Compressor Rotor Blade Research Articles

Deposition of thick nitrides and carbonitrides for sand erosion protection

Thick nitrides (ZrN, CrN and TiN) and carbonitrides (ZrSiCN and TiSiCN) have been deposited using a Plasma Enhanced Magnetron Sputtering (PEMS) technique. The technique combines a conventional magnetron sputtering and an independently generated plasma from which high current density can be obtained. By using heavy ion bombardment prior to and during deposition to increase the coating adhesion and limit columnar growth, single-layered thick nitrides of ZrN, CrN, and TiN coatings up to about 80 μm and thick carbonitride coatings of ZrSiCN and TiSiCN about 30 μm have been obtained. In this paper, we will discuss the deposition technology and the properties of the thick coatings. Scanning Electron Microscopy (SEM) with Energy Dispersive Spectroscopy (EDS), and X-Ray Diffractometry (XRD) were used to study the microstructure and morphology of these coatings. Nanoindentation was conducted to determine the hardness and Young's modulus, while sand erosion tests were conducted to rank the erosion resistance of the coatings. It was observed that TiSiCN rendered the best erosion resistance — nearly 25 times higher than the uncoated stainless steel or Ti–6Al–4V and about 5–10 times higher than all other nitrides. The technology may be applied to protect turbine engine compressor blades, vanes and rotor blades in advanced aircraft and fluid pump impellers as well as piston rings for heavy-duty diesel engines.

ANALYSIS OF FLOW INSIDE OF A COMPRESSOR BLADE ROW(Fluid Machinery)

Three-dimensional blading has been used in the process of turbomachine designs. To meet the need for efficient turbine blade designs, CFD predictions of a complex 3D flow field in turbine blade passages had been used. Because the numerous advantages of 3-D CFD have been reported in the open literature, many industries already use 3D blading in their turbomachines. In addition, blade lean and sweep have been implemented to increase the blade row efficiency. Experimental studies have shown the advantages of these features. However, most of the experimental results combined other features together, making it difficult to determine the effects of individual features. The development of CFD techniques has made it possible to do three-dimensional turbulent flow analyses in a very short time. In this study, numerical studies are presented to study the sweep effects on a transonic compressor air foil. The focus of the paper is to investigate the sweep effects without changing other compressor blade features i.e. keeping the blade out flow angle the same for all cases. The purpose of the study is to enhance the understanding of the sweep in a transonic compressor rotor blade. The results showed that the sweeps redistributes the flow reducing the secondary flow loss, depending on the baseline. It was shown that forward sweep reduces the tip loading in terms of the static pressure coefficient.

Parametric investigation of entire annular recess casing treatment on compressor stable operation

Application of casing treatment in compressor has been found to give substantial improvement in compressor stable operating range. An experimental investigation has been carried out with one variant of casing treatment, an entire annular recess (EAR) over the compressor rotor blades. Various tip clearances and two groups of annular recesses (without and with penetration of rotor blade tips into casing) were used. The results for 20 such configurations are presented in the proposed paper, analyzed with respect to relative tip clearance and recess relative volume influence on compressor stable operating range.

Failure Analysis of Compressor Blade of vpical Fighter-class Aero-engine-A Case Study

Failure of second-stage compressor rotor blade of a typical Russian fighter-class aero-engine has resulted in few accidents and many incidents. Apart from this, during manufacturing stage there was high rejection rate of this rotor blade in vibrational fatigue tests. Metallurgical investigations of the failed rotor blade reveals that the mode of failure'is fatigue. The various reasons of this fatigue crack are micro-inclusion, silica embedment and hydrogen attack. This was analysed from the three different approaches, namely, theoretical stress analysis, basic design and quality control. To contain the possibility of silica embedment and hydrogen attack, sand electro-corrundum blasting and measure-replaced blasting have been taken to control hydrogen embedment.

Numerical Comparative Study of Compressor Rotor and Stator Blade Deposition Rates

Compressor fouling is generally accepted to be an important factor when monitoring the efficiency of an engine’s operation. However, there are not many studies related to the local fouling behavior of the individual components of the compressor. In the present paper, the CFD-ACE software package is used for the flow field calculation and the results are utilized to calculate the deposition rates on the blade surfaces of a conventional compressor stator and rotor. The deposition model takes into account the particle and surface material properties and the energy balance at the point of impact. A discussion is presented regarding the various mechanisms that produce the final deposition rate distribution and how the flow field and blade geometry affect it.

Three-Dimensional Inverse Method for Turbomachinery Blading Design

An iterative procedure for three-dimensional blade design is presented, in which the three-dimensional blade shape is modified using a physical algorithm, based on the transpiration model. The transpiration flux is computed by means of a modified Euler solver, in which the target pressure distribution is imposed along the blade surfaces. Only a small number of modifications is needed to obtain the final geometry. The method is based on a high-resolution three-dimensional Euler solver. An upwind biased evaluation of the advective fluxes allows for a very low numerical entropy generation, and sharp shock capturing. Non-reflecting boundary conditions are applied along the inlet/outlet boundaries. The capabilities of the method are illustrated by redesigning a transonic compressor rotor blade, to achieve, for the same mass flow and outlet flow angle, a shock-free deceleration along the suction side. The second example concerns the design of a low aspect ratio turbine blade, with a positive compound lean to reduce the intensity of the passage vortices. The final blade is designed for an optimized pressure distribution, taking into account the forces resulting from the blade lean angle.

High-temperature fatigue of titanium alloys

The development of high-temperature titanium alloys has been dominated by the requirements of aerospace industries. Initially, improvements in tensile strength and creep properties were the main interest. Gas inlet temperatures in aero-engines have increased continuously and, nowadays certain components in gas turbine engines that are made out of titanium alloys, such as compressor discs and rotor blades, approach operating temperatures as high as 550°C. As a result of alloy development, near-a alloys capable of operating temperatures up to 600°C are commercially available. These alloys are intended to partially replace heavier nickel-base superalloys, e.g. in the compressor of gas turbine engines and much work has been directed towards understanding high-temperature fatigue, creep-fatigue interaction and oxidation behaviour of titanium alloys. It has been recognized that alloy composition, heat treatment and microstructure all influence strongly the high-temperature properties of titanium alloys. This review focusses on the relationship between microstructure and high-temperature fatigue behaviour of titanium alloys. It will cover mainly conventional (α + β) alloys and near-α alloys which have been optimized with respect to both creep and high-temperature fatigue properties.

Incompressible Viscous Flow Analysis of Axial Compressor Blade Cascade.

In the present work, incompressible high Reynold's number flows around a compressor blade cascade at low speeds (Mach number 0.25) are analyzed by solving the N-S equations using the finite difference method. A generalized coordinate system is used so that a sufficient number of grid points are distributed in the boundary layer and in the wake region. A third-order upwind scheme is used for the convective term, which suppresses the non-linear instability for calculation of the high Reynold's number flow used in the present computation. The experimental-flow Reynold's number defined on the basis of blade chord length and cascade inlet velocity is 7×105. Grid topology is thought to be as important as flow solver, therefore calculations are carried out using H-type and C type grids and the results of the calculations are compared with the experimental data obtained for a controlled diffusion-type axial compressor rotor blade cascade for three different inlet flow conditions, which included design and off-design conditions. The study led to a clear understanding of the importance of the grid topology on the prediction of experimental flow behavior.

Rotor Blade Unsteady Aerodynamic Gust Response to Inlet Guide Vane Wakes

A series of experiments is performed in an extensively instrumented axial flow research compressor to investigate the fundamental flow physics of wake-generated periodic rotor blade row unsteady aerodynamics at realistic values of the reduced frequency. Unique unsteady data are obtained that describe the fundamental unsteady aerodynamic gust interaction phenomena on the first-stage rotor blades of a research axial flow compressor generated by the wakes from the inlet guide vanes. In these experiments, the effects of steady blade aerodynamic loading and the aerodynamic forcing function, including both the transverse and chordwise gust components, and the amplitude of the gusts, are investigated and quantified.

軸流圧縮機のワイドコード化による失速限界改善に関する研究

This paper shows an improved performance of a five stages axial flow compressor and a two stages fan applying a wide chord blade in some of their rotor stages. The estimation procedure of maximum loading capability on the wide chord rotor blade is also described in this paper based on the analogy between a compressor rotor blade and a diffuser passage. A procedure for estimating the maximum pressure rise potential and minimum flow rate boundary was formulated on the basis of meanline parameters, such that it can be used during the preliminary design early phase in the planning. A good agreement is demonstrated between the predicted and tested stall pressure and stall flow rate limits for the wide chord compressor and fan.

Three-Dimensional Boundary Layer on a Compressor Rotor Blade at Peak Pressure Rise Coefficient

A comprehensive study of the three-dimensional turbulent boundary layer on a compressor rotor blade at peak pressure rise coefficient is reported in this paper. The measurements were carried out at various chordwise and radial locations on a compressor rotor blade using a rotating miniature V configuration hot-wire probe. The data are compared with the measurement at the design condition. Substantial changes in the blade boundary layer characteristics are observed, especially in the outer 16 percent of the blade span. The increased chordwise pressure gradient and the leakage flow at the peak pressure coefficient have a cumulative effect in increasing the boundary layer growth on the suction surface. The leakage flow has a beneficial effect on the pressure surface. The momentum and boundary layer thicknesses increase substantially from those at the design condition, especially near the outer radii of the suction surface.

Experimental Cascade Analysis of a Transonic Compressor Rotor Blade Section

A transonic compressor rotor blade cascade was tested in order to elucidate the flow behavior in the transonic regime and to determine the performance characteristic in the whole operating range of a rotor blade section. The experiments have been conducted in a transonic cascade wind tunnel, which enables tests even at sonic inlet velocities. The influence of the upstream Mach number between 0.8 and 1.1 and the inlet flow angle between choking and stalling of the blade row was investigated. The effect of the axial velocity density ratio (AVDR) could be studied by applying an endwall suction device. Furthermore, the level of the shock losses was determined from a wake analysis. A final comparison of cascade losses and those of the corresponding rotor blade element shows good agreement which underlines the applicability of the cascade model in the design of axial flow turbomachines.

Casing Wall Boundary-Layer Development Through an Isolated Compressor Rotor

Detailed measurements were made of the casing wall boundary layer development across a large-scale, low-speed axial compressor rotor blade row. An important feature of the work was the use of blading which allowed the tip clearance to be varied. A conventional pressure probe was used to obtain time-averaged measurements of the outer-wall boundary layer downstream of the rotor whilst a hot-wire anenometry technique yielded the three-dimensional, blade to blade structure of the flow. The downstream boundary layer was found to thicken as the rotor loading and blade-end clearance were increased, with fluid tending to accumulate towards the pressure side of the passage. By its pronounced effects upon wall boundary layer development, tip clearance had a deleterious effect upon the performance of the compressor.

Characteristics of Mean Velocity in the Tip Region of Turbomachinery Rotor Exit

This paper reports the experimental study of the three-dimensional characteristics of the mean velocity in the annulus-wall boundary layer and wake mixing region of a moderately loaded compressor rotor blade. The measurements reported were obtained using a three-sensor hot-wire probe at the exit of the rotor. They consisted of tangential surveys at several axial and radial locations downstream of the compressor rotor blade passage. The variation of the streamwise, normal, and radial velocity profiles and their decay characteristics are derived from this data. The measurements reveal the highly complex and three-dimensional nature of the annulus-wall flowfield and the nature of the interaction of the annulus-wall boundary layer, the blade boundary layers and wake, and the secondary flow associated with the tip region.

Interaction of Compressor Rotor Blade Wake With Wall Boundary Layer/Vortex in the End-Wall Region

This paper reports the experimental study of the three-dimensional characteristics of the mean velocity of the rotor wake inside the annulus- and hub-wall boundary layers. The measurements were taken with a rotating three-sensor hot wire behind the rotor. This set of measurements probably represents the first set of comprehensive measurements taken inside the annulus- and hub-wall boundary layers. The wake was surveyed at several radial locations inside the boundary layer region and at several axial locations. Interaction of the wake with the annulus-wall boundary layer, secondary flow, tip-leakage flow, and the trailing vortex system results in slower decay and larger width of the wake. The presence of a strong vortex and its merger with the wake is also observed. The end-wall boundary layers and the secondary flow were found to have a substantial effect on both the decay characteristics and the profile of the wake. These and other measurements are reported and interpreted in this paper.

Blade Loading and Spanwise Effects on Wake Characteristics of Compressor Rotor Blade

An experimental investigation of the blade loading and spanwise effects on the rotor wake is presented. The investigation was limited to a study of a low subsonic and incompressible wake flow found downstream of a lightly loaded rotor. Measurements were made with a triaxial hot wire probe mounted in the stationary frame of reference at six radial and nine axial positions. At each measurement location, the rotor was run at different operating conditions to discern the effects of blade loading on the wake. Nearand far-wake measurements are given, including mean velocity and turbulence intensity characteristics. The loading and spanwise effects on rotor-wake characteristics were found to be substantial.

Rotor Wake Mixing Effects Downstream of a Compressor Rotor

An experimental study of rotor wake was conducted in the trailing-edge and near-wake regions of a moderately loaded compressor rotor blade using a rotating triaxial hot-wire probe in a rotating frame of reference. The flow-field was surveyed very close to the trailing-edge as well as inside the annulus and hub-wall boundary layers. The large amount of data acquired during this program has been analyzed to discern the decay effects as well as the spanwise variation of three components of velocity, three components of intensities and three components of shear stresses. The data set also include extensive information on the variation of the flow properties downstream. The other derived quantities include wake momentum thickness and deviation angles at various spanwise and downstream locations. These data are presented and interpreted, with emphasis on the downstream mixing as well as endwall-wake interaction effects.

Analysis of Turbulent Boundary Layer on Cascade and Rotor Blades of Turbomachinery

The momentum integral technique for predicting the boundary-layer growth in three-dimensi onal flow has been extended to include the entrainment equation as the closure model. The numerical solution is compared with the cascade, inducer, compressor, and fan rotor blade data from various sources. The agreement is found to be excellent in all cases, with the exception of the separated flow. Both the momentum thickness and the limiting streamline angle predicted from this analysis compare well with the measured data for a rotor blade. The technique is extremely useful in engineering design, analysis, and performance prediction.

Freestream turbulence effects on the development of a rotor wake

The effect of freestream turbulence on the development of a three-dimensional wake of a compressor rotor blade was studied experimentally. The turbulence level at the inlet of a rotor was varied systematically using grids upstream of the rotor. The rotor wake was measured with inlet turbulence intensities of 0.5, 3, and 5%. The experimental results indicate that the maxium change in the mean velocity defect is 4% over the range of inlet turbulence levels employed, while the turbulence structure in the wake is altered more substantially. The freestream turbulence effect was also analyzed, numerically, using the modified Reynolds stress closure model. The comparison between numerical prediction and experimental data shows that the freestream turbulence effect can be represented successfully with the turbulence closure model employed in this paper.

Asymmetric Distortion Generation in a Variable Height Annulus

POROUS screens were placed in the compressor intake to create quasisteady pressure maldistributed inlet flows. This study was aimed at creating a single-lobe circumferential sine-wave velocity profile with no radial components at the compressor inlet. A method relating screen blockage to predicted velocity levels was devised. The method yields velocity levels close to those for which the screen was designed. Contents Quasisteady pressure maldistributed compressor inlet flows are normally created during testing by the inclusion of porous screens in the compressor intake, but interactions between a row of compressor rotor blades and upstream placed screens in such experiments can alter the shape of the generated distortion.l Suppression of this effect may be realized by remote positioning of the screen upstream of the compressor leading-edge plane. Often, an intake may not be a constant annulus, so flow differences between the compressor leadingedge plane and the plane of the screen call for differences in the screen design whose generated distortion is itself modified in passing through the annulus to the compressor. The aim of this study was to present to a compressor inlet a single-lobe circumferential sine-wave velocity profile with no radial components, the distortion being generated remotely upstream in a duct dissimilar to the annulus at the compressor intake. Based upon small perturbation assumptions, a method of screen design is proposed which superimposes a generated distortion upon the undistorted potential flow solution of the developed annulus. The example used was a bellmouth and nose-bullet sited coaxially close to the compressor leadingedge plane and separated only by a short length of the constant height annulus. The screen was set at the bellmouth intake, perpendicular to the axial direction. While this example used symmetrical ducting, the principles involved could be applied to a nonsymmetrical installation.