number of empirical models of turbine isentropic efficiency and mass flow rate are available. Most of them, however, are for turbocharger, automotive and marine engine turbines, and air-cycle refrigeration turbines used in aircraft environmental control systems, and those for aircraft engine turbines are scarce. This paper presents a comprehensive review on empirical models of turbine isentropic efficiency and mass flow rate, and their applicability to aircraft engine turbines is evaluated against the turbine performance maps of five different aircraft engine turbines. It is found that the accuracy of the models varies with turbines. The results show that it is very likely feasible to use empirical models developed for other types of turbines for the energy analysis, performance simulation, and control- and diagnostics-oriented dynamic simulation of an aircraft engine turbine if proper models are chosen, and that it remains a research topic to develop a more accurate empirical model for turbine performance of aircraft engines. [ABSTRACT FROM AUTHOR]

Flow fields around a supersonic aircraft with installed engine nacelles are simulated to study engine installation effects on the aerodynamic performance and ground sonic boom signature of the aircraft. The engine nacelle is a dual-stream nacelle composed of a relaxed isentropic compression supersonic inlet and an annular bypass duct around the core engine. Two different bypass duct shapes are considered here; one is a full three-dimensional configuration consisting of struts and a gearbox fairing, and the other is a simplified axisymmetric duct with equivalent area distribution to the original full three-dimensional configuration. The full 3-D and axisymmetric ducts are found to produce almost identical external shock patterns as expected, which confirms our previous vane shape design results. Representative airframe-propulsion interference cases are suggested, and configurations with different engine locations are simulated and identified according to the interference cases. [ABSTRACT FROM AUTHOR]

This paper addresses the extended Kalman filtering approach to aircraft engine gas path analysis with control loop parameter uncertainty, which results from faulty sensors and actuators. A novel approach of gas path performance anomaly detection is proposed during the life course, which relies on combinations of recovered and actual measurements. A couple of adaptive onboard engine models (AOBEMs) are designed independently tuned to the real engine with state buffers, and sensor and actuator fault isolation (SAFI) logic is developed using hybrid matrices of the AOBEM couple. Neighborhood measurements with prior thermodynamic analysis reconstruct the faulty sensor signal once the control loop fault is recognized by the SAFI logic. The contribution of this paper is to develop engine performance monitoring against uncertain parameters over time in a control loop, and the gas path fault false alarm rate is independent of sensor and actuator faults. The tests involve abrupt fault and degradation datasets of the aircraft engine life cycle in the flight envelope, and the results demonstrate the superiority of the proposed methodology in the uncertain engine control loop. [ABSTRACT FROM AUTHOR]

The assembly accuracy of aero-engine rotor parts is strictly required in the docking processes. The application of a multi-degrees-of-freedom (multi-DOF) numerical control (NC) motion platform could potentially replace traditional manual docking operations for its better performance. A key issue is ensuring the assembly accuracy qualified by using the new mechanism. In this paper, a systematic study is carried out on modeling and simulation of the docking assembly accuracy of aero-engine rotors using a specially designed novel multi-DOF NC motion platform. A simulation algorithm based on the Monte Carlo method is proposed, to simulate large quantities of rotor assemblies under different docking conditions. To construct the algorithm, new general models are proposed to indicate the geometric errors and assembly accuracy in typical rotor parts. A motion error model of a novel multi-DOF NC motion platform is established using the differential transformation method. Then, simulation results are analyzed to guide the precision design of the six driving systems in the motion platform, and verified by experiments. The results show that the optimized precision setting of the motion platform achieves the qualified rates of assembly accuracy exceeding 92.0% under most docking conditions. For the six items representing assembly accuracy, the deviation rates between the simulation and experimental results are less than 23.0%. The modeling and simulation study has referential significance for both theory and practice. [ABSTRACT FROM AUTHOR]

Recently, the growing amount of unmanned aerial vehicles (UAVs) has brought a huge threat to the safety management of manned aircraft operation, in which the UAV airborne collision is an incident that would lead to serious damage to the manned aircraft and will affect its operational safety significantly. In the present paper, the bird strike data over the year of 1990∼2019 is analyzed, which demonstrates that the engine is the most vulnerable component under bird strike, and the most severe hazard would happen during the flight phases of take-off, climb and approach. The dynamic response of UAV airborne collision with the manned aircraft engine is simulated based on the combination of FEM (Finite Element Method) and CFD (Computational Fluid Dynamics) simulations. Not only the damage of fan blades but also the thrust loss of the engine core caused by the damage in the compressor core is taken into account. The damage severity level of the engine under UAV airborne collision is studied by considering different collision configurations, different collision positions and different flight phases. Both the damage of fan blades and the percentage of thrust loss are considered to reflect the influence of UAV airborne collision on the aircraft operation. It is expected that this study can be used to guide the airborne safety assessment of UAV airborne collision. [ABSTRACT FROM AUTHOR]

With the development of modern aero-engine, the compressor is required to achieve higher performance such as transonic operation, high stage pressure ratio and efficiency. At present, the traditional blade design methods have the disadvantage of highly relying on designer's experience and heavy computational burden. In order to reduce design variables used to define blade geometry, the present work introduces the Free-Form Deformation technique. It can achieve high degree-of-freedom deformation and update mesh and geometry simultaneously. To further alleviate computational cost, the Support Vector Regression surrogate model is adopted to replace the time-consuming numerical simulation. It performs well in small sample learning problems. On this basis, a surrogate-based optimization design framework is established by combining the Advanced Latin hypercube sample and NSGA-II multi-objective genetic algorithm. Then an aerodynamic optimization of the transonic NASA Rotor 37 is conducted as a validation. The results show that the pressure ratio and isentropic efficiency are increased by 4.2% and 2.5%, respectively. The optimized shape sweeps forward with a slight increase in twist angle, and shock loss is reduced effectively. Compared with traditional optimization method, the proposed framework can improve the optimization efficiency by decreasing design variables and training samples. [ABSTRACT FROM AUTHOR]

To avoid the use of specific conversion coefficients with high expense and unacceptable prediction accuracy, a probability-based prediction method is proposed by considering probabilistic feature parameters, to predict the service safety life (SSL) of aeroengine turbine blades. The direct correlation between laboratory remaining life (LRL) and SSL was firstly established by considering probabilistic feature parameters. By conducting Combined high and low Cycle Fatigue (CCF) tests of turbine blades, the effectiveness of the developed method was validated based on the failure event. The proposed method was further verified by predicting the SSL of treated blades with certain operation time. In respect of the studies, it is illustrated that (1) the SSL of turbine blade can be reasonably reflected by the LRL in respect of probabilistic feature parameters; (2) the prediction errors of the raw and treated blades are 2.2% and 12.7%, respectively, indicating that the developed probability-based prediction method has acceptable prediction precision and is an effective method in the SSL prediction of aeroengine turbine blades; (3) the developed method needs less samples than the specific conversion coefficients method, indicating that the SSL prediction of turbine blade needs fewer time and costs. The efforts of this study provide a promising approach for the SSL prediction of turbine blades, offer a useful guidance for the service life management of aeroengine turbine blades to reduce the cost of expense and time and enhance the safety of aeroengine operation. [ABSTRACT FROM AUTHOR]

In order to increase propulsive efficiency, and hence reduce fuel consumption, future aero-engines are expected to operate with higher bypass ratios and larger fan diameters relative to current in-service engines. As such, propulsion systems are likely to be more closely-coupled with the airframe which is expected to accentuate detrimental aerodynamic interference effects between the engine and airframe. It is therefore crucial that the design of future aero-engine exhaust systems is considered as part of an engine-airframe configuration in order to ensure that the expected benefits of high BPR engines are realised. This work presents the aerodynamic performance and evaluation of a set of novel exhaust systems within complete engine-airframe configurations. The introduction of non-axisymmetric exhaust systems was shown to mitigate the aerodynamic penalties associated with closely-coupled propulsion systems at cruise conditions. Relative to an axisymmetric baseline configuration, the introduction of non-axisymmetric bypass and core nozzles were found to increase the net vehicle force of the engine-airframe configuration by 0.8% and 0.6% respectively. As a result of this work, it can be concluded that non-axisymmetric exhaust systems represent a viable method for reducing aircraft cruise fuel burn. [ABSTRACT FROM AUTHOR]

In this paper, a novel multiple-fault diagnosis (MFD) scheme using radial basis function neural network (RBFNN)-based observers is presented for a spacecraft attitude control system (ACS) in the presence of external disturbances and nonlinear uncertainties. Based on dynamic and kinematic models, robust fault detection observers (FDOs) are designed to detect the simultaneous occurrence of actuator, gyro, and star sensor faults. Then, a series of RBFNN-based fault isolation observers (FIOs) are designed to decouple the faults of different components completely. This complete decoupling will guarantee that the diagnosis result of one component is not affected by the faults of other components; thus, multiple faults can be diagnosed simultaneously. To improve the accuracy of fault detection and reconstruction, disturbance compensation observers (DCOs) based on the RBFNN are also designed to compensate for the external disturbances. It is worth noting that the developed fault diagnosis scheme can be used to detect and isolate small faults. Finally, simulation results are presented to show the effectiveness and feasibility of the proposed method. [ABSTRACT FROM AUTHOR]

Commercial aircraft operating in some regions of the world are subject to harsh atmospheric conditions that can affect the efficiency and integrity of modern aero-engines. Conditions that are investigated in this contribution are dusty atmospheres following sand storms. Specifically, the dust ingested by the engines of an Airbus A380-841 (Rolls-Royce Trent 900) operating out of Doha airport (IATA: HIA) is estimated by simulating climb-out trajectories over three separate calendar months during which a number of sand storms have been identified. The atmosphere model incorporates dust concentration hind-casts from the European Centre for Medium-Range Weather Forecasts (ECMWF) near real-time Copernicus Atmosphere Monitoring Service database. A total of 365 flights were considered. The average dust mass ingested into the engine core per flight is estimated as ∼8.5 g. Using a compressor fouling prediction model, the dust ingestion reduces the asymptotic deterioration rate time constant by a factor ∼8. The worst dust storm within the time frame considered is selected for further examination. This flight is flagged as a dust-flight , during which the total mass of dust ingested into the engine core exceeds the monthly mean by two standard deviations. A peak dust ingestion rate of 22 mg/s occurs shortly after take-off, midway through the first climb segment. A second peak presents as the aircraft transitions into the climb from 3,000 feet to 10,000 feet. Finally, the dust ingested during a 20-minute holding pattern over the Persian Gulf is estimated as ∼8 g, which is approximately the same as the dust ingested during climb-out on a very dusty day. The results suggest that a mission-based approach may be more useful for determining aircraft engine durability in dusty environments. [ABSTRACT FROM AUTHOR]

Due to increases in fossil fuel prices worldwide, aero-engine manufacturers in the aerospace industry are introducing innovative designs that emphasize mass reduction, improving efficiency and lowering operating costs. However, this phenomenon has increased the exposure of aero-engines to vibration transfer, due to internal excitation loadings produced by issues such as rotor system unbalance. Due to these circumstances and the general advancement of aero-engine design technology, the demand for reductions in interior fuselage noise and vibration has increased. In this research, bond graph transfer path analysis (TPA) is extended to apply to aero-engines, with the goal of analyzing vibration propagation through the aero-engine's structural pathways. Bond graph TPA is a reliable and valuable method that can be utilized to efficiently diagnose and minimize noise and vibration problems during the design phase, prior to prototyping. One of the main advantages of bond graph TPA is its analytical algorithm, which attempts to circumvent the empirical limitations imposed by previously available TPA models such as operational path analysis and operational path analysis with exogenous inputs. The result is superior time efficiency when analyzing vibration propagation in structures such as aero-engines, especially as compared to earlier TPA methods which demand vast and extensive amount of experimental measurements which could be an inefficient approach at times In this study, the bond graph transfer path analysis methodology is explored and implemented on a reduced lumped model of an aero-engine in order to demonstrate this novel transmissibility analysis approach. Using this unique process, the concepts of global and direct transmissibility and path contribution in aero-engines were investigated based on the TPA methodology, and vibration reduction strategies proposed that will help to minimize vibration transfer to aircraft fuselages that originates from rotor's unbalanced forces. [ABSTRACT FROM AUTHOR]

Future civil air vehicles are likely to feature propulsion systems which are more closely integrated with the airframe. For a podded underwing configuration, this close coupling is expected to require non-axisymmetric design capabilities for the aero-engine exhaust system. This work presents the development of a novel parametric representation of non-axisymmetric aero-engine exhaust system geometries based on Intuitive Class Shape Transformation (iCST) curves. An exhaust design method was established and aerodynamic analyses of a range of non-axisymmetric configurations was demonstrated. At typical flight conditions, the introduction of non-axisymmetric separate-jet nozzles was shown to increase the engine net propulsive force by 0.12% relative to an axisymmetric nozzle. [ABSTRACT FROM AUTHOR]