High-pressure Compressor Stage Research Articles

Aerodynamic Aspect of the Problem Relating to the Reduction of the Vibration Strain in the Axial Compressor Blades of the Gas Turbine Engine

Consideration was given to the design methods available in practice for an increase in the vibration reliability of the blades of axial compressor. The problem was formulated to show the possibility of the reduction of the strain in the blade nib by the control of the exciting aerodynamic forces. When it is impossible to eliminate the resonance completely and the measures taken to increase structural damping fail to give positive effect it is reasonable to make use of the possibility of the strain reduction by the control of the excitation source. Consideration was given to the aerodynamic aspect of the problem relating to an increase of the vibration reliability of the blades of axial compressor. The aerodynamic loads affecting the axial compressor blades ranging from the trace to potential effects of adjacent shrouds were determined numerically. The program package SUnFlow with the realized in it numerical integration of the Reynolds-averaged Naiver-Stocks equations was used as an investigation tool. The high-pressure compressor stage containing 120 blades of the input guide unit, 83 operating blades and 104 blades of the guide system was used as a test object. The computations were done for the initial structure of the compressor stage and for the stage structure with embedded measures in the form of the variable- pitch input guide unit. Nonstationary computations of the options were done by including the three stage rings into the computation domain of total circumferences. Based on the obtained computation data, the integral and distributed loads onto operating blade were analyzed for comparable stage options. For the obtained distributed loads, the bend strains were calculated using the rod theory. Based on the obtained experimental data, the conclusion was made that the strains can be reduced in the blade nib by the control of the exciting aerodynamic forces.

IMPROVEMENT IN DURABILITY OF WELDED DRUMS COMPRESSOR ROTORS BY TREATMENT IN FLUDIZED BED OF ABRASIVE MATERIAL

Purpose. Improvement in durability of welded drums of compressor rotors of the gas turbine engines by treatment in fluidized bed of abrasive material.
 Methods of research. The investigations were carried out on the welded drums of rotors of high pressure compressor (HPC) of the D-36 aero engine. The drums were treated in the fluidized bed of abrasive material of the АПС-600 installation without the nozzles and with the use of special air-blast nozzles.
 Test specimens for checking durability were cut of HPC stages 1 and 4 discs after carrying out various machining processes.
 Durability testing of test specimens cut of HPC stage 4 disc was carried out at the УМЭ-1 ТМ machine at temperature Т = 400 °С.
 Durability testing of test specimens cut of HPC stage 1 disc was carried out at the ЭДЦ-20 machine at temperature Т = 20 °С.
 Results. It was found that averaged cyclic life of the test specimens made of titanium alloy ВТ-9, treated in the fluidized bed of abrasive material being a part of the drum with the use of air-blast nozzles, is 2.2 times higher than the averaged cyclic life of the test specimens treated in the fluidized bed of abrasive material without the use of air-blast nozzles at temperature Т = 400 °С.
 When treating the drum of the HPC rotor in the fluidized bed of abrasive material without the use of air-blast nozzle, better treatment of the bottom of the slot and corners of its end faces is ensured.
 Low-temperature annealing of the drum at temperature Т = 550 °С is preferable than at Т = 750 °С, since cyclic life of the test specimens annealed at Т = 550 °С is somewhat higher.
 Annealing at Т = 750 °С removes completely strengthening effect obtained by treatment in the fluidized bed of abrasive material.
 Successive treatment of the disc as a part of the drum in the fluidized bed restores its durability to the initial state.
 Scientific novelty. It was demonstrated that treatment of the welded drums of the compressor rotors in the fluidized bed of abrasive material with the use of special air-blast nozzles improves the quality of treatment and cyclic life if compared to the treatment without air-blast nozzles.
 Annealing temperature for the compressor rotor drum, which ensures high average cyclic life, was established.
 Practical value. There were offered technological procedure and modes of treatment of the welded drums of the HPC of the D-36 aero engine in the fluidized bed of abrasive material with the use of special air-blast nozzles, ensuring durability 2.2 times higher than without the use of the air-blast nozzles.

Aerodynamic simulation of a high-pressure compressor stage using the lattice Boltzmann method

PurposeThe present paper aims at evaluating the lattice Boltzmann method (LBM) on a high-subsonic high-pressure compressor stage at nominal regime.Design/methodology/approachThe studied configuration corresponds to the H25 compressor operated in a closed-loop test rig at the von Karman Institute. Several operating points are simulated with LBM for two grids of successive refinements. A detailed analysis is performed on the time-averaged flow predicted by LBM, using a comparison with experimental and existing RANS data.FindingsThe finest grid is found to correctly predict the mean flow across the machine, as well as the influence of the rotor tip gap size. Going beyond time-averaged data, some flow analysis is performed to show the relevance of such a high-fidelity method applied to a compressor configuration. In particular, vortical structures and their evolution with the operating points are clearly highlighted. Spectral analyses finally hint at a proper prediction of tonal and broadband contents by LBM.Originality/valueThe application of LBM to high-speed turbomachinery flows is very recent. This paper validates one of the first LBM simulations of a high-subsonic high-pressure compressor stage.

Analisis Pengaruh High Pressure Compressor Rotor Clearance Terhadap Exhaus Gas Temperature Margin pada CFM56-7

Exhaust Gas Temperatue is an parameter where the hot gases’s temperature leave the gas turbine. Exhaust gas temperature margin is the difference between highest temperature at take off phase with redline on indicator (???????????? ???????????????????????? °????=???????????? ????????????????????????????−???????????? ???????????????? ????????????). EGTM is one of any factor to determine engine performance. A good perfomance of an engine when it has a big margin (EGTM), during operation of an engine the EGTM could decrease untill 0 (zero). So many factors could affect EGTM deteroration there are: distress hardware such as airfoil erosion, leak of an airseals, and increase of clearance between tip balde and shroud. Increase of clearance happens in high pressure compressor rotor clearance. In CFM56-7 have 9 stage(s) of high pressure compressor and each stage give the EGT Loses. The calculation of EGT Effect/Losses is actual celarance – minimum clearance x 1000 x EGT Effect °C, where actual clearance define by the substraction of outside diameter’s rotor with inside diameter’s shroud, minimum clearance define in the manual, 1000 is adjustment from mils/microinch to inch, and EGT Effect is temperature that define in the manual. The analysist had done with 6 (six) engine serial number and proceed by corelation that shown linkage between clearance and EGT Effect, the corelation is strong shown the result of corelation (r) is 0.994275999 or nearest 1.

Loss Analysis of Unsteady Turbomachinery Flows Based on the Mechanical Work Potential

Abstract Loss analysis is a valuable technique for improving the thermodynamic performance of turbomachines. Analyzing loss in terms of the “mechanical work potential” (Miller, R.J., ASME Turbo Expo 2013, GT2013-95488) provides an instantaneous and local account of the thermal and aerodynamic mechanisms contributing to the loss of thermodynamic performance. This study develops the practical application of mechanical work potential loss analysis, providing the mathematical formulations necessary to perform loss analysis using practical Reynolds-averaged Navier–Stokes (RANS) or large eddy simulations (LES). The analysis approach is demonstrated using RANS and LES of a linear compressor cascade, both with and without incoming wakes. Spatial segmentation is used to attribute loss contributions to specific regions of the flow, and phase-averaging is performed in order to associate the variation of different loss contributions with the periodic passage of wakes through the cascade. For this un-cooled linear cascade, viscous dissipation is the dominant source of loss. The analysis shows that the contribution of the viscous reheat effect depends on the operating pressure of the compressor stage relative to the ambient “dead state” pressure—implying that the optimal blade profile for a low-pressure compressor stage may be different from the optimal profile for a high-pressure compressor stage in the same engine, even if the operating conditions for both stages are dynamically similar.

Sensitivity Analysis of Rotor/Stator Interactions Accounting for Wear and Thermal Effects within Low- and High-Pressure Compressor Stages

In the current design of turbomachines, engine performance is improved by reducing the clearances between the rotating components and the stator, which allows for loss decrease. Due to these clearance reductions, contact events may occur between the rotor and the stator. An abradable coating is deposited along the stator circumference as a sacrificial material to lower the contact severity. However, experiments highlighted the occurrence of rotor/stator interactions with high wear depth on the abradable coating as well as high temperature increases within the abradable coating following contacts. This work focuses on the sensitivity analysis of rotor/stator interactions with respect to the rotor angular speed and the initial clearances between the rotor and the stator, taking into account thermal effects within the abradable coating. Convergence analyses are first conducted to validate the numerical model. Then, after a calibration of the thermal model of the abradable coating based on two experimental test cases, the numerical model is used to investigate the cross effects of the angular speed and the initial clearances on the obtained rotor/stator interactions.

Viscous flow and performance issues in a 6:1 supersonic mixed-flow compressor with a tandem diffuser

The advancement of multi-dimensional and viscous computational tools has eased the accessibility and overall effort for thorough analysis of complex turbomachinery designs. In this paper, we computationally evaluate a high-pressure ratio supersonic mixed-flow compressor stage designed using an in-house mean-line code. Objective is to include three dimensionalities, viscous flow and compressibility effects including the shock wave systems into account. As mixed-flow compressors are advantageous especially for small jet engine applications we choose mass flow rate, stage total pressure ratio and maximum diameter as the main design constraints. This computational analysis is the second paper of a two-part series explaining strategy for designing a high-pressure ratio mixed-flow compressor stage. The high-pressure ratio and small diameter requirements push this compressor for a highly-loaded supersonic ‘shock-in rotor’ design with supersonic stator/diffuser.The used RANS based computational fluid dynamics model is thoroughly assessed for its ability to predict compressor performance using existing well-established experimental data. NASA Rotor 37 and RWTH Aachen supersonic tandem stator are chosen as the test cases for exhibiting similar flow characteristics to present design. The computational approach helps to shed light upon the mixed-rotor and supersonic-stator 3D shock structures and viscous/secondary flow. Stage performance map, pressure and velocity distribution of this high-pressure ratio mixed-flow compressor is obtained. Areas of design optimization are highlighted to further improve performance and efficiency. The in-house mean-line design code predicted a pressure ratio of 6.0 with 75.5% efficiency for a mass flow rate of 3.5 kg/s. The mean-line code obviously lacked to fully represent three-dimensionality effects due to its inherent over-simplifying assumptions thus, inclusion of RANS based computations improves the fidelity of mixed-flow compressor design performance calculations at a great rate. Comprehensive computational analysis of the stage shows that our design goal is met with a stage total pressure ratio of ΠTT = 5.83 with an efficiency of ηIS = 77% for a mass flow rate of m˙ = 3.03 kg/s. A total pressure ratio of 6.12 at 75.5% efficiency is reached with a 3.5% increase in design rotational speed.

Некоторые особенности реализации расчетной модели высоконапорной центробежной компрессорной ступени с входным направляющим аппаратом

Некоторые особенности реализации расчетной модели высоконапорной центробежной компрессорной ступени с входным направляющим аппаратом

Comparison of sensitivities to geometrical properties of front and aft high pressure compressor stages

In modern aviation, the engine-related proportion of the direct operating costs (DOC) is about one quarter. Thereby, the engine-related costs are devided into three roughly equivalent parts: depreciation/financing, fuel and maintenance and overhaul (Rupp in DLRK 2001, 2001). Consequently, an efficient and high technical quality maintenance has a great impact on the DOC. Nowadays, the transition from the exhaust gas temperature based to the condition based maintenance is developing. Thereby, the maintenance, repair and overhaul (MRO) companies provide tailored maintenance actions for each jet engine, depending on the engine history and operation conditions. Obviously, a detailed knowledge of the influence of the different engine components and their features on the engine performance is necessary to accomplish a condition based maintenance. In order to illustrate this, the example of the high pressure compressor (HPC) is shown: The HPC has a strong impact on the gas turbine efficiency and the specific fuel consumption (SFC). Nevertheless, the HPC blading is classified coarsely into serviceable, scrap and repairable. The repairable blades are repaired by a few repair processes, without evaluating the blade geometry in detail. With a detailed knowledge of the geometric main parameters to the aerodynamic performance and their interactions, future repair processes could maintain suited maintenance actions as well as tailored blade sets for the HPC. Therefore, a more efficient maintenance could be achieved to meet future requirements of the costumers. This paper contributes a part to this maintenance development. Therefore, a design of experiments and a sensitivity analysis will be carried out for a HPC-aft stage and compared with previous gained results of a HPC-front stage. Therefore, 700 different stage geometries will be analyzed for their aerodynamic performance and imported to a Kriging Method to generate a meta-model. This meta-model allows a sensitivity analysis for individual geometric properties. The detected sensitivities will be compared to those of the front stage. Changes of the behavior of the stages through the HPC could help the MRO companies to focus on appropriate geometric properties and, therefore, suitable repair processes for each stage and airfoil.

Numerical Investigation of the Impact of the Compressor Operation Mode on Working Process of the Combustion Chamber

The method of integrated compressor/combustor simulation was used to investigate the impact of flow distortion, appeared due to compressor blades, during the combustion chamber workflow. The method was improved in terms of generating a common grid and of principles of the boundary conditions settings. The geometric model includes four geometric volume bodies: guide vanes of the penultimate stage of high-pressure compressor, the impeller and guide vanes of the last stage and the flow path of combustion chamber. The calculation was carried out for some operation mode of the engine (nominal, 0.7 of nominal and 0.5 of nominal regimes) with and without compressor. The results were compared with the results of combustion chamber simulation without the compressor. Simulations showed that blade wakes extend up to the flame tube head. These wakes influence on the flame tongue, pressure field, temperature and velocity in the recirculation-mixing zone. It can influence on combustion efficiency, ecological performance and on temperature field at the combustor outlet. Thus, the simulations, which take into account combustion chamber and compressor, are more fully represent the characteristics of the working process of the combustion chamber and increase the efficiency of the design of new products.

Comparison of Two Methods for Sensitivity Analysis of Compressor Blades

This paper will compare two approaches of sensitivity analysis, namely (i) the adjoint method which is used to obtain an initial estimate of the geometric sensitivity of the gas-washed surfaces to aerodynamic quantities of interest and (ii) a Monte Carlo type simulation with an efficient sampling strategy. For both approaches, the geometry is parameterized using a modified NACA parameterization. First, the sensitivity of those parameters is calculated using the linear (first-order) adjoint model. Since the effort of the adjoint computational fluid dynamics (CFD) solution is comparable to that of the initial flow CFD solution and the sensitivity calculation is simply a postprocessing step, this approach yields fast results. However, it relies on a linear model which may not be adequate to describe the relationship between relevant aerodynamic quantities and actual geometric shape variations for the derived amplitudes of shape variations. Second, in order to better capture nonlinear and interaction effects, a Monte Carlo type simulation with an efficient sampling strategy is used to carry out the sensitivity analysis. The sensitivities are expressed by means of the coefficient of importance (CoI), which is calculated based on modified polynomial regression and therefore able to describe relationships of higher order. The methods are applied to a typical high-pressure compressor (HPC) stage. The impact of a variable rotor geometry is calculated by three-dimensional (3D) CFD simulations using a steady Reynolds-averaged Navier–Stokes model. The geometric variability of the rotor is based on the analysis of a set of 400 blades which have been measured using high-precision 3D optical measurement techniques.

ДО ОПТИМАЛЬНОГО ПРОЕКТУВАННЯ ОСЬОВОГО БАГАТОСТУПІНЧАСТОГО КОМПРЕСОРА

The work is devoted to the development of an integrated approach to the aerodynamic improvement of the blade vanes of multistage axial compressors of gas turbine engines. One of the features of the studies carried out to optimize the blade vanes was to take into account the mixing effects by various sources of losses. Thus, minimization of losses sources in compressor blade vanes during their profiling is achieved due to the account of the complex three-dimensional vortex nature of the flow and the use of modern software systems for solving the Navier-Stokes equations. Calculations of the three-dimensional viscous flow in the work are performed with ANSYS software. The profiling of the multi-stage compressor blades is carried out with “Kompress 2.0” software package, the feature of which is the use of parametric models of rotor and stator blade vanes, adapted for their aerodynamic optimization, as well as taking into account the radial unevenness of the incident flow when selecting the tangential blades shape of the stator vanes. As a result, a complex methodology for the optimal design of the blades of a multi-stage compressor with the use of "controlled" diffusivity, S-shaped profiles, lean, and chord variation is developed, with simultaneous variation of geometric parameters according to special plans and the use of the compressor efficiency as an integral efficiency criterion, “Kompress 2.0” for the blades profiling, ANSYS for the calculation of threedimensional viscous flow. Multi-stage axial compressors have been designed, numerically and experimentally investigated. Based on the results of experimental studies of multistage axial compressors, new generalizing relationships are established, that establish a relationship between the geometric parameters of the profiles and their gas dynamic characteristics to take into account the features of the spatial shape of the flow at the stage when solving the inverse design problem. The results of the research were used in the design and improvement of the elements of the flow paths of the ZM multistage compressors. Test aerodynamic optimization of the blade systems of experimentally studied multistage compressors with the use of the proposed approaches made it possible to obtain the efficiency of high-pressure compressor stages at the level of 0.92 ... 0.93 and to increase the efficiency of multistage axial compressors by 1.0-2.4%. The efficiency of the 10-stage compressor is 0.8% (to the level of 0.853), the efficiency of the 6-stage compressor is 1.1% (to the level of 0.825), the efficiency of the 3-stage compressor is 2.4% (to the level of 0.885), and the efficiency of an 8-stage axial compressor is 1.6% (to the level of 0.879).

The Impact of Blade Loading and Unsteady Pressure Bifurcations on Low-Pressure Turbine Flutter Boundaries

The three major aeroelastic issues in the turbomachinery blades of jet engines and power turbines are forced response, nonsynchronous vibrations, and flutter. Flutter primarily affects high-aspect ratio blades found in the fan, fore high-pressure compressor stages, and aft low-pressure turbine (LPT) stages as low natural frequencies and high axial velocities create smaller reduced frequencies. Often with LPT flutter analyses, physical insights are lost in the exhaustive quest for determining whether the aerodynamic damping is positive or negative. This paper underlines some well-known causes of the LPT flutter in addition to one novel catalyst. In particular, an emphasis is placed on revealing how local aerodynamic damping contributions change as a function of unsteady (e.g., mode shape, reduced frequency) and steady (e.g., blade torque, pressure ratio) parameters. To this end, frequency domain Reynolds-averaged Navier–Stokes (RANS) CFD analyses are used as computational wind tunnels to investigate how aerodynamic loading variations affect flutter boundaries. Preliminary results show clear trends between the aerodynamic work influence coefficients and variations in exit Mach number and back pressure, especially for torsional mode shapes affecting the passage throat. Additionally, visualizations of qualitative bifurcations in the unsteady pressure phases around the airfoil shed light on how local damping contributions evolve with steady loading. Final results indicate a sharp drop in aeroelastic stability near specific regions of the pressure ratio, indicating a strong correlation between blade loading and flutter. Passage throat shock behavior is shown to be a controlling factor near the trailing edge, and as with critical reduced frequency, this phenomenon is shown to be highly dependent on the vibratory mode shape.

Riblet Application in Compressors: Toward Efficient Blade Design

Nowadays, modern axial compressors have already reached a very high level of development. The current study is focused on the question, if the application of riblets on the surfaces of a highly efficient modern compressor blade can be a further step toward more efficient blade design. Therefore, a highly loaded compressor cascade has been designed and optimized specifically for low Reynolds number (LRN) conditions, as encountered at high altitudes and under consideration of the application of riblets. The optimization was performed at a Mach number of 0.6 and a Reynolds number of 1.5 × 105. Two objective functions were used. The aim of the first objective function was to minimize the cascade losses at the design point and at incidence angles of +5 and −5 deg. The intention of the second objective function was to achieve a smooth distribution of the skin friction coefficient on the suction side of the blade by influencing the blade curvature in order to apply riblets. The MISES flow solver as well as the DLR optimizer “AutoOpti” was used for the optimization process. The developed compressor cascade was investigated in the transonic cascade wind tunnel of DLR in Cologne, where the Reynolds number was varied in the range of 1.5 × 105–9.0 × 105. Furthermore, the measurements were carried out at a low turbulence level of 0.8% and at a high turbulence level of 4%, representative for high pressure compressor stages. The measurement program was divided into two parts. The first part consisted of the investigation of the reference cascade. In the second part of the study, riblets were applied on suction and pressure side of the cascade blades; two different manufacturing techniques, a rolling and a coating techniques, were applied. The rolling technique provides riblets with a width of 70 μm and the coated riblets (CRs) have a width of 50 μm. The wake measurements were performed using a three-hole probe at midspan of the cascade in order to determine the resulting losses of the reference blade and the blades with applied riblets. The detailed analysis of the measurements shows that the riblets have only a slight influence on the viscous losses in the case of the compressor application in this study. Finally, these results are discussed and assessed against the background of feasibility and effort of riblet applications within the industrial design and manufacturing process.

A parallelized multidomain compact solver for incompressible turbulent flows in cylindrical geometries

We present an efficient parallelized multidomain algorithm for solving the 3D Navier–Stokes equations in cylindrical geometries. The numerical method is based on fourth-order compact schemes in the two non-homogeneous directions and Fourier series expansion in the azimuthal direction. The temporal scheme is a second-order semi-implicit projection scheme leading to the solution of five Helmholtz/Poisson equations. To handle the singularity appearing at the axis in cylindrical coordinates, while being able to have a thinner or conversely a coarser mesh in this zone, parity conditions are imposed at r=0 for each flow variable and azimuthal Fourier mode. To simulate flows in irregularly shaped cylindrical geometries and benefit from a hybrid OpenMP/MPI parallelization, an accurate perfectly free-divergence multidomain method based on the influence matrix technique is proposed. First, the accuracy of the present solver is checked by comparison with analytical solutions and the scalability is then evaluated. Simulations using the present code are then compared to reliable experimental and numerical results of the literature showing good quantitative agreements in the cases of the axisymmetric and 3D unsteady vortex breakdowns in a cylinder and turbulent pipe flow. Finally to show the capability of the algorithm to deal with more complex flows relevant of turbomachineries, the turbulent flow inside a simplified stage of High-Pressure compressor is considered.

Air-cooled gas turbine cycles – Part 1: An analytical method for the preliminary assessment of blade cooling flow rates

It is well known that, for a given compressor technology, gas turbine efficiency increases with the turbine inlet temperature (TIT): both modern aeronautical and land-based gas turbines operate at very high temperatures (1500–2000K) –and correspondingly high pressure ratios. As the TIT increases, the heat transferred from the expanding gas to the turbine blade also increases, and the need to extend the operational life make it necessary to adopt internal air cooling to reduce blade creep, oxidation and low-cycle fatigue. The cooling medium is usually air extracted from the high-pressure compressor stages, and since this extraction decreases the thermal efficiency and power output of the engine, it is important to bleed the minimum amount of coolant to attain a prescribed maximum material temperature in the blade with the maximum possible uniformity (lower thermal stresses): thence the need to properly model the cooling system for a given turbine blade geometry under realistic engine operating conditions. In the preliminary design of the first statoric and rotoric blading, it is essential for designers to rely on simple models that often neglect the small scales effects on the external flows and also by force adopt a much simplified treatment of the internal ones, and as a result attain a substantially lower degree of approximation than that offered by more complex and expensive numerical simulations. The goal in the design of a lumped model is therefore to make it both sufficiently general and accurate to analyze blade shapes and cooling channels structures that can be further refined by means of more accurate, but also more computationally intensive, models.This paper presents a simple, globally lumped thermodynamic model of blade cooling whose most important feature is its being analytical, so that the solution is devoid of numerical approximations and leads to closed-form expressions that can be easily manipulated to accommodate for different process parameters, blade arrangements, cooling channel structure and fluid properties. The model predicts the mass-averaged gas temperature along the channel and the required cooling air flowrate, under the assumption of a constant metal temperature along and inside of the blade: therefore, thermo-mechanical effects are completely neglected.The results are validated against some published data, and the agreement is found to be quite satisfactory: when discrepancies arise, a phenomenological explanation is offered.

Investigation of the internal flow field for a subsonic stage of high pressure compressor through model testing

A large-scale four-stage low-speed research compressor was designed to represent the flow field structure in a subsonic stage, which is the seventh stage of a high pressure compressor. The aim of the design work was to undertake a careful high-to-low speed transformation. The low-speed compressor was designed as an “approximate repeating-stage” configuration, in which the third stage is the model stage and is slightly different from the other stages, while the first two stages and the last stage set up inlet and outlet conditions for it, respectively. The tip–hub ratio, solidity, and aspect ratio of the low-speed research compressor were restricted to be as close as possible to those of high pressure compressor. The blade sections along the span were designed to match the distributions of high-speed surface static pressure coefficient; furthermore, stacking rules of high pressure compressor apply to low-speed blades. The low-speed blades have been tested at the 4-Stage Research Compressor Facility of Nanjing University of Aeronautics and Astronautics. Detailed traverse flow measurements were conducted in the stator blade passages of the third stage and between blade rows, along with static pressure measurements on stator blade surface and particle image velocimetry measurements in the rotor passage. Experiment results show that due to the boundary layer separated below the mid-span of the rotor blade, the total pressure loss increases at the hub region of both the rotor and stator. In addition, the experimental results in the stator blade passage reflect the characteristics of the flow field in the bowed stator passage. The results indicated that the seventh stage of high pressure compressor needs to be optimized by inhibiting the boundary layer separation mainly on the rotor blade surface below the mid-span, which is our further work.

Erratum: “Low-Speed Model Testing Studies for an Exit Stage of High Pressure Compressor” [ASME Journal of Engineering for Gas Turbines and Power, 2014, 136(11), p. 112603, DOI: 10.1115/1.4027637

In the Inlet Flow Field subsection of the Results and Discussion section, the blockage factor values of “0.12” and “0.115” should be replaced by “0.93” and “0.9315” respectively, in consideration of the general definition of blockage factor. The text is revised as:…of blockage factor for HPC exit stage is 0.93. By various numerical attempts, the length of the inlet duct between cowling end and compressor inlet is determined, which is equal to 0.8 m, and the measured result show that the blockage factor for design operating point is 0.9315, which agrees very well with the design value. It…Due to the same reason, the boundary layer parameter values of “3.23,” “3.28,” “3.96” and “3.16” are to be replaced by “3.05,” “3.11,” “2.25” and “1.86” respectively, in Table 2 on page 5. The revised table is below:Figures 12 and 13 on page 6 of the original manuscript contain some errors. The correct figures are as follows:

Low-Speed Model Testing Studies for an Exit Stage of High Pressure Compressor

This paper discusses detailed experimental studies of a low-speed large-scale axial compressor, which is typical of an exit stage of HPC. Numerous measuring techniques were performed, and detailed experimental results were obtained, including inlet boundary layer total pressure distributions, overall compressor and model-stage performance, traverse flow field between blade rows and inside the stator for the model stage, static pressure on the stator blade and casing dynamic pressure of the rotor. The objective of the study is to assess the low-speed model compressor design and verify 3D computational fluid dynamics (CFD) code. Results show that inlet endwall blockage requirement of HPC exit stage is achieved; the low-speed model compressor design is fundamentally successful; the flow rate and pressure rise requirements are met at the design operating point, although the flow loss is relatively larger than design values for the lower half span, which can be attributed to a certain hub-corner separation. Furthermore, the reliability of adopted 3D commercial CFD code is validated. It is proved that the low-speed model testing technique is still a prospective way for the design of high performance HPC.

Some Exergetic Measures of a JT8D Turbofan Engine

In this article, some exergetic measures are calculated for a JT8D turbofan engine at takeoff. Selected exergetic measures in this study are as follows: fuel depletion ration, productivity lack ratio, fuel exergy factor, product exergy factor and improvement potential rates. The engine has low-pressure compressor (LPC) stages, high pressure compressor (HPC) stages, a single HP turbine (HPT), and finally three LPT stages. The exergetic assessment of the JT8D turbofan components provided here should be helpful for designing turbofan engines. Results from this study also evaluate effects of the maximum power setting on the exergetic measures of the engine components commonly used in medium range commercial aircrafts. 