Integrated design of compressor transition ducts with swirling flow and aerodynamic lifting struts

Integrated design of compressor transition ducts with swirling flow and aerodynamic lifting strut

An experimental, aerodynamic evaluation of design choices for a low-pressure compressor transition duct

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 1 ½ 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

Experimental investigation of the effect of bleed on the aerodynamics of a low-pressure compressor stage in a turbofan engine [GT2023-102260]

The compression system in modern turbofan engines is split into several stages linked by s-shaped transition ducts. Downstream of the low-pressure system, a handling bleed is often required for off-design performance and/or to extract ice/water and foreign debris prior to the air entering the highpressure compression stages. The inclusion of this bleed and various structural vanes can introduce unwanted component interactions and compromise the aerodynamic performance of the upstream low-pressure compressor stage and downstream transition duct. This paper presents an experimental investigation of the aerodynamic performance of a compressor transition duct and bleed for a very high bypass ratio turbofan. A fully annular, low-speed test facility incorporating a 1½ stage axial compressor was used to examine the mean and unsteady flow in the last stage of a low-pressure compressor and the downstream transition duct. The transition duct incorporated load bearing struts, including a so-called King strut with twice the thickness of the regular struts. The bleed utilized a 360° annular slot located on the casing immediately downstream of the low-pressure rotor and upstream of the outlet guide vane. The results showed that the King strut, caused a similar flow distortion and redistribution in the OGV like the Regular struts, and had otherwise imposed a negligible effect on overall performance over a range of rotor flow coefficients. The addition of bleed had a more notable effect, generating an increasing outboard bias in the rotor efflux, as the flow migrated towards the offtake. At the design flow operating point, the OGV were relatively insensitive to this until the highest bleed rate (18%) where evidence of stall was observed. At a lower operating point, the change of rotor swirl and additional OGV incidence caused earlier onset of stall and a full OGV stall was observed above 10% bleed. Increasing bleed was observed to cause a gradual increase in duct loss up to the point of OGV stall when losses increased more rapidly.</p

Experimental investigation of the effect of bleed on the aerodynamics of a low-pressure compressor stage in a turbofan engine

The compression system in modern turbofan engines is split into stages linked by transition ducts. Downstream of the low-pressure system, a handling bleed is often required and, in conjunction with structural vanes, can introduce component interactions which compromise aerodynamic performance. In this paper a fully annular, low-speed test facility incorporating a 1½ stage axial compressor is used to examine the flow in the last stage of a low-pressure compressor and the downstream transition duct. The transition duct incorporated load bearing struts, including a socalled king strut with twice the thickness of the regular struts. The bleed utilized a 360o annular slot located on the casing immediately downstream of the low-pressure rotor and upstream of the outlet guide vane. The results showed that both the regular and king strut caused a similar flow distortion in the vane row but overall imposed a negligible effect on overall performance. The addition of bleed had a larger effect, generating an increasing outboard bias at rotor exit as the bleed flow migrated towards the offtake. At the design operating point, the outlet guide vanes were relatively insensitive to this until the highest bleed rate (18%) where evidence of stall was observed. At a lower operating point, a modification to the rotor swirl caused additional incidence onto the vanes resulting in earlier onset of stall; a full stall was observed above 10% bleed. Increasing bleed caused a gradual increase in duct loss up to stall when losses increased rapidly.  </p

Experimental investigation of the effect of bleed on the aerodynamics of a low-pressure compressor stage in a turbofan engine [GT2023-102260]

The compression system in modern turbofan engines is split into several stages linked by s-shaped transition ducts. Downstream of the low-pressure system, a handling bleed is often required for off-design performance and/or to extract ice/water and foreign debris prior to the air entering the highpressure compression stages. The inclusion of this bleed and various structural vanes can introduce unwanted component interactions and compromise the aerodynamic performance of the upstream low-pressure compressor stage and downstream transition duct. This paper presents an experimental investigation of the aerodynamic performance of a compressor transition duct and bleed for a very high bypass ratio turbofan. A fully annular, low-speed test facility incorporating a 1½ stage axial compressor was used to examine the mean and unsteady flow in the last stage of a low-pressure compressor and the downstream transition duct. The transition duct incorporated load bearing struts, including a so-called King strut with twice the thickness of the regular struts. The bleed utilized a 360° annular slot located on the casing immediately downstream of the low-pressure rotor and upstream of the outlet guide vane. The results showed that the King strut, caused a similar flow distortion and redistribution in the OGV like the Regular struts, and had otherwise imposed a negligible effect on overall performance over a range of rotor flow coefficients. The addition of bleed had a more notable effect, generating an increasing outboard bias in the rotor efflux, as the flow migrated towards the offtake. At the design flow operating point, the OGV were relatively insensitive to this until the highest bleed rate (18%) where evidence of stall was observed. At a lower operating point, the change of rotor swirl and additional OGV incidence caused earlier onset of stall and a full OGV stall was observed above 10% bleed. Increasing bleed was observed to cause a gradual increase in duct loss up to the point of OGV stall when losses increased more rapidly.</p

Experimental investigation of the effect of bleed on the aerodynamics of a low-pressure compressor stage in a turbofan engine

The compression system in modern turbofan engines is split into stages linked by transition ducts. Downstream of the low-pressure system, a handling bleed is often required and, in conjunction with structural vanes, can introduce component interactions which compromise aerodynamic performance. In this paper a fully annular, low-speed test facility incorporating a 1½ stage axial compressor is used to examine the flow in the last stage of a low-pressure compressor and the downstream transition duct. The transition duct incorporated load bearing struts, including a so-called king strut with twice the thickness of the regular struts. The bleed utilized a 360° annular slot located on the casing immediately downstream of the low-pressure rotor and upstream of the outlet guide vane. The results showed that both the regular and king strut caused a similar flow distortion in the vane row but overall imposed a negligible effect on overall performance. The addition of bleed had a larger effect, generating an increasing outboard bias at rotor exit as the bleed flow migrated towards the offtake. At the design operating point, the outlet guide vanes were relatively insensitive to this until the highest bleed rate (18%) where evidence of stall was observed. At a lower operating point, a modification to the rotor swirl caused additional incidence onto the vanes resulting in earlier onset of stall; a full stall was observed above 10% bleed. Increasing bleed caused a gradual increase in duct loss up to stall when losses increased rapidly.</p

The influence of fan root flow on the aerodynamic of a low-pressure compressor transition duct

To reduce fuel-burn and emissions there is a drive towards 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 characterised 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 CFD have been used to examine 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 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 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

Experimental investigation of secondary flows and length reduction for a low-pressure compressor transition duct

The need to reduce fuel-burn and CO2 emissions, is pushing turbofan engines towards geared architectures with very high bypass-ratios and small ultra-high-pressure ratio core engines. However, this increases the radial offset between compressor spools and leads to a more challenging design for the compressor transition ducts. To minimise weight, these ducts must achieve the radial turning in as short a length, but this leads to strong curvature induced pressure gradients, increased aerodynamic loading and likelihood of flow separation. For the duct connecting the low-pressure fan to the engine core this is further complicated by the poor-quality flow generated at the fan hub which is characterised by low total pressure and large rotating secondary flow structures. In a previous paper the authors numerically designed modifications to an existing test facility such that the rotor would produce these large structures. The current paper presents an experimental evaluation of the new rotor design and examines the effect of the increased loss cores on the performance of a set of engine sector stators (ESS) or outlet guide vanes (OGV) and an engine representative compressor transition duct. Aerodynamic data were collected via miniature five-hole probes, for the time-averaged pressure and velocity field, and phaselocked hot-wire anemometry to capture the rotating secondary flows. Analysis of the experimental data showed that these structures promoted mixing through the ESS increasing the momentum exchange between the core and boundary layer flows. Measurements within the duct showed a continued reduction in the hub-wall boundary layer suggesting that the duct has been moved further from separation. Consequently, a more aggressive duct with 12.5% length reduction was designed and tested with the data confirming that the more aggressive duct remained fully attached. Total pressure loss data suggested a slight increase in loss over the vane row but that was offset by a reduced loss in the duct due to improved flow quality and reduced length. Overall, the 12.5% length reduction represents a significant cumulative effect in terms of reduced fuel burn and CO2 over the operational life of an engine

The influence of fan root flow on the aerodynamic of a low-pressure compressor transition duct

To reduce fuel-burn and emissions there is a drive towards 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 characterised 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 CFD have been used to examine 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 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 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

Experimental investigation of secondary flows and length reduction for a low-pressure compressor transition duct

The need to reduce fuel-burn and CO2 emissions, is pushing turbofan engines towards geared architectures with very high bypass-ratios and small ultra-high-pressure ratio core engines. However, this increases the radial offset between compressor spools and leads to a more challenging design for the compressor transition ducts. To minimise weight, these ducts must achieve the radial turning in as short a length, but this leads to strong curvature induced pressure gradients, increased aerodynamic loading and likelihood of flow separation. For the duct connecting the low-pressure fan to the engine core this is further complicated by the poor-quality flow generated at the fan hub which is characterised by low total pressure and large rotating secondary flow structures. In a previous paper the authors numerically designed modifications to an existing test facility such that the rotor would produce these large structures. The current paper presents an experimental evaluation of the new rotor design and examines the effect of the increased loss cores on the performance of a set of engine sector stators (ESS) or outlet guide vanes (OGV) and an engine representative compressor transition duct. Aerodynamic data were collected via miniature five-hole probes, for the time-averaged pressure and velocity field, and phaselocked hot-wire anemometry to capture the rotating secondary flows. Analysis of the experimental data showed that these structures promoted mixing through the ESS increasing the momentum exchange between the core and boundary layer flows. Measurements within the duct showed a continued reduction in the hub-wall boundary layer suggesting that the duct has been moved further from separation. Consequently, a more aggressive duct with 12.5% length reduction was designed and tested with the data confirming that the more aggressive duct remained fully attached. Total pressure loss data suggested a slight increase in loss over the vane row but that was offset by a reduced loss in the duct due to improved flow quality and reduced length. Overall, the 12.5% length reduction represents a significant cumulative effect in terms of reduced fuel burn and CO2 over the operational life of an engine