Ernesto Benini

Airfoil Data Fitting Using Multivariate Smoothing Thin Plate Splines

Smoothing thin plate splines, a fitting technique based on a rigorous roughness penalty approach, have been recently investigated as a promising tool for bivariate interpolation of aerodynamic data. In this paper, this technique is implemented and extended to multivariate fitting. In particular, the method is applied for estimating the aerodynamic polars of well-known two-dimensional symmetrical and nonsymmetrical airfoils as functions of some geometric parameters describing the airfoil shape and a further variable defining the flow regime (either the Mach or the Reynolds number). Therefore, the simultaneous influence of five independent variables on three responses (lift, drag, and pitching moment coefficients) is investigated. To this purpose, a large database is generated via numerical simulations (using a validated flow solver) containing all information required to build a reliable response surface. Then, the model is built and its performance validated by performing queries on complete aerodynamic polars at various flow regime conditions of a series of airfoils not included into the database. Results show a very good matching between predicted and calculated curves, thus demonstrating the remarkable predictive capability of the implemented tool.

State-of-Art of Transonic Axial Compressors

Transonic axial flow compressors are today widely used in aircraft engines to obtain maximum pressure ratios per single-stage. High stage pressure ratios are important because they make it possible to reduce the engine weight and size and, therefore, investment and operational costs. Performance of transonic compressors has today reached a high level but engine manufacturers are oriented towards increasing it further. A small increment in efficiency, for instance, can result in huge savings in fuel costs and determine a key factor for product success. Another important target is the improvement of rotor stability towards near stall conditions, resulting in a wider working range. Important analytical and experimental researches in the field of transonic compressors were carried out since 1960s (e.g. Chen et al., 1991; Epstein, 1977; Freeman & Cumpsty, 1992; Konig et al., 1996; Miller et al., 1961; Wennerstrom & Puterbaugh, 1984). A considerable contribution for the new developments and designs was the progress made in optical measurement techniques and computational methods, leading to a deeper understanding of the loss mechanisms of supersonic relative flow in compressors (e.g. Calvert & Stapleton, 1994; Hah & Reid, 1992; Ning & Xu, 2001; Puterbaugh et al., 1997; Strazisar, 1985; Weyer & Dunker, 1978). Fig. 1 shows the low pressure and high pressure compressors of the EJ200 engine as examples for highly loaded, high performance transonic rotors of an aero engine. A closer look at the current trend in design parameters for axial flow transonic compressors shows that, especially in civil aircraft engines, the relative flow tip Mach number of the rotor is limited to maintain high efficiencies. A typical value for the rotor inlet relative flow at the tip is Mach ≈ 1.3. The continuous progress of aerodynamics has been focused to the increase in efficiency and pressure ratio and to the improvement in off-design behaviour at roughly the same level of the inlet relative Mach number. Today’s high efficiency transonic axial flow compressors give a total pressure ratio in the order of 1.7-1.8, realized by combining high rotor speeds (tip speed in the order of 500 m/s) and high stage loadings (2Δh/u2 in the order of 1.0). The rotor blade aspect ratio parameter showed a general trend towards lower values during past decades, with a current asymptotic value of 1.2 (Broichhausen & Ziegler, 2005). The flow field that develops inside a transonic compressor rotor is extremely complex and presents many challenges to compressor designers, who have to deal with several and concurring flow features such as shock waves, intense secondary flows, shock/boundary layer interaction, etc., inducing energy losses and efficiency reduction (Calvert et al., 2003; Cumpsty, 1989; Denton & Xu, 1999; Law & Wadia, 1993; Sun et al., 2007). Interacting with secondary flows, shock waves concur in development of blockage (Suder, 1998), in corner

On the Use of Synthetic Jets in Transonic Compressors

Several fluidic-based methods have been investigated by compressor designers with the aim of improving the aerodynamic behavior of compressor blades. Recently, synthetic jets have been proposed as a promising fluidic-based solution for reducing flow separation inside turbomachines. With the characteristic of zero net mass flux and non-zero momentum flux, synthetic jets could be very effective because at the suction part of the cycle the low-momentum fluid is sucked into the device, whereas in the blowing part a high momentum jet accelerates it. However, to the authors’ knowledge, the use of synthetic jets has never been experimented in transonic compressors. In these machines, synthetic jets could have a great impact on the thickness and stability of the blade suction side boundary layer arising after the interaction with the shock. With the aim to investigate that, a rake of 600 Hz sinusoidal wave form synthetic jets (with no radial velocity component) positioned into a thin radial slot on the blade suction side of the NASA Rotor 37 has been modeled and simulated. Simulations have been carried out using a validated 3D CFD (Computational Fluid Dynamics) unsteady model.The study gave a clear understanding of positive and negative impacts of the simulated device on the local flow field during the entire jets working cycle. The shock structure and boundary layer resulted significantly affected in the proximity of the synthetic jets, with different effects depending on slot location and jet channel angle. The local perturbation then affected the overall flowfield with impacts on the overall rotor performances. Despite the reported cases were selected for the purpose of the analysis only, and not for the aerodynamic enhancement of the rotor, slight benefits in terms of overall performance were observed for almost all the cases.Copyright

Advances in Aerodynamic Design of Gas Turbines Compressors

Aerodynamic design techniques of gas turbine compressors have dramatically changed in the last years. While traditional 1D and 2D design procedures are consolidated for preliminary calculations, emerging techniques have been developed and are being used almost routinely within industries and academia. The compressor design still remains a very complex and multidisciplinary task, where aerothermodynamic issues, traditionally considered prevalent, now become part of a more general design approach, where aeromechanical, technological, structural, noise-related concerns and many other matters have to be taken into account simultaneously, thus leading to a very challenging problem for designers. To this respect, designer experience still plays a decisive role; however, the complexity evidenced above claims for a more structured and organized way of handling the problem, where mathematical and statistical tools are implemented and used as a decisive support in the decision making process. Nowadays, interesting and alternative options are in fact available for compressor 3D design, such as new blade shapes for improved on-off design efficiency, endwall contouring and casing treatments for enhanced stall margin and many others. For this reason, while experimental activity remains decisive for ultimate assessment of design choices, numerical design optimization techniques, along with Computational Fluid Dynamics (CFD) are assuming more and more importance for the detailed design and concrete evaluation of options. In this chapter, a contribution to illustrating the roadmap for modern compressor design technique development, as well as an updated picture of available procedures, is given. Starting from the precise formulation of the design problem and the choice of the so-called design or nominal condition, both conventional and advanced techniques will be presented and discussed.