Georg Kröger

Theory and Application of Axisymmetric Endwall Contouring for Compressors

Engine operating range and efficiency are of increasing importance in modern compressor design for heavy duty gas turbines and aircraft engines. These highly challenging objectives can only be met if all components provide high aerodynamic performance and stability. The aerodynamic losses of highly loaded axial compressors are mainly influenced by the leakage flow through clearance gaps. Especially the leakage flow due to the radial clearances of rotor blades affects negatively both, the efficiency and the operating range of the engine. Recent publications showed that the clearance flow and the clearance vortex can be influenced by an additional static pressure gradient at the outer casing, which is created by an axisymmetric wavy casing shape. A notable performance increase of up to 0.4% stage efficiency at design point conditions was reported for high pressure stages with large tip clearance heights [1] as well as for a transonic stage with a relatively small radial clearance gap [2]. An analytic approach to predict the effects of axisymmetric casing contouring has been developed at DLR, Institute of Propulsion Technology, and is outlined in the first part of this work. The characteristic behavior of the clearance vortex in an adverse pressure gradient is discussed by means of an inviscid vortex model [3]. The critical vortex parameters are isolated and related to the static pressure increase due to the casing contour. The second part illustrates the application of an axisymmetric endwall contour. A three dimensional optimization of the outer casing and the corresponding blade tip airfoil section of a typical gas turbine high pressure compressor stage with a high number of free variables is presented. The optimization led to a significant increase in aerodynamic performance of about 0.8% stage efficiency and to a notable reduction of the endwall blockage at ADP conditions. Furthermore, an improved off-design performance was found and a simple design rule is given to transfer both, the casing contour and the blade tip section modification on similar high pressure compressor blades. Based on these design rules the results of the optimized stages were applied to the rear stages of a Siemens gas turbine compressor CFD model. An increase of 0.3% full compressor performance was reached at design point conditions.© 2011 ASME

Optimised Aerodynamic Design of an OGV With Reduced Blade Count for Low Noise Aircraft Engines

In order to meet the ACARE environmental goals further noise reduction from aircraft operation is required relative to that achieved to date. To support this aim the European Commission Seventh Framework Programme promotes the “Optimisation for low Environmental Noise impact AIRcraft (OPENAIR)” programme. Within the OPENAIR programme Rolls-Royce leads the Integrated Propulsion System Design sub platform containing a task to produce multi disciplinary optimized Fan System outlet guide vane (OGV) designs. As a partner in this task, DLR have undertaken an aerodynamic optimized design for a novel OGV concept as the first part within this task.As a starting point, Rolls-Royce specified a high bypass ratio turbo fan engine style and a novel OGV number and space envelope within the engine. Within the EU FP6 VITAL programme a lower number of OGVs compared to current practice was investigated [1] and Rolls-Royce had previously undertaken rig tests of a range of fan / OGV blade ratios and showed potential noise penalties but also benefits, particularly for the higher tone harmonics and broadband noise for very low numbers of OGVs [2]. The advent of multidisciplinary optimization techniques for noise, aerodynamics and mechanical constraints led Rolls-Royce to specify a very low number of OGVs to investigate the benefits and the mitigation of the penalties using these optimization techniques. The multidisciplinary design optimization was planned to be undertaken in two parts. The aim of the first part, to be reported in this paper, was to achieve a good aerodynamic design for the novel OGV concept. Controlling the aerodynamic secondary losses was expected to be a challenge for the novel concept. The second part of the optimisation will also include noise.As part of the initial design optimization study within OPENAIR, DLR present a multi objective aerodynamic design optimization for the novel very low OGV number and space envelope specified. The blade number of the OGV was reduced from 42 to 14. The first design phase covered the two dimensional airfoil section design for the reduced blade count OGV. In order to achieve an enhanced aerodynamic performance with a reduced blade number an automated process chain was used to couple the DLR in house optimiser AutoOpti [3, 4] with the two dimensional flow solver MISES [5]. Three operating points were considered in the aerodynamic design covering a wide range of the compressor operating map.It has been found that the design of a low blade count OGV is feasible by means of good aerodynamic performance. Additionally the aerodynamic losses could be reduced in all operating points further and small flow separation areas on the suction surface closed to the hub and tip endwalls implied a further improvement potential. Therefore, the second design phase was focused on the three dimensional endwall optimisation including a variable flow duct and fillet geometry. An automated process chain was established for this purpose. As a result the flow separation was prevented in all operating points. In the last step of the DLR work a high fidelity URANS simulation of the configuration with the newly designed OGV was conducted to assess its aero-acoustical behavior at this stage of the aerodynamically optimized OGV design and the emitted sound power level downstream the OGV for the optimised low blade number configuration is presented in this work.Copyright

Rotor Casing Contouring in High Pressure Stages of Heavy Duty Gas Turbine Compressors With Large Tip Clearance Heights

Clearance leakage losses of axial compressor rotors and stators have a major impact on the overall compressor performance. The clearance heights in the last stages (high pressure stages) of a gas turbine compressor are very large in comparison to the low pressure stages due to mechanical constraints and small blade heights. The reduction of clearance leakage losses in a high pressure stage still holds an important potential for the overall performance improvement at design point conditions. In the following work, a method for tip clearance loss reduction by circumferential casing contouring above a high pressure stage rotor with a constant clearance height is presented. The subsonic compressor blade provides Siemens HPA-Family [1, 2, 3] airfoils. Starting over with a 3DOptimization of the mentioned rotor casing the work additionally refers to the aerodynamic effects and the off design performance of the optimized geometry. It has been found that an optimized casing and blade tip contour lead to a smaller overall clearance mass flow and lower pressure loss coefficient of the clearance flow so that the endwall blockage is reduced and the stage performance is improved by about 0.35% at design point conditions. Furthermore it was found that the performance improvement drops with increasing exit pressure to about 0.1% close to stall conditions. At lower exit pressure values the optimized geometry provides an additional performance improvement in comparison to the baseline configuration.

Influence of a Hub-Groove on the Unsteady Losses of a Transonic Compressor Stage

In this paper the influence of a non-symmetrical sidewall-contour on the flow field of an axial compressor is investigated. In earlier studies, carried out at the DLR Institute of Propulsion Technology in Cologne, it has been shown that it is possible to influence secondary flow regimes by a non-symmetrical sidewall contour. With the help of the contour a vortex is generated that serves as an aerodynamic barrier and thus reduces the flow transport orthogonal to the main flow direction. In this paper the described method of flow manipulation is investigated for the compressor rotor at the hub wall. It turns out that despite the unsteady inflow conditions the vortex is stably generated and works as an aerodynamic separator. The rotor hub region increases its flow capacity. A higher pressure ratio is achieved at a highly loaded operating point. The efficiency at this point remains unchanged.