bluuebunny/crossref_metadata_2025 huggingface dataset:
The Influence of Blade Sweep Technique in Linear Cascade Configuration
Authors: Spataro Rosario , D’Ippolito Gabriele , Dossena Vincenzo
Published: May 2012

One key issue in the advanced aerodynamic optimization of turbomachinery involves the application of 3D blade design techniques. The complex shape of the resulting blades is often a combination of simple techniques such as sweep. Such a blade arrangement is often imposed to the designer by structural constrains, space reduction needs, diameter optimization or spanwise blade loading control. This work aims to study the aerodynamic effects produced on turbine passages by blade sweep; with this term we refer to a configuration where the flow mainstream direction and the blade stacking axis are not orthogonal. A linear cascade of turbine blades, obtained by stacking the same profile with a sweep angle of 20 degrees, was investigated in a blow down facility at an isentropic downstream Mach number of 0.65. Standing the low aspect ratio of the cascade, the blade was built by simply shifting in axial direction the 2D profile originally used in the reference prismatic blade. The choice to build the swept blade keeping the same 2D section parallel to the incident flow was considered taking into account the blade low aspect ratio. Measurements were performed by means of blade surface pressure taps and five holes probe traversing downstream of the cascade; oil and dye flow visualizations were also performed to study the effects on the secondary vortices evolution inside of the passage. Moreover, a commercial CFD code was applied to provide information on the flow field all along the passage. The same profile was already extensively investigated both by measurements and CFD calculations [1, 2] in order to clarify the effects of blade lean and bowing. This additional paper gives a final contribution addressed to deeply understand the aerodynamic effects produced on turbine cascade flow field by the separate application of each one of the typical 3D design techniques. Results from both the experimental and computational investigations are presented and discussed in the paper where a phenomenological approach has been preferred. Measurements of the blade surface pressure distribution, performed at several blade heights, support the analysis of the pressure field inside of the passage which is mainly based on numerical results. In particular, the paper shows the influence of pressure contours shape on streamwise vorticity inside and downstream of the passage focusing the attention on secondary structures. The downstream vorticity field is then discussed together with the loss distribution in the same region to provide a more exhaustive description.
The Influence of Leading Edge Diameter on Stagnation Region Heat Transfer Augmentation Including Effects of Turbulence Level, Scale, and Reynolds Number
Authors: Gandavarapu P. , Ames F. E.
Published: May 2012

Stagnation region heat transfer measurements have been acquired on two large cylindrical leading edge test surfaces having a four to one range in leading edge diameter. Heat transfer measurements have been acquired for six turbulence conditions including three grid conditions, two aero-combustor conditions, and a low turbulence condition. The data have been run over an eight to one range in Reynolds numbers for each test surface with Reynolds numbers ranging from 62,500 to 500,000 for the large leading edge and 15,625 to 125,000 for the smaller leading edge. The data show augmentation levels of up to 110% in the stagnation region for the large leading edge. However, the heat transfer results for the large cylindrical leading edge do not appear to infer a significant level of turbulence intensification in the stagnation region. The smaller cylindrical leading edge shows more consistency with earlier stagnation region heat transfer results correlated on the TRL parameter. These results indicate that the intensification of approaching turbulence is more prevalent with the more rapid straining of the smaller leading edge. The downstream regions of both test surfaces continue to accelerate the flow but at a much lower rate than the leading edge. Bypass transition occurs in these regions providing a useful set of data to ground the prediction of transition onset and length over a wide range of Reynolds numbers and turbulence intensity and scales.
The Influence of Inlet Contraction on Vane Aerodynamic Losses and Secondary Flows With Variable Turbulence and Reynolds Number
Authors: Chowdhury N. H. K. , Dey P. K. , Ames F. E.
Published: May 2012

Exit survey measurements comparing exit losses for a constant inlet height cascade and a contracting inlet cascade have been acquired in a low speed wind tunnel facility. The measurements were taken for both a low and a high turbulence level across a four to one range in chord Reynolds number (500,000, 1,000,000, and 2,000,000). The high intensity turbulence has been generated using a simulated aero-combustor for both cascades. Exit survey measurements have been acquired for the two cascades using a five-hole cone probe at stations representing an exit location of 1/4 axial chord downstream from the trailing edge plane. Measurements detail total pressure loss, turning angle, and secondary velocities. At the low turbulence condition, the contracting inlet cascade shows a significant reduction in losses compared with the constant height inlet cascade. The contracting inlet cascade also features an aft loaded vane profile which may have some impact on the development of secondary flows. The constant cross-section cascade uses a fully loaded vane profile. The difference in losses between the two cascades is also significant at the high turbulence condition. However, at the high turbulence condition, the losses for the contracting inlet cascade are greater than the constant cross-section cascade. These increased losses are believed to be due to the more aggressive turbulence of the contracting inlet case. The contracting inlet effectively moves the vane leading edge plane forward into the exit of the combustor where the vane passage is subjected to more aggressive turbulence levels.
Turbo-Machinery Monitoring Measures for Propulsion Safety and Affordable Readiness
Authors: Millar Richard C.
Published: May 2012

The US Department of Defense initiative for propulsion safety and affordable readiness (P-SAR) has been engaged by all three services as a means to mitigate the operational and environmental hazards to aircraft propulsion system integrity, dependability and life cycle cost in military service. This paper focuses on the turbo-machinery of military propulsion systems and addresses the options available to manage turbo-machinery health and life cycle cost. Three primary turbo-machinery integrity and performance degraders are identified: foreign object damage, erosion and corrosion, and high cycle fatigue due to aero-elastic response of turbo-machine blades & vanes. A variety of sensors have been proposed as measures to monitor and mitigate the hazards created by such degradation. Many of these were developed as tools for use in component models & design verification, such as non-intrusive stress measurement systems [NSMS] — an indirect means to supplement direct on-blade strain measurements. Others tools have been defined specifically for in service monitoring, such as the use of miniature radar to detect blade motion. This paper follows a systems engineering analysis identifying alternative functional hazard mitigations and a qualitative reliability centered maintenance (RCM) view to assess alternative approaches to mitigation of each hazard. Finally, synergies across the three priority degraders are considered to recommended topics for research.
Enhanced Static-Dynamic Pressure Transducer for the Detection of Acoustic Level Flow Instabilities in Gas Turbine Engines
Authors: Hurst Adam M. , Goodman Scott , Kochman Boaz , Ned Alex
Published: May 2012

The push to advance the performance and longevity of gas turbine engines requires better characterization of flow instabilities within the compressor and most importantly the combustor. Detecting the earliest onset of these flow instabilities can help engineers either manipulate the flow to restabilize it or make informed design changes to the engine. The pressures within gas turbine engines are typically composed of an undesired, low-level oscillatory pressure of less than 1kPa to several kPa superimposed on top of a large, relatively constant pressure of several thousand kPa [1–7]. The high-pressure transducers used to measure the pressures within these environments are often unable to resolve these low-level oscillatory pressures that characterize the flow instabilities because the signal output for such pressures is often the same level as the noise within the sensor-data acquisition system. This paper presents an engine test ready, high temperature, combined static and dynamic pressure transducer that uses static pressure compensation in order to measure these low-level dynamic pressures with an excellent signal to noise ratio and, at the same time, captures the overall static pressure within a gas turbine [8–10]. Test bench experiments demonstrate the static-dynamic transducer’s unique ability to capture both large static or quasi-static pressures of 1,380kPa or greater and simultaneously measure the acoustic-level dynamic pressures superimposed on top of these pressures. The static-dynamic transducer achieves this advanced sensitivity through the use of a low-pass acoustic filter that passes the large static pressure to the reference port of a high sensitivity dynamic pressure sensor within the transducer such that the overall static pressures cancel out and the sensor measures all acoustic-level dynamic pressures. These bench tests additionally demonstrate the transducer’s ability to operate reliably when exposed to the harsh, high temperature environment (up to 500°C) within a gas turbine [8–10].
Design Optimization of a Retractable Holder for Compressor Discharge Brush Seal
Authors: Samudrala Omprakash , Kumar Siddarth , Wolfe Christopher E. , Chupp Raymond E.
Published: May 2012

In many industrial gas turbines, a portion of the compressor discharge air is extracted through a secondary flow path to aid the cooling of critical turbine components as well as to supplement purge flow for preventing hot gas ingestion in the first forward turbine bucket wheel space. GE has developed advanced brush seals for controlling the amount of cooling/purge flow passing through this secondary flow path (also called the high pressure packing (HPP) circuit) and has successfully implemented them in the field in a variety of E, F & H class gas turbines. During turbine shutdown, due to a lag in thermal response between the rotor and the stator, interference can result between brush seal bristles and the rotor surface causing significant amounts of wear. This wear can accumulate over several start up / shut down cycles resulting in an increased secondary flow through the HPP circuit and thus a loss in turbine efficiency and power output. In order to alleviate this situation, a seal holder has been designed to passively retract the HPP brush seal, from a low clearance position to a high clearance position, during turbine shut down and thus prevent seal interference/wear. This paper delves into the design and optimization of a retractable seal. An analytical model was developed to predict the seal motion during startup and shutdown of the turbine. Critical geometry and design parameters affecting seal closure and retraction behavior were identified. In addition, criteria for stability of seal motion were developed and the design was optimized to meet these requirements. Seal wear during turbine shutdown is avoided by ensuring that the seal retracts faster than the rate of thermally induced interference. The effect of design variables was minimized to ensure seal closure and retraction behavior does not vary significantly over the operating life of the seal. Model predictions were validated by subscale rig tests performed in the laboratory.
Gas Turbine and Driven Machinery Management and Diagnostics
Authors: Maalouf Mel , Eldridge Thomas
Published: May 2012

The size and complexity of gas turbines has evolved tremendously over the years and the controls, instrumentation and diagnostics tools have kept pace with the advances. This paper discusses the progress of the tools to keep these complex machines running for continued reliability, efficiency, emissions compliance and power output. The technology to enable the user to manage their machinery on site and remotely will be discussed in this paper along with the benefits added by the technologies.
On the Nonlinear Dynamics of Rotor-Foil Bearing Systems: Effects of Shaft Acceleration, Mass Imbalance and Bearing Mechanical Energy Dissipation
Authors: San Andre´s Luis , Ryu Keun
Published: May 2012

Gas foil bearing (GFB) technology has reached great maturity as per engineered design and construction and its system integration into rotating machinery. Empirical research has gone beyond showing a few instances of acceptable mechanical performance, to demonstrate GFB multiple-cycle repeatable performance in spite of persistent large amplitude sub synchronous whirl motions. A GFB is a forgiving mechanical element whose engineered resilient underspring structure contains and ameliorates large rotor excursions. Analyses, however, fail to distinguish the hardening stiffness from the FB underspring structure, which under conditions of large force excitations due to imbalance, produces a complex rotordynamic behavior, rich in sub harmonic motions when operating at super critical speeds. This paper extends an earlier analysis of a rigid rotor-GFB system that dispenses with the gas film component to predict the effect of shaft rotation acceleration/deceleration on rotor amplitudes of motion and whirl frequency content. For operation above the system critical speed and as the rotor accelerates, large amplitude whirl motions appear with a main subsynchronous frequency tracking rotor speed, first at 50% speed and later bifurcating into at 33% whirl frequency. Rotor imbalance awakens and exacerbates the system nonlinear response. Slow rotor accelerations result in responses with more abundant subsynchronous whirl patterns, increased amplitudes of whirl, and accompanied by a pronounced mechanical hysteresis when the rotor decelerates. Large rotor imbalances produce both jump phenomenon and a stronger hysteresis during slow acceleration and deceleration cases. Material damping (dry friction) in the FB aids to reduce and delay the nonlinear response, eventually eliminating the multiple frequency behavior. The results bring to attention rotordynamic issues during start up and shut down events that can result from an inadequate FB technology or an unacceptable rotor imbalance grade condition.
Theoretical and Experimental Investigation on Nonlinear Characterization of Metal Rubber
Authors: Hong Jie , Zhu Bin , Ma Yanhong
Published: May 2012

Metal Rubber (MR) can be widely used in many aspects such as damping of blade and support, pipe and equipment in air space technology, vehicle and ship. A theoretical method was performed to describe hysteretic properties and nonlinear stiffness and damping characteristics of Metal Rubber component. Spiral wire was considered as the micro-element structure of MR material by analyzing the manufacture process of MR component. Based on the material mechanics and coulomb friction theory, a mechanical model of spiral wire was established which is combined with the cylindrical compression coil spring theory. It was easy to explain the mechanism of hysteresis loop and the nonlinear stiffness and damping characteristic of MR component by means of analyzing contact conditions of micro-element. The quasi-static experiments were conducted to MR component with different material parameters. The influencing laws of material parameters on the performance of MR component were studied. The research was valuable for the analysis of the material mechanics and the design of MR component. It provides theoretical support for the further engineering application of MR material in the field of vibration reduction.
Development of the 5MW Power Generation Gas Turbine Engine
Authors: Kim Sooyong , Lee Sungryong , Ryu Jewook , Spitsyn V. E.
Published: May 2012

Gas turbine engine has been applied to the aircraft and ship propulsion with its advantages of compactness and comparatively short starting time. With a significant improvement in gas turbine efficiency with development of super alloy materials and advancement in cooling technologies in the second half of 1990s, its importance as a source of base load as well as peak load power generation has been increasing. However, with increased demand in nuclear power and renewable energy in the 21st century, there seems to be speculations among the power generation industries that gas turbine will take more or less a buffering role supplementing the irregular inflow of electricity to the grid rather than acting as a base load power source. With the shift in the role of gas turbine from base to supplementary, CHP application based on small powered gas turbine utilizing biogas or syngas as its fuel is expected to increase in the future. In this context, this paper describes the development result of 5MW gas turbine engine for CHP application. It can be operated with LNG or syngas of low LHV fuel. Originally, the engine was designed for LNG as its primary fuel, but since the importance of syngas power generation market will be increasing in the future, a complementary work for modification of combustor part has been carried out and has been tested. However, this paper deals with the parts developed with the use of LNG fuel. The test result of emission characteristics meets the standards required in Korea. The development has been made through the cooperation of Doosan Heavy Industry (DHI, Korea) and Zory-Mashproekt (Ukraine).
Validation of an Infrared Extinction Method for Fuel Vapor Concentration Measurements Towards the Systematic Comparison Between Alternative and Conventional Fuels for Aviation
Authors: Fritzer Johannes , Strzelecki Alain , Giuliani Fabrice , Bodoc Virginel
Published: May 2012

Due to an increasing oil price and the obvious influence of the combustion of fossil fuel-derivatives on climate change on one hand and the steady growth of transportation needs on the other, it is necessary to develop alternatives to oil for aviation. For this purpose a specific research program on the investigation of adequate alternative fuels for aviation has been founded by the European Commission’s Framework Program. The project Alfa Bird (Alternative Fuels and Bio-fuels in Aircraft Development) focuses on an identification of possible alternative fuels to kerosene, the investigation of the adequacy of the selected ones, an evaluation of the environmental and economical impact of those and finally the creation of a future perspective for the industrial use of the “best” alternative. The main part of the investigation activities at TU Graz, in cooperation with ONERA Centre de Toulouse and Fauga-Mauzac on these specific topics consists of the analysis of the evaporation of the previously chosen fuel types in comparison to Fully Synthetic Jet Fuel (FSJF). Therefore qualitative measurements to obtain vapor concentration gradients will be done using the Infrared Extinction (IRE) measurement method. Based on a simplified Beer-Lambert-Law the integral vapor concentrations can be obtained. The main hypothesis is that if the line-of-sight extinction due to Mie-scattering is similar for both infrared and visible wavelengths because of the presence of the spray, only infrared light will be absorbed by the fuel vapor, being transparent to visible light. This contribution focuses on the validation of the infrared measurement technique on a well characterized spray. The tests are performed under controlled boundary conditions. Therefore an existing IRE test arrangement at ONERA Toulouse using an ultrasonic atomizer injecting n-octane at atmospheric conditions has been analyzed. Error sources related to misalignments in the hardware have been considered and an iterative alignment method of the laser beams followed by a beam diameter and diffraction analysis have been performed. Optimizing the setup to obtain a stable operation point has been successful. Improved experimental results at this operation point were compared with existing simulation results for the evaporation of the used ultrasonic atomizer. The achieved data has shown good accordance to the existing simulation results. This work has been supported by the Eccomet project (Efficient and Clean Combustion Experts Training) in the framework of Alfa Bird.
Rotor Life Prediction and Improvement for Steam Turbines Under Cyclic Operation
Authors: Checcacci Damaso , Cosi Lorenzo , Sah Sanjay Kumar
Published: May 2012

The evolution of the energy market is leading to a general increase in demand for cyclic operation and rapid startup capability for steam turbines utilized in power utility plants. As a consequence, turbine manufactures must optimize designs to minimize transient stress and make available to plant operators the necessary understanding of the impact of operating conditions on parts life. In addition, if continuous duty operation is not economical for an existing plant, operators considering switching to the cyclic mode need to take into account the cost associated with reduced maintenance intervals and parts replacement. This paper presents the methodologies applied to assess and optimize steam turbine rotor life. The discussion stems from the case analysis of a 60 MW steam turbine that was operated almost uninterrupted for 10 years in a combined cycle plant and was then expected to switch to cyclic operation with approx 250 startups/year. The effects of different rotor geometries on transient thermal stress/strain conditions are presented along with the consequences of startup sequence modifications for rotor life vs. on-line time. The discussion is supported by modeling details and results from transient thermomechanical FEM analyses. The possibility of a simplified approach in the form of approximate models for the analysis of such behavior on a project basis is also addressed.
Rotor Dynamic Analysis of Tie-Bolt Fastened Rotor Based on Elastic-Plastic Contact
Authors: Peng He , Liu Zhansheng , Wang Guilong , Zhang Min
Published: May 2012

The tie-bolt fastened rotor which is assembled by rods distributed circumferentially is modeled and analyzed by finite element method with the consideration of elastic-plastic contact between discs. Based on elastic-plastic contact model between an elastic hemisphere and a rigid plane, the contact between discs is investigated by the statistical contact model of rough surfaces, and the contact stiffness is derived. The equivalent bending stiffness between discs is acquired. With the increase of the load between the two contact surfaces, the difference between the contact stiffness of purely elastic contact and elastic-plastic model is compared. With the obtained contact stiffness, the equation of motion for the tie-bolt fastened rotor system is formed and the critical speeds are calculated. It indicates that the contact stiffness between discs increases as the load increases. The contact stiffness of elastic-plastic contact model is lower than that of the elastic contact model, and the difference between the two models increases with load. With the stiffness of elastic-plastic contact, the critical speeds of tie-bolt fastened rotor are lower than that of the pure elastic contact situation.
Development of State of the Art Dynamic Stress Prediction Methodology for Steam Turbines
Authors: Chen Tie , Rogerson John , Yang Fang , Singh Gurnam , Hemsley Phil
Published: May 2012

To predict the dynamic stresses due to forced response of steam turbine blades, a commercial FE solver ABAQUS has been linked with an in-house CFD solver TF3D-VIB, in the time domain in both one-way and two-way coupling. Both methods have been applied to analyse a freestanding subsonic turbine stage excited by upstream flow perturbations. Over a frequency range the peak responses are very similar, but the peak response of two-way coupling is shifted to a lower frequency, due to the aerodynamic coupling effect of fluid-structure interaction. That means a speed / frequency sweep is necessary to search for the peak response in two-way coupling. However, in one-way coupling, the frequency shift can be derived from the vibration induced modal force, and only one calculation is needed to predict the response over a range of frequency ratio using the classic single degree-of-freedom equation. One calculation using two-way coupling typically takes seven times more computing time than one-way coupling. The total computing time for two-way coupling to define the response characteristic is therefore much higher; more calculations are needed and each calculation takes much longer. Thus a one-way coupling method including the frequency shift correction is much more practical and suitable for blade design iterations. The blade forced response is also limited by damping. In the case of low damping such as material only damping, this can be well represented in the harmonic ABAQUS calculation. However, high values of nonlinear damping can be deliberately introduced by managing the friction forces at blade root attachment. The nonlinear damping can be simulated directly by ABAQUS/Explicit method, convergence criteria often lead to excessive runtimes. Therefore a simple mass/spring model has been developed, which applies an exciting force to a system comprising two masses and springs to represent the blade and the root respectively and includes modelling of both the stick and slip forces of the root due to friction. Both the masses and their spring stiffness are chosen to produce either the sticking natural frequency (with infinite friction) or sliding natural frequency (zero friction). Using the simple two-mass model, the significant nonlinear response pattern is demonstrated. The resulting pattern has been verified against the ABAQUS/Explicit method. This allows the blade forced response prediction from the one way coupling to be further corrected to account for the nonlinear friction damping effect.
Aeroelastic Stability Assessment of an Industrial Compressor Blade Including Mistuning Effects
Authors: Zhai Yaoguang , Bladh Ronnie , Dyverfeldt Go¨ran
Published: May 2012

This paper presents a comprehensive investigation into the aeroelastic stability behavior of a transonic front blade in an industrial compressor when operating outside its normal range of service parameters. The evolution of the airfoil’s aeroelastic stability in the first flexural mode is studied as the front blade operation progresses towards choked flow conditions. First, linearized 3D flutter computations representing today’s industry standard are performed. The linearized calculations indicate a significant, shock-driven flutter risk at these off-design flow conditions. To further explore the aeroelastic behavior of the rotor and to find a viable solution toward flutter risk elimination, two parallel investigations are undertaken: (i) flow perturbation nonlinearity effects and potential presence of limit-cycle oscillation; and (ii) effects of blade mistuning and flutter mitigation potential of intentional mistuning, including its impact on forced response behavior. The nonlinear harmonic analyses show that the minimum aerodynamic damping increases rapidly and essentially linearly with blade oscillation amplitude beyond the linear regime. Thus, a state of safe limit-cycle oscillation is predicted for the fully tuned blade. Additionally, it is found that intentional, realizable blade frequency offsets in an alternating pattern efficiently stabilize the blade. Finally, it is verified that alternating mistuning has a beneficial effect versus the inevitable random mistuning also in the forced response.
Forced Response of Bladed Disks With Damping Mistuning
Authors: Khemiri Oualid , Martel Carlos , Corral Roque
Published: May 2012

The effect of mistuning on the vibration of bladed disks has been extensively studied in the past 30 years. Most of these analysis typically cover the case of small variations of the elastic characteristics (mass and stiffness) of the blades. In this work we study the not so common case of the forced response of a stable rotor with damping mistuning. The Asymptotic Mistuning Model (AMM) is used to analyze this problem. The AMM methodology provides a simplified model that describes the effect of blade to blade damping variation, and gives precise information on the underlying mechanisms involved in the action of damping mistuning.
Aerodynamic Shape Design Optimization for Turbomachinery Cascade Based on Discrete Adjoint Method
Authors: Zhang Chaolei , Feng Zhenping
Published: May 2012

Achieving higher aerodynamic performance in terms of efficiency, pressure ratio or stable operation range has been of interest to both researchers and engineers in the field of turbomachinery. The design of optimal shaped aerodynamic configurations based on Computational Fluid Dynamics (CFD) and predefined targets can be obtained by using deterministic search algorithms, which need to calculate the first and second order sensitivities of the objective function with respect to the design variables. With the characteristics of quick and exact sensitivity analysis, as well as less computational resource requirement, the adjoint method has become a research focus in aerodynamic shape design optimization over the past decades. In this paper, a discrete adjoint solver was developed and validated based on an in-house flow solver code. Moreover, a turbomachinery cascade optimization design system was established by coupling the flow solver, the discrete adjoint solver, the parameterization technology, the grid generation technology and the gradient-based optimization algorithms. During the development process of the discrete adjoint solver, the automatic differentiation tool was used in order to ease the construction of the discrete adjoint system based on the flow solver code. However, in order to save the memory requirement and to reduce the computational cost, the automatic differentiation tool was used selectively to build the fundamental subroutines. The top-most module of the discrete adjoint solver was established based on the discrete adjoint theory and the automatic differentiation technology manually. The treatments of the discontinuity in the flow field, such as strong shocks, and the imposition of strong boundary conditions which were implemented in the adjoint solver were discussed in detail. At the same time, several technologies were used to accelerate convergence. Based on the optimization system, a typical 2D transonic turbomachinery cascade was optimized under the viscous flow environment. The optimization results were analyzed in detail. The validity and efficiency of the present optimization design system were proved.
Analysis of Unsteadiness on Casing Treatment Mechanisms in an Axial Compressor
Authors: Legras G. , Gourdain N. , Tre´binjac I. , Ottavy X.
Published: May 2012

Control devices based on casing treatments have already shown their capability to improve the flow stability in compressors. However their optimization remains complex due to a partial understanding of the related physical mechanisms. The present paper proposes to use a budget analysis of the Navier Stokes equations to support the understanding of such flow phenomena. Based on the original work of Shabbir and Adamczyk (2005), the strength of the present contribution is to generalize the flow analysis method to all Navier-Stokes equations, including unsteady terms. A high-pressure multistage compressor equipped with circumferential casing grooves is chosen to demonstrate the potential of this approach. Steady and unsteady Reynolds-Averaged Navier-Stokes (URANS) equations are solved with a structured multi-blocks solver. Results are then briefly compared to experimental data to validate the numerical method. The analysis of the unsteady axial momentum equation for configurations with and without casing treatment points out some of the mechanisms responsible for the stability improvement. The analysis also indicates that the flow unsteadiness generated by upstream stator wakes (stator/rotor interaction) reduces viscous efforts and increases convective forces, significantly modifying the compressor stability. Finally, the proposed post processing method shows very interesting results for the understanding of circumferential grooves and it should be also used for non-axisymmetric casing treatment configurations.
Minimising Loss in a Heat Exchanger Installation for an Intercooled Turbofan Engine
Authors: Kwan Pok-Wang , Gillespie David R. H. , Stieger Rory D. , Rolt Andrew M.
Published: May 2012

An intercooled turbofan engine has been proposed within NEWAC (New Aero Engine Core Concepts, an European Sixth Framework Programme) using lightweight heat exchangers. The requirement for compactness has led to the need for zigzag heat exchanger arrangement where the heat exchanger matrices are inclined to the cooling flows approaching them, but such an arrangement creates non-uniform mass flows through the cold fluid side intercooler ducting and the intercooler heat exchanger matrices. Design guidelines aimed at minimizing aerodynamic losses caused by the flow mal-distribution in such ducting is reported. Minimising the loss has the effect of optimising the heat transfer performance. Flow velocities and pressure distributions were measured experimentally in a simplified model of a heat exchanger and simulated in Computational Fluid Dynamics (CFD). Good agreement was found between measurement and predictions of the flow distribution in the cold fluid side intercooler ducting downstream of the heat exchanger matrices. A dominant jetting flow in the centre of each exit passage was identified as a source of aerodynamic loss. The CFD simulation has also shown that the main source of aerodynamic loss arises from the severe flow mal-distribution within the heat exchanger matrices. From these results, design guidelines are presented in this paper for the ducting, based on CFD studies on a series of simplified heat exchanger arrangement geometries.
An Investigation Into Two Alternative Approaches for the Identification of SFD Bearings for Aeroengine Analysis
Authors: Groves Keir , Bonello Philip
Published: May 2012

Identification techniques provide a means of efficiently implementing complex nonlinear bearing models in practical turbomachinery applications. This paper considers both identification from an advanced numerical model and identification from experimental tests. Identification from numerical models is essential at the design stage, where rapid simulation of the dynamic performance of a variety of designs is required. Experimental identification is useful to capture effects that are difficult to model (e.g. geometric imperfections, compressibility and its effect on cavitation). With regard to identification from a numerical model, it was shown in a previous paper that the numerical solution of the incompressible Reynolds equation may be replicated using Chebyshev polynomial fits. Tests were performed on a simple rotor-bearing configuration incorporating an advanced numerical bearing model. The identified model was found to be able to match the accuracy of the numerical solution to the Reynolds equation while requiring a fraction of the computation time. In the present work the SFD identification scheme is applied to a realistically-sized representative whole-aeroengine model. It is shown that using recently introduced nonlinear solvers combined with the identified high accuracy bearing models it is possible to run full engine rotor-dynamic simulations, in both the time and frequency domains, at a fraction of the previous computational cost. One major drawback of the Chebyshev technique is that it is not amenable to experimental identification of actual bearings. For this reason, a second identification approach, involving the use of neural networks, is considered in this paper. A test rig that enables empirical identification of SFD forces has been constructed and details of the building and operation of the test rig is presented. The method used to ascertain the training data required by the neural network identification scheme, is also described.
A Comparison of Advanced Numerical Techniques to Model Transient Flow in Turbomachinery Blade Rows
Authors: Connell Stuart , Braaten Mark , Zori Laith , Steed Robin , Hutchinson Brad , Cox Graham
Published: May 2012

Computational predictions of the transient flow in multiple blade row turbomachinery configurations are considered. For cases with unequal numbers of blades/vanes in adjacent rows (“unequal pitch”) a computation over multiple passages is required to ensure that simple periodic boundary conditions can be applied. For typical geometries, a time accurate solution requires computation over a significant portion of the wheel. A number of methods are now available that address the issue of unequal pitch while significantly reducing the required computation time. Considered here are a family of related methods (“Transformation Methods”) which transform the equations, the solution or the boundary conditions in a manner that appropriately recognizes the periodicity of the flow, yet do not require solution of all or a large number of the blades in a given row. This paper will concentrate on comparing and contrasting these numerical treatments. The first method, known as “Profile Transformation”, overcomes the unequal pitch problem by simply scaling the flow profile that is communicated between neighboring blade rows, yet maintains the correct blade geometry and pitch ratio. The next method, known as the “Fourier Transformation” method applies phase shifted boundary conditions. To avoid storing the time history on the periodic boundary, a Fourier series method is used to store information at the blade passing frequency (BPF) and its harmonics. In the final method, a pitch-wise time transformation is performed that ensures that the boundary is truly periodic in the transformed space. This method is referred to as “Time Transformation”. The three methods have recently been added to a commercially-available CFD solver which is pressure based and implicit in formulation. The results are compared and contrasted on two turbine cases of engineering significance: a high pressure power turbine stage and a low pressure aircraft engine turbine stage. The relative convergence rates and solution times are examined together with the effect of non blade passing frequencies in the flow field. Transient solution times are compared with more conventional steady stage analyses, and in addition detailed flow physics such as boundary layer transition location are examined and reported.
Evaluation of Total Engine Performance Degradation Based on Modular Efficiencies
Authors: Ebmeyer Carsten , Friedrichs Jens , Wensky Tobias , Zachau Uwe
Published: May 2012

The current maintenance and overhaul of large civil jet engines is completely based on-condition and is widely customized to the individual requirements of the operation. Therefore, a very important factor for an effective and economic engine maintenance program is the investigation and appreciation of the current engine condition, as well as its individual deterioration mechanism. This paper is introducing a method to analyze the engine performance deterioration between two typical off-wing maintenance events (shop visits) so as to draw conclusions for maintenance planning and operation. In order to perform a precise evaluation the performance analysis is conducted on a modular level. Therefore the engine is divided into the following major modules: FAN, LPC, HPC, combustor, HPT, LPT and exhaust nozzle. The basis for the evaluation is the overhauled engine condition after a shop visit (pass-off test run) and the deteriorated engine condition after operation (incoming test run). These two points in the engine life cycle provide specific engine conditions that are to be analyzed by scientific and commercial software, and combined with a self-developed engine performance model in order to obtain the desired results: The individual engine deterioration during operation demonstrated by the differences of the modular performance between incoming test run and the last pass-off test run. In addition, to ensure the continuous monitoring of the performance status between the two test runs, it is important to analyze the “on-wing operation”. This is done using MTU’s Engine Trend Monitoring (ETM) system, which generates performance data based on the available in-flight data. In this paper an analysis example is used to present the analytic method and the obtained results. Reasons of deterioration are evaluated separately in reference to different environmental influences from specific geographical regions. In summary this paper introduces a solution to track the total engine performance based on modular evaluation values, starting at improvements for pass-off and incoming test runs as well as performance degradation during the on-wing time.
Probabilistic Life Assessment of Turbine Vanes
Authors: Arkhipov Alexander N. , Krasnovskiy Yevgeny E. , Putchkov Igor V.
Published: May 2012

Life of a gas turbine vane generally depends on different factors such as scatter of material properties, load variation and manufacturing tolerances. However, deterministic finite element (FE) life analysis gives just a discrete value typically based on the nominal or worst case conditions. It precludes considering sensitivity to the input parameters and obtaining the expected life range. To consider the possible variations of the input parameters from their nominal values, a probabilistic approach has been applied to compute the LCF (Low Cyclic Fatigue) and creep life distributions for the uncooled vane. The deterministic 3D FE life assessment of the gas turbine components is based on the input data such as physical and mechanical properties of the base material and coating at operating temperatures, nominal geometry of the component, thermal and mechanical loadings. Each of the above mentioned inputs has its own scatter band characterized either by average and minimum values of mechanical properties (tensile strength, LCF, creep) or by variations of manufacturing tolerances; thermal boundary conditions and gas pressure distribution. The probabilistic life analysis has been performed in order to assess individual impact of each input on vane’s life scatter. LCF and creep life distributions as well as variation of the base metal oxidation layer thickness have been obtained for each scatter factor and for their overall contribution. It is seen from results that LCF and creep lives of the analyzed vane have been influenced mainly by material properties and secondarily by OTDF (hot gas temperature variation in the circumferential direction) and uncertainties of thermal boundary conditions, which depended on the operation conditions of the engine. Manufacturing tolerances and alternation of ambient air temperature in the compressor intake have the lowest impact. The derived model is useful for the risk analysis or maintenance planning. For instance, it has been shown how probability of small fatigue crack indication in one vane can be extended onto the overall probability for the failure detection of n vanes at the stator stage during regular inspection. The probability of micro crack growth due to creep after the determined amount of operating hours for the single vane may be also redefined into the overall stage probability for the detection of n such vanes. To perform validation, normalized field data have been used for comparison with the analytical predictions. Good correlations between the field data and analytical predictions have been shown.
Failure Investigation of a 6th Stage Compressor Blade on a ETF40B Gas Turbine Engine
Authors: Slogar Gerrick , Prus Gary , Canilang Benjamin
Published: May 2012

Results from an engineering investigation, including findings, root cause and corrective action of a ETF40B compressor incident, are presented in this paper. Vericor’s ETF40B is used to power US NAVY Landing Craft Air Cushion (LCAC) vehicles that have undergone the Service Life Extension Program (SLEP). This paper discusses the initial review of the failed engine, the results of the engineering investigation that led to identifying root cause and the successful validation testing providing fleet corrective action. This paper will present the history of the investigation, the metallurgical analysis findings, the finite element stress analysis results performed on the blade, the Fourier analysis performed on the bleed system and engine testing conducted to verify the investigation analysis. Additionally, this paper will discuss what caused the high stress levels in the blade along with the subsequent analysis conducted to define modifications to correct the problem. The paper will conclude with the testing conducted to verify the design solution and how the final field and production compressor changes were incorporated.
Application of an Inverse Cascade Analytical and Numerical Design Method Used in the Design of Axial Fans
Authors: Miclea-Bleiziffer Mihai , Epple Philipp , Smith Henrik , Semel Matthias , Delgado Antonio
Published: May 2012

This contribution presents an inverse design method for axial flow fans validated by 3-D RANS numerical simulations. It is based on an analytical performance derivation for different free-vortex distributions combined with a mean-line pre-design tool and supported by a new 2-D cascade potential flow computational method. This computes the 2-D planar cascade flow field and the inviscid deflection angle of thin arbitrary shape airfoils at different machine radii followed by a numerical radial integration into a quasi-3-D method. For validation, several fan designs were investigated, showing the influence of important design parameters as well as the potential of the present method by redrafting an industrial fan. The 2-D planar potential cascade code was validated against the Weinig’s coefficient diagram for flat plate and circular arc cascades.
Performance Comparison of Aero-Engine Thrust Augmentors Stabilized by Bluff Body Flame Holders and a Flame-Holder-Less Concept
Authors: Tammer Armin , Cutright John T. , Neumeier Yedidia , Zinn Ben T.
Published: May 2012

This study investigates the performance benefits of a flame-holder-less flame stabilization concept for thrust augmentors compared to the common flame holder design. The concept proposes to burn a small portion of the augmentor fuel in a rich mixture with air bled from the compressor to produce a highly reactive partially oxidized fuel-air mixture (POx). The POx mixture is injected into the turbine exit flow to enhance combustion kinetics in order to achieve stable combustion in the augmentor. Thermal efficiency during wet and dry operation is compared, taking into account both the pressure losses due to the flame holders and the reduction of core air for the flame-holder-less concept. Furthermore, the thrust-to-weight ratio and the corresponding flight range have been investigated with respect to the system weight and the induced losses. It was found that the thermal efficiency during dry operation is significantly increased when the pressure losses of the flame holders are eliminated. During wet operation, it was calculated that a flame holder system with only 2% total pressure loss of the flow would operate at the same thermal efficiency as the flame-holder-less concept when 3% air is bled from the compressor. If, for an engine operating at these conditions, the flame-holder-less system could maintain stable combustion using less than 3% bleed air, it would increase the thermal efficiency of that engine during wet operation. The results also suggest that a flame-holder-less system is lighter weight and has the potential to increase engine thrust-to-weight ratio and extend flight range when compared to a flame holder system designed to operate at the same overall engine thermal efficiency.