bluuebunny/crossref_metadata_2025_split huggingface dataset:
Uncertainty Quantification and Missing Data for Turbomachinery With Probabilistic Equivalence and Arbitrary Polynomial Chaos, Applied to Scroll Compressors
Authors: Pepper Nick , Montomoli Francesco , Giacomel Francesco , Cavazzini Giovanna , Pinelli Michele , Casari Nicola , Sharma Sanjiv 
Published: January 2021

Abstract
This work presents a framework for predicting unknown input distributions for turbomachinery applications starting from scarce experimental measurements. The problem is relevant to turbomachinery where important parameters are obtained using indirect measurements. In this paper a scroll compressor is used as example but the suggested framework is completely general and can be used to infer missing data on material composition (carbon fiber properties, laser melted specimens for additive manufacturing etc) or input data (such as the turbine inlet temperature).
Scroll compressors are small devices with a very complex geometry that is difficult to measure. Moreover these compressors are highly sensitive to manufacturing errors and clearances. For these reasons we have chosen this example as an ideal candidate to prove the effectiveness of the framework. An input probability distribution for the scroll height is recovered based on a scarce, synthetic data set. The scroll height is used as an example of a missing distribution for a geometric parameter as it has the highest variance and is challenging to measure experimentally.
The framework consists of two main building blocks: an equivalence in a probabilistic sense and a Non-Intrusive Polynomial Chaos formulation able to deal with scarce data. The probabilistic equivalence is defined by a Probability Density Function (PDF) matching approach in which the statistical distance between probability distributions is quantified by either the Kolmogorov-Smirnov (KS) distance or the Kullback-Leibler (KL) divergence. By representing the missing inputs with a generalised Polynomial Chaos Expansion (gPCE) the back-calculation problem can be recast as an optimisation problem in which an arbitrary Polynomial Chaos (aPC) formulation was used to propagate the uncertain input distributions through a computational model of the system and generate a probability distribution for the Quantity of Interest (QoI).
The framework has been tested with multiple non-Askey scheme distributions to prove the generality of the proposed approach.
Fast Model Predictive Control for Aircraft Engine Based on Automatic Differentiation
Authors: Wang Yunlai , Wang Xi 
Published: January 2021

Abstract
Nonlinear model predictive control (NMPC) is a strategy suitable for dealing with highly complex, nonlinear, uncertain, and constrained dynamics involved in aircraft engine control problems. Because of the complexity of the algorithm and the real-time performance of the predictive model, it has thus far been infeasible to implement model predictive control in the realtime control system of aircraft engine. In most nonlinear model predictive control, nonlinear interior point methods (IPM) are used to calculate the optimal solution, which iterate to the optimal solution based on the Jacobian and Hessian matrix. Most nonlinear IPM solver, such as MATLAB fmincon and IPOPT, cannot calculate the Jacobian and Hessian matrix precisely and quickly, instead of using numerical differentiation to calculate the Jacobian matrix and BFGS method to approach to the Hessian matrix. From what has been discussed above, we will 1) improve the real-time performance of predictive model by replacing the time-consuming component level model (CLM) with a neural network model, which is trained based on the data of component level model, 2) precisely calculate the Jacobian and Hessian matrix using automatic differentiation, and propose a group of algorithms to make NMPC strategy quicker, which include making use of the structure of predictive model, and the integrity of weighted sums of Hessian matrix in IPM. Finally, considering input and output constraints, the fast NMPC strategy is compared with normal NMPC. Simulation results with mean time of 19.3% – 27.9% of normal NMPC on different platforms, verify that the fast NMPC proposed can improve the real-time performance during the process of acceleration, deceleration for aircraft engine.
Study on Optimization Design and Flow Control Mechanism of Little Blades in a Compressor Cascade
Authors: Guo Zhengtao , Chu Wuli , Chen Xiangyi 
Published: January 2021

Abstract
In view of the characteristics of flow separation in the compressor cascade corner region, a new flow control method for installing little blades in the front of the cascade passage was proposed, which took into account the flow control advantages of end wall fences and vortex generators. Firstly, the little blades could hinder the cross flow on the end wall and the development of the horseshoe vortex pressure surface branch. Secondly, the little blades could generate induced vortices to take away the low-energy fluid near the end wall and the corner region.
Based on numerical simulations, the effects of different pitchwise positions, stagger angles and heights of the little blades on the aerodynamic performance of the cascade were studied, and the optimal little blades were obtained by NSGA-II using EBF neural network as the agent model. The results show that the little blades have the optimal pitchwise position, stagger angle and height range for improving the aerodynamic performance of the cascade. When the optimized little blades are introduced in the baseline cascade, the stable working range of the cascade is expanded, and the stall type of the cascade changes from the hub-corner stall to the overload of flow separation near the mid-span. At the near stall attack angle of the baseline, the total pressure loss coefficient is reduced by about 10.38% and the static pressure coefficient is increased by about 4.31%. Meanwhile, the loss of the lower span is decreased and the diffuser capacity of the whole span is improved. The passage secondary loss and wake loss are reduced because of the delay of corner separation. Moreover, the strength of the end wall vortex is weakened and the end wall vortex no longer develops as part of the passage vortex. The induced vortex, horseshoe vortex pressure surface branch and initial passage vortex develop into new passage vortex.
Uncertainty Quantification of Aero-Thermal Performance of a Blade Endwall Considering Slot Geometry Deviation and Mainstream Fluctuation
Authors: Tao Zhi , Guo Zhendong , Song Liming , Li Jun 
Published: January 2021

Abstract
With the continuous increase of aerodynamic and thermal load, the endwall of modern gas turbines has become the critical region affected by the uncertainties in the manufacturing and operation process and thus is very likely to suffer performance degradation and thermal failure. Therefore, it is critical to understand and quantify the impacts of uncertainty factors on endwall aero-thermal performance. Based on Kriging surrogate, the frameworks of uncertainty quantification and global sensitivity analysis are constructed for a gas turbine blade endwall. The impacts of slot geometry deviations (slot width, endwall misalignment) and mainstream condition fluctuations (turbulence intensity, inlet flow angle) on endwall aero-thermal performance are quantified and analyzed. Results show that the actual performance of the endwall has a high probability of deviating from its nominal value. With respect to the nominal values, the maximum deviations of aerodynamic losses, averaged film cooling effectiveness and averaged Nusselt number reach up to 0.33%, 45% and 5.0%, respectively. The critical regions which are most sensitive to the input uncertainty parameters are identified. Furthermore, the inlet flow angle is proved to be the most significant parameter affecting the endwall aero-thermal performance through sensitivity analysis. The influence mechanisms of the inlet flow angle on endwall aero-thermal performance are clarified by detailed flow and thermal field analysis. Results show that the inlet flow angle significantly alters the size and strength of the secondary flow structures, resulting in a large variation of endwall aero-thermal performance. Quantitatively, a positive incidence angle of 2 degrees can lead to a 0.1% reduction of total pressure coefficient, a 12% reduction of averaged film cooling effectiveness and a 2% enhancement of averaged Nusselt number.
A Three-Dimensional Unsteady Through-Flow Model for Rotating Stall in Axial Compressors
Authors: Guo Jin , Hu Jun , Wang Xuegao , Xu Rong 
Published: January 2021

Abstract
Rotating stall is a natural limit to the stable operating range of compressors due to the inverse pressure gradient of viscous gas. Effective prediction of compressor stall boundary is an important guarantee for the successful development of aeroengine. In this paper, a three-dimensional unsteady through-flow model based on body force theory is developed to reflect the dynamic stall process of multistage axial compressors with acceptable computational costs. The influence of blade geometric parameters is fully considered in blade force source terms. The source terms are related to the attack angle and Mach number of the blade inlet using the deviation angle and loss model in the through-flow theory. Meanwhile, the temporal lag response of the source terms to the upstream flow conditions is taken into account. Therefore, it can be utilized for predicting the off-design performance and rotating stall characteristics of multistage axial compressors. The developed model is validated on a two-stage low-speed axial compressor. The calculated performance line and stall cell speed are in agreement with the experimental results. The unsteady flow behavior of the compressor during stall is presented by the model. The results indicate that the developed model has the potential to be applied to the preliminary evaluation of compressor stability in design stage.
Effect of Inlet Mach Number on Aerothermodynamics of Sealing Leakage Flow Cooling on Endwall
Authors: Gao Hongyu , Li Xueying , Ren Jing , Jiang Hongde 
Published: January 2021

Abstract
The gas turbine has an assembly gap between the combustion chamber and the first stage vane. The coolant air discharge from the gap can prevent the ingestion of the high temperature gas. This leakage flow also provides a cooling coverage on the vane endwall. Taking the cooling effect of the leakage flow on the endwall into consideration is very important for an efficient cooling design.
In this paper, the cooling effect of leakage flow on endwall is studied by means of experimental and numerical methods. The study included slots at 30°, 45°, and 60° angles, and six blowing ratios of 0.3, 0.6, 1.0, 1.4, 1.7, and 2.0. The experiment and numerical calculation are conducted under the condition that the inlet Mach number is 0.125 and the outlet Mach number is 0.72, which is close to the working Mach number of the real gas turbine.
Under the same slot inclination and blowing ratio, the distribution of endwall adiabatic cooling effectiveness is more nonuniform under the condition of near-real engine Mach number. This is because the passage vortex is weaker under the low Mach number condition, and the leakage flow has a better wall attachment effect. In terms of the spanwise average of endwall adiabatic cooling effectiveness, when the blowing ratio is small, the adiabatic cooling effectiveness is lower under the condition of near-real engine Mach number than that under the condition of low Mach number, but the opposite is true under the condition of large blowing ratio. This is because under the condition of large blowing ratio, the turbulence is stronger under the condition of Mach number of near-real engine. With the reduction of blowing ratio, the turbulent kinetic energy weakens more strongly. In the studied cases, there is a critical blowing ratio of 1.0, and the total endwall cooling adiabatic cooling effectiveness is not significantly affected by the Mach number when it is smaller than M1.0. The average adiabatic cooling effectiveness of the endwall under the condition of near-real engine Mach number is about 7% lower than that under the condition of low Mach number.
It means that the experimental results of leakage flow cooling obtained under the condition of low inlet Mach number need to be corrected by a correction factor, which may be less than 1 to make it engine relevant.
A Degradation Diagnosis Method for Gas Turbine – Fuel Cell Hybrid Systems Using Bayesian Networks
Authors: Mantelli L. , Zaccaria V. , Kyprianidis K. , Ferrari M. L. 
Published: January 2021

Abstract
During the last decades there has been a rise of awareness regarding the necessity to increase energy systems efficiency and reduce carbon emissions. These goals could be partially achieved through a greater use of gas turbine - solid oxide fuel cell hybrid systems to generate both electric power and heat. However, this kind of systems are known to be delicate, especially due to the fragility of the cell, which could be permanently damaged if its temperature and pressure levels exceed their operative limits. This could be caused by degradation of a component in the system (e.g. the turbomachinery), but also by some sensor fault which leads to a wrong control action. To be considered commercially competitive, these systems must guarantee high reliability and their maintenance costs must be minimized. Thus, it is necessary to integrate these plants with an automated diagnosis system capable to detect degradation levels of the many components (e.g. turbomachinery and fuel cell stack) in order to plan properly the maintenance operations, and also to recognize a sensor fault. This task can be very challenging due to the high complexity of the system and the interactions between its components. Another difficulty is related to the lack of sensors, which is common on commercial power plants, and makes harder the identification of faults in the system.
This paper aims to develop and test Bayesian belief network based diagnosis methods, which can be used to predict the most likely degradation levels of turbine, compressor and fuel cell in a hybrid system on the basis of different sensors measurements. The capability of the diagnosis systems to understand if an abnormal measurement is caused by a component degradation or by a sensor fault is also investigated. The data used both to train and to test the networks is generated from a deterministic model and later modified to consider noise or bias in the sensors.
The application of Bayesian belief networks to fuel cell - gas turbine hybrid systems is novel, thus the results obtained from this analysis could be a significant starting point to understand their potential. The diagnosis systems developed for this work provide essential information regarding levels of degradation and presence of faults in gas turbine, fuel cell and sensors in a fuel cell – gas turbine hybrid system. The Bayesian belief networks proved to have a good level of accuracy for all the scenarios considered, regarding both steady state and transient operations. This analysis also suggests that in the future a Bayesian belief network could be integrated with the control system to achieve safer and more efficient operations of these plants.
Study on the Influence of Film Cooling on the Flow Field and Infrared Radiation Characteristics in the Divergent Section of a Nozzle
Authors: Yang Haoqi , Yang Qingzhen , Zhang Saile , Yang Huicheng , He Yubo 
Published: January 2021

Abstract
As the last part of the convergent divergent nozzle, the divergent section is exposed to high temperature and high-speed airflow and thus, it is more easily to be detected by the infrared detector. It is one of the main sources of the infrared radiation in the exhaust system. Film cooling is applied to protect the wall from hot flow and reduce the infrared radiation.
In this paper, the study is conducted on a nozzle with spherical convergence flap in a turbofan engine exhaust system. The effect of film cooling on the internal flow and infrared radiation characteristics of the exhaust system in the divergent section was studied by numerical simulations. The k-ω SST turbulence model was used to simulate the flow field, and the Reverse Monte Carlo Method was employed to calculate the infrared radiation characteristics of the nozzle. Four different kinds of film hole arrangements are involved, they are cylindrical film holes in an in-line pattern, cylindrical film holes in a staggered pattern, converging-expanding film holes in an in-line pattern and converging-expanding film holes in a staggered pattern. The cylindrical film hole and the converging-expanding film hole have a round shape inlet, with an equivalent diameter of d = 5mm on the projection surface perpendicular to the axial direction. Angles between each film hole and the wall surface are 35°. The impact of the heat conduction on the wall was taken into account. The results show that with the given mass flow rate of the coolant, the lengths of the high temperature core zone of the four models with different film cooling structures are slightly shorter than the core zone of the model without cooling structures. However, no significant difference can be found for the length of the core zone of the four models. The average temperature of the wall in the divergent section decreases significantly by using film cooling. No significant difference can be found in the wall temperature distribution for the four models. In the 3∼5μm and 8∼14μm bands, the cooling technique barely affects the infrared radiation of the main exhaust jet flow, while it significantly reduces the infrared radiation of the solid wall in the divergent section, and the decreasing amplitude is from 45% to 51%. Different film hole arrangements result in similar effects on the infrared radiation of the nozzle.
Overall, the usage of film cooling in the divergent section of the nozzle effectively reduces the averaged wall temperature and substantially suppresses the solid infrared radiation on the wall. However, the shape and arrangement of the film holes have no significant influence on the infrared radiation intensity and temperature of the wall in the divergent section.
A Study of Two Variable Cycle Engine Concepts for High Speed Civil Aircraft
Authors: Zhang Jiyuan , Chen Min , Tang Hailong , Liu Xin 
Published: January 2021

Abstract
High speed civil aircraft has become a promising field with the development of globalization. The propulsion system is an indispensable part of the aircraft. Conventional engines have difficulty meeting the performance requirement of the high-speed civil aircraft.
In this article, two variable cycle engines were studied to preliminarily as aircraft propulsion system. Their performance and matching mechanism were analyzed and compared with each other. Firstly, the cycle parameters design was conducted to explain the principle of cycle parameter determination for the high-speed civil aircraft. Secondly, the control law of variable geometry components was studied to optimize engine performance during supersonic cruising. Finally, the throttling process with constant airflow was studied to solve the problem of thrust surplus during subsonic cruising.
According to this study, given same cycle parameters, the engine with variable fan stage can produce equal or slightly higher thrust with slightly less fuel consumption than the engine with core-driven-fan stage. The engine with core-driven-fan stage has advantages in aero-dynamical stabilities. It can also throttle to slightly lower thrust level during subsonic cruising. Considering the advantages in performance and derived development comprehensively, the engine with variable fan stage is a better option for high speed civil aircraft.
Analysis of Friction-Saturated Flutter Vibrations With a Fully-Coupled Frequency Domain Method
Authors: Berthold Christian , Gross Johann , Frey Christian , Krack Malte 
Published: January 2021

Abstract
Flutter stability is a dominant design constraint of modern gas and steam turbines. Thus, flutter-tolerant designs are currently explored, where the resulting vibrations remain within acceptable bounds. In particular, friction damping has the potential to yield Limit Cycle Oscillations (LCOs) in the presence of a flutter instability. To predict such LCOs, it is the current practice to model the aerodynamic forces in terms of aerodynamic influence coefficients, derived for some normal modes of the linearized structural model and fixed oscillation frequency. However, this approach neglects that both the nonlinear contact interactions and the aerodynamic stiffness cause a change in the deflection shape and the frequency of the LCO. This, in turn, may have a substantial effect on the aerodynamic damping. The goal of this paper is to assess the technical importance of these neglected interactions. To this end, a state-of-the-art aero-elastic model of a low pressure turbine blade row is considered, undergoing nonlinear frictional contact interactions in the tip shroud interfaces. The LCOs are computed with a fully-coupled harmonic balance method, which iteratively computes the Fourier coefficients of structural deformation and conservative flow variables, as well as the a priori unknown frequency. The coupled algorithm was tested for various combinations of harmonics in both domains and found to provide excellent computational robustness and efficiency. Moreover, a refinement of the conventional energy method is developed and assessed, which accounts for both the nonlinear contact boundary conditions and the linearized aerodynamic influence. It is found that the conventional energy method may not predict a limit cycle oscillation at all while the novel approach presented here can. Furthermore the refined energy method provides deep understanding of the nonlinear aero-elastic interactions.
Stress Analysis of Fan Blades Subjected to Hail Impact
Authors: Sun Xin , Liu Bo , Han Chang-fu 
Published: January 2021

Abstract
The rapid development of modern aircraft technology is increasingly demanding aero-engines. It not only requires aeroengines to ignite and take off under suitable conditions, but also faces various severe weather and foreign object impact tests during flight. The stress variation caused by the impact of an aero-engine fan blade is studied. Because the compressor guide vane has a small solidity, the hail may not only hit the vane when striking the blade, but also may hit the first-stage rotor blade. The effects of hail impact on the vanes and rotor blades, especially the impact on the blade strength, were studied. Firstly, the UG software is used for geometric modeling based on the existing vane and rotor geometry data. Secondly, the ANSYS/LSDYNA software is used to simulate the process when the hail hits different positions of the blade and rotor at different speeds. To explore the influence of hail impact on the blade. It is found that when the hail hits the guide vane, the maximum effective stress of the vane does not exceed the yield stress of the material. When the hail hits the rotor, the maximum effective stress of the rotor is closely related to the rotor speed, the hail incident velocity, and the relative position of the rotor and the hail. When the impact is severe, the maximum effective stress will exceed the yield stress of the material, and even the blade may be bent and pitted, which has an important impact on the strength and life of the blade.