To reduce fuel-burn and emissions, there is a drive toward higher bypass ratio and smaller high-pressure ratio core engines. This makes the design of the ducts connecting compressor spools more challenging as the higher radius change increases aerodynamic loading. This is exacerbated at inlet to the engine core by fan root flow which is characterized by a hub-low-pressure profile and large secondary flow structures. Additionally, shorter, lighter nacelles mean that the intake may not provide a uniform inlet flow when the aircraft is at an angle of attack or subject to cross winds. Such inlet distortion can further degrade the flow entering the engine. A combination of experiments and computational fluid dynamics (CFD) has been used to examine the effects on the aerodynamics of an engine section splitter (ESS) and transition duct designed to feed the low-pressure spool of a high bypass ratio turbofan. A test facility incorporating a 1½ stage axial compressor was used to compare system performance for a flat rotor exit profile to one with a hub deficient flow. Validated Reynolds averaged Navier–Stokes (RANS) CFD was then used to further investigate the effects of increased inlet boundary layer thickness and bulk swirl distortion at rotor inlet. These changes were seen to have a surprisingly small effect on the flow at the duct exit. However, increased secondary flows were observed which degraded the performance of the ESS and significantly increased loss. Nevertheless, the enhanced mixing delayed separation in the duct suggesting that overall the design was reasonably robust albeit with increased system loss.

To reduce fuel-burn and CO2 emissions from aero gas turbines there is a drive towards very-high bypass ratio and smaller ultra-high-pressure ratio core engine technologies. However, this makes the design of the ducts connecting various compressor spools more challenging as the larger required radius change increases their aerodynamic loading. This is exacerbated for the duct which feeds the engine core as it must accept the relatively low-quality flow produced by the fan root. This is characterised by a hub-low pressure profile and large secondary flow structures which will inevitably increase loss and the likelihood of flow separation. This paper presents a numerical design study aimed at modifying an existing fully annular, low-speed test facility to produce larger, more representative rotor loss cores. The CFD domain comprises a single stage axial compressor and transition duct representative of the low-pressure spool in a high bypass ratio turbofan. The rotating and stationary frames were coupled using a sliding mesh and a uRANS approach with Reynolds Stress closure for the turbulence modelling. The methodology was first validated using existing experimental data before examining a number of parameters such as inlet boundary layer thickness, inlet swirl, tip gap, blade solidity, thickness, lean. It was found that a combination of an increased inlet hub boundary layer and a reduction in solidity sufficiently increased the hub loading and encouraged the development of significantly large loss cores. Preliminary CFD results also showed that the increased rotor loss cores promoted better mixing through the outlet guide which subsequently reduced hub boundary layer thickness and secondary flows. Despite incurring a larger total pressure loss this suggested that the transition duct moved further from stall and could therefore potentially be made shorter.

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To reduce fuel-burn and CO2 emissions from aero gas turbines there is a drive towards very-high bypass ratio and smaller ultra-high-pressure ratio core engine technologies. However, this makes the design of the ducts connecting various compressor spools more challenging as the higher required radius change increases their aerodynamic loading. This is exacerbated for the duct which feeds the engine core as it must accept the relatively low-quality flow produced by the fan root. This is characterised by a hub-low pressure profile and large secondary flow structures which will inevitably increase loss and the likelihood of flow separation. Additionally, the desire for shorter, lighter nacelles means that the engine intake may be unable provide a uniform inlet flow to the fan when the aircraft is at an angle of attack or subject to cross winds. Any inlet distortion this generates can also further degrade the quality of the flow entering the core of the engine. This paper uses a combination of experiments and CFD to examine the effects of the inlet flow on the aerodynamics of an engine section splitter and transition duct designed to feed the low-pressure spool of a high bypass ratio turbofan. A fully annular test facility incorporating a 1½ stage axial compressor was used to compare the system performance of a rotor that produced a nominally flat profile with one that had a notably hub deficient flow. A RANS CFD model, employing a mixing plane between the rotor and Engine Section Stator (ESS) and a Reynolds Stress turbulence model, was then validated and used to further investigate the effects of increased inlet boundary layer thickness and bulk swirl distortion at rotor inlet. Overall, changes to the inlet condition were seen to have a surprisingly small effect on the flow at duct exit — i.e. the flow presented to the downstream compressor. Changes to the inlet did, however, generate increased secondary flows and degrade the performance of the ESS. This resulted in notably increased total pressure loss; in excess of 12% for the hub-low inlet and in excess of 30% at high inlet swirl where the flow in the ESS separated. However, the increased ESS wake structures, and the enhanced mixing, delayed separation in the duct suggesting that, overall the design was reasonably robust, albeit with a significant penalty in system loss.

Heat transfer and aerodynamic measurements are conducted by means of Liquid Crystal Thermography and Particle Image Velocimetry in a two-pass rotating ribbed channel. The channel presents a square cross section, a sharp U-bend connecting the inlet and outlet passes and square ribs placed on two opposite walls, normal to the mean flow. In the heat transfer experimental campaign, the Reynolds (Re) and rotation (Ro) numbers are respectively ranging between 15,000–55,000 and 0–0.26 to investigate their influence upstream and downstream of the bend region. The aerodynamic measurements are taken in the symmetry plane of the channel at Re = 15,000 and Ro = 0 and 0.26, to complement the heat transfer data in the same regions. The results show how the Coriolis forces affect the flow stability and the secondary flow pattern. In the first pass, the behavior with varying Reynolds and rotation numbers is very similar to the one observed in a similar single-pass channel in terms of flow stability, velocity distribution and heat transfer performance. The measurements indicate an increase of the turbulent kinetic energy and the heat transfer downstream of the bend due to the large separation bubble, the high streamline curvature and the Dean vortices. Both the heat transfer and velocity distributions suggest that the interaction of the Dean vortices and Coriolis-induced secondary flows downstream of the bend is also highly dependent on the rotational regime.