The emergency gas turbine generators offered by IHI Power Systems are designed to supply large volumes of electric power instantaneously in the event of power outages. They are installed in data centers, public facilities, waterworks systems, and other facilities, where they play key roles in ensuring the safety and security of people and society. To meet the growing need for highcapacity generators, we have undertaken component development for higher output compressors and turbine blades to boost the output of 2 MW class gas turbine engines by 25%. This paper describes the efforts to develop compressors and turbines required for highpowered emergency gas turbine.

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Experimental investigations of acoustic excitation and noise propagation in turbomachines require specialized rigs with sophisticated instrumentation. As it remains challenging to integrate such measurements into a real engine, sub-scale test rigs are required with acoustically optimized boundary conditions to isolate acoustic sources. Thus, scaling approaches have to be applied to establish similarity. In the present work similarity parameters are derived and methods to geometrically scale machines while maintaining aeroacoustic similarity are proposed. The scaling approach is investigated using a 1.5-stage turbine test rig. Numerical simulations of the original and geometrically scaled rig are performed to validate similarity using both time and frequency domain methods to simulate aerodynamic and aeroacoustic effects. Transient blade pressures and the acoustic fields induced by blade-vane interaction show a good agreement between the two test cases. Aeroacoustic similarity could thus be achieved using the scaling approach presented which allows for better transferability of sub-scale tests.

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Under the NASA Energy Efficient Engine (E3) program, two high pressure turbines (HPTs) were separately developed and tested by General Electric (GE) and Pratt & Whitney (P&W). Despite the corresponding NASA E3 design reports and several subsequent publications related to these HPTs, there is still no uniform and consistent data base, leading to the absence of some essential parameters. Therefore, 0D performance models of the NASA E3 HPTs were generated based on the available literature. The performance results agree well with the literature data and will be presented. Moreover, a well-known turbine cooling modeling approach is calibrated using the 0D performance models and further 1D turbine models, which were created in a collaborative process documented in this work and an accompanying paper. This is crucial since the predicted coolant mass flows significantly affect the gas turbine efficiency. Based on the obtained calibration results, a simplified turbine cooling model is also derived and calibrated in this paper, being applicable for performance studies in the early phase of preliminary design. In order to quantify the error due to simplification, both the original and the simplified turbine cooling model are applied in two parametric studies.

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Accurately estimating turbine cooling requirements at a preliminary design stage is crucial for modeling the overall propulsive system. The operating conditions of compressor, combustor and turbine are significantly influenced by these requirements. Empirical cooling models have thus far provided reliable initial estimations. For next-generation aero-engines this solution becomes inadequate. To address this challenge, an alternative semi-empirical approach based on an established cooling model is built into a collaborative turbine design tool chain. This cooling model is applied to the two cooled high-pressure turbines developed by P&W and GE within the NASA E3 program for both validation and to provide calibrated model parameters for future studies. Finally, sensitivity analysis provide a better understanding on how cooling requirements can be reduced through turbine preliminary design decisions, material and cooling technologies. This work focuses on 1D turbine studies with an accompanying paper on 0D engine modeling.

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A full 3D rotor/stator FE model with CMS-based superelement modal reduction is developed. The method enables a high fidelity and efficient rotordynamic analysis for a newly designed power turbine (PT) with a slim spoke frame stator. In parallel, the stator support structure dynamic stiffness is quickly assessed using the FE simulated frequency response functions (FRF) data on the requirement of API standard. The results of the dynamic stiffness of the stator shows lower than the API standard recommended. The derived support dynamic stiffness is directly applied to the unbalance response analysis of the PT rotor incorporating the bearing characteristics. The comparison of the unbalance response shows that the rotor with FRF representation of stator structure and the full 3D FE superelement model are very well cross validated. Finally, the rotordynamic analysis of full FE model shows that with the slim spoke frame, the PT still meets all the API requirements.

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The circumferential bowing of stator blade is known to possess the ability to suppress the corner separation. To enhance the improvement of the aerodynamic performance of axial compressors by the application of the bowed stator blade, the optimizations of blade shape have been studied by many researchers. In most of these studies, the blade root fillets which exist in actual compressors were neglected. However, it was reported that the stator blade root fillets influence the generation of the corner separation in the compressor which has the conventional stator blades. Therefore, the stator blade root fillets can contribute to the improvement of the aerodynamic performance due to the bowed stator. In this study, the influences of the root fillets and the positive dihedral angle of the bowed stator blade on the overall performances and the internal flow behavior of low-speed single-stage axial compressors were clarified.

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This paper deals with the component level breakdown of performance characteristics of a centrifugal compressor using detailed measurements. The compressor stage used for Heavy-Duty applications consisting of the impeller, vaneless diffuser, and volute has been experimentally tested in a gas stand. Several static pressure measurements combined with total pressure and temperature measurements on the compressor stage have been carried out. Based on the detailed measurements, the component level aerothermal performance parameters are calculated on the entire compressor map that includes tip speeds from subsonic to supersonic speeds. The effect of individual component performance characteristics and their impact on the design of the entire stage is discussed both at design and off-design conditions. A better understanding of the component level performance aids to design of a compressor stage for increased efficiency, range, and better surge margin. However, this component-level detailed breakdown of the compressor stage for HD truck application is unavailable in open literature. In addition to the detailed measurements, a 1D model of the impeller is created to correlate measurements with predictions iteratively to identify the relevant 1D modeling parameters for varied tip speeds on the compressor map.

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This paper presents the development of a gas turbine simulator based on an application of a real turbogenerator used to generate electricity on an offshore oil platform, the configuration is a turboshaft with free power turbine. The compressor, turbines and the control system were developed using specific methodologies. The development of the simulator was done using the Simulink environment in Matlab®. The development was done using blocks to represent each one of the main components in the engine. A stage stacking methodology based on the real geometry for each stage was adopted to create the compressor maps. The map was used in lookup tables blocks with help of auxiliary coordinates, also known as beta lines. To model both turbines were applied an ellipse equation also known as Stodola’s law. The engine simulator model was tested in an open loop and the results evaluated with the manual data from the engine.

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The turbine of an aeroengine is exposed to a high-temperature mainstream, making a sufficiently dimensioned cooling and sealing for all components involved indispensable. A balanced thermal safety and overall efficiency demands a better understanding of the aerodynamic and thermal behavior. This study focuses on the cavity at the bottom of the rotor platform, which is connected to the main flow through the midpassage gap between two adjacent blades mounted on the turbine disc. A numerical approach has been conducted to obtain the flow mechanism, sealing effectiveness and heat transfer characteristic for various purge flow rates and midpassage gap clearances. The results of the steady-state simulations show variations in the flow pattern and in the thermal load parameters, such as the adiabatic cooling effectiveness or the heat transfer coefficient. The distinct sensitivity highlights the necessity of a subtle adjustment of those parameters, to avoid critical hot spots endangering thermal safety.

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In order to improve the efficiency of modern turbomachinery, the simulation setups in design processes become increasingly detailed, and thus computationally more expensive. This paper aims to quantify the numerical influence of squealer tips on the performance of a 4½-stage axial compressor. For this purpose, the performance map, obtained by steady-state calculations using the k-w SST turbulence model and the TRACE flow solver, is simulated for two setups: the reference setup and a setup with squealer tips added. The inclusion of squealer tips enhances the stage-wise compressor performance by reducing secondary flow and flow separation at the casing. This improvement results from the change in blade-tip geometry. The stage total-pressure ratio is thus improved by up to 0:28%. However, a deterioration of the fourth-stage flow as a result of the squealer tips prevents these improvements from enhancing the overall compressor performance. NOMENCLATUR

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The impingement cooling of hydrocarbon fuels represented by kerosene has broad application prospects in the field of hypersonic aircraft and engine thermal protection. However, compared with air and water, the thermal properties of kerosene are more complex and need to be further studied. Using the Reynolds average method and the SST k-ω turbulence model, the superalloy is used as the heat transfer material. By simplifying the model, the effects of different jet structures on the impingement cooling effect under the same inlet conditions are numerically simulated. The wall temperature distribution is used as the evaluation index of heat transfer performance, and the overall design of the stabilizer is based on this. The research shows that the wall temperature of the stabilizer shows a tendency of increasing annularly around the stagnation point. The increase of opening ratio, hole pitch and jet distance will lead to the decrease of heat transfer effect.

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Axial compressors tend to experience significantly increased losses during off-design operation and are prone to flow separation under high aerodynamic loads. Active flow control (AFC) presents a way to counteract this by means of local injection or aspiration, in order to obtain more favourable momentum distributions. It is a strength of active injection methods that the injection rates may be adapted during compressor operation, in order to maximise the benefit at different operating points. This paper investigates the potential of varying the mass-flow rate of a suction-side stator injection, in order to improve the off-design performance.

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The part-load performance of the gas turbine engine can be improved in VCE using variable geometry components. The VANT is one of the variable geometry components of the gas turbine engine to meet the power requirement at part load conditions by controlling the mass flow rate through the engine turbine section. However, the provision of the part clearances near the hub and casing endwall allows the leakage flow to occur. This leakage flow further increases the loss at the exit of the turbine nozzle vane. The provision of the pivot to hold the vane in endwalls creates a blockage to the leakage flow and changes the flow field of the turbine nozzle passage. As the operating Reynolds number of the LPT is of the order of 105, flow experiences a transition between laminar and turbulent flow regimes. Hence, the present study is conducted with two different turbulent models, i.e., widely used SST k-ω and SST γ-Reθ, to analyze the part clearance flow field. The flow field is analyzed in the present study using streamlines plotted at different streamwise planes superimposed on non dimensional vorticity contours, and total pressure loss coefficient (Cpt) at the exit of the turbine nozzle vane.

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