Unsteady Numerical Simulation of Shock Systems in Vaneless Counter-Rotating Turbine

A detailed unsteady numerical simulation has been carried out to investigate the shock systems in the high pressure (HP) turbine rotor and unsteady shock-wake interaction between coupled blade rows in a 1+1/2 counter-rotating turbine (VCRT). For the VCRT HP rotor, due to the convergent-divergent nozzle design, along almost all the span, fishtail shock systems appear after the trailing edge, where the pitch averaged relative Mach number is exceeding the value of 1.4 and up to 1.5 approximately (except the both endwalls). A group of pressure waves create from the suction surface after about 60% axial chord in the VCRT HP rotor, and those waves interact with the inner-extending shock (IES). IES first impinges on the next HP rotor suction surface and its echo wave is strong enough and cannot be neglected, then the echo wave interacts with the HP rotor wake. Strongly influenced by the HP rotor wake and LP rotor, the HP rotor outer-extending shock (OES) varies periodically when moving from one LP rotor leading edge to the next. In VCRT, the relative Mach numbers in front of IES and OES are not equal, and in front of IES, the maximum relative Mach number is more than 2.0, but in front of OES, the maximum relative Mach number is less than 1.9. Moreover, behind IES and OES, the flow is supersonic. Though the shocks are intensified in VCRT, the loss resulted in by the shocks is acceptable, and the HP rotor using convergent-divergent nozzle design can obtain major benefits.

Optimization of Inlet Ring Groove Arrangement for Suppression of Unstable Flow in a Centrifugal Impeller

An optimization of the inlet ring groove arrangement has been pursued in the present study for obtaining better impeller characteristics and a wider operation range at both small and large flow rates in a high specific speed type centrifugal impeller with inducer. The effects of the shape parameters with respect to the inlet ring groove on the impeller characteristic and the flow incidence were analyzed mainly based on numerical simulations, but also compared to the experimental results. At small flow rates, a significant improvement in the impeller characteristic is achieved due to reduction in the excessive-positive flow incidence by optimizing both location and width of the rear groove near the inducer tip throat. On the other hand, the impeller characteristic is improved at large flow rates by implementing the corner radius at the rear groove edge and by placing another front ring groove in the suction pipe. As a result, by the optimized configuration of the front and rear ring grooves, the unstable flow range of the test impeller can be reduced by about 50% without deterioration of the impeller characteristic even at the 125% flow rate.

Turbine Cascade Flow Control Using a “Wake Filling” Pulsed Plasma Actuator

Separation bubbles in high-camber blades under part-load conditions have been addressed via continuous and pulsed jets, and also via plasma actuators. Numerous passive techniques have been employed as well. In this type of blades, the laminar boundary layer cannot overcome the adverse pressure gradient arising along the suction side, resulting on a separation bubble. When separation is abated, a common explanation is that kinetic energy added to the laminar boundary layer speeds up its transition to turbulent. In the present study, a plasma actuator installed in the trailing edge (i.e. “wake filling configuration”) of a cascade blade is used to excite the flow in pulsed and continuous ways. The pulsed excitation can be directed to the frequencies of the large coherent structures (LCS) of the flow, as obtained via a hot-film anemometer, or to much higher frequencies present in the suction-side boundary layer, as given in the literature. It is found that pulsed frequencies much higher than that of LCS reduce losses and improve turning angles further than frequencies close to those of LCS. With the plasma actuator 50% on time, good loss abatement is obtained. Larger “on time” values yield improvements, but with decreasing returns. Continuous high-frequency activation results in the largest loss reduction, at increased power cost. The effectiveness of high frequencies may be due to separation abatement via boundary layer excitation into transition, or may simply be due to the creation of a favorable pressure gradient that averts separation as the actuator ejects fluid downstream. Both possibilities are discussed in light of the experimental evidence.

The Effect of Hot-Streaks on HP Vane Surface and Endwall Heat Transfer: An Experimental and Numerical Study

Pronounced non-uniformities in combustor exit flow temperature (hot-streaks), which arise because of discrete injection of fuel and dilution air jets within the combustor and because of end-wall cooling flows, affect both component life and aerodynamics. Because it is very difficult to quantitatively predict the affects of these temperature non-uniformities on the heat transfer rates, designers are forced to budget for hot-streaks in the cooling system design process. Consequently, components are designed for higher working temperatures than the mass-mean gas temperature, and this imposes a significant overall performance penalty. An inadequate cooling budget can lead to reduced component life. An improved understanding of hot-streak migration physics, or robust correlations based on reliable experimental data, would help designers minimise the overhead on cooling flow that is currently a necessity. A number of recent research projects sponsored by a range of industrial gas turbine and aero-engine manufacturers attest to the growing interest in hot-streak physics. This paper presents measurements of surface and end-wall heat transfer rate for an HP nozzle guide vane (NGV) operating as part of a full HP turbine stage in an annular transonic rotating turbine facility. Measurements were conducted with both uniform stage inlet temperature and with two non-uniform temperature profiles. The temperature profiles were non-dimensionally similar to profiles measured in an engine. A difference of one half of an NGV pitch in the circumferential (clocking) position of the hot-streak with respect to the NGV was used to investigate the affect of clocking on the vane surface and end-wall heat transfer rate. The vane surface pressure distributions, and the results of a flow-visualisation study, which are also given, are used to aid interpretation of the results. The results are compared to two-dimensional predictions conducted using two different boundary layer methods. Experiments were conducted in the Isentropic Light Piston Facility (ILPF) at QinetiQ Farnborough, a short duration engine-size turbine facility. Mach number, Reynolds number and gas-to-wall temperature ratios were correctly modelled. It is believed that the heat transfer measurements presented in this paper are the first of their kind.

Review and Understanding on Sweep in Axial Compressor Design

As a means of improving axial compressor performance, sweep technique has been investigated for over half a century and gained wide uses, in the past one decade. However, there is still diverse controversy about the roles of sweep in axial compressor design. In this paper, historical remarks about the sweep are presented firstly. Then, an understanding about the role of sweep is put forward. That is, the sweep is a degree of freedom (DOF) of blade design that emphasizes on matching the aerodynamic loading of every blade element along the whole span within the full operation range. The present understanding about the role of sweep may lead it a more sophisticated use.

Experimental Study of Boundary Layer Control Through Tip Injection on Straight and Swept Compressor Blades

Recent research on the use of three dimensional blade designs incorporating sweep and casing boundary layer control through tip injection have shown improved efficiency, stage loading and stable operating range Both methods are known to stabilize the flow at the tip trailing edge (Stall Inception Point) of the compressor blade under lower than design mass flow conditions and improve the compressor performance. This study, performed on a low speed axial flow fan, is aimed to improve the understanding and to assess the improved stall margin due to tip injection on swept and straight blades. Injection through flush mounted nozzles on the casing before the rotor blades is of practical importance because of the non-interference of the nozzles with the main flow. Benefits in terms of efficiency and operating range have already been established and swept blades were found to perform better than the baseline straight blades. The injection mechanism involves 12 symmetrically located flush mounted nozzles before the rotor blades in the casing. The number of injection nozzles used is an additional parameter of the study. Although there are manufacturing limitations in the air injection angle, two injection angles (10° and 30°) have been tried. The improvement in stall margin is more pronounced in the straight blades than in the swept blades and the 10° injection angle with 6 active injection holes gives the best performance in comparison to other combinations.

Some Studies on Low Solidity Vaned Diffusers of a Centrifugal Compressor Stage

Numerical simulation of impeller and low solidity vaned diffuser (LSD) of a centrifugal compressor stage is performed individually using CFX- BladeGen and BladeGenPlus codes. The tip mach number for the chosen study was 0.35. The same configuration was used for experimental investigation for a comparative study. The LSD vane is formed using standard NACA profile with marginal modification at trailing edge. The performance parameters obtained form numerical studies at the exit of impeller and the diffuser have been compared with the corresponding experimental data. These parameters are pressure ratio, polytropic efficiency and flow angle at the impeller exit where as the parameters those have been compared at the exit of diffuser are the static pressure recovery coefficient and the exit flow angle. In addition, the numerical prediction of the blade loading in terms of blade surface pressure distribution on LSD vane has been compared with the corresponding experimental results. Static pressure recovery coefficient and flow angle at diffuser exit is seen to match closely at higher flows. The difference at lower flows could be due to the effect of interaction between impeller and diffuser combinations, as the numerical analysis was done separately for impeller and diffuser and the effect of impeller diffuser interaction was not considered.

Large-Eddy Simulation of Unsteady Surface Pressure Over a LP Turbine Blade Due to Interactions of Passing Wakes and Inflexional Boundary Layer

The unsteady pressure over the suction surface of a modern low-pressure (LP) turbine blade subjected to periodically passing wakes from a moving bar wake generator is described. The results presented are a part of detailed Large-Eddy Simulation (LES) following earlier experiments over the T106 profile for a Reynolds number of 1.6×105 (based on the chord and exit velocity) and the cascade pitch to chord ratio of 0.8. The present LES uses coupled simulations of cylinder for wake, providing four-dimensional inflow conditions for successor simulations of wake interactions with the blade. The three-dimensional, time-dependent, incompressible Navier-Stokes equations in fully covariant form are solved with 2.4×106 grid points for the cascade and 3.05×106 grid points for the cylinder using a symmetry-preserving finite difference scheme of second-order spatial and temporal accuracy. A separation bubble on the suction surface of the blade was found to form under the steady state condition. Pressure fluctuations of large amplitude appear on the suction surface as the wake passes over the separation region. Enhanced receptivity of perturbations associated with the inflexional velocity profile is the cause of instability and coherent vortices appear over the rear half of the suction surface by the rollup of shear layer via Kelvin-Helmholtz (K-H) mechanism. Once these vortices are formed, the steady-flow separation changes remarkably. These coherent structures embedded in the boundary amplify before breakdown while traveling downstream with a convective speed of about 37 percent of the local free-stream speed. The vortices play an important role in the generation of turbulence and thus to decide the transitional length, which becomes time-dependent. The source of the pressure fluctuations on the rear part of the suction surface is also identified as the formation of these coherent structures. When compared with experiments, it reveals that LES is worth pursuing as an understanding of the eddy motions and interactions is of vital importance for the problem.

LES of the Flow and Acoustics Generated by an Aircraft Fan Running in the Vicinity of the Ground

An important component of the aircraft generated noise, especially ahead of it, is the fan noise created by the high velocity variation near the blades and the interaction of the rotating fan with the fluid. In order to predict the sound, the method used involves the acoustical analogy developed by Ffowcs Williams and Hawkings. Computation of the flow field is performed in the incompressible LES framework, while the noise is evaluated using a non-homogeneous wave equation. In this work the influence of the angle between the fan and ground, on the total sound spectrum, the noise generation and the noise propagation is investigated. It is found that the near field is dominated by the blade passage frequency and an upward inclination of the fan will produce a slightly different sound pressure level than a horizontal or downward inclination.

Effects of the Inlet Boundary Layer Thickness on the Flow and Loss Characteristics in an Axial Compressor

A three-dimensional computation was conducted to understand effects of the inlet boundary layer thickness on the internal flow and the loss characteristics in a low-speed axial compressor operating at the design condition (φ = 85%) and near stall condition (φ = 65%). At the design condition, independent of the inlet boundary layer thickness, flows in the axial compressor show similar characteristics such as the pressure distribution, size of hub corner-stall, tip leakage flow trajectory, limiting streamlines on the blade suction surface, etc. But, as the load is increased, for the thick inlet boundary layer at hub and casing, the hub corner stall grows to make a large separation region between the hub and suction surface, and the tip leakage flow is more vortical than that observed in the case with thin inlet boundary layer and has the critical point where the trajectory of the tip leakage flow is suddenly turned to the downstream. For the thin inlet boundary layer, the hub corner stall decays to form the thick boundary layer from hub to midspan on the suction surface owing to the blockage of the tip leakage flow and the tip leakage flow leans to the circumferential direction more than at the design condition. In addition to these, the severe reverse flow, induced by both boundary layers on the blade surface and the tip leakage flow, can be found to act as the blockage of flows near the casing, resulting in a heavy loss. As a result of these differences of the internal flow made by the different inlet boundary layer thickness, the spanwise distribution of the total loss is changed dramatically. At the design condition, total pressure losses for two different boundary layers are almost alike in the core flow region but the larger losses are generated at both hub and tip when the inlet boundary layer is thin. At the near stall condition, however, total loss for thick inlet boundary layer is found to be greater than that for thin inlet boundary layer on most of the span except the region near the hub and casing. In order to analyze effects of inlet boundary layer thickness on total loss in detail, total loss is scrutinized through three major loss categories available in a subsonic axial compressor such as profile loss, tip leakage loss and endwall loss.