Compressor Stator Blade Research Articles

Research on the influence of spanwise cross-flow on the boundary layer transition of compressor cascade

The cross-flow perpendicular to the inviscid main flow in the boundary layer has potential instability, causing the transition from laminar flow to turbulent flow. In order to explore the mechanism of cross-flow in the blade boundary layer on transition, this paper studies the rectangular cascade of a certain compressor stator blade. Large eddy simulation calculations and flow display experiments for six attack angles with end wall cascades were carried out. It is found that the disturbance is dominated by the two-dimensional Kelvin–Helmholtz (K–H) instability. The transition begins at the position where the separation bubble begins to fall off into a two-dimensional K–H vortex and is completed where the K–H vortex breaks. The closer to the blade root, the later the transition occurs and the smaller the total pressure loss. The cross-flow velocity develops alternately between positive and negative, showing severe instability with more than 4 inflection points. The study on variable angles of attack shows that there is a superposition of two mechanisms, namely, separation bubble transition and cross-flow transition, at an angle of attack from −4° to 10°. In summary, although the separation bubble transition is dominated by K–H vortices, the occurrence of cross-flow instability is closely related to the transition position.

Proof of concept study on a self-driven pulsed jet on a compressor stator blade by numerical simulation

This study presents a new concept of self-driven pulsed jet flow control on a compressor stator blade. This passive unsteady flow control method has the advantage that neither external flow nor electrical source is needed. This study’s preliminary proof-of-concept study on a low-speed compressor stator blade is performed using numerical simulation. When the pulsed jet frequency is 100 Hz, the optimum control performance and control efficiency are reached, and the total pressure loss coefficient is reduced by 8.9%. As the valve’s rotational speed increases, the pulsed jet’s momentum coefficient decreases gradually. The analysis of the unsteady characteristics of the self-driven pulsed jet shows that the jet velocity is close to a periodic square wave signal, and the typical reduced jet velocity ranges from approximately 0.15 to 0.8. Moreover, the time-averaged driving torque on the valve depends on the rotational speed and is relevant to the self-starting and self-driven characteristics of this passive flow control method. Under different bearing resistance torque, the self-driven valve behaves differently in three cases that can self-start and be self-driven, cannot self-start but can be self-driven, and cannot self-start nor be self-driven.

Experimental and Numerical Study of Nonlinear Vibration in Compressor Stator Blade with Holding Ring Structure Considering Contact Friction

At present, there are very few vibration test designs for compressor stator blades in damped blade test research, and there is a lack of consideration of the holding ring structure of the blade installation. In recent years, compressor designs have become more elaborate, and their structures have become more compact. For improving compressor stator blade design, it is necessary to study the blade vibration response considering contact and friction. In this paper, a compressor stator blade vibration test system with a holding ring structure is designed and built. The composition of the system and the flow of vibration are tested. Then, the nonlinear vibration rules of the model stator blades are experimentally studied through the built test bench. The experimental conditions include excitation forces of 1 N, 2 N, 3 N and 4 N, respectively. At the same time, the normal load of the outer holding ring and the blade contact surface includes 10 N, 20 N, 30 N and 40 N working conditions. The nonlinear vibration law of the test model blade considering contact friction is analyzed using a numerical method and experimental data. The experimental data agree well with the numerical analysis results, and the relative error of measurement is basically less than 5%. It is concluded that the damping vibration rule of the test model blade is related to the exciting force applied by the exciter and the contact surface normal load applied by the adjusting bolt. When these two parameters change, the movement behavior of the contact surface will switch between macro-slip, fretting slip and even a viscous state. Therefore, the amplitude of the blade will change and show nonlinear vibration characteristics. The experimental data in this paper can provide a reference for research on stator ring vibration safety in compressors.

Numerical Study on Vibration Response of Compressor Stator Blade Considering Contact Friction of Holding Ring

There are many kinds of assembly structures in heavy-duty gas turbine compressor stator blades which have significant influence on the complex damping vibration characteristics. At present, compressor design is becoming more and more compact, so it is very meaningful to accurately obtain stator blade vibration characteristics of structures with contact damping. Firstly, a fretting slip friction dynamic model was introduced, and then a vibration analysis model of the compressor stator blade with outer ring structure was established based on the slip friction theory. Then, the vibration response of the compressor stator blade was obtained according to various working conditions, and the main factors affecting the vibration characteristics of the stator blade were revealed. Finally, the vibration response of the blade under a particular exciting force condition was simulated. The results show that the damping vibration characteristics of the compressor stator blades were affected by the excitation force and the normal load of the contact surface. The vibration response curve of the stator blade and the equivalent stiffness coefficient of the contact surface were analyzed, and the friction motion of the contact surface changed with the change of the working condition. The model can simulate the nonlinear vibration characteristics of stator blades. This method provides a reference for the vibration safety analysis of compressor stator blades.

Prediction Method of Unsteady Flow Load of Compressor Stator under Working Condition Disturbance

Due to the complexity of the compressor operating conditions and the existence of various disturbances and unsteady effects in the flow field, the analysis of compressor stator vibration characteristics becomes particularly critical. The convolutional neural network model combined with a transient CFD method was introduced to solve the difficulty of analyzing the flow load of the compressor stator blade. This paper mainly focuses on two key points: the complex change of the aerodynamic load and the accurate prediction of the blade excitation. Considering the stator–rotor interference, the unsteady effects, and the variable working condition characteristics, the random disturbance analysis model of the flow field boundary was generated to simulate the unsteady flow excitation of the stator under complex working conditions. By establishing the neural network of boundary disturbance and flow excitation characteristics, the prediction model was trained and generated under the support of large-scale data. The most important role of the model was to establish the end-to-end data mapping between the disturbance condition and the aerodynamic load of the stator blade. The conclusions demonstrate that the introduction of an airflow disturbance is helpful to obtain the excitation characteristics of the stator under complex working conditions. The model established in this paper based on 1000 groups of disturbed working condition data can effectively predict the aerodynamic load of the blades under complex working conditions. In addition, the construction of the model is beneficial for saving a lot of computing resources, and the prediction accuracy also reaches a good level. The method presented in this paper provides a reference for the vibration analysis of the compressor stator.

Heat Transfer, Thermal Stress and Failure Inspection of a Gas Turbine Compressor Stator Blade Made of Five Different Conventional Superalloys and Ultra-High-Temperature Ceramic Material: A Direct Numerical Investigation

Heat Transfer, Thermal Stress and Failure Inspection of a Gas Turbine Compressor Stator Blade Made of Five Different Conventional Superalloys and Ultra-High-Temperature Ceramic Material: A Direct Numerical Investigation

Study of the stability of the geometric dimensions of wax models of GTE blades using rapid prototyping technology

The work presents the results of a study of the stability of geometric dimensions of wax models of compressor stator blades of the 15th stage, manufactured using rapid prototyping technology. In the process of performing the work, the algorithm for assigning the total shrinkage was studied. It is established that the deviation of geometric dimensions, both in the greater and in the lesser direction relative to the nominal size of the blade is caused by the amount of volumetric shrinkage of the material used. Measuring the linear dimensions of wax models of blades allowed us to determine the nature of the geometry deviation for further accounting and compensation.

Aerodynamic Design Optimization of an Axial Flow Compressor Stator Using Parameterized Free-Form Deformation

AbstractThis paper describes an aerodynamic design optimization of a highly loaded compressor stator blade using parameterized free-form deformation (FFD). The optimization methodology presented utilizes a B-spline-based FFD control volume to map the blade from the object space to the parametric space via transformation operations in order to perturb the blade surface. Coupled with a multi-objective genetic algorithm (MOGA) and a Gaussian process-based response surface method (RSM), a fully automated iterative loop was used to run the optimization on a fitted correlation function. A weighted average reduction of 6.1% and 36.9% in total pressure loss and exit whirl angle was achieved, showing a better compromise of objective functions with smoother blade shape than other results obtained in the open literature. Data mining of the Pareto set of optimums revealed four groups of data interactions of which some design variables were found to have skewed scatter relationship with objective functions and can be redefined for further improvement of performance. Analysis of the flow field showed that the thinning of the blade at midspan and reduction in camber distribution were responsible for the elimination of the focal-type separation vortex by redirecting the secondary flow in an axially forward direction toward the midspan and near the hub endwall downstream. Furthermore, the reduction in exit whirl angle especially at the shroud was due to the mild bow shape which generated radial forces on the flow field thereby reducing the flow diffusion rate at the suction surface corner. This effect substantially delayed or eliminated the formation of corner separation at design and off-design operating conditions. Parameterized FFD was found to have superior benefits of smooth surface generation with low number of design variables while maintaining a good compromise between objective functions when coupled with a genetic algorithm.

Optimization of geometry blade for modern high pressure compressor

An integration of CFD-3D codes and modern efficient methods of optimization is essential for the contemporary compressors improvement. In order to solve optimization tasks the insular information exchange circle was created. The data circulates between the mathematical model of the stream in a compressor, the algorithm of compressor’s geometry construction based on parameterization and algorithm of optimization itself. The mathematical model of the compressor covers procedures of mesh generation and calculations using CFD code. This model has been verified by experimental data.A high-pressure compressor of multi regime gas turbine engine for supersonic civil aircraft has been considered as the object of our study. The goal of the research was to increase the efficiency in the design point with the ensured preset gas dynamic parameters under other modes. The variables for optimization task were installation angles and the shape of the middle line of the profile in three sections along the height of the compressor rotor and stator blades. The total amount of independent variables was 72. As a result of the study the optimal geometry for the stator and rotor blades in supersonic aircraft engine has been obtained. The initial project efficiency had been improved by 1-2%, while stability margin remained constant.

Phase-averaged flow statistics in compressors using a rotated hot-wire technique

A technique based on a rotated hot wire has been developed to characterise the unsteady, three-dimensional flow field between compressor blade rows. Data are acquired from a slanted hot wire rotated through a number of orientations at each measurement point. Phase-averaged velocity statistics are obtained by solving a set of sensor response equations using a weighted, non-linear regression algorithm. The accuracy and robustness of the method were verified a priori by conducting a series of tests using synthetic data. The method is demonstrated by acquiring a full set of phase-averaged flow statistics in the wake of a compressor stator blade row. The technique allows three components of phase-averaged velocity, six components of phase-averaged deterministic stress, and six components of phase-averaged Reynolds stress to be recovered using a single rotated hot-wire probe.

DETERMINATION OF SUITABLE ELECTROLYTE FOR MAXIMUM MRR AND THE BEST SURFACE FINISH FOR MACHINING TITANIUM ALLOYS BY ECM

The rapid developments in recent years of the advanced technological industries like aerospace, nuclear power etc. has been accompanied by an increasing use of difficult-to-machine like titanium and nimonic. Some Aerospace industries manufacture R-29B engine which is fitted into MIG-27M aircraft. In this engine, six out of eleven stages of compressor rotor blades and compressor stator blades upto fifth stage are made of titanium alloys. A great deal of problems is encountered in machining these blades conventionally because of some inherent mechanical properties.These problems are multiplied due to the complexity of the blade aerofoil surface. All these problems result in frequent resharpening of tools and cams, higher machining times due to lower speed, feed & depth of cuts, higher inventory of tools and incorporation of an intermediate belt grinding process to give required shape, size and finish to the blade. One of the methods of conveniently machining them is by electro chemical machining. (ECM). In this paper, several experiments have been carried out to find out suitable electrolyte(s) among four electrolytes, namely Sodium Chloride (NaCl), Sodium Nitrate (NaNa3), Potassium Bromide (KBr) & Sodium Sulphate (Na2So4) for machining titanium alloys by ECM. Experiments have been conducted by using ANOVA”to optimize the results.

Failure Analysis of a Compressor Blade of Gas Turbine Engine

Abstract The stage II compressor stator blade of a developmental gas turbine engine was found damaged during dismantling of the engine after test run. A portion of the blade was found fractured from the hub region at leading edge. A crack was also observed extending from the fractured surface towards the centre of the airfoil region of the blade. Low magnification stereo-binocular observation revealed presence of beach marks on the fractured surface indicating the blade failure in progressive mode. This observation was further confirmed by scanning electron microscopy. The crack origin was at the blade hub-stem junction on the leading edge side. Presence of machining/filing marks appeared to be the reason for the fatigue crack initiation from this region. No metallurgical abnormalities were present at the crack origin. However, deep filing/machining lines were observed at the stem region of the blade attributing to the cause of failure.

Fracture of a Compressor Stator Blade in a Gas Turbine Engine

Abstract On the occasion of the revision of a gas turbine power plant in Northern Germany, a surface crack was detected in one of the 44 stage 0 compressor stator blades during a routine non-destructive testing of the blades. By then the engine had completed more than 38,000 h of operation. The damaged blade was precision drop-forged and copy-milled of X20Cr13 and for some time was wearing a subsequently applied corrosion protection coating containing aluminium pigments. The blade's condition before the subsequent coating was applied is unknown. The metallurgical investigation found that the cause of failure was corrosion fatigue.

Pitting Corrosion Induced Fatigue Fracture on a Gas Turbine Compressor Blade

Abstract According to VDI Guideline 3822-3 “this type of corrosion is … caused by the formation of anodic regions of slight local expansion on the electrolyte-wetted surface. The presence of protective layers (passive layers) is a prerequisite for the occurrence of pitting corrosion.” Both requirements are met on gas turbine compressor blades if water condensates on the turbine blades when the machine is at standstill and/or if the compressor temperature remains low enough in a running machine to allow aqueous media to persist, i.e. for the condensation not yet to have evaporated. This is the case with the front row of blades, generally up to stage 6. Passive layers are present on the gas turbine compressor blades, since these parts are generally manufactured from martensitic 12–16% chrome steel, as also the case for steam turbine blades and discussed in detail in other case studies in the book. The failed compressor stator blade in this case study is made of X15Cr13, material no. 1.4024. Pitting corrosion on the front row compressor blades was not uncommon before the introduction of high temperature corrosion protection coatings containing aluminium pigments on these blades. After the introduction of protective coatings, this corrosion mechanism was absent on the coated blades. The coating systems contain aluminium spherules which, given the correct coating method, are in contact with each other to form a continuous electrically conducting layer between the surface of the blade and the protective coating, creating a sacrificial anode resulting in the less noble aluminium – as compared to steel – preferentially corroding to protect the blade material. The damaged part discussed in this case study originates from an older model gas turbine commissioned in 1955 already, having accumulated almost 163000 hours of operation when the damage was diagnosed. Depending on the type of operation, this represents a period of operation of 15 to 25 years. Such compressor blades were not coated back then.

Topological Studies of Three-dimensional Flows in a High Pressure Compressor Stator Blade Row without and with Boundary Layer Aspiration

This paper presents a numerical study of the flow topologies of three-dimensional (3D) flows in a high pressure compressor stator blade row without and with boundary layer aspiration on the hub wall. The stator blade is representative of the first stage operating under transonic inlet conditions and the blade design encourages development of highly complex 3D flows. The blade has a small tip clearance. The computational fluid dynamics (CFD) studies show progressive increase of hub corner stall with the increase in incidence. Aspiration is implemented on the hub wall via a slot in the corner between the hub wall and the suction surface. The CFD studies show aspiration to be sensitive to the suction flow rate; lower rate leads to very complex flow structures and increased level of losses whereas higher rate renders aspiration effective for control of hub corner separation. The flow topologies are studied by trace of skin friction lines on the walls. The nature of flow can be explained by the topological rules of closed separation. Furthermore, a deeper analysis is done for a particular case with advanced criterion to test the non-degeneracy of critical points in the flow field.

Lifetime evaluation of a thick thermal barrier coated superalloy used in turbine blade

This paper deals with evaluation of lifetime of a thick thermal barrier coated (TBC), AE-437A Ni base superalloy mostly employed for manufacturing compressor and stationary stator blades in aero turbines from accelerated creep tests. Severe high-temperature oxidation is likely the cause of the reduced lifetime of the bare substrate while accelerated creep experiments are carried out in an oxidizing environment. The bond coat was capable of supporting some load and the apparent equivalence of 800 °C stress rupture properties of thickly and thinly coated AE-437A was not anticipated since the thick TBC coating with its 160 μm bond coat represents an ∼20% increase in cross-sectional area for a 3 mm × 4 mm sample; whereas thin TBC with a 100 μm bond coat has a ∼12% increase in cross-section for the same initial sample size. Delamination of top coat layer and bond coat from the substrate occurred at very high creep stress, whereas delamination of TBC from the bond coat was evident at intermediate creep stress level. It was found that furnace thermal cycling (10–50, 1-h thermal cycles) on the TBC samples prior to creep loading at 70 MPa/800 °C, drastically reduces the lifetime or rupture life of the TBCs. The damage generated by this micro-cracking is expected to be a primary life-limiting factor.

Evaluation of the Lifetime and Effect of Furnace Thermal Cycling on Stress Rupture Behaviour of a Thermal Barrier Coated Superalloy Used in a Turbine Blade

This paper highlights the trends recently observed while evaluating the lifetime of a thermal barrier coated (TBC) AE-437A Ni base superalloy mostly employed for manufacturing compressor and stationary stator blades in aero turbines from accelerated creep tests in air at 800°C. Based on accelerated stress rupture properties, the TBC specimens had superior life in terms of rupture hours compared to that of the bare substrate as severe high temperature oxidation is likely the cause of the reduced lifetime of the bare substrate while accelerated creep experiments are carried out in an oxidizing environment. The effect of furnace thermal cycling on accelerated creep properties of the TBC samples at 70 MPa and 800°C was also studied. It was observed that furnace thermal cycling (10 to 50, one hour thermal cycles) prior to creep loading, drastically reduces the rupture life of the TBCs. The life of thermal barrier coating prior to delamination and spallation is dominated by micro-cracking in both the thermally grown oxide and the yttria stabilized zirconia top coat. The damage generated by this micro-cracking is expected to be a primary life limiting factor. Delamination of top coat layer and bond coat from the substrate occurred at very high creep stress, whereas delamination of TBC from the bond coat was evident at intermediate creep stress level.RésuméCet article souligne les tendances observées récemment lors de l'évaluation de la durée de vie du superalliage AE-437A à base de Ni, revêtu d'une barrière thermique (TBC), utilisé principalement dans la fabrication d'aubes de compresseur et de stator stationnaire dans les turbines éoliennes, à partir d'essais de fluage accéléré à l'air, à 800°C. D'après les propriétés de fracture sous contrainte accélérée, les échantillons de TBC avaient une durée de vie supérieure en termes d'heures à la rupture par rapport à celle du substrat nu puisque l'oxydation sévère à haute température est probablement la cause de la durée de vie réduite du substrat nu alors que les expériences de fluage accéléré sont effectuées dans un environnement oxydant. On a également étudié l'effet du cycle thermique en fournaise sur les propriétés de fluage accéléré des échantillons de TBC à 70 MPa et 800°C. On a observé que le cycle thermique en fournaise (10 à 50 cycles thermiques d'une heure), avant le chargement en fluage, réduisait radicalement la durée de vie de rupture des TBCs. La durée de vie d'un revêtement de barrière thermique avant le délaminage et l'éclatement est dominée par la microfissuration tant dans l'oxyde de croissance thermique que dans le revêtement supérieur de zircone stabilisée à l'oxyde d'yttrium. On s'attend à ce que le dommage engendré par cette microfissuration soit un facteur primaire de limitation de la durée de vie. Le délaminage de la couche supérieure de revêtement et de la couche d'ancrage du substrat se produisait à des charges très élevées de fluage, alors que le délaminage du TBC de la couche d'ancrage était évident à un niveau intermédiaire de charge de fluage.

Mechanical property and characterization of a NiCoCrAlY type metallic bond coat used in turbine blade

This paper highlights some of the results obtained while studying Ni20Co18Cr12.5Al0.6Y (NiCoCrAlY) type metallic bond-coat properties of a thermal-barrier coated (TBC), AE-437A Ni base superalloy mostly employed for manufacturing compressor and stationary stator blades in aero turbines. Experiments were mainly focused in the area of evaluation of microstructure, residual stress, shear strength, hardness and with special emphasis in establishing the ductile to brittle transition temperature (DBTT) of the bond coat by using acoustic emission technique during room temperature and high temperature tensile tests. Results reveal that the residual stress was tensile in nature in the TBC layer and compressive in the bond coat as well as in the substrate. The DBTT of this bond coat is around 650 °C, which is in close proximity to the value reported in literature for CoCrAlY type of bond coat. Finite element technique was used to analyze the equivalent stresses in the bond coat, the result of which revealed the highest order of equivalent stress 800 °C, as the bond coat is ductile above 650 °C. Shear strength of the bond coat is in close proximity with that of the bond strength reported in literature for CoCrAlY and Ni22Co17Cr12.5Al0.6Y types of bond coat.

Unsteady Transition Phenomena at a Compressor Blade Leading Edge

Wake-induced laminar-turbulent transition is studied at the leading edge of a C4-section compressor stator blade in a 1.5-stage axial compressor. Surface hot-film sensor observations are interpreted with the aid of numerical solutions from UNSFLO, a quasi-three-dimensional viscous-inviscid flow solver. The passage of a rotor wake, with its associated negative jet, over the stator leading edge is observed to have a destabilizing effect on the suction surface boundary layer. This leads to transition closer to the stator leading edge than would have occurred under steady flow conditions. The strength of this phenomenon is influenced by the rotor-stator axial gap and the variability of individual rotor wake disturbances. A variety of transition phenomena is observed near the leading edge in the wake path. Wave packets characteristic of Tollmien-Schlichting waves are observed to amplify and break down into turbulent spots. Disturbances characteristic of the streaky structures occurring in bypass transition are also seen. Examination of suction surface disturbance and wake-induced transitional strip trajectories points to the leading edge as the principal receptivity site for suction surface transition phenomena at design loading conditions. This contrasts markedly with the pressure surface behavior, where transition at design conditions occurs remotely from leading-edge flow perturbations associated with wake chopping. Here, the local receptivity of the boundary layer to the wake passing disturbance and turbulent wake fluid discharging onto the blade surface may be of greater importance.

Three-dimensional inverse design of axial compressor stator blade using neural-networks and direct Navier–Stokes solver

In this article we describe a new method for the aerodynamic optimisation and inverse design problem resolution. This method is based on the coupling of a classical optimiser with a neural-network. A Navier–Stokes flow solver is used for an accurate computation of the objective function. At first the neural-network, which has been trained by an initial small database, is used to obtain, by the interpolation of the design sensitivities, a new design point, which is then computed by the Navier–Stokes solver in order to update the neural-network training database for further iterative step. Since the neural-network provides the optimiser with the derivatives, the objective function has to be evaluated only once at every step. By this method, the computational effort is significantly reduced with respect to the classical optimisation methods based on the design sensitivities, that are computed directly by the flow solver. The method proposed has been positively tested on the inverse design of a three-dimensional axial compressor blade, and a summary of the results is provided.