Compressor Blades Research Articles

Vortex dynamics and entropy generation in separated transitional flow over a compressor blade at various incidence angles

The transition process within a Laminar Separation Bubble (LSB) that formed on a compressor blade surface was investigated using Large Eddy Simulations (LESs) at a Reynolds number of 1.5 × 105 and incidence angles of 0°, +3°, and +5°. The vortex dynamics in the separated shear layers were compared at various incidence angles and its effects on the loss generation were clarified through entropy analysis. Results showed that transition onset, which was accurately identified by the Linear Stability Theory (LST), was significantly promoted at the increased incidence angle. As such, the development of LSB was suppressed and the relative role of viscous instability played in the transition process was weakened. At the incidence angle of 0°, two-dimensional spanwise vortices detached from the blade surface and rolled up periodically, which were further stretched and eventually evolved into large-scale hairpin vortices. As time passed, the fully developed hairpin vortices broke down into small-scale eddies. Meanwhile, the flow near the wall reversely ejected into the outer separated shear layers and a sweeping process happened subsequently, forcing the separated shear layers to reattach and accelerating the generation of turbulent fluctuations. By comparison, the strength of vortex rolling-up was weakened at higher incidence angles, and the vortex pairing and breakdown of large-scale vortices were less pronounced. Therefore, the level of turbulent fluctuations that generated in the separated shear layers was reduced. Detailed entropy analysis showed that the turbulent dissipation effect related to the Reynolds shear stresses determined the largest amount of positive entropy generation, which declined to a lower level as the incidence angle increased from 0° to +5°. Correspondingly, the profile loss was reduced by 50.4%.

An improved multiple per revolution-based blade tip timing method and its applications on large-scale compressor blades

Blade Tip Timing (BTT) is one of the most effective technologies for blade vibration measurement of turbomachinery. The blade vibration measurement indicates the operational status of the blades, which is of great significance for blade health management and fault diagnosis. However, the approaches and applications of BTT technologies are still facing enormous challenges. This study, based on the traditional once per revolution (OPR) methods, the measurement error of blade vibration induced by the speed fluctuation per revolution (SFPR) is investigated. Considering the effect of the SFPR, an improved multiple per revolution-based BTT method (MPR) for calculating the blade vibration is proposed. Furthermore, numerical simulation and experimental work of the MPR method are executed. Compared with the results of different technologies, the feasibility and reliability of the MPR method are verified by theoretical analysis and experimental verification. The MPR technology is also applied to blade vibration measurement of a large-scale Centrifugal Compressor and a large-scale Axial-Centrifugal Compressor. The conclusions of applications are well practical.

Low-Cost Tool Design for Pulse-Reverse Electrochemical Machining of Aerospace Alloys via Multiphysics Simulation

Electrochemical machining (ECM) is a manufacturing technology that precisely removes material from metallic workpieces via electrochemical means, typically oxidation, with subsequent dissolution/dispersal of the machining products into a working electrolyte. ECM is able to produce parts with complicated/intricate geometries, and is especially well suited for workpieces to be fabricated from “difficult to cut” materials, as no mechanical interaction occurs between tool and workpiece. In ECM, an electrochemical cell is established with the workpiece as the anode and a shaped tool as the cathode; a suitable electrical potential is then applied while the shaped tool is advanced into the workpiece, resulting in formation of a near-mirror image of the tool in the workpiece. Compared to mechanical or thermal machining processes, where material is subtracted by cutting or by electric/laser discharges, respectively, ECM does not suffer from tool wear or introduce a heat-affected zone to the workpiece. As a result, ECM has strong utility as a manufacturing technology for a wide variety of workpiece materials and ultimate part geometries, and encompasses machining, deburring, boring, radiusing and polishing processes, among others. As noted, ECM provides particular value when applied to high-strength/tough, brittle, galling-prone and/or work-hardening materials such as high strength steels, nickel alloys (e.g., IN718), titanium alloys (e.g., Ti-6Al-4V), cobalt-chrome alloy (Stellite 25), tantalum-tungsten alloy (Ta10W), molybdenum and tungsten.One challenge in practical execution of ECM is the tooling design process, which can require many costly and slow prototyping iterations and which often represents a significant barrier to entry for smaller manufacturing operations that might otherwise benefit from implementation of ECM. Multiphysics simulation has the potential to replace a significant fraction of the initial experimental prototyping effort with rapid, low-cost in-silico methods. This talk will present recent work by Faraday demonstrating a pulse-reverse ECM approach for fabrication of shaped Ni-alloy (IN625/IN718) and Ti-alloy (Ti64/Ti6242) workpieces, along with comparison of the experimentally-obtained results to simulated machined forms produced using the COMSOL Multiphysics® software package. For these preliminary simulation efforts, a 3D primary current distribution model was used to solve for the time-dependent electric field between a moving cathode tool and a workpiece surface, simulating deformation owing to metal dissolution. The simulated and experimental workpiece surfaces showed a satisfactory match for the complex, non-axisymmetric feature geometry embodying a notional form relevant to aerospace turbine and compressor blade manufacture, especially given the preliminary nature of the simulations implemented to date. A cost and lead-time analysis comparing a design process involving all-experimental tool prototyping iterations with one in which initial prototyping is performed computationally reveals the potential for significant cost/time reductions in the latter scenario. Thus, introduction of multiphysics modeling prototyping iterations to the tool design process has the potential to significantly reduce the barrier to entry for newcomers to ECM fabrication. Figure Caption Left to right: Non-axisymmetric tool fabricated by additive-manufacturing methods. Photograph and optical profilometry scan of indentation obtained by pulse-reverse waveform ECM. Indentation form predicted by multiphysics simulation. Figure 1

Damage evolution behavior of TiN/Ti multilayer coatings under high-speed impact conditions

The erosion of compressor blades by solid particles brings huge safety hazards to aero engines. The application of erosion-resistant hard coating has become a good solution to protect titanium alloy blades. In this work, TiN monolayer and TiN/Ti multilayer coatings were prepared on Ti-6Al-4V alloy using a magnetic filtered cathode vacuum arc system (FCVA) and a metal vapor vacuum arc (MEVVA) ion implantation device. The mechanical properties of the coatings were evaluated by nanoindentation and scratch tests. The damage evolution behavior and mechanism of TiN monolayer and TiN/Ti multilayer coatings are studied by single-particle impact test and finite element simulation. The results show that the fracture toughness of TiN monolayer coating is lower than that of TiN/Ti multilayer coating; TiN/Ti multilayer coatings exhibit good erosion resistance. The main reason lies in that the Ti layer dissipates part of impact energy by plastic deformation and reduces the tensile stress. The interfaces in the multilayer coating inhibited the propagation of cracks.

Effects of Specimen Size and Welded Joints on the Very High Cycle Fatigue Properties of Compressor Blade Steel KMN-I

The effects of specimen size and welded joints on the very high cycle fatigue properties of compressor blade steel KMN-I were studied by ultrasonic fatigue testing. It was found that the S-N curve of large specimens had a slow decline above 107 cycles, and fatigue failure still occurred in the very high cycle regime (>107 cycles), while the very high cycle fatigue characteristics of welded specimens was less obvious, and the fatigue limit was observed. Metallographic observation and SEM analysis were carried out on the fracture of the specimens. The results showed that surface fractures were mostly observed in the large specimens, and only a small number of cracks initiated from non-metallic inclusions above 107 cycles. The cracks of welded specimens initiated from the surface below 107 cycles and initiated from the internal matrix above 107 cycles. In addition, the formation mechanism of GBF (granular bright facet) was analyzed by the “dispersive decohesion of spherical carbide” theory, and the fatigue strength and fatigue life were predicted, which was consistent with the experimental results.

Dependence of thermogasodynamic parameters of operation of a helicopter gas turbine engine on its operation and dust in the atmosphere

By scanning the blades of the impeller and the guide devices of the full-scale compressor of the helicopter gas turbine engine, a solid-state design model of the compressor with a tunable geometry in the height of the blades was developed. The empirical dependencies of the values of nonlinear wear of the blade of the 1st stage of the compressor on the operating time and the concentration of dust in the air are presented. A block diagram of the procedure for calculating the characteristics and parameters of a helicopter gas turbine compressor is presented. When calculating the flow in the gas–air path of the compressor, the numerical solution of the Navier–Stokes equations averaged by Reynolds, the finite element method in combination with the establishment of patterns of erosive wear of the blades depending on the operating conditions of the engine is used. Mathematical dependencies of the thermogasodynamic parameters of the helicopter gas turbine compressor operation on its operating time and the dustiness of the atmosphere are obtained using modeling. The obtained results can be used in the development of a methodology for automated monitoring of the wear condition of the compressor blades of a helicopter gas turbine engine according to its thermogasodynamic parameters, applied to various climatic operating conditions.

Effect of Laser Surface Texturing on Coating Adherence and Tribological Properties of CuNiIn and MoS2 Coating

Copper nickel indium (CuNiIn) and Molybdenum disulphide (MoS2) duplex coating is used in aero engine compressor blades of Ti6Al4V to prevent fretting failure. Surface before coating is roughened by grit blasting process for coating adhesion. In present study, in place of grit blasting, laser surface texturing (LST) process is used for surface preparation and compared with grit blasted surface. Four different elliptical and square patterns are created by LST on Ti6Al4V. Difference of surface topography between LST and grit blasted surface is analyzed by SEM and white light interferometery. CuNiIn is deposited on Ti6Al4V by atmospheric plasma spray. MoS2 lubricating coating is deposited on top of CuNiIn by spray painting and curing. Effect of difference in surface preparation on coating microstructure as well as on tribological properties is studied. The results showed that geometry and dimensions of LST pattern influence the coating adherence and wear performance. LST process can be optimized for better performance and explored as an alternative surface preparation process in industry for thermal spray coatings.

Multiaxial fatigue damage and reliability assessment of aero-engine compressor blades made of TC4 titanium alloy

This paper discusses the multiaxial fatigue damage and reliability evaluation of aero-engine compressor blades using an improved strength degeneration scheme based on Chaboche damage model combined with stress-strength interference (SSI) approach. Firstly, the multiaxial proportion and non-proportion fatigue tests of TC4 (Ti-6Al-4V) alloy under a series of loading paths are carried out. The nonlinear fatigue damage calculation and life prediction of TC4 alloy is described by combining the Chaboche nonlinear damage criterion and critical plane method with proposed damage control parameter. Secondly, a nonlinear strength degradation model based on accumulative damage is further developed, and it is proved that the model can get greater precision under different loading conditions. Finally, a revised time-varying reliability model is established by viewing mechanical characteristic of compressor blade set as interconnected system with variant load component on each blade. According to the results of Monte Carlo simulation, the proposed revised model has a better prediction precision, with the error controlled less than 5%.

A Parametric Blade Design Method for High-Speed Axial Compressor

The blade geometry design method is an important tool to design high performance axial compressors, expected to have large design space while limiting the quantity of design variables to a suitable level for usability. However, the large design space tends to increase the quantity of the design variables. To solve this problem, this paper utilizes the normalization and subsection techniques to develop a geometry design method featuring flexibility and local adjustability with limited design variables for usability. Firstly, the blade geometry parameters are defined by using the normalization technique. Then, the normalized camber angle f1(x) and thickness f2(x) functions are proposed with subsection techniques used to improve the design flexibility. The setting of adjustable coefficients acquires the local adjustability of blade geometry. Considering the usability, most of the design parameters have clear, intuitive meanings to make the method easy to use. To test this developed geometry design method, it is applied in the design of a transonic, two flow-path axial fan component for an aero engine. Numerical simulations indicate that the designed transonic axial fan system achieves good efficiency above 0.90 for the entire main-flow characteristic and above 0.865 for the bypass flow characteristic, while possessing a sufficiently stable operation range. This indicates that the developed design method has a large design space for containing the good performance compressor blade of different inflow Mach numbers, which is a useful platform for axial-flow compressor blade design.

Combined flow control with full-span slot and end-wall boundary layer suction in a large-camber compressor cascade

To investigate combined flow control strategies with full-span slot and end-wall (EW) boundary layer suction, a large-camber compressor cascade was treated with a full-span slot configuration firstly. Then, two combined flow control schemes with different EW boundary layer suction slots were set up based on the full-span slot configuration. Finally, the performance of the baseline cascade and the slotted cascades was evaluated and compared numerically. The results show that the full-span slotted scheme can eliminate the trailing edge (TE) separation near the blade midspan and weakens the corner separation. However, the EW secondary flow still develops on the blade suction surface (SS) before the full-span slot outlet and separates to form the wall vortex (WV). In contrast, the two combined flow control schemes can effectively inhibit the development of the EW secondary flow on the blade SS, contributing to the smaller total pressure loss and stronger static pressure recovery capacity. Between them, the combined flow control scheme with full-chord-length EW suction slot behaves better. It can effectively eliminate the concentrated shedding vortex, WV and corner vortex in the cascade passage, and the passage vortex is also weakened to some extent. Therefore, the TE separation and the corner separation are basically eliminated. The total pressure loss, flow turning angle and static pressure coefficient can be increased by an average of −38.4%, 3.1° and 16.2% in the combined flow control cascade with full-chord-length EW suction slot respectively, which is beneficial for improving the stability margin and static pressure recovery capacity of compressor blades.

Numerical and Experimental Assessment of the Effect of Residual Stresses on the Fatigue Strength of an Aircraft Blade.

The work presents the results of numerical fatigue analysis of a turbine engine compressor blade, taking into account the values of initial stresses resulting from surface treatment-shot-peening. The values of the residual stresses were estimated experimentally using X-ray diffraction. The paper specifies the values of the residual stresses on both sides of the blade and their reduction due to cutting through the blade-relaxation. The obtained values of the residual stresses were used as initial stresses in the numerical fatigue analysis of the damaged compressor blade, which was subjected to resonant vibrations of known amplitude. Numerical fatigue ε-N life analysis was based on several fatigue material models: Manson’s, Mitchell’s, Baumel-Seeger’s, Muralidharan-Manson’s, Ong’s, Roessle-Fatemi’s, and Median’s, and also on the three models of cyclic hardening: Manson’s, Xianxin’s, and Fatemi’s. Because of this approach, it was possible to determine the relationship between the selection of the fatigue material ε-N model and the cyclic hardening model on the results of the numerical fatigue analysis. Additionally, the calculated results were compared with the results of experimental research, which allowed for a substantive evaluation of the obtained results. These results are of great scientific and practical importance. The problem of determining the fatigue life of blades with defects operating under resonance vibrations is one of the original tasks in the field of fracture mechanics and experimental mechanics. The results obtained are of great importance in the aviation industry and can be used during engine maintenance and inspections to assess the suitability of blades with defects in terms of the needs of further work. This aspect of engineering maintenance is of great importance from the aircraft safety point of view.

Performance Study of Novel Compressor Blades in a Two-Dimensional Cascade—Transonic Regime

Performance Study of Novel Compressor Blades in a Two-Dimensional Cascade—Transonic Regime

Метод оптимизации многоступенчатых компрессоров

The paper presents an algorithm for seeking an optimal blade configuration for multistage axial-flow compressors. The primary tool behind the algorithm is 3D CFD simulation, augmented by commercial optimisation software. The core of the algorithm involves feeding an initial data vector to the parametric simulation module so as to form a "new" blade geometry, which is then transferred to 3D computational software. The results obtained are further processed in a program that implements the algorithm for seeking the optimum and forms a new input data vector to achieve the set goal. We present a method of parametrically simulation the blade shape, implemented in a software package, making it possible to describe the shape of the compressor blade profiles using a minimum number of variables and to automatically change the shape in the optimisation cycle. The algorithm developed allows the main parameters of compressor operation (efficiency, pressure ratio, air flow rate, etc.) to be improved by correcting the profile shape and relative position of the blades. The algorithm takes into account various possible constraints. We used the method developed to solve practical problems of optimising multistage axial compressors of gas turbine engines for various purposes, with the number of compressor stages ranging from 3 to 15. As a result, the efficiency, pressure ratio and stability margin of gas turbine engines were increased

Fabrication and Testing of Rocket Engine Construction Elements by Addictive Production Approach

For the first time, using the technology of selective laser sintering, prototypes of rocket engine compressor blades were manufactured with subsequent analysis of the strength, technological, physical and mechanical characteristics of the product. The physical and mechanical properties of the manufactured blades were investigated, it was found that the short-term strength limit at 20 °C is 1450 MPa, and at 300 °C the ultimate strength is 1300 MPa, thus thermal losses in deformation resistance are no more than 12%, which allows the material to be used in aircraft construction, including for supersonic aircraft

Vision-based damage detection of aircraft engine’s compressor blades

1. Masłyk M, Obrocki W, Setkowicz A, Sieniawski J. Experimental fatigue strength determination of damaged aircraft engine blades. Advances in Manufacturing Science and Technology. 2016;40(4):55:66. https://doi.org/10.2478/amst-2.... CrossRef Google Scholar

Experimental Investigation of Wash Fluid Preheating on the effectiveness of online Compressor Washing in a Suction Tunnel Compressor Cascade

Online compressor washing is a promising method of preventing/ recovering the effects of fouling on the compressor blades. However, proper strategies need to be implemented for online compressor washing to be effective since it is conducted when the engine is in operation. This study presents an experimental investigation of wash fluid preheating on the effectiveness of online compressor washing in a suction wind tunnel compressor cascade. Crude oil was uniformly applied on the compressor cascade blade surfaces using a roller brush, and carborundum particles were ingested into the tunnel to create accelerated fouled blades. Demineralized water was preheated to 500C using the heat coil provided in the tank. Washing of the fouled blades were conducted using single flat fan nozzle, where the preheated and non preheated demineralized water were used separately to wash the fouled blades. When fouled blades washed with preheated demineralized and the non-preheated were compared, it was observed that there was little or no difference in terms of total pressure loss coefficient and exit flow angle. However, when the fouled and washed cases were compared, there was a significant difference in total pressure loss coefficient and exit flow angle.

Dimensionality-reduction-based surrogate models for real-time design space exploration of a jet engine compressor blade

Design space exploration (DSE) is a systematic analysis of designs based on parameters that are varied. DSE, however, often involves computationally expensive analyses that require large amounts of time and computer resources. To overcome this obstacle, tools and methods have been developed that allow the designer to quickly explore the design space. One such tool is surrogate modeling. A surrogate model's predictions are based off the relationship between the inputs and the outputs. If this relationship is sufficiently nonlinear, a nonlinear dimensionality reduction technique may unfold the data and create a linear relationship by discovering the data's true dimensionality. The resultant surrogate has the potential to be more accurate due to the simplified relationship between the inputs and outputs. This paper uses surrogate models based on dimensionality reduction to predict finite element analysis (FEA) nodal properties of a jet engine compressor blade. Four different dimensionality reduction methods are compared, namely PCA, KPCA, ISOMAP, and LLE. The results show that nonlinear dimensionality-reduction-based surrogate models can reduce surrogate error in sufficiently nonlinear data spaces. To measure the linearity of the manifold, a new term, called the manifold distance ratio (MDR), is introduced. Nonlinear dimensionality-reduction-based surrogate models can have less error than PCA-based surrogate models when the MDR >1.7. For our application, ISOMAP-based surrogate models decreased the mean error of the surrogate for predicting nodal stresses by 35.7% compared to PCA. The lowest error in predicting the nodal coordinates was the inverse-transform-based PCA. While all errors in nodal coordinates were small, less than 0.23%, the other linear and nonlinear algorithms had between 2.4 and 3.2 times more error.

Дослідження напруженого стану міжпазових виступів диска робочого колеса компресора з урахуванням впливу відцентрових сил та розладу частот коливань лопаток

The authors present the results of the numerical experiments on the determination of the effect of compressor blade mistuning on the stress state of a blade-disk joint under the action of the centrifugal forces and their comparison with results of full-scale experiments. The analysis of the results of the previously performed experimental and numerical experiments on the determination of the possible sources of blade-disk joints failure shows their stress state strongly depends on the disk rim thickness and the blade vibration mode and their mistuning. It indicates the need for further studies to determine the causes of cracks and possible fracture of blade-disk joints in the compressor of a gas-turbine engine. The task is solved in the variation of the dovetail-rim region thickness of the disk of the 1st stage of the low-pressure compressor of the gas-turbine engine D-36, which has blades without shrouds. The finite element models of the blade assembly, are used to perform the investigation. Obtained results of the vibration stress analysis of blade-disk joints in the compressor both modifications under the action of the centrifugal forces have good agreement with the data of full-scale experiments and numerical data using rod models. The presented results of the numerical experiments indicate the possibility of reducing vibration stresses by analyzing and choosing parameters that affect the rigidity of the coupling of the blades, which is of practical importance for increasing the vibration reliability of the compressors of gas-turbine engines.

Failure Analysis of a Fighter Jet Engine Compressor Blade Due to Foreign Object Damage and High-Cycle Fatigue

Failure Analysis of a Fighter Jet Engine Compressor Blade Due to Foreign Object Damage and High-Cycle Fatigue

Development of an accelerated test for pitting corrosion of CUSTOM 450 by rDHM

Abstract A key application of CUSTOM 450 alloy is in the construction of gas turbine compressor blades. The study of pitting corrosion can prevent the failure of many gas turbine compressor blades. In this study, a reflective digital holography microscopy method was employed to investigate the growth of pitting corrosion in depth. To this end, a constant potential of 350 mVSCE in a 3.5 wt.% NaCl solution was applied to the specimen. The generated pits were simulated in three dimensions, and it was indicated that pitting corrosion rate was decreased as time passed. Comparing the obtained experimental data with the data gathered from the real industrial environment surrounding a compressor installation, an accelerated test was proposed. By the proposed accelerated test, it is possible to produce a pit similar to the one that will be initiated and propagated at any time in the future in real conditions.