As part of the design of a new particle accelerator at CERN, a research is conducted to study the challenges and opportunities of multi-stage turbocompressor machines operating with light gases and more specifically with a mixture of helium and neon. First, a 1D stage performance prediction model is implemented and coupled with a genetic algorithm in order to generate an impeller database. Then, a stacking method is developed considering design philosophies and technological limitations observed in the industry. This model is coupled with a second loop of the same genetic algorithm, which provides multi-stage architectures optimised for either compactness, i.e. number of stages, or efficiency. For both objectives, an ideal number of stages can be determined which increases significantly as the operating gas becomes lighter. The impellers diversity within the database also plays an important role on the overall machine architecture. Finally, in alignment with potential technological improvements, the motor maximum rotational speed is varied to study the achievable reduction in the required number of stages.

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In this study, we aim to elucidate the effect of a forward-swept rotor on the stall margin and flow field at the distorted inflow conditions. The rig in this research is a low-speed, single-stage axial compressor, which has two types of rotor blades: the radially stacked blade (Radial) and the forward-swept blade (Sweep). The distortion screen that circumferentially generates distorted inflow is located upstream of the rotor. The stall margin of Sweep was found to be larger than that of Radial. Sweep was considered to improve the flow fields at the distorted inflow conditions. From the results of the study, it was observed that Sweep suppressed the circumferential expansion of the high-load regions and the spike-type disturbances generated at the distorted sector. Therefore, Sweep enlarged the stall margin more than to Radial.

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This paper presents a review of aeroelasticity research concerning fan blades in modern civil aircraft engines. It summarises the research carried out at the Rolls-Royce Vibration University Technology Centre (VUTC) at Imperial College over the past 25 years. The purpose of this paper is to gather information on all the aeroelastic issues observed for civil aero-engine fan blades into one document and provide a useful synopsis for other researchers in the field. The results presented here are based on numerical methods but wherever possible data from experiments are used to verify the numerical findings. For cases where such datasets do not exist fundamental principles, engine observations and engineering judgement are used to support the numerical results. Numerical methods offer a cheaper alternative to rig tests, especially in cases of blade failure, and can also provide more information about the nature of instabilities, which can be useful in the design of future civil aircraft engines. In fact, in cases such as crosswind testing that use smaller rig-scale blades, such results can even be more representative of real engine flows.

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The use of additive manufacturing (AM), for example Selective Laser Melting (SLM), is poised to spark a revolution in the way high-temperature components for gas turbines are designed, but a number of grave uncertainties remain. These lie mainly with the materials sciences, but some questions with regard to manufacturing and operating SLM-parts as hot gas path components and the demands on the tolerances of the cooling features associated therewith remain as well. In order to quantify the impact of these uncertainties, Nozzle Guide Vanes (NGVs) with a geometry that would normally be investment-cast were produced with SLM. A back-to-back comparison of vanes from the two manufacturing processes was performed. The design of the SLM-vanes will be described and the SLM-manufacturing process of the NGVs will be touched upon, especially the use of MAR M-509, which is seldom used for SLM. In addition, characterization of the NGVs with 3D-scans of the outer geometry and the pin-fin matrix shall be discussed. The NGVs were operated for approximately 70 hours at relevant load conditions in a highly-instrumented test engine on a test bed at the Oberhausen plant of MAN. The temperatures of the AM and investment-cast vanes were measured using Thermal History Paints (THPs); a comparison between these different kinds of parts will be drawn.

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Electrification of aircraft has been realizing improvement in efficiency, reliability, and safety of the aircraft by substituting hydraulic and mechanical system with electric system. On the other hand, in future electric aircraft partially replacing the fan driving engines with motors, the thermal management of heat generation from the electric system will become crucial problem to be solved. In this study, for the future electric aircraft, thermal management in oil-cooling motors and air-cooling motor controllers was considered. Three-dimensional steady thermal network analysis (TNA) was performed for analyzing temperature field in an oil cooler for motor cooling and a heat sink for motor controller air-cooling. The present numerical procedure was verified by comparing the results of TNA with those of three-dimensional fluid-solid conjugate heat transfer analysis (3D-CFD). After the verification, TNA was performed for several aircraft flight scenarios, and the optimum geometry of the oil cooler and the heat sink was investigated under the constraints of allowable pressure loss of air flow and outlet oil temperature for oil cooler or maximum local wall temperature for heat sink using the weight (mass) as the object function to be minimized. Furthermore, the water-mist injection to the air flow was considered for lowering the air temperature and the weight of the heat sink.

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In the present study, the preliminary design optimization of an axial-compressor is considered. The axisymmetric flow solver is developed to simulate the flow field within the compressor to predict quickly some important drawbacks of the conceptual design and treat them in preliminary design phase where the conceptual design is made by the mean-line method and the free-vortex assumption is utilized to find the radial distribution of the flow deflection angles. The finite volume scheme is used in this numerical procedure of the inviscid flow simulation where the advection upstream splitting method is used to calculate the fluxes. The focus is on the axial velocity changes along the compressor, and the optimization target of the preliminary design is to increase the minimum axial velocity with the criteria of keeping the mass flow rate and the total pressure ratio constant. The results demonstrate that this target is achieved by minor modification in flow deflection angles to improve the variation of the axial velocity, which can be more important especially in off-design performance of the compressor.

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This paper describes an experimental investigation on a state-ofthe- art compressor airfoil with three different leading edges at high subsonic flow conditions. In addition to a conventional circular and elliptical geometry which possess curvature discontinuities at the blend points, a continuous curvature leading edge is studied. The investigation considers the performance at design incidence, as well as the impact of off-design incidences. Pressure spikes near the leading edge can lead to early transition associated with higher profile losses. Goodhand and Miller [1] showed that in low subsonic conditions the avoidance of curvature discontinuities can diminish pressure spikes and therefore reduce the profile losses and enlarge the working range. In this paper, measurements are conducted to assess the potential of this concept for a high-pressure jet engine compressor airfoil operated at high subsonic conditions (M1 = 0:7, Red=2 = 20;000). The results show that, at design incidence, the total pressure loss coefficient of the continuous curvature leading edge reduces by up to 15:4%compared to the circular leading edge and by up to 3:1 % for the elliptical geometry. At off-design incidence, the reduction can be up to 40:2 % at maximum positive incidence under consideration.

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This paper presents an analysis of the vibration-induced effects on the aerofoil aerodynamics and boundary-layer development of a low-pressure–turbine blade. Large-eddy simulations of an MTUT161 low-pressure–turbine blade with imposed sinusoidal rigidbody oscillations were conducted for frequencies of 50 and 100 Hz as well as for a fixed reference blade. The oscillations are shown to impact both the time-averaged flow field and unsteady velocity fluctuations. These changes appear most markedly as a reduction in the stagnation-point pressure and a partial suppression of the separation bubble on the suction side of the aerofoil. The results suggest that the deterministic velocity fluctuations introduced by the oscillating blade promote transition on the suction side and expedite the generation of turbulence.

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