Large Scale Turbine Rig Research Articles

Investigation of the Interaction Between Tip Leakage and Main Annulus Flow in the Large Scale Turbine Rig: Comparison of Different Rotor Tip Geometries

Abstract Shroudless rotor blades are state-of-the-art in modern high pressure turbines. Tip leakage flow has a crucial impact on turbine efficiency. Specific blade tip designs are a key factor to handle tip leakage losses by controlling tip leakage flow and its re-entry into the rotor passage. Comparative measurements of a cavity squealer type tip and a notch type tip have been conducted at the Large Scale Turbine Rig at Technical University of Darmstadt. The test rig has been operated at the blade tips design point. Experimental data have been acquired at rotor inlet and outlet as well as within the rotor passage. For cavity squealer tips, a tip leakage vortex develops at the suction side as the tip leakage flow is rolled-up and further mixed with main annulus flow. The tip leakage vortex determines the blockage of main annulus flow at the blade tip. The design of the suction side of the notch tip benefits a jet-like re-entry of tip leakage flow into the passage. Results are a tip leakage vortex system with smaller sized vortices and a more homogeneous mass flow redistribution in the outer annulus of the rotor. The zone of affected main annulus flow at the blade tip increases through the dominant tip leakage jet.

Experimental Investigation of Rotor Tip Film Cooling at an Axial Turbine with Swirling Inflow Using Pressure Sensitive Paint

Rotor tip film cooling is investigated at the Large Scale Turbine Rig, which is a 1.5-stage axial turbine rig operating at low speeds. Using pressure sensitive paint, the film cooling effectiveness η at a squealer-type blade tip with cylindrical pressure-side film cooling holes is obtained. The effect of turbine inlet swirl on η is examined in comparison to an axial inflow baseline case. Coolant-to-mainstream injection ratios are varied between 0.45% and 1.74% for an engine-realistic coolant-to-mainstream density ratio of 1.5. It is shown that inlet swirl causes a reduction in η for low injection ratios by up to 26%, with the trailing edge being especially susceptible to swirl. For injection ratios greater than 0.93%, however, η is increased by up to 11% for swirling inflow, while for axial inflow a further increase in coolant injection does not transfer into a gain in η .

The Influence of Combustor Swirl on Pressure Losses and the Propagation of Coolant Flows at the Large Scale Turbine Rig (LSTR): Experimental and Numerical Investigation

The aerothermal interaction of the combustor exit flow on the first vane row has been examined at the Large Scale Turbine Rig (LSTR) at Technische Universität Darmstadt (Darmstadt, Germany). A baseline configuration of axial inflow and a variation of swirling combustor inflow have been studied. The nozzle guide vane (NGV) featured endwall cooling, airfoil film cooling and a trailing edge slot ejection as well as NGV-rotor wheel space purge flow. CO2 is injected for coolant flow tracing. The results are compared to five hole probe (5HP) measurements. The experiments for the baseline configuration are accompanied by numerical simulations using a passive scalar tracking method to validate the results and study the propagation of the coolant flow. The endwall coolant injection is detected to influence the pressure losses in the NGV. It has an impact on the Trailing Edge (TE) coolant ejection as well. For swirling combustor inflow, increased NGV pressure losses and increased mixing of Rear Inner Discharge Nozzle (RIDN) coolant and main flow is observed. An influence of the clocking position of the swirler to the vane is detected. Additional losses within the NGV row can be assigned to the swirler by means of flow tracing.

Influence of Combustor Swirl on Endwall Heat Transfer and Film Cooling Effectiveness at the Large Scale Turbine Rig

At the large scale turbine rig (LSTR) at Technische Universität Darmstadt, Darmstadt, Germany, the aerothermal interaction of combustor exit flow conditions on the subsequent turbine stage is examined. The rig resembles a high pressure turbine and is scaled to low Mach numbers. A baseline configuration with an axial inflow and a swirling inflow representative for a lean combustor is modeled by swirl generators, whose clocking position toward the nozzle guide vane (NGV) leading edge can be varied. A staggered double-row of cylindrical film cooling holes on the endwall is examined. The effect of swirling inflow on heat transfer and film cooling effectiveness is studied, while the coolant mass flux rate is varied. Nusselt numbers are calculated using infrared thermography and the auxiliary wall method. Boundary layer, turbulence, and five-hole probe measurements as well as numerical simulations complement the examination. The results for swirling inflow show a decrease of film cooling effectiveness of up to 35% and an increase of Nusselt numbers of 10–20% in comparison to the baseline case for low coolant mass flux rates. For higher coolant injection, the heat transfer is on a similar level as the baseline. The differences vary depending on the clocking position. The turbulence intensity is increased to 30% for swirling inflow.

Robustness of a Turbine Endwall Film Cooling Design to Swirling Combustor Inflow

Modern low-emission combustor concepts, such as lean combustion, pose challenges to turbine designers, among them a flatter temperature profile with increased thermal load to the endwalls and a three-dimensional inflow. The complexity of the flowfield make numerical investigations and rules for turbine thermal design under swirling inflow a delicate task. An experimental investigation is conducted to evaluate a film-cooling design for a stator endwall in respect to its robustness to swirling combustor inflow and its limitations. The experiments were conducted at the Large Scale Turbine Rig at the Technische Universität Darmstadt, operating at low-Mach-number, cold-flow conditions. A swirler module is used to generate an engine-realistic whirl-angle distribution with averaged values of 14 deg in the upper main annulus and 11 deg in the lower annulus. Spatially resolved Nusselt numbers and film-cooling effectiveness data are acquired by infrared thermography with the auxiliary wall method for swirling inflow. The comparison to the axial inflow baseline case highlights the robustness of the design. Swirling inflow increases Nusselt numbers until at 2.9% injection the level remains similar to the baseline, whereas film-cooling effectiveness is reduced by 30%.

Experimental Analysis of the Interaction Between Rim Seal and Main Annulus Flow in a Low Pressure Two Stage Axial Turbine

The spoiling effects of rim seal flow are studied at the Large Scale Turbine Rig (LSTR) at Technische Universität Darmstadt. Detailed flow field measurements and efficiency measurements were performed for various ingress and egress setups and will be presented in this paper. Efficiency measurements show an efficiency decrease as the rim seal mass flow is increased. Five hole probe measurements upstream and downstream of the second stator row show that an increasing rim seal mass flow leads to an increased pressure loss across the stator, to altered incidence angles and to an intensification of secondary flow structures within the lower 50% span. Static pressure taps at the stator profile primarily show altered aerodynamic loading with increased rim seal air. In addition, the end wall profile pressure was measured at the stator 2 hub. It can be seen that seal air injection causes increased pressure fluctuations on the platform. Temperature measurements with a temperature difference between rim seal and main annulus flow show that rim seal air primarily enters the passage vortex.