In modern gas turbines hot section components, the over tip leakage (OTL) flow that occurs between the stationary casing and rotating tip of a shroudless HP turbine is still a considerable source of loss of performance. The principal means of reducing this loss have been to minimise the tip gap and/or to apply a rotating shroud to the rotor. Tip clearance control systems continue to improve, but a practical limit on tip gap remains. Winglets have been identified by a number of researchers as having potential, but none have yet to enter commercial service. Harvey & Ramsden [1] analysed a novel design of one, which indicated that it could significantly reduce OTL loss. This paper presents the design of such a winglet as applied to the rotor blade of a research high pressure turbine carried out as part of the ANTLE (Advanced Near Term Low Emissions) technology demonstrator programme. The use of Computational Fluid Dynamics (CFD) calculations in the design process is discussed. In particular, the use of coarse meshes and idealised geometries, for computational speed, did involve some compromise with accuracy. Results from high speed model rig testing of this research turbine are presented. The turbine efficiency was measured for three different tip gaps over a range of conditions. In addition detailed measurements of the flow field were taken, principally exit area traverses and rotor surface static pressures. These experimental results are very encouraging and show a high potential for further development. Part II of this paper presents a post-test re-analysis of the rig results using the state of the art Rolls-Royce in-house CFD code HYDRA, good agreement being found between the two.

For the small to medium thrust range of modern aero engines, highly loaded single stage HP turbines facilitate an attractive alternative to a more conventional 2-stage HPT architecture. Whereas the potential benefits of reductions in component length and part count, hence, in weight and cost do motivate their application, the related risks are in maintaining associated losses of supersonic flows at low values as well as managing the interaction losses between HPT and the downstream sub-component to arrive at competitive levels of component efficiencies. This paper focuses on fundamental aerodynamic concept studies and related cascade experiments in support of a future highly loaded high-pressure turbine architecture. Starting with some general remarks on low-loss supersonic aerodynamic concepts for high-pressure turbines, results from development efforts towards 2D airfoil concepts viable for high-pressure turbine airfoils are shown. In particular, CFD based design approaches are compared against experimental data taken at DLR Go¨ttingen in un-cooled cascade tests and at engine representative levels of Mach and Reynolds numbers. For the airfoils investigated, it turns out that there is indeed a supersonic Mach number range were loss levels are comparable to high Mach number subsonic values, thereby enabling a competitive aerodynamic design concept for a 3D high-pressure turbine stage.

This paper introduces a new 2-stage high-pressure turbine rig for aerodynamic investigations. It is operated by DLR Göttingen (Germany) and installed in DLR’s new testing facility NG-Turb. The rig’s geometrical size as well as the non-dimensional parameters are comparable to a modern engine in the small to medium thrust range. The turbine rig closely resembles engine hardware and features all relevant blade and vane cooling as well as secondary air-system flows. The effect of variations of each individual flow and different tip clearances on overall turbine efficiency will be studied. While the first part of the testing program will be based on uniform inlet conditions the second part will be run with a combustor simulator, which is based on electrical heaters and delivers a flow field similar to a rich-burn combustor. In order to find the optimum relative position for maximum turbine efficiency the combustor simulator can be rotated relative to the HPT inlet (clocking). For the same reasons the stators can also be clocked. The paper gives a brief overview of the testing facility and from there on focuses on the HPT rig features such as aerodynamic design, cooling and sealing flows. The aerodynamic optimisation of the stator vanes and shroudless rotor blades will be outlined. Further topics are the aerodynamic design of the combustor simulator, a comparison with engine combustors as well as the implementation in the rig. The paper also describes the rig instrumentation in the stationary and rotating system which most importantly focuses on measurements of efficiency and capturing of traverse data. The topic of blade and vane manufacturing via direct metal laser sintering will be briefly covered. The discussion of test results and comparison with numerical simulations will be the subject of a follow-up paper.

Very-low NOx combustion concepts require a high swirl number of the flow in the combustion chamber to allow for lean burn combustion. This article deals with the influence of the resulting combustor exit swirl on the turbine aerodynamics of the first stage. This investigation is based on numerical simulations. According to the literature research additional insight into combustor-turbine interaction is achieved by taking into account a fully two dimensional inlet boundary condition. Up to now published results on combustor-turbine interaction were mostly restricted to the inhomogeneous temperature distribution at the turbine inlet. The investigations are carried out on a real engine geometry — the E3E Core 3/2 — a research project of Rolls-Royce Deutschland on lean combustion. Calculations are conducted by means of the Rolls-Royce plc code Hydra. The swirled inlet boundary condition is further scaled to test rig conditions to check for the transferability between the test rig and the real engine geometry. The results show a significant impact of the inhomogeneous turbine inflow on the stage efficiency and the thermal load. The optimization potential due to the clocking position of the combustor swirl is analyzed. The impact on the secondary flow mechanisms is analyzed with a novel visualization technique. A frequency spectrum analysis is carried out to investigate the effects of the 2D inlet boundary condition on the rotor row.