In a multistage intermediate pressure compressor an efficiency benefit may be gained by reducing reaction in the rear stages, and allowing swirl to persist at the exit. This swirl must now be removed within the transition duct that is situated between the intermediate and high pressure compressor spools, in order to present the downstream compressor with suitable inlet conditions. This paper presents the numerical design and experimental validation of an initial concept which uses a lifting strut to remove tangential momentum from the flow within an S-shaped compressor transition duct. The design methodology uses an existing strut profile with the camber line modified to remove a specified amount of the inlet tangential momentum. A linear strut loading was employed in the meridional direction with a nominally constant loading in the radial direction. This approach was applied to an existing aggressive S-duct configuration in which approximately 12.5° of swirl remains at OGV exit. 3D CFD predictions were used for preliminary assessment of duct loading and to determine how much swirl could be removed. Consequently, a fully annular test facility incorporating a 1½ stage axial compressor was used to experimentally evaluate four configurations; an unstrutted duct, a non-lifting strut and lifting struts designed to remove 50% and 75% of the inlet tangential momentum. Despite the expected large increase in loss caused by the introduction of struts there was not a significant additional loss measured with the inclusion of turning. No evidence of flow separation was observed and the data suggested that it may be possible to remove more swirl than was attempted. Although the turning struts did not remove the entire targeted swirl due to viscous deviation the data still confirm the feasibility of using a lifting strut/duct concept which has the potential to off-load the rear stages of the upstream compressor.

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 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.

The S-shaped duct which transfers flow from the low-pressure fan to the engine core of modern large civil turbofans presents a challenging design problem. Aerodynamically it must accommodate a spatially and temporarily non-uniform inlet in conjunction with a complex flow field which will develop under the combined influence of pressure gradients and streamline curvature. It must also allow for the transfer of structural loads and services across the main gas path. This necessitates the use of structural vanes which can compromise aerodynamics, introduce unwanted component interactions and erode performance. Furthermore, this must all be achieved with minimum length/weight and without flow separation. This paper presents a comprehensive aerodynamic evaluation of three options for a low-pressure compressor transition duct containing (i) a long-chord, structural compressor outlet guide vane, (ii) a more aerodynamically optimal but non-structural outlet guide vane in conjunction with a small number of discrete radial load bearing struts and (iii) a fully integrated outlet guide vane and strut design. Evaluation was performed using a fully annular, low-speed test facility incorporating a 1 ½ stage axial compressor and transition duct representative of an engine design. Aerodynamic data were produced from miniature five-hole probe area traverses conducted at several locations with compressor/duct. The data suggest that all the options were viable. However, the aerodynamic vane and discrete struts produced the lowest system loss with the other two options being comparable. The performance of the structural vane was seen to be sensitive to off-design conditions producing a notably increased loss at a low flow coefficient. The more optimized aerodynamic vanes were much less sensitive to off-design conditions whilst the fully integrated design showed only very small changes in loss.

The S-shaped duct which transfers flow from the low-pressure fan to the engine core in large civil turbofans presents a challenging problem. Aerodynamically it has a spatially and temporarily varying inlet flow combined with a complex flow field which develops under the combined influence of pressure gradients and streamline curvature. It must also accommodate the transfer of structural loads and services across the main gas path. This necessitates the use of structural vanes which can compromise aerodynamics, introduce unwanted component interactions, and erode performance. This must all be achieved with minimum length/weight and without flow separation. This paper presents a comprehensive aerodynamic evaluation of three design options for a transition duct containing (i) a long-chord, structural compressor outlet guide vane, (ii) a aerodynamically optimal but non-structural outlet guide vane in conjunction with a small number of load bearing struts and (iii) a fully integrated outlet guide vane and strut design. Evaluation was performed using a low-speed test facility incorporating a 1½ stage axial compressor and engine representative transition duct. Measured data suggest that all the options were viable. However, the aerodynamic vane and discrete struts produced the lowest system loss with the other two options being comparable. The performance of the structural vane was sensitive to off-design conditions producing a notable increase in loss at a low flow coefficient. The optimized aerodynamic vanes were much less sensitive to off-design conditions whilst the fully integrated design showed only very small changes in loss.