Stall inception in high stage loading low pressure ratio transonic fans

There is an economical and ecological incentive for jet engine manufacturers to reduce fuel consumption of its engines. Reducing the pressure ratio of the fan system is one route through which fuel consumption can be reduced by increasing propulsive efficiency. However, reducing the fan total pressure ratio encourages a move to increase the stage loading, which can enable engine weight reduction. This moves the fan into an area of the design space that is highly challenging in terms of stall. Stall inception of a modern, high stage loading, low pressure ratio transonic was investigated using single-passage steady and full annulus unsteady computational fluid dynamics. The computational method was validated using experimental data and approached as a blind test case. A key post-processing method used in this work to expose the route to stall in the full annulus unsteady results was the circumferential averaging of incidence and blockage for each passage separately. Sudden growth of the suction surface separation behind the shock around the mid-span of the mis-staggered blade passage has been found to trigger stall. The baseline fan geometry has a mid-loaded radial work distribution, so a combination of high work loading and high incidence made the suction surface boundary layer behind the strong shock sensitive to separation. After stall has been triggered around mid-span, the stall cell embryo moved to the casing and develops into a stable stall cell. Two design alterations to the baseline fan geometry confirmed that stall was triggered around mid-span: the effect of changing tip gap is small and the introduction of a slip patch on the suction surface between 25-75\% span increased stall margin by 25\%. These showed there is significant potential to increase stall margin by delaying the onset of the suction surface boundary layer separation separation behind the shock around the mid-span of the baseline rotor geometry. It was shown that redesigning the baseline fan rotor geometry between 25-75\% span only can be beneficial for stall margin. The rotor blade leading edge was recambered in the direction to decrease incidence near the stability limit relative to the baseline. This decreased the pre-shock Mach number by 13\% near stall, which reduced the suction surface boundary layer separation behind the shock and in this way increased the stall margin by 6.7\%. This design change had minimal effect on the design point efficiency, which stayed within 0.1\% of the datum geometry. At off-design rotational speeds, the behaviour of the suction surface boundary layer changed. At 100\% speed, the blade relative Mach number increased by 9\%, which caused the radial work distribution of the fan to increase towards the tip of the rotor blade compared to design speed. The trigger of stall inception, which was still driven by suction surface separation behind the shock, followed this radial distribution change and moved away from mid-span towards the tip region of the rotor blade. At part speed conditions, the tip relative Mach number dropped to become subsonic, which altered the behaviour of the interaction between the shock and suction surface boundary layer. The suction surface boundary layer now separated close to the leading edge near stall (x/c_x < 1.5%) and induced axial oscillations prior to stall. Eventually, these stable periodic oscillations became unstable and pushed the fan into stall. The implications of this work is that a fan designer who aims to maximise flow range for a high stage loading low pressure ratio transonic fan needs to pay close attention to the behaviour of the suction surface boundary layer and its interaction with the shock along the whole span of the rotor blade since increasing stage loading coefficient make the suction surface boundary layer more prone to separate behind the shock. It has been shown that the radial location of the stall trigger changes concurrently with changes in pre-shock Mach number and radial work loading distribution along the span. Therefore, it is insufficient to solely focus on the tip region when investigating stall inception in high stage loading low pressure ratio transonic fans.

Name dc:type.qualificationname

Doctor of Philosophy (PhD)

Level dc:type.qualificationlevel

Doctoral

Grantor dc:publisher.institution

University of Cambridge

Year dc:date.issued

2020

Author dc:creator

• Verschueren, Hans

Advisor dc:contributor.advisor

• Hall, Cesare

• fluid dynamics
• aerodynamics
• transonic fan
• turbomachinery
• jet engine design

Licence

• https://www.repository.cam.ac.uk/bitstreams/4d5be06b-0c80-4c0c-b86a-9da0f41fd618/download
• https://www.rioxx.net/licenses/all-rights-reserved/

Language dc:language

eng

Author Identifier

0000-0003-2741-5487

OAI identifier oai:identifier

oai:www.repository.cam.ac.uk:1810/321427