The demand of further increased bypass ratio of aero engines will lead to low pressure turbines with larger diameters, which rotate at lower speed. Therefore, it is necessary to guide the flow leaving the high pressure turbine to the low pressure turbine at larger diameters minimizing the losses and providing an adequate flow at the low pressure (LP)-turbine inlet. Due to costs and weight, this intermediate turbine duct has to be as short as possible. This would lead to an aggressive (high diffusion) s-shaped duct geometry. It is possible to shorten the duct simply by reducing the length but the risk of separation is rising and losses increase. Another approach to shorten the duct and thus the engine length is to apply a so called integrated concept. These are novel concepts where the struts, mounted in the transition duct, replace the usually following LP-vane row. This configuration should replace the first LP-vane row from a front bearing engine architecture where the vane needs a big area to hold bearing services. That means the rotor is located directly downstream of the strut. This means that the struts have to provide the downstream blade row with undisturbed inflow with suitable flow angle and Mach number. Therefore, the (lifting) strut has a distinct three-dimensional design in the more downstream part, while in the more upstream part, it has to be cylindrical to be able to lead through supply lines. In spite of the longer chord compared with the base design, this struts have a thickness to chord ratio of 18%. To apply this concept, a compromise must be found between the number of struts (weight), vibration, noise, and occurring flow disturbances due to the secondary flows and losses. The struts and the outer duct wall have been designed by Industria de Turbopropulsores. The inner duct was kept the same as for the base line configuration (designed by Motoren und Turbinen Union). The aim of the design was to have similar duct outflow conditions (exit flow angle and radial mass flow distribution) as the base design with which it is compared in this paper. This base design consists of a single transonic high pressure (HP)-turbine stage, an aggressive s-shaped intermediate turbine duct, and a LP-vane row. Both designs used the same HP-turbine and were run in the continuously operating Transonic Test Turbine Facility at Graz University of Technology under the same engine representative inlet conditions. The flow field upstream and downstream the LP-vane and the strut, respectively, has been investigated by means of five hole probes. A rough estimation of the overall duct loss is given as well as the upper and lower weight reduction limit for the integrated concept.

The demand for a further increased bypass ratio of aero engines will lead to low pressure turbines with larger diameters, which rotate at a lower speed. Therefore, it is necessary to guide the flow leaving the high pressure turbine to the low pressure turbine at a larger diameter without any loss generating separation or flow disturbances. Due to costs and weight, this intermediate turbine duct (ITD) has to be as short as possible. This leads to an aggressive (high diffusion) and, furthermore, to a super-aggressive s-shaped duct geometry. In order to investigate the influence of the blade tip gap size on such a high diffusion duct flow a detailed test arrangement under engine representative conditions is necessary. Therefore, the continuously operating Transonic Test Turbine Facility (TTTF) at Graz University of Technology has been adapted: An super-aggressive intermediate duct is arranged downstream of a transonic high pressure (HP)-turbine stage providing an exit Mach number of about 0.6 and a swirl angle of –15 deg. A second low pressure (LP)-vane row is located at the end of the duct and represents the counter-rotating low pressure turbine at a larger diameter. A following deswirler and a diffuser are the connection to the exhaust casing of the facility. In order to determine the influence of the blade tip gap size on the flow through such a super-aggressive s-shaped turbine, duct measurements were conducted with two different tip gap sizes, a 1.5% span (0.8 mm) and a 2.4% span (1.3 mm). The aerodynamic design of the HP-turbine stage, ITD, LP-vane, and the de-swirler was done by MTU Aero engines. In 2007 at the ASME Turbo Expo, the influence of the rotor clearance size onto the flow through an aggressive ITD was presented. For the present investigation, this aggressive duct has been further shortened by 20% (super-aggressive ITD) so that the flow at the outer duct wall is fully separated. This paper shows the influence of the rotor tip clearance size on to this separation. The flow through this intermediate turbine duct was investigated by means of five-hole-probes, static pressure taps, boundary layer rakes, and oil flow visualization. The oil flow visualization showed the existence of vortical structures within the separation where they seem to be imposed by the upstream HP-vanes.

For this study, measurements were carried out in a one-and-a-half stage test turbine. In order to characterize the flow field and to obtain steady flow quantities, five-hole probes are used in the rig in several measurement planes. Although a very high resolution in circumferential and radial direction is possible with five-hole probes, only certain sectors of the flow field can be measured due to quite long measurement times. For this reason, total pressure rakes, which can be traversed 360 degrees in a circumferential direction, were built for this test configuration. During foregoing measurements, differences between the total pressure measured with the rakes and the total pressure measured with a five-hole probes were observed, especially in the measurement plane downstream of the rotor where the flow is highly unsteady. It was found that the measured value for total pressure is particularly influenced by the stochastic pressure fluctuations and turbulence and can therefore deviate significantly from the “real” value.

Modern low pressure turbines (LPT) are designed in order to fulfil a various number of requirements such as high endurance, low noise, high efficiency, low weight, and low fuel consumption. Regarding the reduction of the emitted noise, different designs of LPT exit guide vanes (aerodynamically and/or acoustically optimized) of the turbine exit casing (TEC) were tested, and their noise reduction capabilities and aerodynamic performance were evaluated. In particular, measurements of TEC-losses were performed, and differences in the losses were reported. Measurements were carried out in a one and a half stage subsonic turbine test facility at the engine relevant operating point approach. This work focuses on the study of the unsteady flow field downstream of an unshrouded LPT rotor. The influence on the upstream flow field of a TEC design including acoustically optimized vanes (inverse cut-off TEC) is investigated and compared with a second TEC configuration without vanes (Vaneless TEC), by means of fast response aerodynamic pressure probe (FRAPP) measurements. The second configuration served as a reference concerning the influence of turbine exit guide vanes (TEGVs) onto the upstream located LPT rotor. The interactions between the stator and rotor wakes, secondary flows, and the TEGVs potential effect are identified via modal decomposition according to the theory of Tyler and Sofrin. The main structures constituting the unsteady flow field are detected, and the role of the major interaction effects in the loss generation mechanism and in the acoustic emission is analyzed. This study based on the modal analysis of the unsteady flow field offers new insight into the main interaction mechanisms and their importance in the assessment of the aerodynamic and aeroelastic performance of modern LPT exit casings.