Volume 1 of the FCC Feasibility Report presents an overview of the physics case, experimental programme, and detector concepts for the Future Circular Collider (FCC). This volume outlines how FCC would address some of the most profound open questions in particle physics, from precision studies of the Higgs and EW bosons and of the top quark, to the exploration of physics beyond the Standard Model. The report reviews the experimental opportunities offered by the staged implementation of FCC, beginning with an electron-positron collider (FCC-ee), operating at several centre-of-mass energies, followed by a hadron collider (FCC-hh). Benchmark examples are given of the expected physics performance, in terms of precision and sensitivity to new phenomena, of each collider stage. Detector requirements and conceptual designs for FCC-ee experiments are discussed, as are the specific demands that the physics programme imposes on the accelerator in the domains of the calibration of the collision energy, and the interface region between the accelerator and the detector. The report also highlights advances in detector, software and computing technologies, as well as the theoretical tools/reconstruction techniques that will enable the precision measurements and discovery potential of the FCC experimental programme. The content and structure of this report are guided by the scope and priorities defined in the mandate of the FCC Feasibility Study. It is therefore not intended to serve as an exhaustive review of the full physics potential of FCC. Several topics, already covered in earlier reports such as the FCC CDR, are not reiterated here or are addressed only briefly, in alignment with the study’s focus. This volume reflects the outcome of a global collaborative effort involving hundreds of scientists and institutions, aided by a dedicated community-building coordination, and provides a targeted assessment of the scientific opportunities and experimental foundations of the FCC programme.

Accurate structural modeling of blisk mistuning is critical for the analysis of forced response in turbomachinery. Apart from intentional mistuning, mistuning can be due to the manufacturing tolerances, corrosion, foreign object damage, and in-service wear in general. It has been shown in past studies that mistuning can increase the risk of blade failure due to energy localization. For weak blade to blade coupling, this localization has been shown to be critical and higher amplitudes of vibration are expected in few blades. This paper presents a comparison of three reduced order models (ROMs) for the structural modeling of blisks. Two of the models assume cyclic symmetry, while the third model is free of this assumption. The performance of the reduced order models for cases with small and large amount of mistuning will be examined. The benefits and drawbacks of each reduction method will be discussed.

Blade flutter in the last stage is an important design consideration for the manufacturers of steam turbines. Therefore, the accurate prediction method for blade flutter is critical. Since the majority of aerodynamic work contributing to flutter is done near the blade tip, resolving the tip leakage flow can increase the accuracy of flutter predictions. The previous research has shown that the induced vortices in the tip region can have a significant influence on the flow field near the tip. The structure of induced vortices due to the tip leakage vortex cannot be resolved by unsteady Reynolds-averaged Navier–Stokes (URANS) simulations because of the high dissipation in turbulence models. To the best of author’s knowledge, the influence of induced vortices on flutter characteristics has not been investigated. In this paper, the results of detached-eddy simulation (DES) and URANS flutter simulations of a realistic-scale last-stage steam turbine are presented, and the influence of induced vortices on the flutter stability has been investigated. Significant differences for the predicted aerodynamic work coefficient distribution on the blade surface, especially on the rear half of the blade suction side near the tip, are observed. At the least stable interblade phase angle (IBPA), the induced vortices show a destabilizing effect on the blade aeroelastic stability. The motion of induced vortices during blade oscillation is dependent on the blade amplitude, and hence, the aerodynamic damping is also dependent on the blade vibration amplitude. In conclusion, the induced vortices can influence the predicted flutter characteristics of the steam turbine test case.

In this paper, a method to influence the vibratory blade stresses of mixed flow turbocharger turbine blade by varying the local blade thickness in spanwise direction is presented. Such variations have an influence on both the static and the vibratory stresses and therefore can be used for optimizing components with respect to high-cycle fatigue (HCF) tolerance. Two typical cyclic loadings that are of concern to turbocharger manufacturers have been taken into account. These loadings arise from the centrifugal forces and from blade vibrations. The objective of optimization in this study is to minimize combined effects of centrifugal and vibratory stresses on turbine blade HCF and moment of inertia. Here, the conventional turbine blade design with trapezoidal thickness profile is taken as baseline design. The thicknesses are varied at four spanwise equally spaced planes and three streamwise planes to observe their effects on static and vibratory stresses. The summation of both the stresses is referred to as combined stress. In order to ensure comparability among the studied design variants, a generic and constant excitation order-dependent pressure field is used at a specific location on blade. The results show that the locations of static and vibratory stresses, and hence the magnitude of the combined stresses, can be influenced by varying the blade thicknesses while maintaining the same eigenfrequencies. By shifting the maximum vibratory stresses farther away from the maximum static stresses, the combined stresses can be reduced considerably, which leads to improved HCF tolerance.

A novel control technique for radial turbines is under investigation for providing turbine performance controllability, especially in turbocharger applications. This technique is based on replacing the traditional spiral casing with a multi-channel casing (MC). The MC divides the turbine rotor inlet circumferentially into a certain number of channels. Opening and closing these channels controls the inlet area and, consequently, the turbine performance. The MC can be distinguished from other available control techniques in that it contains no movable parts or complicated control mechanisms. Within the casing, this difference makes it practical for a broader range of applications. In this investigation, a turbocharger featuring a turbine with MC has been tested on a hot gas test stand. The experimental test results show a reduction in the turbine operating efficiency when switching from full to partial admission. This reduction increases when reducing the admission percentage. To ensure the best performance of the turbine featuring MC while operating at different admission configurations, it becomes crucial to investigate its internal flow field at both full and partial admission to understand the reasons for this performance reduction. A full 3D computational fluid dynamics (CFD) model of the turbine was created for this investigation. It focuses on identifying the loss mechanisms associated with partial admission. Steady and unsteady simulations were performed and validated with available test data. The simulation results show that operating the turbine at partial admission results in highly disturbed flow. It also detects the places where aerodynamic losses occur and which are responsible for this performance reduction. This operation also shows flow unsteadiness even when operating at steady conditions. This unsteadiness depends mainly on the admission configuration and percentage.

In order to change the blade count (BC) of axial rotor designs, a scaling technique can be applied where the blades are scaled in axial and circumferential directions while maintaining the solidity. This technique allows to adjust the blade count without changing the steady-state aerodynamics, but the influences on the aerodynamic damping are unknown. The present study is focused on the investigation of the change in aerodynamic damping of a subsonic axial compressor rotor, if the blade count is changed between 13 and 25 blades. The investigation is focused on the first bending mode family and the influences of the hub geometry on the modes are neglected. First, a comparison between the influence coefficient (IC) method and the traveling wave mode (TWM) method is conducted, which shows that the application of the IC method in combination with harmonic balance simulations offers a fast way to compute the aerodynamic damping without introducing significant errors compared with the TWM method simulations. Regarding the aerodynamic damping of the different rotor geometries, it can be noted that the amplitude of the work per cycle influence coefficient of the center blade scales linearly with the blade scaling factor and an increase in aerodynamic damping for an increase in blade count is observed. Furthermore, a simplified analytic theory is established, which explains the phase angle change of the blade influence coefficients due to the blade scaling. In the last part of the paper, an extrapolation method is proposed for the investigated geometry and mode family, which allows for the estimation of damping S-curves for scaled rotor geometries based on only one IC method simulation.