Investigation of flow separation in a transonic-fan linear cascade using visualization methods
An extensive experimental study into the nature of the separated flows on the blade suction surface of modern transonic fans is described in this paper. The study was a subtask of a larger experimental effort focused on blade flutter excited by flow separation in the blade tip region. The tip sections of airfoils on transonic fan blades are designed for precompression and consequently they differ from sections on the rest of the blade. The blade tip section was modeled by a low aspect ratio blade and therefore most of the blade tested was exposed to the secondary flow effects. The aim of this work was to supply reliable data on flow separation on transonic fan blades for validation of future analytical studies. The experimental study focused on two visualization techniques: surface flow visualization using dye oils and schlieren (and shadowgraph) flow visualization. The following key observations were made during the study. For subsonic inlet flow, the flow on the suction surface of the blade was separated over a large portion of the blade, and the separated area increased with increasing inlet Mach number. For the supersonic inlet flow condition, the flow was attached from the leading edge up to the point where a bow shock from the upper neighboring blade imposed on the blade surface. Downstream, there was a separated flow region in which air flowed in the direction opposite the inlet flow. Finally, past the separated flow region, the flow reattached to the blade surface. For subsonic inlet flow, the low cascade solidity resulted in an increased area of separated flow. For supersonic flow conditions, the low solidity resulted in an improvement in flow over the suction surface.
Unsteady Pressure Measurements on a Biconvex Airfoil in a Transonic Oscillating Cascade
Flush-mounted dynamic pressure transducers were installed on the center airfoil of a transonic oscillating cascade to measure the unsteady aerodynamic response as nine airfoils were simultaneously driven to provide 1.2 deg of pitching motion about the midchord. Initial tests were performed at an incidence angle of 0.0 deg and a Mach number of 0.65 in order to obtain results in a shock-free compressible flow field. Subsequent tests were performed at an angle of attack of 7.0 deg and a Mach number of 0.80 in order to observe the surface pressure response with an oscillating shock near the leading edge of the airfoil. Results are presented for interblade phase angles of 90 and −90 deg and at blade oscillatory frequencies of 200 and 500 Hz (semichord reduced frequencies up to about 0.5 at a Mach number of 0.80). Results from the zero-incidence cascade are compared with a classical unsteady flat-plate analysis. Flow visualization results depicting the shock motion on the airfoils in the high-incidence cascade are discussed. The airfoil pressure data are tabulated.
Stall Flutter Prediction Techniques for Fan and Compressor Blades
A review is presented of stall flutter prediction methodology including empirical, analytical, and semiempirical approaches. Because of the lack of accurate and efficient unsteady aerodynamic models for stall flutter predictions, interim semiempirical methods are developed to assist turbomachinery designers during the development of current advanced fan and compressor designs for both military and commercial applications. A semiempirical approach that builds on an empirical stall flutter database combined with a classical flat plate linearized unsteady aerodynamic model that accounts for chordwise deformations has been developed. This model is shown to provide good correlations when compared to isolated airfoil and cascade data for standard configurations. When applied to low aspect ratio fans, this approach provided excellent correlations with experimental results.