As bypass-ratio in modern aero engines is continuously increasing over the last decades, the radial offset between low pressure compressor (LPC) and high pressure compressor (HPC), which needs to be overcome by the connecting s-shaped intermediate compressor duct (ICD), is getting higher. Due to performance and weight saving aspects the design of shorter and therefore more aggressive ducts has become an important research topic. In this paper an already aggressive design (with respect to current aero engines) of an ICD with integrated outlet guide vane (OGV) is used as a baseline for an aerodynamic optimization. The aim is to shorten the duct even further while maintaining it separation free. The optimization is broken down into two steps. In the first optimization-step the baseline design is shortened to a feasible extent while keeping weak aerodynamic restrictions. The resulting highly aggressive duct (intermediate design), which is shortened by 19 % in axial length with respect to the baseline, shows separation tendencies of low momentum fluid in the strut/hub region. For the second step, the length of the optimized duct design is frozen. By implementing new design features in the process of the optimizer, this optimization-step aims to eliminate separation and to reduce separation tendencies caused by the aggressive shortening. In particular, these features are: a nonaxisymmetric endwall contouring and parametrization of the strut and the OGV to allow for changes in lift and turning in both blade designs. By comparison of the three designs: Baseline, intermediate (separating flow) and final design, it can be shown, that it is possible to decrease length of the already aggressive baseline design even further, when adding a nonaxisymmetric endwall contouring and changes in blade shape of the strut and OGV. Flow separation can be eliminated while losses are kept low. With a more aggressive and therefore shorter duct the engine length and weight can be reduced. This in turn leads to lighter aircrafts, less fuel consumption and lower CO2 and NOx emissions.

A rotor blisk of a supersonic space turbine has previously been designed to allow for free flutter to occur in an air test rig (Groth Mårtensson, and Edin, 2010, “Experimental and Computational Fluid Dynamics Based Determination of Flutter Limits in Supersonic Space Turbines,” 132(1), p. 011010). Flutter occurred at several operating conditions, and the flutter boundary for the test turbine was established. In this paper the rotor blisk is redesigned in order to inhibit flutter. The design strategy chosen is to introduce a mistuning concept. Based on aeroelastic analyses using a reduced order model a criterion for the required level of mistuning is established in order to stabilize the lower system modes. Proposals in literature suggest and analyze mistuning by varying blade mode frequencies in random patterns or by modifying blades in an odd-even pattern. Here a modification of sectors of the blisk is introduced in order to bring a sufficient split of the system mode frequencies. To verify that the redesigned blisk efficiently could inhibit flutter, an experiment similar to that in the work of Groth et al. is performed with the mistuned rotor blisk. By running the redesigned blisk at operating conditions deep into the unstable region of the tuned blisk, it is demonstrated that a relative low level of mistuning is sufficient to eliminate rotor flutter.

A finite volume method for blade flutter analyses, using moving grids is presented and partly validated. The method which solves the unsteady three-dimensional Euler equations is formulated in the four-dimensional time-space domain. An algebraic grid generation technique based on transfinite interpolation is used to move and deform the grid to conform to the blade motion. Fluxes are calculated using a third-order upwind-biased scheme. For time marching both an explicit three-stage Runge-Kutta scheme and a Crank-Nicolson scheme is used. Internal and external flows are calculated using the present method. Calculated results agree well with the corresponding experiments and with results obtained using other methods.

A rotor blisk of a supersonic space turbine has previously been designed to allow for free flutter to occur in an air test rig [1]. Flutter occurred at several operating conditions and the flutter boundary for the test turbine was established. In this paper the rotor blisk is redesigned in order to inhibit flutter. The design strategy chosen is to introduce a mistuning concept. Based on aeroelastic analyses using a reduced order model (ROM) a criterion for the required level of mistuning is established in order to stabilize the lower system modes. Proposals in literature suggest and analyze mistuning by varying blade mode frequencies in random patterns or by modifying blades in an odd-even pattern. Here a modification of sectors of the blisk is introduced in order to bring a sufficient split of the system mode frequencies. To verify that the redesigned blisk efficiently could inhibit flutter an experiment similar to that in [1] is performed with the mistuned rotor blisk. By running the redesigned blisk at operating conditions deep in to the unstable region of the tuned blisk it is demonstrated that a relative low level of mistuning is sufficient to eliminate rotor flutter.

Turbines operating at high pressure in high velocity flow are susceptible to flutter. As reduced frequencies become sufficiently low, negative aerodynamic damping will be found in some modes. Ensuring that the total system damping is positive over the entire turbine operating envelope for all modes is of utmost importance in any design since flutter in a turbine often causes blade failures. This is in contrast to the normal engineering approach, which is to require a positive aerodynamic damping. A unique test campaign with a 1.5 stage supersonic space turbine has been performed. The turbine was operated at simulated running conditions over a large operating envelope in order to map out flutter limits. During the test, flutter was intentionally triggered at seven different operating conditions. Unique data have been obtained during the test that supports validation of design tools and enables better understanding of flutter in this type of turbine. Based on the data the flutter boundary for the turbine could be established. Using computational fluid dynamics (CFD) tools flutter was predicted at all operating points where the flutter limit was crossed. Both in predictions and as evidenced in test the two nodal diameter backward traveling mode was the most unstable. In addition to this predicted values of aerodynamic damping at flutter agreed well with damping estimated from measured amplitude growth.