Compressor Rotor Blade Research Articles

Multi-Disciplinary Surrogate-Based Optimization of a Compressor Rotor Blade Considering Ice Impact

Multi-Disciplinary Surrogate-Based Optimization of a Compressor Rotor Blade Considering Ice Impact

Improved non-contact vibration measurement via acceleration-based blade tip timing

The conventional displacement-blade tip timing (D-BTT) measurements often rely on Once-Per-Revolution (OPR) sensors to compute expected or theoretical time of arrival (TOA) based on rotational speed. However, integrating OPR sensors within the confined space of aero-engines presents challenges, and estimating theoretical TOA introduces computational errors, particularly for low-amplitude, high-frequency blade vibrations. This paper presents the Acceleration-based Blade Tip Timing (A-BTT) method as a solution. Building upon A-BTT, the acquisition of blade tip vibration acceleration serves not only to capture rapid changes in vibration but also to offer richer data for precise measurement of high-frequency blade vibrations. With only three blade tip timing sensors recording actual TOA without an OPR sensor, A-BTT eliminates the need for estimating expected TOA, thus mitigating calculation errors. Furthermore, by analyzing the relationship between maximum blade tip vibration acceleration and displacement during synchronous blade vibration, vibration parameters, such as natural frequency, can be directly identified without prior information such as engine orders. Simulation and experimental results validate the efficacy of the A-BTT method, demonstrating its successful application to specific aero-engine compressor rotor blades.

Numerical and experimental study on the critical geometric variation based on sensitivity analysis on a compressor rotor

Abstract Different types of geometric variations often appear coupledly in manufactured blades. It is desired to identify the ones that have the strongest impact on the performance. In this paper, the influence of multiple geometric variations on the compressor rotor blade at the design point is studied. Two hundred and nine varied blades are constructed by adding variation data to the design-intent blade and assessed using steady-state Reynolds-Averaged Navier–Stokes (RANS) simulations. It is shown that the region near the lower and upper end of the blade is more sensitive to geometric variations. Spearman’s rank correlation coefficient is used to measure the sensitivity of rotor performance to geometric variations. The results show that the stagger angle variation is the most influencing variation. Local sensitivity analysis at different spans also shows that the stagger angle variation is an important geometric variation that needs more attention during blade manufacturing. Also, the influence of geometric variations can reach further than nearby regions. Consequently, experiments on linear cascades at different spans of the rotor are carried out to study the effect of stagger angle on flow characteristics. Results show that the stagger angle variation could lead to different changes in performance depending on the specific cascade profile.

Influence of trailing edge radius of intake struts on excitation force and vibration of rotor blades

Adjusting the trailing edge radius of the intake struts may be advantageous in suppressing the vibration of compressor rotor blades. To explore the impact of the intake struts trailing edge radius on the excitation force and vibration, this study focuses on a compressor with intake struts, and fluid and vibration calculations were conducted. The study revealed that decreasing the trailing edge radius led to a reduction in separation near the strut’s trailing edge, resulting in a narrower wake and weakened mixing near the trailing edge. This caused the total pressure loss region induced by the wake to become more elongated in the low-frequency component, leading to a nonlinear decrease in total pressure amplitude with the reduction of the trailing edge radius. As the trailing edge radius decreased, the excitation force and vibration amplitude of the first-stage compressor rotor blades decreased, while the excitation force distribution showed little variation. The research findings provide insights for the design of intake struts.

CFD Analysis and Validation of Axial Compressor Rotor Blade

This research paper employs computational simulations to investigate the consequences of inlet flow distortion, focusing specifically on inlet swirl and total pressure distortion, on the operational characteristics and robustness of an axial transonic compressor. Various permutations of inlet swirl and total pressure distortion are examined, uncovering diverse impacts on the compressor's operational robustness and efficiency. The analysis demonstrates that while co-swirl configurations marginally enhance the operational robustness of the compressor, counter-swirl configurations tend to diminish it. The combination of the effects of individual distortion patterns can provide insights into certain aspects of how a composite pattern of distortion would affect the compressor's efficiency, yet this approach falls short in accurately evaluating the compressor's operational robustness when subjected to a composite distortion pattern.

Identification of the Aeromechanical Parameters of a Compressor Rotor from Blade Tip Timing Data

Abstract This article presents a maximum likelihood approach to identifying the mistuning, aerodynamic forcing functions, and aerodynamic influence coefficients of compressor rotor blades from noisy experimental measurements of blade deflection. The maximum likelihood estimation approach leads naturally to a nonlinear least squares optimization problem that can be solved using numerical optimization techniques such as quasi-Newton methods. An eigenmode decomposition approach is presented that permits the optimization objective function and gradient to be efficiently calculated so that the optimization is computationally feasible, even for experiments involving very large data sets. Simulation results show that the method has the potential to accurately identify the mistuning and aerodynamic nodal diameter forcing functions, but obtaining accurate results for influence coefficients may not be possible if there is significant disk flexibility. We demonstrate the method using both simulated and experimental blade vibration data from rotor 2 of the Purdue University three-stage axial compressor research facility collected using a light probe tip timing system as the rotor is slowly accelerated or decelerated through a Campbell diagram crossing.

Experimental and numerical investigations on non-synchronous vibration and frequency lock-in of transonic compressor rotor blades

Non-synchronous vibration (NSV) problems in axial compressors are challenging and frequently result in high vibration stress or structural failure within the frequency lock-in region. A deeper understanding of the lock-in phenomenon in NSV is necessary to improve blade reliability and safety. In this study, an experiment was conducted on a multistage transonic axial compressor to obtain frequency lock-in characteristics. The enforced motion of the flexible blade simulation method was used to study the frequency lock-in mechanism between blade vibration and tip unstable flow. The results showed that the aerodynamic frequency characteristic changed from a broadband spectrum to a distinct frequency peak as the NSV onset. The nonlinear vibration pattern could maintain the limited cycle oscillation for 26 s. The vibration stress of the frequency locked-in case was approximately 19 times that of the frequency unlocked one. The pressure fluctuations generated by the large-scale radial separation vortex structures were stronger than those of the tip leakage flow, and the third-order harmonic frequency of the circumferential instability flow was close to the first-order bending model natural frequency. The periodic shedding and reattachment process of flow separation provided the initial non-synchronous aerodynamic excitation source. The frequency lock-in phenomenon occurred at a maximum vibration amplitude of 2 % and 3 % tip chord length. The tip unstable flow characteristics were consistent and entirely dominated by blade vibration.

Effects of tip clearance on non-synchronous aerodynamic excitation of compressor rotor blades

The unstable tip flow of compressor rotor may cause the aeroelastic problems, inducing the non-synchronous vibration with high amplitude level, which has a significant effect on blade high cycle fatigue and safety. In order to investigate the effects of tip clearance on non-synchronous aerodynamic excitation of compressor rotor blades, the full annular numerical simulations were conducted for the unsteady flow field of a 1.5-stage compressor under near stall conditions. The frequency characteristics of non-synchronous aerodynamic excitation and tip unstable flow structure were analysed for different tip clearances (0.5%C, 1.0%C, 2.0%C, and 3.0%C). The results showed that the frequency and amplitude are strongly influenced by the tip clearance size. For 0.5%C tip clearance, the main frequency of non-synchronous aerodynamic excitation is the low-frequency aerodynamic excitation frequency, which is mainly caused by tornado-like vortex in the tip suction side, the dominant aerodynamic disturbance mode number is 19. For 2.0%C and 3.0%C tip clearance, the dominant aerodynamic disturbance mode number is 47, and the main aerodynamic excitation frequency is high frequency non-synchronous aerodynamic excitation frequency, mainly caused by the shedding vortices propagating circumferentially at the blade tip, exhibiting the feature of “rotating instability”.

Three-dimensional optimization of squealer-tip for a transonic axial-flow compressor rotor blade

This research represents an innovative method for geometry generation of the squealer-tip in axial compressor rotors and its exploit in a numerical optimization process to obtain a better stage performance. For this purpose, the NASA Rotor-67 transonic compressor rotor blade is used as a test case to study the aerodynamic performance using computational fluid dynamics. The validation was performed for the characteristic map at the design speed and the comparison with the experimental results indicates excellent matching and high adaptability of the numerical method. An ingenious method of producing squealer tip for an axial compressor rotary blade is presented in this article, which is capable for locally shaping both suction and pressure surface geometry at a desired spanwise location simultaneously, while keeping the tip clearance at its value of the baseline NASA Rotor-67 geometry. In this method, control points are used to produce the starting spanwise location of the squealer, and modify the depth of the squealer geometry. The L-27 orthogonal array of the Taguchi method as the Design of Experiment (DOE) has been used to investigate the sensitivity of the aerodynamic results in three performance points of the choke, design and near stall regions, in relation to the design variables of the squealer. The generated database in the sensitivity analysis was used to train artificial neural networks to replace the CFD solutions with overwhelming run time. By coupling the genetic algorithm to the aforementioned neural networks and by applying penalties to maintain the minimum performance of the Rotor-67, enhancement of total pressure ratio, adiabatic efficiency, mass flow rate and even the surge margin was achieved. The main effect of the squealer is to modify the shape of blade tip vortices, and by more dissipation of energy in blade tip area and reduced equivalent flow area in this region, finally results in improved overall mass flow rate, total pressure ratio, adiabatic efficiency and surge margin by 0.58 %, 0.36 %, 0.19 % and 4.81 % respectively, at design point.

Applications of design and reverse engineering for the development of digital and smart tools for composite additive manufacturing

This paper presents a study dedicated to the development of Digital and Smart Tools, based on solved applications of design engineering and reverse engineering. This approach is justified by the fact as well as the need for preparations prior to Composite Additive Manufacturing. Future integration of Digital Smart Tools with Composite Additive Manufacturing will significantly contribute to the efficient and effective support of the green economy, the active, responsible, safe and resilient protection of environment, life and climate. The research involved in this paper contributes to develop Digital and Smart Tools Applications, intended for integrated digital design, development, manufacturing and further predictive maintenance and services, based on robotic systems, extended automatic control, Artificial Intelligence and Machine Learning, including Mathematical Modeling and Numerical Simulations for Performance Prediction at Design Regime and Off-Design Regimes for jet engines, with the best capabilities to generate and integrate improvements, optimizations and potential innovative solutions. This paper presents significant applications of design, concept engineering development and reverse engineering design, as: 1/ the design of a transonic axial compressor rotor blade, 2/ the design of a swept axial compressor rotor, 3/ concept design engineering developments in case of a swept fan rotor blade, 4/ concept design developments and reverse engineering in case of a HP axial compressor rotor blade, part of Spey 512-14 DW turbofan engine, 5/ reverse engineering design of a Cessna 182 Skylane N223IF light aircraft wheel cover. In line with Europe´s vision for sustainable aviation, this research study and INCAS´ TGA Project "Technological Development Platform for "Green" Technologies in Aviation and Ecological Manufacturing with Superior Added Value; TGA - Technologies for Green Aviation" will significantly contribute to the Green Deal, as a production center using "Green" Technologies in aviation and ecological manufacturing, as well as collaborative developer of Digital and Smart Tools for Composite Additive Manufacturing.

Dynamic Stress Test of Aero-engine Compressor Rotor Blade

To evaluate the effect of the compressor rotor blade on compressor blade fatigue, a non-contact dynamic stress measurement scheme of one aero-engine eighth stage compressor rotor blade was designed, and the dynamic stress test was carried out. Meanwhile, modal analysis of the eighth stage compressor rotor blade was done by finite element analysis software. Results show that between the range of the idle speed and the highest speed, there is no resonance for the eighth stage compressor rotor blade and the maximum vibration amplitude is less than 65 μm, which cannot result in high cycle fatigue failure.

Influence of the chordwise distribution of tip winglets on the stability of a high-load compressor stage

Compressor rotor blade tip winglet technology has been confirmed to be a passive flow control method that is effective in increasing the operating stability of compressors. As revealed by extensive research, the structural parameters of tip winglets are capable of affecting their stability expansion effect, whereas their mechanism and influence law of the compressor stage remains unclear. In this study, the effect exerted by the chordwise distribution of tip winglets on the stability of a high-load compressor stage was investigated. As indicated by the result of this study, the tip winglet exhibiting a proper widest position and an adequate chordwise length can effectively increase the stall margin. To be specific, the widest position was located at 25% chord, and the stall margin was increased by up to 25.39%. A winglet exhibiting a chordwise length of 50% achieved the stability expansion ability equivalent to winglet with a full chordwise length, which can be increased by 22.58%. The tip leakage flow turned out to be weakened as the axial momentum was increased, and the circumferential momentum of the flow that arose from the tip winglet was reduced. The result of this study also suggested that tip winglets can play a certain role in combing the inflow of the stator, such that the blockage of the stator channel can be reduced.

Static aeroelasticity analysis of a rotor blade using a Gauss-Seidel fluid-structure interaction method

The present study introduces a Gauss-Seidel fluid-structure interaction (FSI) method including the flow solver, structural statics solver and a fast data transfer technique, for the research of structural deformation and flow field variation of rotor blades under the combined influence of steady aerodynamic and centrifugal forces. The FSI method is illustrated and validated by the static aeroelasticity analysis of a transonic compressor rotor blade, NASA Rotor 37. An improved local interpolation with data reduction (LIWDR) algorithm is introduced for fast data transfer on the fluid-solid interface of blade. The results of FSI calculation of NASA Rotor 37 show that when compared with the radial basis function (RBF) based interpolation algorithm, LIWDR meets the interpolation accuracy requirements, while the calculation cost can be greatly improved. The data transmission time is only about 1% of that of RBF. Moreover, the iteration step of steady flow computation within one single FSI has little impact on the converged aerodynamic and structural results. The aerodynamic load-caused deformation accounts for nearly 50% of the total. The effects of blade deformation on the variations of aerodynamic performance are given, demonstrating that when static aeroelasticity is taken into account, the choke mass flow rate increases and the peak adiabatic efficiency slightly decreases. The impact mechanisms on performance variations are presented in detail.

Solid Particle Erosion Performance and Damage Mechanism of AlTiN Coating

Aircraft, when operating in a sandy environment, can experience severe solid particle erosion damage to gas turbine engine components, such as compressor rotor blades, vanes, and impeller blisks/wheels, due to sand particle ingestion into the engine. Material removal from these components as the result of erosion not only leads to significant aerodynamic losses, but also results in their structural weakening. Applying erosion-resistant coatings on airfoil surfaces has been proven effective in extending the serviceable life of engine components. In this study, cathodic arc evaporated AlTiN was explored for its potential as an erosion-resistant coating for the compressor blades and helicopter impellers applications. AlTiN coating, along with baseline TiN coating, was subjected to gas jet sand erosion testing under different particle impingement angles and velocities. The erosion rates of the AlTiN coating were found to be much lower than those of the TiN coating in the entire range of the tested conditions. While the TiN coating mainly suffered from severe brittle erosion damage through coating chipping, the AlTiN coating exhibited only mild erosion damage through microcutting by sand particles. To include the aerodynamic effect on airfoil erosion in the consideration, dynamic wind tunnel sand erosion testing was also performed on AlTiN- and TiN-coated propeller blades. The results of these tests also confirmed superior erosion resistance of the AlTiN coating over that of the TiN coating.

A New Approach for Deviation Modeling in Compressors: Sensitivity-Correlated Principal Component Analysis

Studies on the geometry variation-related compressor uncertainty quantification (UQ) have often used dimension reduction methods, such as the principal component analysis (PCA), for the modeling of deviations. However, in the PCA method, the main eigenmodes were determined based only on the statistical behavior of geometry variations. While this process can cause some missing modes with a small eigenvalue, it is much more sensitive to blade aerodynamic performances, and thereby reducing the reliability of the UQ analysis. Hence, a novel geometry variation modeling method, named sensitivity-correlated principal component analysis (SCPCA), has been proposed. In addition, by means of the blade sensitivity analysis, the weighting factors for each eigenmode were determined and then used to modify the process of the PCA. As a result, by considering the covariance of geometry variations and the performance sensitivity, the main eigenmodes could be determined and used to reconstruct the blade samples in the UQ analysis. With 98 profile samples measured at the midspan of a high-pressure compressor rotor blade, both the PCA and SCPCA methods were employed for the UQ analysis. The results showed that, compared to the PCA method, the SCPCA method provided a more accurate reconstruction of sensitive deviations, leading to an 11.8% improvement in evaluating the scatter of the positive incidence range, while also maintaining the accuracy of the uncertainty assessment for other performances.

Mechanism analysis and uncertainty quantification of blade thickness deviation on rotor performance

The blade thickness distribution is an important geometric parameter of a compressor rotor blade. It determines the aerodynamic performance of a rotor, which should be carefully designed. However, the manufacturing inaccuracies cause the blade thickness distribution to deviate from the ideal design. These geometric deviations alter the flow field near the blade surface, which affects the aerodynamic performance of the rotor. Therefore, it is of great significance to quantitatively investigate the effects of blade thickness deviation on the aerodynamic performance. In this paper, the influence of blade thickness deviation on the flow field is analyzed based on a transonic compressor rotor, and an uncertainty quantification process is performed to study the effects of blade thickness deviation on the aerodynamic performance. Cases with different geometry features are checked in the current study using 3-dimensional Reynolds-averaged Navier–Stokes simulations. Results show that the blade thickness deviation leads to changes in the intensities of the expansion wave in the upper part of the blade suction and the passage shock wave, which affects the supercharging process of the rotor. The influence of thickness deviation on the strength of the tip leakage vortex changes the isentropic efficiency of the rotor. The results of the uncertainty quantitative analysis indicate that the linear correlation between the variation of rotor performance and the thickness deviation is strong when the machining accuracy is high, whereas the linear correlation between the two is weakened when the tolerance range is larger.

Application of Machine Learning to Forced Response Predictions of an Industrial Axial Compressor Rotor Blade

Abstract Machine learning has gone way beyond a ground-breaking technology a few decades ago to now taken for granted in many day-to-day activities. It is now providing new ways for manufacturing, assembling, operating, monitoring, and maintaining products. Typical application areas include performance optimization, quality improvements, fault detection and predictive maintenance. In this paper application of machine learning algorithms to forced response prediction during the design and analysis of a typical gas turbine compressor blade is reported. The forced response prediction process typically involves utilizing harmonic or time domain computational fluid dynamics (CFD) analyses to compute the forcing and the aero damping, to calculate reserve factors that represent the high cycle fatigue life of the blade. This time-consuming process is generally limited to the later phases of the design cycle and can lead to hundreds of calculations if one must address all the resonances in a typical twin shaft running range. A neural network trained using historical data is used to directly predict the reserve factor with high confidence without the need for costlier high fidelity CFD by using just the finite elements (FE) predicted parameters. This allows to perform high-fidelity aero-mechanical assessment at an early stage in the design process. Further, application of image recognition using a convoluted neural network to aid in the identification of FE predicted Modeshapes is also demonstrated, which can also improve the quality of the reserve factor predictions.

Dynamic Bayesian network model for comprehensive risk analysis of fatigue-critical structural details

Dynamic Bayesian network (DBN) models are widely used for structural risk analysis because of their powerful parameter learning ability and capability of simplifying the problem by parsing it using nodes and arcs. However, these models update the reliability after inspection and maintenance (I&M) in a manner different than that specified by the aircraft structural integrity program (ASIP). In this study, a DBN model is developed to correctly represent the ASIP method. The model updates the crack length distribution after I&M based on the nondestructive testing (NDT) reliability and repair crack length distribution. The nodes for inequality and equality data are explicitly represented in the DBN, which correspond to the crack length vs. signal amplitude and noise characteristics of the NDT system. The proposed model overcomes drawbacks of existing models—initial overestimation and final underestimation—by appropriately considering the repair crack length distribution. Specifically, a decision node is used, which records the fraction of the crack length distribution removed after inspection. Furthermore, a method for constructing conditional probability tables is presented. The proposed model is applied to the in-service fatigue problem for a J85 engine compressor rotor blade. The findings demonstrate that the proposed model can be used in a wide range of applications.

A Methodology for a Coupled Structural - CFD Analysis of Compressor Rotor Blades Subjected to Ice Impact with Uncertain Impactor Parameters

Abstract A method for a coupled structural - Computational Fluid Dynamics (CFD) analysis of a compressor rotor blade subjected to an ice impact scenario is investigated to assess the impact related blade deformations from a structural and fluid-dynamics perspective. On the basis of a probabilistic approach, in total 50 impact scenarios are derived for this study. In a first step the numerical structural model based on Finite Elements (FE) is discussed, including several parameter variations like impact location, ice diameter, ice density and rotor speed. Different analysis steps are subsequently carried out using LSDYNA implicit / explicit on a high performance computing (HPC) cluster. Resulting blade deformations are evaluated in terms of local plastic deformation, cup size and modal parameters in comparison to the undamaged reference structure. The resultant post-impact blade geometry is extracted from the result data and passed to the CFD simulation setup in a fully automated manner. Based on this deformed structural mesh data, the fluid mesh is morphed via a radial basis function (RBF) approach and analysed with CFD. Finally, an uncertainty quantification study is performed to assess the variability of results with regard to the definition of the ice impactor.

Structural and Thermal Analysis of F Class Gas Turbine Compressor Blade

In this study, the structural and thermal analysis of the final stage compressor rotor blades, which are currently used on-site, is performed by using the ANSYS program. For that reason, a 3D model of the existing compressor blade has been done. Later on, this model was transferred to the ANSYS program and analyzed. In the analysis, the parameters formed in the blade geometry were determined. By creating a design geometry with the selected parameters, the stress of the existing blade under operating conditions was examined. All external parameters in this study were taken from an F-class gas turbine operating under real field conditions. The stresses of the obtained values in different regions on the blade were examined. This study is also a reverse engineering study and the effects of real field conditions on the compressor blades have been analyzed.