Multi-Objective Design and Optimisation of Bypass Outlet-Guide Vanes

This paper describes the development of an automated design optimization system that makes use of a high fidelity Reynolds-Averaged CFD analysis procedure to minimize the fan forcing and fan BOGV (bypass outlet guide vane) losses simultaneously taking into the account the down-stream pylon and RDF (radial drive fairing) distortions. The design space consists of the OGV’s stagger angle, trailing-edge recambering, axial and circumferential positions leading to a variable pitch optimum design. An advanced optimization system called SOFT (Smart Optimisation for Turbomachinery) was used to integrate a number of pre-processor, simulation and in-house grid generation codes and postprocessor programs. A number of multi-objective, multi-point optimiztion were carried out by SOFT on a cluster of workstations and are reported herein.

Tip Leakage Flows Near Partial Squealer Rims in an Axial Flow Turbine Stage

This paper deals with an experimental investigation of aerodynamic characteristics of full and partial-length squealer rims in a turbine stage. Full and partial-length squealer rims are investigated separately on the pressure side and on the suction side in the “Axial Flow Turbine Research Facility” (AFTRF) of the Pennsylvania State University. The streamwise length of these “partial squealer tips” and their chordwise position are varied to find an optimal aerodynamic tip configuration. The optimal configuration in this cold turbine study is defined as the one that is minimizing the stage exit total pressure defect in the tip vortex dominated zone. A new “channel arrangement” diverting some of the leakage flow into the trailing edge zone is also studied. Current results indicate that the use of “partial squealer rims” in axial flow turbines can positively affect the local aerodynamic field by weakening the tip leakage vortex. Results also show that the suction side partial squealers are aerodynamically superior to the pressure side squealers and the channel arrangement. The suction side partial squealers are capable of reducing the stage exit total pressure defect associated with the tip leakage flow to a significant degree.

Effects of Freestream Turbulence on the Losses of a Highly-Loaded Compressor Stator Blade

Very little published literature documents the effects of different freestream turbulence intensities on compressor flows at realistically high Reynolds numbers. This paper presents a study of these effects on a transonic, linear, compressor stator cascade. The cascade consisted of high turning stator airfoils that had the camber of 55 degrees. The effects of freestream turbulence intensities of approximately 0.1% (baseline) and 1.6% were examined. Inlet Mach numbers to the cascade were tested from 0.55 to 0.89. Reynolds numbers, based on the inlet conditions and blade chord, varied between 1.0–2.0×106. Inlet flow angles to the cascade ranged from a choking to a stall condition. For the baseline cases, at most positive incidence angles to the cascade, surface oil flow visualization and Schlieren pictures showed a significant flow separation on the suction surface of the blade. Under these conditions, the increase in freestream turbulence from 0.1% to 1.6% significantly reduced the flow losses of the cascade (by as much as 57% in some cases). In other test conditions where no evidence depicted flow separation on the blade, there were no measurable effects on the losses due to the increase in freestream turbulence intensity. In addition, the increase of freestream turbulence intensity also improved the effective operating range of the cascade significantly (e.g., by 46% or higher).

A CAD-Based Blade Geometry Model for Turbomachinery Aero Design Systems

A CAD-centric approach for constructing and managing the blade geometry in turbomachinery aero design systems is presented in this paper. Central to the approach are a flexible CAD-based parametric blade model definition and a set of CAD-neutral interfaces which enable construction and manipulation of the blade solid model directly inside the CAD system’s geometry kernel. A bottleneck of transferring geometry data passively via a file-based method is thus eliminated, and a seamless integration between the CAD system, aero design system, and the larger design environment can be achieved. A single consistent CAD-based blade model is available at all stages of the aero design process, forming the basis for coupling the aero design system to the larger multi-disciplinary design environment. The blade model construction is fully parameterized so that geometry updates can be accurately controlled via parameter changes, and geometric sensitivities of the model can be easily calculated for multidisciplinary interaction and design optimization. A clear separation of the parameters that control the three-dimensional shape of the blade (such as lean and sweep) from the parameters that control the elemental profile shape allows any blade profile family or shape definition to be utilized. The blade model definition, construction interface, and implementation approach are described. Applications illustrating solid model construction, parametric modification and sensitivity calculation, which are key requirements for automated aerodynamic shape design, are presented.

Design, Test and Analysis of a High-Pressure-Ratio Transonic Fan

Aircraft gas turbine engines require compression systems with high performance and low weight and cost. There is therefore a continuing drive to increase compressor stage pressure ratios, particularly for military fans. To meet this need, a technology acquisition programme has been carried out by QinetiQ and Rolls-Royce. Firstly, the stage matching issues for an advanced two-stage military fan were investigated, including the effects of employing variable inlet guide vanes. From this, the requirements for the first stage together with key operating conditions for the blading were defined. The blade profiles were then designed to satisfy the range of aerodynamic conditions using a quasi-3D calculation system. A satisfactory compromise between the aerodynamic and mechanical design requirements was reached in which a blisk construction was employed for the rotor, machined from a single piece of titanium. The new stage was manufactured and tested successfully, and it achieved its target flow, pressure ratio and efficiency on the first build. Detailed measurements of the internal flows using laser anemometry and high response pressure transducers were taken. Finally, these data have been analysed and used to calibrate current 3D multi-row CFD methods.

Adaptive Importance Latin Hypercube Sampling

Probabilistic methods currently require many function evaluations or do not provide a mathematically robust confidence interval. The proposed method searches to find the Most Probable Point (MPP) using a Hasofer-Lind-Rackwitz-Fiessler (HLRF) algorithm, and then estimates reliability with Latin Hypercube Sampling (LHS) evaluating only those points outside of the MPP. Repeated samples provide several estimates of the reliability, which are aggregated to find a reliability estimate with a confidence interval. The computational efficiency is much better than standard LHS sampling and improves as the failure probability decreases. The method is applied to two example problems, each showing a statistically significant reduced confidence interval.

Tailored Structural Design and Aeromechanical Control of Axial Compressor Stall: Part II — Evaluation of Approaches

A theoretical evaluation was conducted delineating how aeromechanical feedback control can be utilized to stabilize the inception of rotating stall in axial compressors. Ten aeromechanical control methodologies were quantitatively examined based on the analytical formulations presented in the first part of this paper (McGee et al, 2003a). The maximum operating range for each scheme is determined for optimized structural parameters, and the various schemes are compared. The present study shows that the most promising aeromechanical designs and controls for a class of low-speed axial compressors were the use of dynamic fluid injection. Aeromechanically incorporating variable duct geometries and dynamically re-staggered IGV and rotor blades were predicted to yield less controllability. The aeromechanical interaction of a flexible casing wall was predicted to be destabilizing, and thus should be avoided by designing sufficiently rigid structures to prevent casing ovalization or other structurally-induced variations in tip clearance. Control authority, a metric developed in the first part of this paper, provided a useful interpretation of the aeromechanical damping of the coupled system. The model predictions also show that higher spatial modes can become limiting with aeromechanical feedback, both in control of rotating stall as well as in considering the effects of lighter, less rigid structural aeroengine designs on compressor stability.

Numerical Investigation of Wake-Shock Interactions and Clocking in a Transonic HP Turbine

A quasi-three-dimensional, blade-to-blade, time-accurate, viscous solver was used for the clocking optimization of a modern transonic heavy-duty, two stage gas turbine. Both stators and rotors operate in a transonic regime with fish-tail shock systems at the blade row exit. These shock systems interact with both stator and rotor wakes. A sensible reduction in the strength of shock waves was observed due to the upstream blade row wake passing. Such wake-shock interactions occur in the inter-blade gap, around locations which are fixed in the frame of reference of the downstream blade-row. The exploitation of such an effect to optimize the axial/circumferential position of blade rows is still compatible with the axial gap values commonly used for these kinds of stages. The results of the clocking investigation will be presented and discussed in terms of unsteady blade loading and efficiency variations.

The Characteristics of Radial Turbines for Small Gas Turbines

Inward flow radial and mixed flow turbines are effectively utilized in both small gas turbine auxiliary power units (APU’s) and turbochargers, where moderately high levels of efficiency can be readily attained with simple cast components, less sensitive to blade end-gap clearances than axial turbines. This paper provides an overview of radial turbine performance characteristics for small gas turbine applications as basically influenced by specific speed, velocity ratio, exit flow coefficient, and rotor tip to exducer root mean square (RMS) diameter ratio. Since turbine rotor mass and inertia play important roles in structural integrity and engine acceleration characteristics, the importance of turbine velocity ratio selection upon rotor tip diameter, and cycle performance are discussed. The effects of rotor reaction on radial turbine flow versus pressure characteristics are examined pertinent to engine matching requirements. Engine transient performance is addressed, as influenced by turbine operation towards and beyond runaway conditions.

Leading Edge Modification Effects on Turbine Cascade Endwall Loss

Experimental results are presented which provide area-averaged total pressure loss coefficient measurements for four different turbine airfoil leading edge configurations. A baseline (Langston) configuration, two leading edge bulbs, and a leading edge fillet were tested in a large-scale, low aspect ratio, high turning linear cascade. Results show that both the small bulb and fillet geometries each reduced area averaged total loss by 8%, while the large bulb exhibited a slight increase in total loss. Contour plots for all geometries are presented and the major differences between each are discussed. Through investigation of pitch averaged loss profiles it is found that the area of greatest reduction differs between the small bulb and fillet, leading to the possibility that the mechanisms through which each is affecting the flow may be different. This provides hope that the best features of each may potentially be combined to determine an optimum shape for secondary flow loss reduction.