Impact of Individual High-Pressure Turbine Rotor Purge Flows on Turbine Center Frame Aerodynamics

This paper presents an experimental study of the impact of individual high-pressure turbine purge flows on the main flow in a downstream turbine center frame duct. Measurements were carried out in a product-representative one and a half stage turbine test setup, installed in the Transonic Test Turbine Facility at Graz University of Technology. The rig allows testing at engine-relevant flow conditions, matching Mach, Reynolds, and Strouhal number at the inlet of the turbine center frame. The reference case features four purge flows differing in flow rate, pressure, and temperature, injected through the hub and tip, forward and aft cavities of the high-pressure turbine rotor. To investigate the impact of each individual cooling flow on the flow evolution in the turbine center frame, the different purge flows were switched off one-by-one while holding the other three purge flow conditions. In total, this approach led to six different test conditions when including the reference case and the case without any purge flow ejection. Detailed measurements were carried out at the turbine center frame duct inlet and outlet for all six conditions and the post-processed results show that switching off one of the rotor case purge flows leads to an improved duct performance. In contrast, the duct exit flow is dominated by high pressure loss regions if the forward rotor hub purge flow is turned off. Without the aft rotor hub purge flow, a reduction in duct pressure loss is determined. The purge flows from the rotor aft cavities are demonstrated to play a particularly important role for the turbine center frame aerodynamic performance. In summary, this paper provides a first-time assessment of the impact of four different purge flows on the flow field and loss generation mechanisms in a state-of-the-art turbine center frame configuration. The outcomes of this work indicate that a high-pressure turbine purge flow reduction generally benefits turbine center frame performance. However, the forward rotor hub purge flow actually stabilizes the flow in the turbine center frame duct and reducing this purge flow can penalize turbine center frame performance. These particular high-pressure turbine/turbine center frame interactions should be taken into account whenever high-pressure turbine purge flow reductions are pursued.

Effect of Active Modulation of Through-Casing Coolant Injection on Turbine Efficiency

This paper presents a computational investigation into the impact of cooling air injected through the stationary over-tip turbine casing on overall turbine efficiency. The high work axial flow turbine is representative of the high pressure turbine of a civil aviation turbofan engine. The effect of active modulation of the cooling air is assessed, as well as that of the injection locations. The influence of the through-casing coolant injection on the turbine blade over-tip leakage flow and the associated secondary flow features are examined. Transient (unsteady) sliding mesh simulations of a one turbine stage rotor-stator domain are performed using periodic boundary conditions. Cooling air configurations with a constant total pressure air supply, constant mass flow rate and actively controlled total pressure supply are assessed for a single geometric arrangement of cooling holes. The effects of both the mass flow rate of cooling air and the location of its injection relative to the turbine rotor blade are examined. The results show that all of the assessed cooling configurations provided a benefit to turbine row efficiency of between 0.2 and 0.4 percentage points. The passive and constant mass flow rate configurations reduced the over-tip leakage flow, but did so in an inefficient manner, with decreasing efficiency observed with increasing injection mass flow rate beyond 0.6% of the mainstream flow, despite the over-tip leakage mass flow rate continuing to reduce. By contrast, the active total pressure controlled injection provided a more efficient manner of controlling this leakage flow, as it permitted a redistribution of cooling air, allowing it to be applied in the regions close to the suction side of the blade tip which more directly reduced over-tip leakage flow rates and hence improved efficiency. Cooling air injected close to the pressure side of the rotor blade was less effective at controlling the leakage flow, and was associated with increased aerodynamic loss in the passage vortex.

High-Fidelity Simulations of a Linear HPT Vane Cascade Subject to Varying Inlet Turbulence

The effect of inflow turbulence intensity and turbulence length scales have been studied for a linear high-pressure turbine vane cascade at Reis = 590,000 and Mis = 0.93, using highly resolved compressible large-eddy simulations employing the WALE turbulence model. The turbulence intensity was varied between 6% and 20% while values of the turbulence length scales were prescribed between 5% and 20% of axial chord. The analysis focused on characterizing the inlet turbulence and quantifying the effect of the inlet turbulence variations on the vane boundary layers, in particular on the heat flux to the blade. The transition location on the suction side of the vane was found to be highly sensitive to both turbulence intensity and length scale, with the case with turbulence intensity 20% and 20% length scale showing by far the earliest onset of transition and much higher levels of heat flux over the entire vane. It was also found that the transition process was highly intermittent and local, with spanwise parts of the suction side surface of the vane remaining laminar all the way to the trailing edge even for high turbulence intensity cases.

Influence of Gas-to-Wall Temperature Ratio on By-Pass Transition

An investigation of thermal effects on bypass transition was conducted on the highly-loaded turbine guide vane LS89 in the short-duration isentropic Compression Tube (CT-2) facility at the von Karman Institute for Fluid Dynamics (VKI). Measurements from high response surface-mounted thin films coupled with analog circuits provided the time-resolved wall heat flux history whereas pneumatic probes, differential pressure transducers and thermocouples allowed the accurate definition of the inlet and outlet flow conditions. The gas-to-wall temperature ratio, ranging from 1.11 to 1.55, was varied by changing the inlet total temperature. The isentropic exit Mach number ranged from 0.90 to 1.00 and the global freestream turbulence intensity value was set at 0.8, 3.9 and 5.3%. The isentropic exit Reynolds number was kept at 106. The onset of transition was tracked through the wall heat flux signal fluctuations. Within the present operating conditions, no significant effect of the gas/wall temperature ratio was put in evidence. At the present (design) transonic exit conditions, the local free-stream pressure gradient appears to remain the main driver of the onset of transition. A wider range of operating conditions must be considered to draw final conclusions.

Numerical Efforts of Aerodynamic Re-Design in a Single-Stage Transonic Axial Compressor: Part 1 — Stator Design

In many aerodynamic design parameters for the axial-flow compressor, three variables of tailored blading, blade lean and sweep were considered in the re-design efforts of a transonic single stage which had been designed in 1960’s NASA public domains. As Part 1, the re-design was limited to the stator vane only. For the original MCA (Multiple Circular Arc) blading, which had been applied at all radii, the CDA (Controlled Diffusion Airfoil) blading was introduced at midspan as the first variant, and the endwalls of hub and casing (or tip) were replaced with the DCA (Double Circular Arc) blading for the second variant. Aerodynamic performance was predicted through a series of CFD analysis at design speed, and the best aerodynamic improvement, in terms of pressure ratio/efficiency and operability, was found in the first variant of tailored blading. It was selected as a baseline for the next design efforts with blade lean, sweep and both combined. Among 12 variants, a case of positively and mildly leaned blades was found the most attractive one, relative to the original design, providing benefits of an 1.0% increase of pressure ratio at design flow, an 1.7% increase of efficiency at design flow, a 10.5% increase of the surge margin and a 32.3% increase of the choke margin.

Effect of Blade Aspiration Slot Configuration on the Aerodynamic Performance of a Highly Loaded Aspirated Compressor Cascade

As a promising active flow control method, boundary layer suction (BLS) can be used to enhance the aerodynamic performance of the highly-loaded compressor effectively, and due to this reason, extensive studies have been carried out on it. However, contrast to those abundant studies focusing on the flow control effects of BLS, little attention has been paid on the design method of the aspiration flow path. This work presents a 3-D steady numerical simulation on a highly-loaded aspirated compressor cascade. The aspiration slot is implemented at its best location based on the previous experimental studies and the aspiration flow rate is fix to 1.5% of the inlet massflow. The plenum configuration follows the blade shape and remains unchanged. One-side-aspiration manner is adopted to simplify the aspiration devices. Two critical geometry parameters, slot angle and slot width, are varied to study the effects of blade aspiration slot configuration on the cascade loss, radial distribution of the aspiration flow rate and inner flow structures within the aspiration flow path. Results show that the slot configuration does affect the cascade performance. In comparison with the throughflow performance, it is especially true once the flow loss caused by the aspiration flow path is also taken into account, and higher flow loss will be generated within the aspiration flow path if an inappropriate scheme is adopted. In the present investigation, apart from the cases with larger negative slot angle, a wider slot is more preferable to a narrower one, since it could enhance the aspiration capacity near the endwall regions and lower the dissipation loss within the aspiration flow path. In terms of the slot angle, a larger negative value, i.e., the slot direction more aligned with the incoming flow, is not beneficial to improve the throughflow performance, while concerning the flow loss yield by the aspiration flow path, a proper negative slot angle is always optimal.

Next Generation Turbine Testing at DLR

Axial turbines for aircraft engines and power plants have reached a very high level of development. Further improvements, in particular in terms of higher efficiency and reduced number of blades and stages, resulting in higher loads, are possible, but can only be achieved through a better understanding of the flow parameters and a closer connection between experiment and numerical design and simulation. An analysis of future demands from the industry and existing turbine research rigs shows that there appears a need for a powerful turbine test rig for aerodynamic experiments. This paper deals with the development and built up of a new so called Next Generation Turbine Test Facility (NG-Turb) at the German Aerospace Center (DLR) in Göttingen. The NG-Turb is a closed-circuit, continuously running facility for aerodynamic turbine investigations, allowing independent variation of engine relevant Mach and Reynolds numbers. The flow medium (dry air) is driven by a 4-stage radial gear compressor with a high pressure ratio and a wide inlet volume flow range. In a first stage the NG-Turb test section will allow investigations on single shaft turbines up to 2½ stages. In a further expansion stage the NG-Turb will be equipped with a second independent shaft system, then enabling experiments with configurations of high and low (or intermediate) pressure turbines and in particular offering the possibility for investigations at counter rotating turbines. Secondary air for cooling investigations can be provided by auxiliary screw compressors. Mass flow through the Turbine is determined redundantly with an uncertainty of about ±0.3%, using well calibrated Venturi nozzles upstream and downstream of the test section. The operation concept and main design features of the NG-Turb will be described and an overview of the applied standard measurement and data acquisition technics capturing efficiency, traverse data etc. will be given. Thermodynamic cycle calculations have been performed in order to simulate the flow circuit of the NG-Turb and to access whether turbine operating points can be driven within the performance map of the compressor system. Finally the calibration procedure for the Venturi nozzles, which has been conducted during the commissioning phase of the NG-Turb by applying a special calibration test section, is explained and some results will be shown.

Loss Generation in Transonic Turbine Blading

The measured loss characteristic in a high-speed cascade tunnel of two turbine blades of different designs showed distinctly different trend with exit Mach number ranging from 0.8 to 1.4. Assessments using steady RANS computation of the flow in the two turbine blades, complemented with control volume analyses and loss modelling, elucidate why the measured loss characteristic looks the way it is. The loss model categorizes the total loss in terms of boundary layer loss, trailing edge loss and shock loss; it yields results in good agreement with the experimental data as well as steady RANS computed results. Thus RANS is an adequate tool for determining the loss variations with exit isentropic Mach number and the loss model serves as an effective tool to interpret both the computational and experimental data. The measured loss plateau in Blade 1 for exit Mach number of 1 to 1.4 is due to a balance between a decrease of blade surface boundary layer loss and an increase in the attendant shock loss with Mach number; this plateau is absent in Blade 2 due to a greater rate in shock loss increase than the corresponding decrease in boundary layer loss. For exit Mach number from 0.85 to 1, the higher loss associated with shock system in Blade 1 is due to the larger divergent angle downstream of the throat than that in Blade 2. However when exit Mach number is between 1.00 and 1.30, Blade 2 has higher shock loss. For exit Mach number above around 1.4, the shock loss for the two blades is similar as the flow downstream of the throat is completely supersonic. In the transonic to supersonic flow regime, the turbine design can be tailored to yield a shock pattern the loss of which can be mitigated in near equal amount of that from the boundary layer with increasing exit Mach number, hence yielding a loss plateau in transonic-supersonic regime.

Experimental and Numerical Study of Honeycomb Tip on Suppressing Tip Leakage Flow in Turbine Cascade

In this paper, the effect of a novel honeycomb tip on suppressing tip leakage flow in a highly-loaded turbine cascade has been experimentally and numerically studied. The research focuses on the mechanisms of honeycomb tip on suppressing tip leakage flow and affecting the secondary flow in the cascade, as well as the influences of different clearance heights on leakage flow characteristics. In addition, two kinds of local honeycomb tip structures are pro-posed to explore the positive effect on suppressing leakage flow in simpler tip honeycomb structures. Based on the experimental and numerical results, the physical processes of tip leakage flow and its interaction with main flow are analyzed, the following conclusions can be obtained. Honeycomb tip rolls up a number of small vortices and radial jets in regular hexagonal honeycomb cavities, increasing the flow resistance in the clearance and reducing the velocity of leakage flow. As a result, the structure of honeycomb tip not only suppresses the leakage flow effectively, but also has positive effect on reducing the associated losses in cascade by reducing the strength of leakage vortex. Compare to the flat tip cascade at 1%H gap height, the relative leakage flow in honeycomb tip cascade reduces from 3.05% to 2.73%, and the loss at exit section is also decreased by 10.63%. With the increase of the gap height, the tip leakage flow and loss have variations of direct proportion with it, but their growth rates in the honeycomb tip cascade are smaller. Consider the abradable property of the honeycomb seal, a smaller gap height is allowed in the cascade with honeycomb tip, and that means honeycomb tip has better effect on suppressing leakage flow. Two various local honeycomb tip structures has also been discussed. It shows that local raised honeycomb tip has better suppressing leakage flow effect than honeycomb tip, while local concave honeycomb tip has no more effect than honeycomb tip. Compare to flat tip cascade, the leakage flow in honeycomb tip cascade, local concave tip cascade and local raised honeycomb tip cascade decrease by nearly 17.33%, 15.51% and 30.86% respectively, the losses at exit section is reduced by 13.38%, 12% and 28.17% respectively.

Experimental Investigations on the Efficiency of Active Flow Control in a Compressor Cascade With Periodic Non-Steady Outflow Conditions

We present results of experiments on a periodically unsteady compressor stator flow of the type which would be expected in consequence of pulsed combustion. A Reynolds number of Re = 600000 was used for the investigations. The experiments were conducted on the two-dimensional low-speed compressor testing facility in Berlin. A choking device downstream the trailing edges induced a periodic non-steady outflow condition to each stator vane which simulated the impact of a pressure gaining combuster downstream from the last stator. The Strouhal number of the periodic disturbance was Sr = 0.03 w.r.t. the stator chord length. Due to the periodic non-steady outflow condition, the flow-field suffers from periodic flow separation phenomena, which were managed by means of active flow control. In our case, active control of the corner separation was applied using fluidic actuators based on the principle of fluidic amplification. The flow separation on the centre region of the stator blade was suppressed by means of a fluidic blade actuator leading to an overall time-averaged loss reduction of 11.5%, increasing the static pressure recovery by 6.8% while operating in the non-steady regime. Pressure measurements on the stator blade and the wake as well as PIV data proved the beneficial effect of the active flow control application to the flow field and the improvement of the compressor characteristics. The actuation efficiency was evaluated by two figures of merit introduced in this contribution.