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Investigation of boundary layer transition for transonic flow over compressor and turbine airfoils
KTH, Superseded Departments, Energy Technology.
1998 (English)Doctoral thesis, comprehensive summary (Other scientific)
Abstract [en]

This thesis considers the laminar-turbulent transition overcompressor and turbine airfoils at high subsonic and transonicflow conditions.

First, experiments on the laminar-turbulent transition on anisolated airfoil in high subsonic and transonic flow wereperformed. The behaviour of the transition process was studiedfor varying incidence angles (0 - 6°) and free-streamturbulence levels (0.2 - 4.9%) for inlet Mach numbers up to0.7. Experimental investigations concerning transition at thishigh inlet Mach numbers are not common in open literature, whythe presented results contributes to extending the dataavailable for validation and development of numericalprediction codes and transition models. The transition onsetwas shown to be strongly affected by the flow conditions,moving rapidly towards the leading edge on the suction side forincreasing incidence angles, Mach numbers and free-streamturbulence levels. On the pressure side, an increased Machnumber and free-stream turbulence level also shift thetransition onset towards the leading edge, while an increasedincidence moves the onset towards the trailing edge. The lengthof the transition zone is decreasing with increasingfree-stream turbulence and somewhat with increasing Machnumbers on both sides of the airfoil. On the suction side, anincreased incidence decreases the transition zone length, whilethe opposite is valid for the pressure side. This is consistentwith what has been found by other researchers for lower Machnumbers and other profiles.

The second experimental study is aimed at studying thetransition in a turbine cascade at transonic flow. Somewhatmore experimental data than for isolated airfoils, concerningtransition in transonic turbine cascades, is available inliterature but there is still a need for additional studies athigher turbulence levels. The VKI-1 cascade used for this studyhas been extensively tested at various wind tunnels in Europeand was chosen for the investigation with the scope to extendthe existing database with boundary layer studies and with morevariation in exit Mach numbers (0.42 - 1.22). For designconditions and an inlet free-stream turbulence level of 1% theVKI-1 profile is known from other investigations to have alaminar boundary layer with a shock induced separation on thesuction side. This was also found in the presented study, eventhough the free-stream turbulence level was higher (3.5% to4.0% at design flow conditions). Increasing the pressure ratiobeyond design conditions moved the shock towards the trailingedge, as expected. Increasing the free-stream turbulence levelshowed how the suction side transition zones for subsonic exitMach numbers moved slightly upstream. On the pressure sideshort transition zones were seen also for supersonic exit Machnumbers. Varying the Reynolds number independently from theMach number showed that the transition onset moved towards theleading edge for increasing Reynolds number and that the lengthof the transition zone decreased. The Reynolds numbersdependence has been shown by other studies for supersonic exitMach numbers, but not for subsonic exit Mach numbers. Thefavourable effects on transition for increasing Mach andReynolds numbers, as well as for increasing free-streamturbulence levels, have also been shown by otherresearchers.

Finally, the boundary layer and the aerodynamicloads on thesuction side of an oscillating airfoil was investigated forincidence angles near stall. The study was performed withreduced frequencies between 0.1 and 0.6, and an inlet Machnumber of 0.5. The combination of such high reduced frequenciesand Mach number have not been found in other experimentalinvestigations in open literature. It was found that even closeto and above the steady state stall limit, the boundary layerin the leading edge region can contain transitional flow duringparts of the oscillation cycle. Increasing the mean angle ofattack caused the blade to become more excited from theaerodynamic loads, while the effect of varying the amplitudeseemed to depend on whether the mean incidence was below orabove the steady state stall limit. It was noted that theboundary layer experienced higher fluctuations for an increasedfrequency but no clear trends concerning the flutter stabilitycould be seen. Other researchers have found that an increasedfrequency will have a stabilising effect on the flutter andthat an increased mean incidence will increase the aerodynamicexcitation. The amplitude effect being dependent on whether themean incidence is below or above the steady state stall limithas not been found in any other public investigation.

Besides the transition studies, investigations of themeasurement equipment used for the experiments have beenperformed, including discussions of measurement accuracy aswell as comparisons with some other measurement techniques. Forexample the hot film technique was compared with liquidcrystals in a linear cascade at high subsonic exit It isbelieved by the author that the presented work will increasethe understanding of the transition process for compressibleflow conditions and will extend existing data for validationsand development of numerical prediction

Keywords:laminar-turbulent transition, transonic flow,turbomachinery, isolated airfoil, linear cascade, steady state,time-dependent state, hot film measurements, pressuremeasurements

Place, publisher, year, edition, pages
Stockholm: Energiteknik , 1998. , 56 p.
Trita-KRV, 1998:1
URN: urn:nbn:se:kth:diva-2734ISBN: 91-7170-348-9OAI: diva2:8444
Public defence
NR 20140805Available from: 2000-01-01 Created: 2000-01-01Bibliographically approved

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