bluuebunny/crossref_metadata_2025 huggingface dataset:
Effect of Radial Location of Nozzles on Performance of Pre-Swirl Systems
Authors: Lewis Paul , Wilson Mike , Lock Gary , Owen J. Michael
Published: August 2009

This paper investigates the effect of the radial location of the inlet nozzles on the performance of a direct-transfer pre-swirl system in a rotor-stator wheel-space. A commercial code is used to solve the Reynolds Averaged Navier Stokes (RANS) equations using a high-Reynolds-number k-ε / k-ω turbulence model with wall functions near the boundary. The 3D steady state model has previously been validated against experimental results from a scale model of a gas turbine rotor-stator system. Computations are performed for three inlet-to-outlet radius ratios, rp/rb = 0.8, 0.9 and 1.0, a range of pre-swirl ratios, 0.5 < βb < 2.0, and varying flow parameter, 0.12 < λT < 0.36. The rotational Reynolds number for each case is 106. The flow structure in the wheel-space and in the region around the receiver holes for each inlet radius is related to the swirl ratio. The performance of the system is quantified by two parameters: the discharge coefficient for the receiver holes (Cd,b) and the adiabatic effectiveness for the system (Θb,ad). As in previous work, the discharge coefficient is found to reach a maximum when the rotating core of fluid is in synchronous rotation with the receiver holes. As the radius ratio is increased this condition can be achieved with a smaller value for pre-swirl ratio βb. A simple model is presented to estimate the discharge coefficient based on the flow rate and swirl ratio in the system. The adiabatic effectiveness of the system increases linearly with pre-swirl ratio but is independent of flow rate. For a given pre-swirl ratio, the effectiveness increases as the radius ratio increases. Computed values show good agreement with analytical results. Both performance parameters show improvement with increasing inlet radius ratio, suggesting that for an optimum pre-swirl configuration an engine designer would place the pre-swirl nozzles at a high radius.
Experimental Study of Internal Heat Transfer Coefficients in a Rectangular, Ribbed Channel Using a Non-Invasive, Non-Destructive, Transient Inverse Method
Authors: Heidrich Peter , von Wolfersdorf Jens , Schnieder Martin
Published: August 2009

This paper describes a non-invasive, non-destructive inverse measurement method that allows one to determine heat transfer coefficients in internal passages of real turbine blades experimentally. For this purpose, a test rig with a fast responding heater was designed to fulfill the requirement of a sudden increase in the air temperature within the internal cooling passages. The outer surface temperatures of the specimen were measured using an infrared camera. To suggest the spatial distribution of the internal heat transfer coefficients from the transient characteristics of the outside surface temperature the inverse heat transfer problem was solved. Differing from former studies which made a thin wall assumption, the conduction inside a finite wall was modelled. Based on a one-dimensional forward solution the best fitting optimization method, the Levenberg-Marquardt algorithm, was chosen. This was verified with artificial data including random noise with positive results. Experimental data were measured for a rectangular H/W = 1:4 aspect ratio channel made of stainless steel with parallel 90° and 45° ribs at Reynolds numbers from 25,000 to 80,000. Results of 90° ribs were compared with simultaneously acquired data using the transient liquid crystal technique. Furthermore the influence of Reynolds number on pitch averaged heat transfer results were evaluated for both rib configurations. These results based on infrared data were compared with earlier studies. It is concluded that the presented experimental measurement method using the transient inverse method could be used to quantitatively determine heat transfer coefficients in internal passages of real turbine blades.
Development and Testing of the Upgraded High Pressure Turbine Section for an Industrial Gas Turbine
Authors: Aschenbruck E. , Frank D. , Korte T. , Mu¨ller R. , Orth U.
Published: August 2009

As part of an ongoing development program to increase power output and efficiency of the THM 1304 gas turbine, modifications were made to the high pressure turbine. The modifications include but are not limited to blade and vane aerodynamics, cooling system and clearance control, mechanical design and materials. The development was to achieve the following goals: • Intensified blade and vane cooling to permit higher turbine inlet temperatures and to further extend service lifetime; • Improved aerodynamic performance; • Blades with pre-loaded tip shrouds to achieve low vibration amplitudes in a broad operating speed range; • Rotor design modifications to simplify assembly and disassembly; • Modified vane carrier and casing designs for optimal tip clearance control and turbine performance. The improved high pressure turbine was extensively tested in MAN TURBO’s full-load gas turbine test facility. Test results verified that component temperatures were within the expected range and design targets have been achieved. The first production gas turbine equipped with the upgraded high pressure turbine was installed in May 2004 as a gas compressor driver. To date a total of 11 units have gone into operation including units for power generation. Dry low emission technology is used on all engines. Every unit is monitored by an online data monitoring system and visually inspected in shorter intervals to verify the behavior in the field. Operation of the fleet is flawless at this time.
Identification and Correction of Rotor Instability in an Oil-Free Gas Turbine
Authors: Lubell Daniel R. , Wade Jonathan L. , Chauhan Navjot S. , Nourse John G.
Published: August 2009

The direction of advanced gas turbines and other turbomachinery has been towards oil-free designs, enabled by the significant improvements of high temperature foil bearings. The advantages of oil-free gas turbines have been studied and shown to be realistic. However, the oil-free technology is still at an early stage in its development relative to conventional oil lubricated turbomachinery systems which have been studied and manufactured for about 100 years, and the bearings even longer. Oil-free gas turbines are most successful as a system design initiated with oil-free bearings. Making these successful designs requires knowledge of the strengths and weaknesses of integrating oil-free bearings. A common example is foil bearings, the type typically considered for oil-free gas turbines. These bearings are lower in damping than their oil lubricated counterparts. Therefore special considerations are made by the experienced oil-free gas turbine designer early in the design process. Knowledge of the opportunities for instability that are not as common in conventional turbomachinery provides value to the final design. This paper presents the identification and correction of rotor instability in an oil-free microturbine of a 65 kW system. The manufacturer put significant effort into identifying the root cause of the seemingly random occurrences of rotor instability, in order to improve yield for acceptance tests. Through the application of conventional rotordynamics theory and techniques, combined with 3-D imaging of complex cast parts, the root cause was identified as an Alford’s-type force at the turbine driven by critical machined and cast features of the turbine wheel that would not have been important in a conventional oil lubricated turbomachine. A successful corrective process has been put in place, providing final confirmation of the root cause.
Redesigning Annular Turbine Blade Rows Using a Viscous-Inviscid Inverse Design Method
Authors: Pa´scoa J. C. , Mendes A. C. , Gato L. M. C.
Published: August 2009

This paper presents the results of the aerodynamic redesign of an annular turbine blade row. The inverse method herein applied is an extension to 3D of an iterative inverse design method based on the imposition of the blade load, thickness distribution and stacking line. We define a mass-averaged mean tangential velocity over one blade pitch, ru¯θ, as the main design variable, since its derivative is related to the aerodynamic load. A time-lagged formulation for the 3D camber surface generator is given in order to include the blade thickness distribution into the design algorithm. The hybrid viscous-inviscid design code comprises three main components: the blade update algorithm; a fast inviscid 3D Euler code; and a viscous analysis code. The blade geometry and flow conditions are typical of LP turbine nozzle guide vanes. The design method will demonstrate its ability to redesign blade rows that achieve lower flow losses and a more uniform exit flow angle distribution. The performance of the new blades is checked by means of a Navier-Stokes computation using the κ–ε turbulence model. The presented results show a minor decrease in the losses and a better redistribution of the exit flow angle.
Static and Dynamic Analysis of 2.5 MW Power Generation Gas Turbine Unit
Authors: Kim SooYong , Choe SeungJoo , Kovalevskiy Valeri P. , Kim Dong Hwa
Published: August 2009

Development of a numerical mathematical model to calculate both the static and dynamic characteristics of a multi-shaft gas turbine consisting of a single combustion chamber with advanced cycle components such as intercooler and regenerator is presented in the paper. The numerical mathematical model is based on simplified assumptions such that quasi-static characteristic in turbo-machine and injector is used, pressure loss and heat transfer relation for static calculation neglecting fuel transport time delay can be employed. For study of static and dynamic regime, a control algorithm with following assumptions was made. When the inlet temperature of HPT exceeds the preset value, a signal to stop or open the fuel valve is issued and with a surge margin coefficient Ks<1.1, signal to open the compressor anti surge valve with proportional deflection is issued. At Ks<1.18 or as in starting, a maximum control signal will be issued. Static characteristic of the gas turbine engine in terms of rotation speed ω at ambient temperature of ta = 15°C is carried out. The rotation speed of power turbine is decreased from 20,000 to 4500 rpm, that means 4.4 times, whereas the rotation speed of gas generator is decreased only 2 ∼ 2.5 times. The air flow rate at this time decreased about 3 times. HPC had enough surge margins at all operation ranges. LPC starts to operate with the power turbine speed of 15,000 rpm (approximately 1 MW power output, that means, 40% of the nominal power) and reaches minimal permissible surge margin for safe operation bypassing the air through supercharging channel. Calculation of heat balance showed the difference did not exceed more than 0.06%, validating the accuracy of the applied method. The accuracy of each calculation is confirmed by monitoring mass and energy balance for different time steps of integration.
Aerodynamically Coupled Flutter of Multiple Blade Rows
Authors: Namba Masanobu , Kubo Ayumi
Published: August 2009

This paper deals with the aeroelastic instability of vibrating multiple blade rows under aerodynamic coupling with each other. A model composed of three blade rows, e.g., rotor-stator-rotor, in which blades of the two rotor cascades are simultaneously vibrating, is considered. The generalized aerodynamic force on a vibrating blade consists of the component induced by the vibrating motion of the blade itself and those induced by vibrations not only of other blades in the same cascade but also of blades in another cascade. To evaluate the aerodynamic forces, the unsteady lifting surface theory for the model of three blade rows is applied. The equations describing motions of blades are coupled via the aerodynamic forces. The so-called k method is applied to determine the critical flutter conditions. A numerical study has been conducted. The flutter boundaries are compared with those for a single blade row. It is shown that the effect of the aerodynamic coupling significantly modifies the critical flutter conditions.
Overspray and Interstage Fog Cooling in Compressor Using Stage-Stacking Scheme: Part 1—Development of Theory and Algorithm
Authors: Wang Ting , Khan Jobaidur R.
Published: August 2009

Compressor intercooling has traditionally been employed to reduce compressor work and augment gas turbine output power. Conventional intercooling schemes are usually applied through non-mixed heat exchangers between two compressor stages or by cooling the outside of the compressor casing. Any cooling schemes that may affect the flow field inside the compressors have not been favorably considered due to concerns of any disturbance that might adversely affect the compressor’s performance stability. As the inlet fog cooling scheme has become popular as an economic and effective means to augment gas turbine output power on hot or dry days, consideration has been given to applying fog cooling inside the compressors by injecting fine water droplets between stages (i.e. interstage fogging). This paper focuses on developing a stage-by-stage wet-compression theory for overspray and interstage fogging that includes the analysis and effect of pre-heating and pre-cooling at each small stage of the overall compressor performance. An algorithm has been developed to calculate the local velocity diagram and allow a stage-by-stage analysis of the fogging effect on airfoil aerodynamics and loading with known 2-D meanline rotor and stator geometries. Thermal equilibrium model for water droplet evaporation is adopted. The developed theory and algorithm are integrated into the system-wise FogGT program to calculate the overall gas turbine system performance.
Numerical Simulation of Inter-Turbine Burner (ITB) Flows With the Inclusion of V-Gutter Flame Holders
Authors: Lin Cheng-Xian , Sekar Balu , Zelina Joseph , Holder Richard Jack , Thornburg Hugh
Published: August 2009

In this paper, a three-dimensional numerical simulation has been performed to study the complex reactive flows during the combustion in an inter-turbine burner (ITB) with the inclusion of V-gutter flame holders. The ITB configuration, which has straight radial vanes (SRV), was based on the innovative, high efficiency, high-g Ultra-Compact Combustor (UCC) concept developed at the Air Force Research Laboratory (AFRL). The V-gutter’s angle of attack was varied from −10 to 10 degrees at fixed JET-A fuel and air injections. The turbulent flow in the ITB was modeled with a RANS-based realizable k-ε turbulence model, while the spray combustion is modeled with an eddy-dissipation model on an unstructured grid. Numerical results indicate that the V-gutter not only generates vortices behind itself, but also alters the turbulent flow feature and mixing behavior between main air flow and the circumferential and SRV cavity flows within the ITB. The exit temperature profile of the ITB can be modified substantially by the inclusion of the V-gutters at different angle of attack. The additional pressure drop incurred by the addition of the V-gutter was found to be less that 1%. Details of the vane cavity dynamics and increased entrainment physics are also discussed in the paper.
Modeling and Experimental Evaluation on Torque Loss in Turbine Test Rig
Authors: Kang Jeong-Seek , Cha Bong-Jun , Ahn Iee-Ki
Published: August 2009

To evaluate the performance of a turbine through turbine rig test, torque or power generated from the test turbine should be measured. This is measured from a dynamometer or torquesensor installed in the test rig. So, to evaluate the performance of turbine with high precision, the accurate measurement of power or torque is necessary. However, there is an intrinsic difficulty as not all the power generated by the turbine is measured by the dynamometer or torquesensor. A small portion of power generated from test turbine is dissipated through bearing loss and windage loss. This dissipated energy is called mechanical loss of test rig. Therefore, it is necessary to measure the mechanical loss of the test rig for the accurate evaluation of the turbine performance. This paper divides mechanical loss as bearing loss, disk windage loss and extra windage loss. Spin down tests are performed in a 1-stage axial turbine test rig to evaluate each losses. It is found that the dominant loss is bearing loss and the total mechanical loss amounts to 0.78∼1.4% of energy generated at the turbine. And the effect of bearing temperature is investigated and it is found that the mechanical loss is dependent on the bearing temperature and that it increases with decreasing bearing temperature.
Successful Oil-Free Version of a Gas Compressor Through Integrated Design of Foil Bearings
Authors: Wade Jonathan L. , Lubell Daniel R. , Weissert Dennis
Published: August 2009

In the pursuit of higher power density turbomachinery the rotor speeds and temperatures have been increased to the limits of conventional oil lubricated bearings. Additionally, with conventional oil lubricated bearings there is a risk of oil contaminating the working fluid; this is unacceptable for some applications. When properly designed and integrated the foil gas bearing is one option that can easily operate at higher temperatures and DN’s than conventional oil lubricated bearings systems. This is a case study of a small variable speed gas compressor that progressed through a variety of bearing configurations. The compressor was initially designed with ball bearings; the ball bearing design did not meet the compressor life targets for some operating regimes, requiring extra maintenance. The second iteration was to move to a foil bearing design; unfortunately because of design constraints an under-sized bearing was selected. The under-sized foil bearing provided only marginally better unit life, and was much more sensitive in the build process and the acceptance test pass rate fell dramatically. Unique field operation experience showed a variety of failures and successes from a marginal design. Finally, a properly sized foil bearing was integrated into the pump, capitalizing on the foil bearing strengths. With the properly sized foil bearings the pumps have seen a 100% acceptance test pass rate, no field failures, and the pumps are exceeding the desired life without any maintenance requirements.
The Primary Discussion of Blade Sections for Application of Helium Compressor
Authors: Liu Yu E. , Sun Yufeng
Published: August 2009

The investigation of the airfoil is a key design issue in the development and deployment of helium compressor for High Temperature Gas Cooled Reactor. As the helium physical properties are different from the air’s, three flow cases are studied to find out the similarities between helium and air using Fluent in this paper. They are flow within the divergent nozzles, flow around cylinders and flow over cascades under same inlet conditions and geometry. The results indicated some similarity relationships between helium and air cascade flow. NACA 65-(12A10)10 compressor blade was used as sample blade for both air and helium cascades. To test the performance of the blades, calculations were carried out under the condition of kinematics similitude, but different working fluid. The numerical simulation results of air cascades are correlated with NACA experiment results of air subsonic cascades very well. The numerical simulation models for the air subsonic cascades were used to test the performance of similar helium cascades. The cascade performance for helium flows were compared with the air flows and the results are presented in this paper. The purpose of numerical studies is to use blades in the current air compressor as the blades for helium compressor, to develop scaling blades in NACA 65-(12A10)10 blade size range, and to identify practical limitations to the scaling of current conventional blades. The analysis and conclusions presented in this paper can be used as a reference to helium compressor design features and the experimental research works.
Loss Reduction in Compressor Cascades by Means of Passive Flow Control
Authors: Hergt A. , Meyer R. , Mu¨ller M. W. , Engel K.
Published: August 2009

Secondary flow effects like the corner stall between the wall and the vane in a compressor stage are responsible for a large part of total pressure losses. An extensive experimental study of flow control in a highly loaded compressor cascade was performed in order to decrease the separation and reduce the losses by means of vortex generators. The vortex generators were attached at the surface of the cascade side walls. These flow control devices produce strong vortices, which enhance the mixing between the main flow and the decelerated boundary layer at the side wall. Thus, the corner flow separation and the total pressure losses could be reduced. The experiments were carried out with a compressor cascade at a high-speed test facility at the DLR in Berlin at minimum loss (design point) and off-design of the cascade at Reynolds numbers up to Re = 0.6 × 106 (based on 40 mm chord) and Mach numbers up to M = 0.7. The cascade consisted of five vanes. The blade profiles are comparable to the hub section of the stator vanes used in the transonic compressor test rig running at Technische Universita¨t Darmstadt. In the range between −2° and +4° angle of incidence the total pressure losses of the cascade could be reduced up to 4.6% by means of vortex generators, whereas the static pressure rise was not influenced. Based on the results of the cascade measurements, the vortex generators were applied in front of the stator row of the single stage axial compressor at Technische Universita¨t Darmstadt. A numerical simulation of the compressor flow provided an indication for the adjustment of the vortex generators at the hub and casing. In the experiments the pressure rise and the efficiency of the axial compressor was measured and it could be shown that vortex generators partially improve the efficiency.
A Numerical Study on Increasing Film Cooling Effectiveness Through the Use of Sister Holes
Authors: Ely Marc J. , Jubran B. A.
Published: August 2009

Film cooling has been the primary focus of turbine blade cooling research for the past half century. However, as engines become more powerful, more effective non-traditional means of cooling become necessary. The current study branches out into a new scheme for film cooling; sister holes. The geometry of the current work makes use of three cylindrical holes inclined at 35° to the horizontal: one primary injectant hole bound by two sister holes. Numerical simulations were run with blowing ratios varying from M = 0.2 to M = 1.5, using the realizable k-ε turbulence model with near-wall modeling. The results were analyzed for both adiabatic thermal effectiveness as well as vortex production due to flow mixing. In general, sister holes offer significant advantages in thermal protection over their single hole counterparts both laterally and along the centre-line, particularly in regions close to the hole. Simulations showed that the laterally averaged adiabatic thermal effectiveness increased by a factor of 1.35 for M = 0.2 up to a factor of 1.62 for M = 1.5. Similarly, the centre-line effectiveness increased by a factor of 1.22 at M = 0.2 up to a factor of 1.68 at M = 1.5. These benefits are heavily weighted by the near-hole region; however, increases are evident throughout the computational domain. This sister hole technique offers significant advantages with minimal penalties, making it a valuable candidate for future blade cooling applications.
On the Numerical Prediction of Finite Length Squeeze Film Dampers Performance With Free Air Entrainment
Authors: Me´ndez Tilmer H. , Ciaccia Marco A. , Torres Jorge E. , Di´az Sergio E.
Published: August 2009

Squeeze Film Dampers (SFDs) are commonly used in turbomachinery to dampen shaft vibrations in rotor-bearing systems. The main factor deterring the success of analytical models for the prediction of SFD’s performance lays on the modeling of dynamic film rupture. Usually, the cavitation models developed for journal bearings are applied to SFDs. Yet, the characteristic motion of the SFD results in the entrapment of air into the oil film, producing a bubbly mixture that cannot be represented by these models. There is a need to identify and understand the parameters that affect air entrainment and subsequent formation of a bubbly air-oil mixture within the lubricant film. A previous model by Diaz and San Andre´s advanced estimation of the amount of film-entrapped air, based on a non-dimensional number that related both geometrical and operating parameters but limited to the short bearing approximation (i.e., neglecting circumferential flow). The present study extends their work to consider the effects of finite length-to-diameter ratios. This is achieved by means of a finite volume integration of the two-dimensional, Newtonian, compressible Reynolds equation combined with the effective mixture density and viscosity defined in the work of Diaz and San Andre´s. A flow balance at the open end of the film is devised to estimate the amount of air entrapped within the film. The results show, in dimensionless plots, a map of the amount of entrained air as a function of the Feed-Squeeze Flow Number, defined by Diaz and San Andre´s, and the Length-to-Diameter Ratio of the Damper. Entrained air is shown to decrease as the L/D ratio increases, going from the approximate solution of Diaz and San Andre´s for infinitely short SFDs down to no air entrainment for an infinite length SFD. The results of this research are of immediate engineering applicability. Furthermore, they represent a firm step to advance the understanding of the effects of air entrapment on the performance of SFDs.
Development and Verification of a Non-Linear Algebraic Model for the Turbulent Heat Fluxes in Rotating Passages
Authors: Younis B. A. , Weigand B. , Mohr F. , Schmidt M.
Published: August 2009

We propose a new approach to modeling the effects of system rotation on the turbulent scalar fluxes. The approach is based on extension to rotating frames of an algebraic model derived using tensor representation theory. The model is formulated to allow for the turbulent scalar fluxes to depend on the details of the turbulence field (via the Reynolds stresses), and on the gradients of both the mean velocity and temperature. Such dependence, which is absent from conventional models, is required by the exact equations governing the transport of the heat fluxes. The model’s performance is assessed by comparisons with results from recent Direct Numerical Simulations of flows in channels rotated about their streamwise, spanwise and wall-normal axes. The results show that the model yields results that are in good correspondence with the DNS results.
A Cost Effective Advanced Emissions Monitoring Solution for Gas Turbines: Statistical Hybrid Predictive System That Accurately Measures Nitrogen Oxides, Carbon Monoxide, Sulfur Dioxide, Hydrocarbon and Carbon Dioxide Mass Emission Rates
Authors: Swanson Brian
Published: August 2009

U.S. Federal regulations under Title IV of the Clean Air Act Amendments promulgated in 1990 require continuous monitoring of nitrogen oxides (NOx) and carbon dioxide emissions from large gas turbines. Local, regional, or State authorities may mandate continuous monitoring for carbon monoxide, sulfur dioxide, volatile organic compounds, and other specific pollutant parameters. U.S. regulations that require continuous emissions monitoring systems (CEMS) also allow for the use of predictive approaches as an alternative providing the installed predictive emissions monitoring system (PEMS) meets rigorous performance specification criteria and the site performs ongoing quality assurance tasks such as periodic audits with portable analyzers and annual accuracy testing. A statistical hybrid predictive emission monitoring system (PEMS) has been deployed at numerous sites in the United States to meet EPA requirements for continuous monitoring of gas turbine pollutant emissions. This paper discusses specific implementations of a unique cost-effective statistical hybrid PEMS on various classes of gas turbines ranging in size from 60kW to 180 MW, both gas-fired and liquid-fired units, in simple cycle and combined cycle mode of operation. The turbines were equipped with a variety of NOx control strategies including dry low NOx, steam and water injection, solid post-combustion catalyst, SoLoNOx™, and selective catalytic reduction. In each instance the predictive engine operated on training data of at least three days and up to ninety days as required to develop a robust empirical model of the emissions. Each model was subsequently evaluated using standard U.S. EPA performance specification test methods. The results of PEMS performance testing on these gas turbines are presented along with additional information regarding the quality assurance and quality control procedures put in place and the costs to support the ongoing operation of the deployed compliance statistical hybrid PEMS.
Radial Swirler Designs for Ultra-Low NOx Gas Turbine Combustion
Authors: Andrews Gordon E. , Escott Nick , Mkpadi Michael C.
Published: August 2009

Four radial swirler vane passage designs were investigated for low NOx lean well mixed combustion of natural gas at simulated gas turbine primary zone conditions in terms of reference velocities and 600–740K air inlet temperatures at atmospheric pressure. Each radial swirler had eight vane passages and the four passage designs were curved, aerodynamic tapered flat bladed, rectangular and circular passage cross sections. The first two designs had an area that decreased towards the passage outlet and the last two designs were constant area vane passages. The swirler exit diameter d (76mm) to the combustor diameter D (140mm) expansion ratio, D/d, was 1.84. Three methods of fuel injection were investigated: eight single fuel holes per radial vane passage injection, swirler outlet throat 8 hole wall injection and 8 hole radially outward central injection. The results showed that the different radial swirler designs did not have a strong influence on the NOx emissions, compared with the stronger influence of the fuel injection location. All radial vane swirler passage designs gave minimum NOx emissions in the 1–2 ppm range. The radial swirler design influenced which fuel injection location, passage or exit throat wall, gave the lowest NOx emissions. The results were compared with the CPMCP ultra-low NOx emissions concept to show that the present rapid fuel and air mixing radial swirler results are equal to the best premixed low NOx designs.
Numerical Investigation on Unsteady Effects of Hot Streak on Flow and Heat Transfer in a Turbine Stage
Authors: An Bai-Tao , Liu Jian-Jun , Jiang Hong-De
Published: August 2009

This paper presents a detailed flow and heat transfer characteristic analysis on gas turbine first stage turbine under hot streak inlet conditions. Two kinds of inlet total temperature conditions are specified at the turbine stage inlet. The first is uniform inlet total temperature, and the second is hot streak 2D total temperature contour. The two kinds of inlet conditions have the same mass-averaged total temperature, and the same uniform inlet total pressure. The hot streak total temperature contours are obtained according to the exit shape of an annular-can combustor. The ratio of the highest total temperature in the hot streak to the mass-averaged total temperature is about 1.23, and one hot streak corresponds to two vane passages and four blade passages. Six hot streak circumferential positions relative to the vane 1 leading edge varied from −2% to 81% pitch are computed and analyzed. Results show that hot streak obviously increases the non-uniform degree of vane heat load in comparison with the uniform total temperature inlet condition. The change of hot streak circumferential position leads to the circumferential parameter variation at stator exit, and also leads to different transient periodic fluctuating characteristics of heat load and pressure on rotor blade surface. The hot streak of relative pitch at 65% obtains similar heat load for the two vanes corresponding to one hot streak and small fluctuation of the averaged heat load on rotor blade.
Near Field of Film Cooling Jet Issued Into a Flat Plate Boundary Layer: LES Study
Authors: Peet Yulia V. , Lele Sanjiva K.
Published: August 2009

We report results from a computational study of film cooling from cylindrical holes inclined at 35 degrees with respect to a flat surface using Large Eddy Simulations (LES). The hole length is L/d = 3.5, distance between the holes is P/d = 3, boundary layer above the flat surface is turbulent with Reθ = 938, density ratio = 0.95, velocity ratio = 0.5. All pertinent components of geometry, namely, supply plenum, film hole and crossflow region above the test surface, are simulated. The simulations are performed using a multicode approach, where a low Mach number code is employed inside the plenum and in the film hole, and a compressible code is used for the flow above the test surface. Flow inside the plenum, film hole and above the test surface is analyzed. Mean velocity and turbulence characteristics in the near field of the jet injection obtained in the simulations are compared to experimental data of Pietrzyk et al. [1]. Adiabatic film cooling effectiveness is estimated and compared with experiments of Sinha et al. [2]. Relation of the coherent vortical structures observed in the flow to film cooling performance is discussed. Advantage of LES over RANS methods for this type of flow is confirmed by showing that spanwise u′w′ shear stress and lateral growth of the jet are predicted correctly in the current LES as opposed to typical RANS computations.
3D Numerical Investigation of Tandem Airfoils for a Core Compressor Rotor
Authors: McGlumphy Jonathan , Ng Wing-Fai , Wellborn Steven R. , Kempf Severin
Published: August 2009

The tandem airfoil has potential to do more work as a compressor blade than a single airfoil without incurring higher losses. The goal of this work is to evaluate the fluid mechanics of a tandem rotor in the rear stages of a core compressor. As such, the results are constrained to shock-free, fully turbulent flow with thick endwall boundary layers at the inlet. A high hub-to-tip ratio 3D blade geometry was developed based upon the best-case tandem airfoil configuration from a previous 2D study. The 3D tandem rotor was simulated in isolation in order to scrutinize the fluid mechanisms of the rotor, which had not previously been well documented. A geometrically similar single blade rotor was also simulated under the same conditions for a baseline comparison. The tandem rotor was found to outperform its single blade counterpart by attaining a higher work coefficient, polytropic efficiency and numerical stall margin. An examination of the tandem rotor fluid mechanics revealed that the forward blade acts in a similar manner to a conventional rotor. The aft blade is strongly dependent upon the flow it receives from the forward blade, and tends to be more three-dimensional and non-uniform than the forward blade.
Thermo Mechanical Fatigue Behaviour of Bare and Coated CMSX-4
Authors: Coppola T. , Riscifuli S. , Tassa O. , Pasquero G.
Published: August 2009

Highly cooled turbine blades undergo very high thermal gradients during rapid engine Idle-Max-Idle cycling. Traditional isothermal fatigue data are often insufficient for predicting service lives. A complete set of high temperature tests, in the range of 750 to 1050°C, was performed on single crystal alloy CMSX-4. The test program comprised tensile, creep, low cycle fatigue (LCF) and thermo-mechanical fatigue (TMF) tests. In particular the cycle time for TMF was 3 min., aiming to simulate the real high-power transient conditions in aircraft engines. Clockwise and counter-clockwise diamond cycle types were applied on bare and coated specimens to investigate their influence on the fatigue limit. The comparison of the results obtained with the available ones from open literature is discussed.
Aerodynamic Optimization Design of Vaned Diffusers for the 100 kW Micro Gas Turbine’s Centrifugal Compressor
Authors: Xi Guang , Wang Zhiheng , Zhang Chunmei , Yuan Minjian
Published: August 2009

In this paper the design optimization of vaned diffuser for the 100kW microturbine’s centrifugal compressor is carried out. The forward-loaded and the conventional airfoil diffusers are respectively redesigned based on the surrogate model and the three dimensional viscous flow analyses. The objective of the optimization is to redesign the diffuser that assures, for a given operating condition of the centrifugal impeller, the stage isentropic efficiency to be highest. Using the surrogate model the optimization process is accelerated and the 3D flow analysis’s application to the practical engineering design is efficiently realized. To validate the optimization result, the compressor stage performance test is performed on a high speed centrifugal compressor test rig with one original diffuser and its redesigned, respectively.
Micro Gas Turbine Test Rig for Hybrid Power Plant Application
Authors: Hohloch Martina , Widenhorn Axel , Lebku¨chner Dominik , Panne Tobias , Aigner Manfred
Published: August 2009

Within the scope of a hybrid power plant project a micro gas turbine test rig was developed and is actually under construction at the DLR Institute of Combustion Technology. The test rig consists of a Turbec T100PH micro gas turbine and the required piping system for the hybrid application. Instead of a real solid oxide fuel cell (SOFC) stack a fuel cell simulator is used to avoid any risks for the sensitive and expensive real device. This simulator is able to emulate the SOFC interface conditions. The present paper reports the underlying pressurized hybrid power plant cycle, the setup of the test rig and the selection of the subsystems. Initially the micro gas turbine, equipped with a detailed instrumentation, was analyzed separately. First experimental data obtained with the micro gas turbine are presented.
Basic Architecture and Sizing of Commercial Aircraft Gas Turbine Oil Feed Systems
Authors: Hart Ken
Published: August 2009

The specific thrust and fuel efficiency of gas turbine engines have improved considerably over the last 20 years primarily due to technology advances in materials science and main gas path thermo-fluid analysis. However, a major influence on the improved engine reliability demonstrated over that same timeframe is the performance of engine mechanical systems which rely on the consistent delivery and recovery of oil. The general architecture of oil systems has changed little over this period. However, detailed attention to how oil system components behave and interact with each other, and with the rest of the engine systems, has helped achieve improved engine mechanical performance. For derivative engines this process is facilitated by a substantial amount of in-service data, reflecting both good and bad experiences, which helps to identify the poor designs, manufacturing anomalies or unusual operational circumstances that are likely to cause problems and the design changes that are needed to solve them. However, new engines being designed and certificated in ever-reducing timescales, do not necessarily have the luxury of such directly relevant, in-service, performance data. Hence it is essential that oil systems are designed with a full appreciation of component and system capabilities and the options available. This knowledge is based fundamentally on a general awareness of how an engine oil system achieves its objectives, why sometimes it doesn’t, and how it integrates with the rest of the engine. This paper aims to identify some of the decisions that need to be made regarding system architecture and component sizing when designing a gas turbine oil system. Full attention to these matters during the engine design phase should lead to reduced development and in-service problems, less use of development innovation to solve problems and ultimately improved engine reliability.
Jet Noise Reduction for High Speed Exhaust Systems
Authors: Martens Steven , Haber Ludwig
Published: August 2009

Jet noise has been an environmental issue since the advent of jet aircraft. The past five decades have seen much research into solving this very difficult challenge for a variety of applications. The Supersonic Transport (SST), High Speed Civil Transport (HSCT), and a variety of supersonic business jet (SSBJ) applications all face significant jet noise challenges. Jet noise from high performance military aircraft has also received growing attention. The continuous drive to higher specific thrust results in increasing jet noise levels. Compounding this is that many military bases, Naval in particular, are located in desirable locations on the coast, and surrounding communities are encroaching closer to these bases. In this paper we will conduct a survey of some jet noise reduction technologies for high-speed exhaust systems investigated in the past, as well as some of the implementation issues associated with them. Specific technologies aimed at changing the mixing characteristics of the jet plume after it leaves the nozzle will be discussed in detail, including chevrons and fluidic injection. Other noise reduction technologies, such as the inverted velocity profile, and fluid shield can also change the mixing characteristics of the jet plume. This includes the added benefit of noise reflection or shielding. Measured data will be presented to show the effect these technologies have on high-speed jets.
Evaluation of Q-Values of a Rotor-Bearing System Using a Modal Open Loop Transfer Function
Authors: Fujiwara Hiroyuki , Oyama Hirot , Anegawa Norihisa , Matsushita Osami
Published: August 2009

One of the key features of vibration characteristics (e.g., critical speed, damping ratio, Q-value, instability margin, etc.) of a rotor-bearing system can be described by their eigenvalues. The eigenvalues are conventionally evaluated by feed forward excitations, e.g., hammering test, harmonic sweep excitation, and unbalance response measurement. However it has been difficult to identify their damping ratios under various conditions. In this paper, we introduce a new method using a modal open loop transfer function. This method provides a more accurate Q-value and can help to identify the damping ratio by considering only the mode which we want to measure. We applied this method to a few models and some test rigs: 1) 3DOF model (simulation), 2) A flexible rotor equipped with active magnetic bearings (experiment). As a result, we concluded: 1) As the number of sensors is increased, the accuracy of the measurement also increases. 2) Our new method is more effective in terms of the accuracy of modal damping ratio than the conventional method is.
Experimental Study on Impeller Blade Vibration During Resonance: Part 2—Blade Damping
Authors: Kammerer Albert , Abhari Reza S.
Published: August 2009

Forming the second part of a two-part paper, the estimation of damping is presented here. Part 1 discusses the experimental approach and the results on blade resonant response measurements. In the study of forced response damping is a crucial parameter which is measured to quantify the ability of a vibrating system to dissipate vibratory energy in response to a given excitation source. The blading of turbomachinery components is particularly prone to forced response excitation which is one of the major causes of high cycle fatigue failure during operation. In turbocharging applications forced response can not be avoided due to a number of factors, i.e. change in speed, inlet bends or obstructions in the flow field. This study aims to quantify the damping parameter for the lightly damped blades of a centrifugal compressor. The impeller geometry is typical of turbocharging applications. As a first step the non-rotating impeller was excited using piezos and the transfer function was derived for a number of pressure settings. Both circle-fit and a curve-fit procedure were used to derive material damping. In the second step, measurements were taken in the test facility where forced response conditions were generated using distortion screens upstream of the impeller. Main blade strain response was measured by sweeping through a number of resonant points. A curve-fit procedure was applied to estimate the critical damping ratio. The contribution of material and aerodynamic damping was derived from a linear curve fit applied to damping data as a function of inlet pressure. Overall, it will be shown that aerodynamic damping dominates the dissipation process for applications with inlet pressure of 1 bar. Damping was found to depend on the throttle setting of the compressor and where applicable CFD results were used to point towards possible causes of this effect.
Design of a Centrifugal Compressor With Low Specific Speed for Automotive Fuel Cell
Authors: Zheng Xinqian , Zhang Yangjun , He Hong , Qiu Zhiling
Published: August 2009

Centrifugal compressors driven by electric motor are the promising type for fuel cell pressurization system. A low specific speed centrifugal compressor powered by an ordinary high-speed (about 25,000rpm) electric motor has been designed at Tsinghua University for automotive fuel cell engines. The experimental results indicate that the designed low specific speed centrifugal compressor has comparatively high efficiency and wide operating range. In the condition of designed speed (24,000rpm), the highest efficiency and pressure ratio of the centrifugal compressor is up to 70% and 1.6, respectively. The designed low specific speed centrifugal compressor can meet the requirement of air systems of automotive fuel cell engines preliminarily. Moreover, the low specific speed centrifugal compressor avoids difficulties of usage of ultra-high-speed electric motors (about 60,000rpm) in high specific speed compressor. Based on the preliminary results of this centrifugal compressor, a new low specific speed centrifugal compressor with higher performances is being developed.
Improving Efficiency of a High Work Turbine Using Non-Axisymmetric Endwalls: Part I—Endwall Design and Performance
Authors: Germain T. , Nagel M. , Raab I. , Schuepbach P. , Abhari R. S. , Rose M.
Published: August 2009

This paper is the first part of a two part paper reporting the improvement of efficiency of a one-and-half stage high work axial flow turbine by non-axisymmetric endwall contouring. In this first paper the design of the endwall contours is described and the CFD flow predictions are compared to five-hole-probe measurements. The endwalls have been designed using automatic numerical optimization by means of an Sequential Quadratic Programming (SQP) algorithm, the flow being computed with the 3D RANS solver TRACE. The aim of the design was to reduce the secondary kinetic energy and secondary losses. The experimental results confirm the improvement of turbine efficiency, showing a stage efficiency benefit of 1%±0.4%, revealing that the improvement is underestimated by CFD. The secondary flow and loss have been significantly reduced in the vane, but improvement of the midspan flow is also observed. Mainly this loss reduction in the first row and the more homogeneous flow is responsible for the overall improvement. Numerical investigations indicate that the transition modeling on the airfoil strongly influences the secondary loss predictions. The results confirm that non-axisymmetric endwall profiling is an effective method to improve turbine efficiency, but that further modeling work is needed to achieve a good predictability.