Aerodynamic characteristics of a transonic airfoil cascade

Abstract

The trend towards higher pressure ratio and compact turbomachines with a reduced number of stages leads to a considerable increase of aerodynamic loading of the airfoil cascades. Thereby the velocities relative to the blades increase to transonic and supersonic speeds and shock-waves occur within the airfoil cascade. These shock waves could interact with the airfoil surface boundary layer and can cause unsteady boundary layer separation. In fact, Shock wave–boundary-layer interactions (SBLIs) occur when a shock wave and a boundary layer converge and since both can be found in almost every supersonic flow, these interactions are commonplace. The most obvious way for them to arise is for an externally generated shock wave to impinge onto a surface on which there is a boundary layer. In the transonic regime, shock waves are formed at the downstream edge of an embedded supersonic region; where these shocks come close to the surface, a SBLI is produced. In any SBLI, the shock imposes an intense adverse pressure gradient on the boundary layer, which causes it to thicken and possibly also to separate. SBLI also causes flow unsteadiness. Shock induced oscillations (SIO), aerodynamic instabilities, high cycle fatigue failure (HCF), non-synchronous vibration (NSV), aeroacoustic noise and so on are the detrimental consequences of this unsteady shock wave boundary layer interaction. On transonic wings, it increases the drag and has the potential to cause flow unsteadiness and buffet. In hypersonic flight, SBLI can be disastrous because at high Mach numbers, it has the potential to cause intense localized heating that can be severe enough to destroy a vehicle. Because of its significance for many practical applications, SBLI is the focus of numerous studies spanning several decades. Many of the investigations have been dealt considering an isolated airfoil in transonic flows. However, little information is available on the aerodynamics of airfoil cascades. The goal of the present research is to analyze and to understand the transonic flow phenomena in a circular arc airfoil cascade using experiments and numerical computation. Experimental tests were conducted to investigate the behavior of passage shock waves with the shock induced boundary layer separation in a supersonic wind tunnel flow facility. Further, a Reynolds averaged Navier- Stokes (RANS) solver was used to provide airfoil surface pressures, overall performance from wake characteristics and so on. Particular attention is to be paid on the embedded shock wave structure and an accurate simulation of the shock boundary layer interaction. The experiment was performed for unstaggered case and the numerical studies were performed for stagger or setting angle 0º to 20º. The results show that the self-excited shock wave oscillations occur in the cascade passage for a pressure ratio of 0.75 in both unstaggered and staggered case. Fluctuating pressure histories are recorded at different locations in the flow field. PSD from FFT calculation of the data is used to find the principal frequency of the unsteady behaviour. A frequency of 976 Hz is found to be the dominating frequency.It is observed that for all the cases, the flow field remains undisturbed from leading edge to x/c=0.50, as in that portion no shock wave is observed on the airfoil surfaces. For different stagger angle, peak RMS of pressure oscillation (prms/q0) is calculated and its location is identified. Flow separation occurs at a distance after the shock wave. Separation points and separation lengths are also calculated. Wave drags contribution is much higher than viscous drag at high speed compressible flow. Wave drag coefficients are also calculated for different flow conditions.