Axial compressors of aero engines should achieve both high efficiency and low weight. Recently, blade-integrated disks (blisks) have been adopted to meet these requirements. Because of this trend, problems of high-cycle fatigue (HCF) associated with forced response vibrations are occurring more frequently these days. Rotor blade fatigue can be caused by not only wakes and potential effects of adjacent stator vanes but also the asymmetric geometries inside compressors. Even the measurement probes can be the cause of severe blade failure. Although, there have been many previous studies on developing simulation approaches for the accurate prediction of blade vibration induced by adjacent stator vanes, studies of vibration induced by the measurement probes have been scarce. Thus, in this study, the vibration induced by measurement probes is studied using fluid structure interaction (FSI) simulation approach and an experimental test rig. The ways to predict and reduce blade vibration are established and validated.

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The effects of the atomizer type on flame stabilization and emissions from a liquid ammonia gas-turbine combustor was studied to explore the impact of enhanced fuel-air mixing. A twinfluid atomizer is expected to enhance the mixing of the atomized fuel and air, which may in turn decrease the amount of unburnt fuel at the combustor outlet due to improved combustion efficiency. Liquid ammonia was used as the fuel, and the combustor pressure was 0.25 MPa abs. The low combustor pressure tends to encourage partial vaporization of the liquid ammonia in the mixing chamber of the twin-fluid atomizer. When the pressure of the mixing chamber was below the design value, it was found that the lower pressure, which may encourage flash boiling and hence rapid atomization, enabled a decrease in the emission of unburnt NH3 as compared with that of the pressure-swirl atomizer. However, when the mixing chamber pressure was as high as the design value, higher levels of unburned ammonia emission was observed. This suggests that flash boiling which promotes atomization may enhance the combustion efficiency.

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The primary focus of this study is to optimize the shape of the exhaust volutes of a particular type of mixed working fluid centrifugal compressor. The objective is to improve the aerodynamic performance of the volutes and enhance the overall performance of the compressor. To achieve this, both the adjoint method and parametric optimization approach are employed, and their results are compared to determine the optimal approach. The results indicate that after 65 iterations of parametric optimization, the total pressure recovery coefficient of the exhaust volute increased by 0.577%. By using 37 steps of adjoint optimization, the optimized total pressure recovery coefficient can be increased by 1.614%. A comparison between the parametric optimization and adjoint optimization reveals that the latter resulted in more noticeable changes in the shape of the discharge volute, ultimately leading to significant improvements in its aerodynamic performance. Faster convergence is achieved by the adjoint optimization approach.

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Recently, in response to the increasing environmental awareness and demands of operational flexibility, gas turbines for power generation must comply with exhaust gas regulations, not only at rated operation but also at partial-load conditions. To improve emission performance of constant speed single-shaft engines at partial-load conditions, compressors should be operated with low mass flow rate to optimize air-fuel ratio of combustors. Flow rate of constant speed compressors can be reduced by controlling angles of variable stator vane (VSV). Under these conditions, flow separation can be induced and it causes flow induced vibration and noise. They limit stable operation range of gas turbines. Thus, in this study, the way to predict the unstable behavior of compressors is investigated using unsteady flow simulation. The accuracy of the simulation is validated by utilizing a gas turbine test-rig. The way to enhance the stable operation range by optimizing the angles of VSV is also validated.

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The high-performance aero engines handle the maximum possible airflow rates for a given frontal area using high-speed compressors. The Mach number in these compressors goes subsonic at the hub to supersonic at the tip, resulting in the complex shock system deteriorating the blade’s aerodynamic performance and structural integrity. The present study focuses on the unsteady nature of the rotor tip shock and evolving flow physics due to its interaction with the core flow at the near-stall condition. The study is performed using steady and unsteady numerical simulations and validated against the experimental data. The analysis showed that high-intensity shock is confined to the tip region, creating shockinduced boundary layer separation. The tip shock keeps oscillating and eventually disappears with the stall onset, leading to fluctuations in the blade loading. Shock-induced leading edge separation near the hub causes radial migration of the flow, which in turn interacts with the tip leakage flow, casing boundary layer, and suction surface blade boundary layer, forming a large re-circulation zone in the blade tip vicinity. The rotor disturbances get convected down to the stator, inducing tip-corner separation. The near-stall unsteady analysis showed the existence of four dominant low frequencies, out of which 0.06×BPF and 0.12×BPF are related to the rotor tip instabilities, and the rest two; 0.44×BPF and 0.84×BPF are related to the rotor blade trailing edge vortex shedding and rotor-stator aerodynamic interaction.

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The industry and the academy are continuously developing new technologies and approaches regarding the gas turbine manufacturing. Logically, sectors of turbomachinery and aerospace engineering are deeply focused on applying newer and even unconventional manufacturing process, aiming on cost reduction, reduced lead times and efficiency. In addition, it is conspicuous that metal additive manufacturing (AM) technologies can provide interesting possibilities for companies seeking to innovate and perfect existing components, with respect to reach better buy-to-fly ratios. In this paper, the authors developed a proposal for additively manufacturing a fuel swirler and evaluated in detail its process of fabrication in order to compare the results with the characteristic of a conventionally manufactured swirler. Furthermore, a dedicated review of the state-of-the-art related to the AM of fuel swirlers were realized to evaluate the relevance of this topic to conclude if the use of AM to fabricate this component can favor the aerospace industry.

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A lifted flame allows partially premixed combustion of hydrogen without flashback for dry low-NOx gas turbine combustors. The use of a small injection hole enhances the shear mixing of hydrogen with air. The objective of this study is to investigate the effect of the hydrogen injection hole diameter on the burning condition of lifted flames. Five hydrogen injectors with hole diameters varying from 0.2 mm to 1.0 mm were used. The hydrogen injection pressure drop, flame stability, flame appearance, NOx emissions, and combustion efficiency were evaluated. As the injection hole diameter decreased, mixing became more uniform, resulting in a decrease in NOx emissions, although the difference in value was small for hole diameters ranging from 0.8 mm to 0.4 mm. On the other hand, both the injection pressure drop and flame stability became issues for the 0.4 mm and 0.2 mm injectors.

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The trend for the engine design goes to smaller core engines to increase the bypass ratio and reduce the weight. With the decrease of core engine size, also the Reynolds Number decrease locally. This leads the focus within the design process of the axial compressor on the accuracy of the numerical models which are used for the simulation. Therefore, an experimental and numerical study was carried out to evaluate the state-of-the art design process for axial compressor bladings concerning the low Reynolds number effects within the flow. As study approach a linear cascade was used. Whereby the experiments were performed at the Transonic Cascade Wind Tunnel TGK of the DLR in Cologne and for the numerical simulations DLR in-house flow solver TRACE was conducted. The investigation was carried out at an inlet Mach number of 0.60 and a Reynolds number of 0.15 x 106. The comparison shows a significant discrepancy which is based on the current weakness of the turbulence and transition modeling at a RANS simulation regarding the viscosity effects at lower Reynolds numbers. Additional simulations were performed at a higher Reynolds number of 0.9 x 106 to substantiate this interpretation. Here, a good agreement with the equivalent measurement results at this Reynolds number is shown.

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