Modern turbofan engines employ a highly loaded fan stage with transonic or low-supersonic velocities in the blade-tip region. The fan blades are often prone to flutter at off-design conditions. Flutter is a highly undesirable and dangerous self-excited mode of blade oscillations that can result in high-cycle fatigue blade failure. The origins of blade flutter are not fully understood yet. Experimental data that can be used to clarify the origins of blade flutter in modern transonic fan designs are very limited.
The Transonic Flutter Cascade Facility at the NASA Glenn Research Center was developed to experimentally study the details of flow mechanisms associated with fan flutter. The cascade airfoils are instrumented to measure high-frequency unsteady flow variations in addition to the steady flow data normally recorded in cascade tests. The test program measures the variation in surface pressure in response to the oscillation of one or more of the cascade airfoils. However, during the initial phases of the program when all airfoils were in fixed positions, conditions were found where significant time variations in the pressures near the airfoil leading edges could be observed.
Fan flow behavior. Left: Shockwave patterns from shadowgraph flow visualization. Right: Time-resolved pressure signal on suction surface past the leading edge. Top row: Subsonic--below Mach 0.90. Middle two rows: Transonic--Mach 0.9 to 1.01. Bottom row: Supersonic--above Mach 1.01.
The variations in flow were observed in shadowgraph images of the tunnel flow and were confirmed by high-frequency electronic pressure sensors embedded in the airfoil surface. Prior to shock waves forming on the airfoils, the flow appeared uniform in the shadow-graph figure and in the surface pressure trace from the sensors. These conditions exist for subsonic inlet flow below an inlet flow Mach number of 0.9. After the Mach number was raised slightly about unity to a value of 1.01, supersonic inlet flow was established, the shadowgraph showed a similar shockwave pattern for each airfoil, and again the pressure trace was uniform in time. While the inlet Mach number was in the low to high transonic range (between 0.90 and 1.01), the shadowgraph showed some airfoils with shock waves and some without. Also, the particular airfoils that had shock waves would vary with time. This was confirmed by the data from the sensors as the pressure level randomly alternated between two pressure levels. The lower pressure level indicated the presence of a shock wave and the higher level the absence.
The observed behavior of the cascade experiment could be very useful in the control and avoidance of stall flutter. If this behavior is an indicator or precursor to a fan entering into an area of operation where stall flutter will occur, it can be used in an active control system to control the fan. At this point, however, it is not clear if the observed behavior is common to all current fans, is only observed in this particular airfoil design, or is an artifact of the experimental arrangement. Study of these preliminary observations will continue.
QSS contact: Dr. Jan Lepicovsky, 216-977-1402, Jan.Lepicovsky@grc.nasa.gov
Glenn contacts: Dr. Eric R. McFarland, 216-433-5915, Eric.R.McFarland@grc.nasa.gov; and Dr. Rodrick V. Chima, 216-977-5919, Rodrick.V.Chima@grc.nasa.gov
Author: Dr. Jan Lepicovsky
Headquarters program office: OAT
Programs/Projects: Propulsion Systems R&T, SEC, UEET, RAC
Last updated: June 2002
For additional information, please contact Cynthia L. Dreibelbis at 216-433-2912 or email@example.com.
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