COMPONENT TESTING AND PROTOTYPE COMMISSIONING OF MAN S NEW GAS TURBINE IN THE 6 MW-CLASS

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1 Proceedings of ASME Turbo Expo 2012 GT2012 June 11-15, 2012, Copenhagen, Denmark GT COMPONENT TESTING AND PROTOTYPE COMMISSIONING OF MAN S NEW GAS TURBINE IN THE 6 MW-CLASS Alexander Wiedermann Member ASME Ulrich Orth Member ASME Emil Aschenbruck Frank Reiss Dietmar Krüger Sven-Hendrik Wiers Member ASME ABSTRACT has developed a new gas turbine in the 6 MW-class for both mechanical drive and power generation applications. The lay-out of the Gas Turbine has been driven by opportunities in current and future markets and the positioning of the competition, and this has determined the characteristics and technical parameters which have been optimized in the 6 MW design. The design makes use of extremely high precision engineering so that the assembly of sub components to modules is a smooth flowing process and can guarantee both the high standards in quality and performance which MAN Diesel & Turbo is aiming for. Individual components have been tested and thoroughly validated. These tests include in particular the compressor of the gas turbine and the combustion chamber. The commissioning of the gas turbine prototype engine had been prepared with a numerous number of measuring probes and carried out at the Oberhausen plant gas turbine test field. Results of component and the gas turbine prototype tests will be presented and discussed. INTRODUCTION SE has established a strong reputation in the field of gas turbines. In 1939 the Swiss division established and entered the first industrial gas turbine to market. In the early 70 s the industrial gas turbines business started to boom, during this period until the 80 s many new products were introduced onto the market. At the beginning of 90 s the gas turbine business at MAN gained momentum due to the incorporation of the THM product line from Snecma and a close cooperation with Pratt & Whitney. Since then, MAN Diesel & Turbo has up-rated and refined these existing product lines but did not introduce any new models to the market. A new phase has started with the development of a completely new gas turbine model which draws upon MAN Diesel & Turbo s longstanding industrial gas turbine experience, in particular from design and operation of aero derivative and THM gas turbine range (1-3). There are many requirements for a newly developed gas turbine. The combustion system requires special attention, since the emissions are strongly focused to satisfy stringent environmental requirements due to common international regulations. Industrial customers demand high robustness and longevity of their investment, since gas turbines usually have a service life of over 30 years. A special emphasis was put on the service friendliness of the machine. The modular structure permits convenient on site servicing and replacement of components. The new model will initially have an output power range between 6 and 7 MW and constitutes a technology platform for further developments to create a new family of gas turbine models covering a wide range of applications. SELECTION OF MAIN DESIGN PARAMETERS The thermodynamic cycle was selected considering all requirements arising from the primary fields of application, and also considering strategic plans of future developments of the company s products. The main features are listed below: 1 Copyright 2012 by ASME

2 Coupling power under ISO conditions beginning with 6.2 MW, developing to 8 MW Potential of future increase in performance High efficiency, good applicability for combined heat and power cycle High efficiency and no operational restrictions also at part load Unrestricted capability to handle sudden load increases and changes Two-shaft design with free power turbine for mechanical and generator drive Reliability and long life time of the parts Flexibility for adaptation to different fuel types and operating modes Ability to meet low emission limit values in the primary markets through combustion with lean premix technology without further exhaust gas treatment The auxiliary gearbox is an integral part of the gas generator and is directly connected with the intake elbow. The mechanical main lube oil pump and the starter motor are located parallel to the machine axis on the rear side of the gearbox. The gearbox is thermo elastically fixed to the base frame and, at the same time, serves as the front-end support for the machine. The size of the inlet casing is strongly restricted because of rotor dynamic constraints and has been optimized based on CFD-calculations to avoid inlet flow non-uniformities in front of the compressor. The front axial and radial bearings, including the instrumentation, are located in the centre of the inlet housing, leading to excellent accessibility of these bearings. The following basic configuration parameters were derived from these requirements: Compact axial flow compressor with multiple adjustable guide vanes Air mass flow rate 27 kg/s under ISO-conditions Compressor pressure ratio around 15 External combustion chambers Two-stage high pressure turbine with air-cooled blades Two-stage power turbine with a wide speed range Optimized secondary air system for cooling and sealing at least possible losses Reliability, long service intervals and long life time of the hot gas components are important requirements. In combination with the turbine inlet temperature target, the chosen pressure ratio is high enough in order to fully exploit the efficiency potential. On the other hand, the pressure ratio is low enough to avoid the need for an additional booster compressor in most cases. Furthermore, use of low-calorific fuel gases, e.g. bio-mass, or fuel gases with high hydrogen content, can be more easily handled than at higher pressure ratios. MACHINE DESCRIPTION With more than 100 years of experience in design and manufacturing of turbo machines as well as some decades of experience with current gas turbine models, MAN Diesel & Turbo has gained expertise to accomplish a successful, market oriented and forward-looking design of a new gas turbine, Fig. 1. The layout of the two-shaft gas turbine has a clear modular structure of main component groups, as shown in Fig. 2. Fig. 1: MAN s new gas turbine Intake Combustion chamber Rear structure Exhaust Auxiliary gearbox Axial compressor HP turbine Power turbine Fig. 2: Main components The casing of the 11-stage axial compressor has a horizontal split line for ease of assembly and service, see Fig. 3. The inlet guide vanes and the following three vane rows are adjustable to achieve good aerodynamic matching at all rotor speeds including part load. 2 Copyright 2012 by ASME

3 The compressor rotor is made of discs which are braced with a central tie rod, Fig. 3. The tie rod is formed such that the high pressure turbine discs can be additionally braced with a second nut. The centering function and the torque transfer between disks are ensured by the use of Hirth-serrations. The compressor disks of the middle and rear part are of a patented multistage-design (multiple compressor stages are integrated in one disk). The combustion chambers are inserted in a single-piece, stable-shaped intermediate casing, Fig. 4. For simplification of inspection and service the combustion chambers, including the transition pieces, are completely removable on the fully mounted machine. The guide vanes and rotor blades of the two-stage high pressure turbine are air-cooled. To reduce tip clearance losses of the free-standing first stage rotor, its blade tip has a rubbing edge. The second stage rotor blades are shrouded and interlocked. Rotor blades are anchored to the discs by fir tree roots. The rear structure supports the rear gas generator bearing and is located downstream of the high pressure turbine, see Fig. 5. This design ensures minimum air leakage into the bearing housing and at the same time a good accessibility to the bearing area. Of the six struts, three support the bearing casing and the other three carry oil ducts for bearing lubrication. Additionally, the rear-end support of the machine, as well as the axial fixation, take place through the rear structure. Fig. 3: Axial compressor rotor Fig. 5: Rear structure Fig. 4: Intermediate casing with combustion chambers Six external combustion chambers (cans) are equally spaced and inclined relative to the centre line. Impingement cooling is applied to the flame tubes and corresponding transition ducts. The burner is a complete modular welding construction. It contains a swirler and two independent fuel injection locations which are utilized in diffusion mode (pilot gas), in premix mode and in a combined premix and diffusion mode, according to the operational requirements. The two-stage low pressure turbine is an overhung design optimized for high efficiency within a very wide speed range, Fig. 6. All blades and vanes of the power turbine are not internally cooled and the rotor blades of both stages are equipped with interlocking tip shrouds. The blades are anchored to the discs by fir tree roots. As for the high pressure turbine, the low pressure turbine discs have Hirth-serrations and are braced with a central tie rod. The turbine stator vane carrier is a one-piece, stable-shaped design. An efficient exhaust gas diffuser is also part of the bearing casing. A view of the low pressure turbine is shown in Fig. 6. The exhaust casing which has to turn then flow away from the axial direction was aerodynamically designed and optimized using CFD and is a welded construction. 3 Copyright 2012 by ASME

4 Fig. 6: Low pressure turbine and outlet diffuser As a matter of principle and in order to ensure a functional and robust construction, all individual parts and assemblies were designed and configured using methods and rules proven in MAN s existing gas turbine range. At the same time, the use of the most advanced manufacturing technologies was taken into account during the design process for all parts. COMPONENT TESTING Compressor: Validation of the aerodynamics and the mechanical integrity was carried out in three phases at the test bed of AneCom Aero Testing (ACAT) using a test compressor (4). The first phase of the test covered the 5 front stages, the second one the entire 11-stage compressor, and in the third phase the effect of inlet distortions on compressor performance and stability was assessed. The flow path of the test compressor was equipped with numerous pressure transducers and thermocouples, Fig. 7. Thermodynamic design parameters and power consumption were both derived from aerodynamic data using rakes at the inlet and exit planes of the compressor as well as from torque measurement on the shaft. In addition several rows of surface static pressure probes were applied on the casing along the entire compressor flow path and the diffuser. Total pressure and temperature sensors were installed at the leading edges of selected stator vanes of all stages in span wise direction. These were used to determine radial flow profiles. For selected stages, pressure transducers were used at pressure and suction sides of stator vanes to estimate local blade loads and incidence angles. The mechanical properties and blade vibrations were measured with an extensive system of strain gauges applied to both rotor blades and stator vanes. The instrumentation plan allowed accurate measurements of higher mode shapes. In addition to the telemetry system vibrations of the front transonic compressor blades were monitored with a tip-timing system which recorded relative motions of each individual compressor blade. Radial clearances had been measured in selected stages in the front, middle and rear sections of the compressor, and a novel control system based on high frequency acoustic sensors was applied to monitor the overall behavior of the rig compressor. Results of the performance map of the 11-stage compressor for use in a two-shaft gas turbine are shown in Figs. 8 and 9. Some operating points were determined at reduced inlet pressure and have been recalculated by a Reynolds number correction. The measured results are compared with predictions of a newly developed through-flow method (5) and a customized version of the 3-D multi-stage Navier-Stokes solver TBLOCK (6) - originally developed by Denton and first applied to turbine flow analysis (7) - which solves the flow fields in all stages of the compressor simultaneously. All these computations were carried out prior to the compressor test. Overall, good agreement of measured and predicted data could be observed. Efficiencies are over predicted for the lower speeds where a certain amount of flow is blown off at the bleed port in front of stage 5 to enhance compressor stability. The comparison of calculated global mean parameters with measured data demonstrates high accuracy levels of the codes applied. Furthermore detailed flow parameters such as the pressure distributions along the blade surfaces are in good agreement with measurements as shown in Fig. 10. Fig. 7: Test compressor 4 Copyright 2012 by ASME

5 referred pressure ratio [-] 1,6 1,4 1,2 1 0,8 0,6 0,4 0,2 N=75% Through flow ACFLOW 3D-CFD Multi Stage (TBLOCK) N=80% N=85% N=90% N=94.4% N=100% N=105% 0 0,4 0,5 0,6 0,7 0,8 0,9 1 1,1 1,2 referred mass flux [-] Fig. 8: Measured and calculated compressor performance map ref. Adiabatic efficiency [-] 1,05 1 0,95 0,9 0,85 0,8 N=75% Through flow ACFLOW 3D-CFD Multi Stage (TBLOCK) N=85% N=80% N=90% N=94.4% N=100% N=105% 0,4 0,5 0,6 0,7 0,8 0,9 1 1,1 1,2 referred mass flux [-] Fig. 9: Measured and calculated compressor efficiency At some speeds the surge limit of the compressor was determined experimentally at reduced inlet pressures. A number of speed lines, including full load were measured at fully un-throttled inlet pressure conditions to obtain guarantee data. During the test, very good thermodynamic and mechanical properties were obtained within the required operating range, Wiedermann et.al. (4). Design target values have been achieved and even exceeded in terms of the stability limit. During the test compressor characteristics were also taken for single-shaft gas turbine operation for a future application of the compressor. More details are given by Wiedermann et.al. (4). For the integration of the compressor into the gas turbine an inlet casing and an exit diffuser were both designed and optimized using CFD. During the last stage of the compressor component test inlet distortion was simulated by adding an inlay device which simulates the effect of radial-axial turning of the flow (4). In addition extensive numerical studies were carried out to optimize the intake flow geometry (8) ,27 0,22 0, ref. axial length [-] Mid span static pressure probes ref. axial len gth [-] IGV Stator 1 0,12-0,05 0,15 0,35 0, 55 0,75 0,95 ref. axial length [-] Experim ents ref. axial length [- ] Stator 2 Stator ref. axial length [- ] Messdaten ref. axial length [- ] Stator 4 Stator 7 Fig. 10: Comparison of measured and calculated pressure distributions in the mid span section of the VGV s, stator 4 and stator 7 (3-D multi-stage CFD TBLOCK) For designing the compressor diffuser its interaction with the combustor cans was studied based on a CFD model for a 60 degree segment. The flow distribution is shown in Fig. 11 and gives insight into flow pattern inside and outside the combustor liner shown by the velocity vector distribution. The colors indicate Mach number levels. 5 Copyright 2012 by ASME

6 For the second high-pressure test additional instrumentation was used. Key points of the test program were: Ignition Enhanced functional test of the combustion chamber (various operating ranges) Extended mapping of combustion performance (pressure losses, emissions, temperatures, extinction limits) at different operating regimes (ignition to full load) for o diffusion mode o combined diffusion and premix-mode o premix-mode Thermal-paint test to determine surface temperature distributions Fig. 11: Fully 3D Navier-Stokes analysis (Ansys-CFX) of the entire flow path stretching from compressor exit to turbine Combustion chamber: In the first phase of the development process the combustion chamber was atmospherically tested at the Helmut Schmidt University in Hamburg. This test was very successful and confirmed the function of the ignition system as well as flawless operation in diffusion, premixed and combined diffusion and premixed modes. The measured emissions corresponded to the design values. These ambient pressure tests were complemented by a series of full pressure tests at the German Aerospace Centre (DLR), Institute of Propulsion Technology in Cologne. Fig. 12 illustrates the experimental setup. Fig. 13 shows a plot of steady-state test points on a typical day of testing where main load parameters were varied. Individual components of the combustion chamber were instrumented with thermocouples and pressure probes. In addition, the acoustic combustion behavior was measured using dynamic pressure probes. Excellent ignition characteristics as previously seen during the atmospheric tests were confirmed, Fig. 14. The switch-over between diffusion and lean premix operation modes at a low pilot gas ratio was tested extensively in a simulated gas turbine load range from 50% to 70%. The operating concept of the combustion chamber requires that the combustion chamber always burns a small amount of fuel as a pilot flame. This gives the gas turbine the capability to safely perform very rapid load changes. Measured NOx emissions were well below 50 mg/nm3, (24 vppm) in the range of % load. Measured CO values in this load range were below 12 mg/nm3 (10 vppm). During the tests, all components of the combustion chamber showed high reliability and availability and the system fully satisfied all requirements of the specification. The combustion system was therefore released for engine prototype testing Load Pilot Fuel Ratio Equivalence Ratio 2,0 Load [%]; Pilot Fuel Ratio [%] ,5 1,0 0,5 Equivalence Ratio [-] Fig. 12: High-pressure test rig installed at DLR test facility The first high-pressure test of the combustion chamber was a basic test under real engine pressure and temperature conditions. The favorable results of atmospheric tests could be confirmed Meassuring Point Fig. 13: Variation of main load parameters, steady-state test points 0,0 6 Copyright 2012 by ASME

7 Air - Fuel - Ratio Ignition shroud Streamline pattern at rotor 2 section ,0 0,2 0,4 0,6 0,8 1,0 1,2 1,4 m air / m air,ref. Fig. 14: Ignition window confirmed during high pressure test Turbine: To confirm the aerodynamic design features of the turbine fully 3-D multi-stage CFD-analysis models of the entire turbine flow path were created covering both the high pressure and power turbines simultaneously. Here, the focus was on the effect of leakage flows because it is well known that the mixing process of the secondary and main flows have to be understood to minimize turbine losses (7). Therefore, geometry of the cavities and the shroud section including seal regions were modeled and the injection of the cooling flow was simulated by an appropriate source distribution derived from the secondary flow model estimation. The calculated relative total temperature distribution is shown along the meridional mean surface in Fig. 15, showing details of the side wall geometries, shrouds and cavities modeled for the 3-D flow analysis. These results give insight in the complex flow features in the sidewall regions and their interaction with the main flow. Unsteady CFD was used to optimize the shape and arrangement of the struts in the rear structure downstream of the high pressure turbine. As mentioned earlier these struts have to be stiff enough to support the bearing. Moreover, the area size has to be large enough to carry lube oil to and from the bearings. Fig. 16 shows snapshots of the entropy fields of a preliminary design pass and the optimized arrangement. Using CFD-based analysis the efficiency of the first stage of the power turbine could be raised about 0.8 %. No dedicated test was scheduled for the isolated turbine section ahead of the full engine test. To verify the cooling system of blades and vanes of the high pressure turbine, through-flow and heat transfer measurements were carried out. The blades have high performance convective cooling and feature internal serpentine cooling channels. The first stator vane is designed with a deflector and impingement cooling. Geometry details of the cooling channels were optimized with CFD, see Fig. 17. Agreement of the measured flow rates with calculations was excellent. Fig. 15: Temperature distribution along the meridional surface of the high-pressure turbine flow path computed with 3-D multi-stage CFD (TBLOCK) Start configuration Adiabatic efficiency 0.5% efficiency Number of steps optimized (final) Optimized (final) Preliminary pass Fig. 16: Design optimization of rear structure strut geometry and arrangement using unsteady CFD (3D CFD TBLOCK) 7 Copyright 2012 by ASME

8 component loadings were determined using both thermocouples and thermal paints. Fig. 17: Calculated flow pattern in the cooling channel of the first blade and comparison of measured and calculated characteristics of through-flow tests (3-D CFD Ansys-CFX) PROTOTYPE COMMISSIONING AND TEST Assembly: After successful completion of design, build and functional testing of major components in individual rig tests, the go-ahead was given for manufacturing and assembly of the first prototype engine which was extensively instrumented for the test run. First testing took place at MAN Diesel & Turbo s Oberhausen test facility during winter 2010/11. Test stand and instrumentation: The prototype engine was equipped with about 500 measuring probes, 7 of these were used for the emissions and 16 for the inflow conditions and the remaining are distributed among the gas generator and the power turbine. Besides measurements within the core engine also the most important parameters of the auxiliary units as e.g. the lube oil system had been taken. For the prototype test a new test stand was erected. Because of security reasons the protecting walls were built of concrete, see Fig. 18, and during the test runs the engine was observed by video cameras. All experimental data can be displayed online. Limits can be allocated, and warning messages were submitted if they were exceeded. The fuel gas composition could be monitored online and was confirmed before and after each test run. Test program: Several goals were defined for the test. Besides the confirmation of the mechanical integrity and safe operation, rotor dynamics, power output and efficiency of the components and the entire gas turbine were tested. One of the main purposes was the performance map of the power turbine. For the combustion chamber the confirmation of the operational concept with regard to ignition and controlled operation in diffusion and premix mode was a major point. In addition, its performance, component temperature, emissions and acoustic had to be proved. In the secondary flow system several pressures and temperatures were taken. Thermal Fig. 18: View into the test cell with instrumented gas turbine prototype The test program consists of the following parts: 1. cold commissioning 2. warm commissioning up to 85% load 3. performance map up to and beyond 100% load 4. thermal paint test run 5. parameter tests with the combustion chamber During the first part instrumentation and data acquisition system as well as the starter motor were checked. The focus of the second part was the determination of the start sequence and confirmation of a safe and fast acceleration up to 85% load. Here, the first successful ignition of the engine was one of the major milestones. Each of the 6 combustors can be ignited individually without being interconnected. The acquisition of the performance map was executed systematically over a longer time range. Steady state operating points were confirmed by holding the operating conditions constant for a certain time. For the thermal paint test runs relevant operating points were considered based on the thermodynamic performance runs. The design temperature at full load had to be achieved and kept constant for three minutes. During this time period changes in the so-called T4 temperature at the exit of the gas generator must not exceed the interval of ±5K. 8 Copyright 2012 by ASME

9 Finally tests were carried out for finding suited configurations of the combustion parameters with focus on the variation of the pilot to main fuel gas ratios at various operating conditions. The goal of these tests was to achieve low emissions for operating points >70% load and to determine the optimum run parameters. The entire test program was lasting more than three months. Borescope checks of the gas turbine had been performed several times between the test runs to guarantee the good and healthy shape of the engine. After completion of the test runs the engine has been disassembled completely and all components investigated carefully. Test execution: After gradual variation of rotational speeds, fuel gas valve positions and ignition time windows the engine could be safely and reproducibly ignited. The starter motor accelerated the gas generator up to 8500 rpm. During the prototype test the compressor was monitored at several measuring stations, and a very good correlation between the data and the results of the compressor rig test could be confirmed. Fig. 19 indicates good agreement of measured and predicted static pressures and temperatures at the entrance planes of the three bleed ports at the casing for all operating conditions taken during the commissioning phase. These data are important boundary conditions for the secondary flow design system. Because of the favorable and quiet behavior of the prototype engine the approach to various operating points up to full load could be realized. In no more than one week after the first firing full load was achieved. The test program could be completed according to the plan. As a consequence a solid basis is set for forthcoming tests. The following thermodynamic data could be confirmed: Pressure ratio of the compressor: up to 15.2 Compressor mass flux up to 29.2 kg/s Operating condition (sea level, ambient temperature T0 =+4 C) Power output: 7.6 MW Thermal efficiency: 34.5% ISO - Operating condition (sea level, T0 = +15 C) Power output: 6.9 MW Thermal efficiency 34.0% Pressure Temperature 2 bar 0% 20% 40% 60% 80% 100% 120% Ref. Aero Speed 100 0% 20% 40% 60% 80% 100% 120% Ref. Aero Speed Fig. 19: Comparison of measured and predicted aerodynamic data at the prototype gas turbine compressor bleed ports in front of stage 5 (blue), stage 9 (purple) and behind diffuser (orange) Performance: A comparison of the measured and predicted low pressure turbine performance chart is shown in Fig. 20. The computations were done with 3-D multi stage TBLOCK with all stages of the turbine, the intermediate diffuser between gas generator and low pressure turbine and the exhaust gas diffuser. In addition to the main flow path also side wall flow cavities, seals and shroud regions had been included in the flow model to achieve a realistic representation of mixing processes of the secondary flow with the main flow, see notes in the previous chapter and Fig. 15. The boundary conditions of the cooling flow discharge were obtained from the results of the hydraulic flow model for the secondary flow system. The computed results show four main operating characteristics at ISO, corresponding to hot day (+40 C), cold day (-10 C) and part load operation (ISO, 70% load). The design point and the load condition which represents the introductory rating are indicated in Fig. 20. Measured and computed results show good agreement of both magnitude and trends of the power output vs. the rotational speed of the power turbine. 9 Copyright 2012 by ASME

10 ref. LPT Power output [-] 1,4 1,3 1,2 1,1 1 0,9 0,8 0,7 OP2: Design OP1: Hot day OP4: Cold Day OP3: Part load Exp (1) Exp (2) LPT Operating points 0,6 0,4 0,5 0,6 0,7 0,8 0,9 1 1,1 1,2 Ref. LPT rotational speed [-] full load introductory rating Fig. 20: Predicted and measured power output of the gas turbine. Simultaneous full 3-D multi stage analysis of high pressure turbine, power turbine and exhaust diffuser with TBLOCK (Exp (1) and (2) refer to measured data of two different days) In the turbine flow path span wise distributions of temperature and pressure profiles were measured with leading edge probes and rakes. Comparisons of predicted and measured total pressure (P4) and temperature slopes (T4) at the exit of the gas generator are shown, where 6 rakes with combined pressure and temperature probes have been installed in lateral direction. A comparison of measured and predicted results is shown in Fig. 21. The agreement is reasonably good, and the predicted slope is almost confirmed by the experimental results. Fraction of span 1 0,9 0,8 0,7 0,6 0,5 0,4 0,3 0,2 0,1 0 Exp. Pos. A Exp. Pos. B Exp. Pos. C Exp. Pos. D Exp. Pos. E Exp. Pos. F TBLOCK - CFD Prototype test HPT exit -- Near Design Pa Total pressure Fig. 21: Predicted and computed results for the total pressure and temperature distributions at the gas generator exit (T4-) plane. Simultaneous full 3-D multi stage analysis of high pressure turbine, power turbine and exhaust diffuser with TBLOCK Also at the power turbine exit plane a good agreement of prediction and experiments can be found, see Fig. 22. At this plane only two combined pressure and temperature rakes were mounted at two selected circumferential positions. Fraction of span 1 0,9 0,8 0,7 0,6 0,5 0,4 0,3 0,2 0,1 0 Exp. Pos. A Exp. Pos. B Exp. Pos. C Exp. Pos. D Exp. Pos. E Exp. Pos. F TBLOCK - CFD Prototype test HPT exit -- Near Design 25 K Total temperature Fig. 22: Predicted and computed results for the total pressure and temperature distributions at the power turbine exit (T5-) plane. Simultaneous full 3-D multi stage analysis of high pressure turbine, power turbine and exhaust diffuser with TBLOCK Emissions: Soon after the first check-outs during the combustion chamber test phase it was already possible to achieve low values for NO X and CO values which are below the limits of common international pollution regulations. Up to now optimization is still pending, in particular for the operational range below 70% load. These are planned for the next test period with additional instrumentation. For this test period also telemetry measurements are considered to obtain additional information of the rotating parts of the gas turbine with emphasis on the turbine section. SUMMARY The new gas turbine design draws upon the long experience of SE in the field of gas turbines and utilizes the latest technology in design, materials, engineering and manufacturing. The targets of efficiency, low emissions and trouble-free operation have been achieved in the test, and the test results hold the promise for high reliability and long service intervals. The new engine can be utilized for mechanical drive as well as power generation applications. The proper function of main components had been verified in individual tests. The fully assembled machine was tested with a high level of instrumentation and according to a rigorous test program. As a result the performance target could be achieved at low emissions. In conclusion we can state that the new gas turbine is a competitive product with further up-rating and growth potential, which will meet or exceed present and future customer expectations. 10 Copyright 2012 by ASME

11 ACKNOWLEDGMENTS The work presented here was supported within the framework of the EU's Objective 2 phase V under grant number T-196 (codeword: BEGIN) and under the NRW objective 2 Programme "Regional competitiveness and employment (ERDF) under grant number DE A (codeword: KONTEST) by funds of the EU and the state of North Rhine-Westphalia. The authors are grateful for the support and permission for publication. The responsibility for the content lies solely with the authors. REFERENCES 1. Aschenbruck, E., Beukenberg, M., Jeske, H.-O., Orth, U.: Increasing Output and Efficiency for Industrial Gas Turbines. MAN forschen planen bauen Aschenbruck, E., Beukenberg, M., Blaswich, M., Bokelmann, H.: The Upgraded Power Turbine for the Industrial Gas Turbine THM 1304 Development and First Operational Experience. ASME-Paper GT , Vienna, Aschenbruck, E.; Frank, D.; Korte, T.; Müller, R.; Orth, U.: Development and Testing of the Upgraded High Pressure Turbine Section for an Industrial Gas Turbine. ASME-Paper GT , Berlin, Wiedermann, A.; Frank, D.; Orth, U.; Beukenberg, M.: Computational and Experimental Analysis of an Industrial Gas Turbine Compressor. ASME-Paper GT , Vancouver, Petrovic, M.; Wiedermann, A.; Banjac, M.: Development and Validation of a Universal Through Flow Method for Axial Compressors. Proc. of the Inst. of Mechanical Engineers (IMechE), Part A: Journal of Power and Energy, pp , Wiedermann, A.: Industrial CFD-based Design System Development for Turbomachines. Newsletter of the ASME-Branch in Japan, pp 4 6, Tokyo, Rosic, B.; Denton, J.D.; Curtis, E.M.: The Influence of Shroud and Cavity Geometry on Turbine Performance, Part 1 and 2 ASME-Paper GT and , Montreal, Hilgert, M.; Boehle, M.; Wiedermann, A.: Stationary Gas Turbine Design and its Impact on a Transonic Compressor Stage. Proceedings of the 9th European Conference on Turbomachinery (ETC), Paper 076, Istanbul Copyright 2012 by ASME